Patent Description:
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.

<CIT> discloses a prior art aircraft as set forth in the preamble of claim <NUM>.

<CIT> discloses a prior art thermal management system with thrust recovery for a gas turbine engine fan nacelle assembly.

<CIT> discloses a prior art shared flow thermal management system.

<CIT> discloses a prior art engine bleed system with a motorized compressor.

From one aspect, there is provided an aircraft as recited in claim <NUM>.

There is also provided a method for providing additional heat capacity in a gas turbine engine as recited in claim <NUM>.

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 <NUM>) as well as propulsion loads (i.e. electrical loads within the engine <NUM>). Increased power demands require the incorporation of additional generators, motors and other electrical components within the engine <NUM>. Incorporation of these additional electrical components causes the generation of parasitic heat loads within the engine <NUM> that must be cooled in order to prevent overheating. In addition, physical space within the engine <NUM> 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 <NUM> 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>, <FIG> schematically illustrates an exemplary engine <NUM> mounted to an aircraft wing <NUM> via a mount pylon <NUM>. 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 <NUM> is not required to drive air to the ECS. In the example architecture, one or more electric motors <NUM> drive an electric compressor <NUM> which provides compressed air to the ECS within the aircraft. The electric motors <NUM> receive generated electricity from a generator <NUM> within the engine <NUM>. The air that is compressed by the electric compressor <NUM> is sourced from any ambient air source, and is not sourced from within the engine core. The utilization of the electric motor <NUM> creates a new heat load that is not present in conventional aircraft.

Due to the utilization of the electric compressor <NUM>, ambient air can be obtained and compressed for utilization in the ECS, and it is not necessary to precool 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 <NUM>. The engine system cooler <NUM> is positioned in the pylon space <NUM>, which is within the bifurcation and also contains the mount pylon <NUM>. The engine system cooler <NUM> includes a first fluid inlet <NUM> and a first fluid outlet <NUM>. The first fluid inlet <NUM> connects an internal passage <NUM> within the cooler <NUM> to an engine system <NUM> such that fluid from the engine system <NUM> can be drawn out of the engine <NUM> and provided to the internal passage <NUM>. The first fluid outlet <NUM> returns the fluid from the internal passage <NUM> to the engine <NUM>. In some cases, the fluid is returned to the same engine system <NUM> that it originated from. In alternative cases, the fluid is returned to a distinct engine system <NUM> as part of a larger coolant cycle.

The cooler <NUM> also includes a second inlet <NUM> configured to receive an airflow <NUM> from a fan section <NUM> of the engine <NUM>. The second inlet <NUM> provides the relatively cold fan air to a second internal passage <NUM> (or set of second internal passages <NUM>) within the cooler <NUM>. A second outlet <NUM> dumps spent air from the cooler <NUM> into a bypass flow and allows the spent air to exit the engine <NUM> structure.

In some examples, the cooler <NUM> is an air-air cooler, and the fluid extracted from the engine <NUM> 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 <NUM> is an air-oil cooler with the first inlet <NUM> withdrawing an oil based lubricant/coolant, cooling the oil based lubricant/coolant using the cooler <NUM>, and returning the cooled oil based lubricant/coolant to the engine <NUM> via the first outlet <NUM>.

With continued reference to <FIG>, and with like numerals indicating like elements, <FIG> schematically illustrates an alternative example engine <NUM> mounted to an aircraft wing <NUM> via a mounting pylon <NUM>. Air for the ECS is acquired via an electrically driven compressor mounted within the aircraft and bleed air from the turbine engine <NUM> is not required to drive air to the ECS. In the example architecture, one or more electric motors <NUM> drive an electric compressor <NUM> which provides compressed air to the ECS within the aircraft. The air that is compressed by the electric compressor <NUM> is sourced from any ambient air source, and is not sourced from within the engine core. The utilization of the electric motor <NUM> creates a new heat load that is not present in conventional aircraft.

Disposed within the pylon space <NUM> (within the bifurcation) are three heat exchangers <NUM>, <NUM>', <NUM>". In some examples, each of the heat exchangers <NUM>, <NUM>', <NUM>" is connected to the same engine system <NUM>. In alternative examples, each heat exchanger <NUM>, <NUM>', <NUM>" is connected to a distinct engine system <NUM>. The heat exchangers <NUM>, <NUM>', <NUM>" each include internal configurations substantially similar to that described above with regards to the heat exchanger <NUM> and illustrated in <FIG>, 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 <NUM>, <NUM> in the pylon space <NUM>, <NUM> provides for an improved heat exchanger efficiency, because being positioned in the engine nacelle bifurcation allows the heat exchanger <NUM>, <NUM> efficiency to benefit from a relatively high source pressure obtained from the ram effect of having the entire inlet <NUM>, <NUM> duct perimeter be entirely in the fan stream <NUM>. 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 <NUM>. By taking advantage of the available ram effect, a higher-efficiency, smaller-sized heat exchanger can be used in the pylon space <NUM> as compared to heat exchangers that are packaged interior to the engine nacelle.

Further, including the engine system cooler <NUM>, <NUM> in the pylon space <NUM>, <NUM> 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 <NUM>, <NUM> to the pylon space <NUM>, <NUM> 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 <NUM>, <NUM> within the pylon space <NUM>, <NUM> 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.

Claim 1:
An aircraft comprising:
at least one gas turbine engine (<NUM>; <NUM>; <NUM>) connected to a wing (<NUM>; <NUM>) via an engine pylon (<NUM>; <NUM>), the gas turbine engine (<NUM>; <NUM>; <NUM>) including an engine core having a compressor section (<NUM>), a combustor section (<NUM>), a turbine section (<NUM>), and at least one actively cooled engine system (<NUM>; <NUM>);
a fan (<NUM>; <NUM>) disposed fore of the engine core and rotatably connected to the engine core via a shaft (<NUM>), the engine pylon (<NUM>; <NUM>) being disposed within a bifurcation that intersects a fan stream (<NUM>) of the fan (<NUM>; <NUM>) and includes a pylon space (<NUM>; <NUM>); and
at least one heat exchanger (<NUM>; <NUM>; <NUM>'; <NUM>") disposed within the pylon space (<NUM>; <NUM>), the heat exchanger (<NUM>; <NUM>; <NUM>'; <NUM>") including a first inlet (<NUM>) connected to the at least one actively cooled engine system (<NUM>; <NUM>) and a first outlet (<NUM>) connected to the at least one actively cooled engine system (<NUM>; <NUM>),
wherein the at least one actively cooled engine system (<NUM>; <NUM>) is interior to the engine core,
characterised in that
the aircraft comprises an electric generator (<NUM>) connected to said engine core such that rotation of the electric generator (<NUM>) is driven by the shaft (<NUM>) and an electric compressor (<NUM>; <NUM>) is driven by the electric generator (<NUM>), wherein the electric compressor (<NUM>; <NUM>) is configured to source air from an ambient air source, and wherein a fluid outlet of the electric compressor (<NUM>; <NUM>) is fluidly connected to an aircraft cabin environmental cooling system (ECS).