Patent Description:
A gas turbine engine may include a compressor section, among other sections. During operation, the compressor section temperatures may increase. Gas turbine engines may undergo extended idle run times following engine operation to stabilize high pressure compressor (HPC) temperatures and to prevent excessive thermal gradients.

<CIT> discloses a system for supplying external air to a compressor and to burner casing.

<CIT>, which is prior art under Article <NUM>(<NUM>) EPC, discloses a system and method of reducing post-shutdown engine temperatures.

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

According to a further aspect of the present invention, there is provided a method as set forth in claim <NUM>.

The detailed description of exemplary embodiments herein makes reference to the accompanying drawings, which show exemplary embodiments by way of illustration. While these exemplary embodiments are described in sufficient detail to enable those skilled in the art to practice the inventions, it should be understood that other embodiments may be realized and that logical changes and adaptations in design and construction may be made in accordance with this invention and the teachings herein.

As used herein, "aft" refers to the direction associated with the tail (e.g., the back end) of an aircraft, or generally, to the direction of exhaust of the gas turbine. As used herein, "forward" refers to the direction associated with the nose (e.g., the front end) of an aircraft, or generally, to the direction of flight or motion. As used herein, "gas" and "air" may be used interchangeably.

Various components included in a high pressure compressor (HPC), including the rotor and case may cool down at different rates, leading to unequal thermal expansion which may physically deform the system. Such thermal inequality may lead to bowed rotors which can cause deflection of the HPC rotor. Starting the engine in this condition can lead to tip strike of the HPC blades against blade outer air seals (BOAS). Cooling systems, as provided herein, may aid in more uniform cooling of gas turbine engine components and may reduce engine maintenance time.

In various embodiments and with reference to <FIG>, a gas turbine engine <NUM> is provided. Gas turbine engine <NUM> may be a two-spool turbofan that generally incorporates a fan section <NUM>, a compressor section <NUM>, a combustor section <NUM> and a turbine section <NUM>. Alternative engines may include, for example, an augmentor section among other systems or features. In operation, fan section <NUM> can drive air along a bypass flow-path B while compressor section <NUM> can drive air along a core flow-path C for compression and communication into combustor section <NUM> then expansion through turbine section <NUM>. Although depicted as a turbofan gas turbine engine <NUM> herein, 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.

Gas turbine engine <NUM> may generally comprise a low speed spool <NUM> and a high speed spool <NUM> mounted for rotation about an engine central longitudinal axis A-A' relative to an engine static structure <NUM> via one or more bearing systems <NUM> (shown as bearing system <NUM>-<NUM> and bearing system <NUM>-<NUM> in <FIG>). It should be understood that various bearing systems <NUM> at various locations may alternatively or additionally be provided including, for example, bearing system <NUM>, bearing system <NUM>-<NUM>, and bearing system <NUM>-<NUM>.

Low speed spool <NUM> may generally comprise an inner shaft <NUM> that interconnects a fan <NUM>, a low pressure (or first) compressor section <NUM> and a low pressure (or first) turbine section <NUM>. Inner shaft <NUM> may be connected to fan <NUM> through a geared architecture <NUM> that can drive fan <NUM> at a lower speed than low speed spool <NUM>. Geared architecture <NUM> may comprise a gear assembly <NUM> enclosed within a gear housing <NUM>. Gear assembly <NUM> couples inner shaft <NUM> to a rotating fan structure. High speed spool <NUM> may comprise an outer shaft <NUM> that interconnects a high-pressure compressor ("HPC") <NUM> (e.g., a second compressor section) and high pressure (or second) turbine section <NUM>. A combustor <NUM> may be located between HPC <NUM> and high pressure turbine <NUM>. A mid-turbine frame <NUM> of engine static structure <NUM> may be located generally between high pressure turbine <NUM> and low pressure turbine <NUM>. Mid-turbine frame <NUM> may support one or more bearing systems <NUM> in turbine section <NUM>. Inner shaft <NUM> and outer shaft <NUM> may be concentric and rotate via bearing systems <NUM> about the engine central longitudinal axis A-A', which is collinear with their longitudinal axes. As used herein, a "high pressure" compressor or turbine experiences a higher pressure than a corresponding "low pressure" compressor or turbine.

The core airflow C may be compressed by low pressure compressor <NUM> and HPC <NUM>, mixed and burned with fuel in combustor <NUM>, then expanded over high pressure turbine <NUM> and low pressure turbine <NUM>. Mid-turbine frame <NUM> includes airfoils <NUM> which are in the core airflow path. Low pressure turbine <NUM> and high pressure turbine <NUM> rotationally drive, respectively, low speed spool <NUM> and high speed spool <NUM> in response to the expansion.

Gas turbine engine <NUM> may be, for example, a high-bypass geared aircraft engine. In various embodiments, the bypass ratio of gas turbine engine <NUM> may be greater than about six (<NUM>). In various embodiments, the bypass ratio of gas turbine engine <NUM> may be greater than ten (<NUM>). In various embodiments, geared architecture <NUM> may be an epicyclic gear train, such as a star gear system (sun gear in meshing engagement with a plurality of star gears supported by a carrier and in meshing engagement with a ring gear) or other gear system. Geared architecture <NUM> may have a gear reduction ratio of greater than about <NUM> and low pressure turbine <NUM> may have a pressure ratio that is greater than about <NUM>. In various embodiments, the bypass ratio of gas turbine engine <NUM> is greater than about ten (<NUM>:<NUM>). In various embodiments, the diameter of fan <NUM> may be significantly larger than that of the low pressure compressor <NUM>, and the low pressure turbine <NUM> may have a pressure ratio that is greater than about <NUM>:<NUM>. The pressure ratio of low pressure turbine <NUM> may be measured prior to inlet of the low pressure turbine <NUM> as related to the pressure at the outlet of low pressure turbine <NUM> prior to an exhaust nozzle. It should be understood, however, that the above parameters are exemplary of various embodiments of a suitable geared architecture engine and that the present disclosure contemplates other gas turbine engines including direct drive turbofans. <FIG> and <FIG> provide a general understanding of the sections in a gas turbine engine, and are not intended to limit the disclosure. The present disclosure may extend to all types of turbine engines, including turbofan gas turbine engines and turbojet engines, for all types of applications.

With respect to <FIG>, elements with like element numbering as depicted in <FIG> are intended to be the same and will not necessarily be repeated for the sake of clarity.

<FIG> illustrates the primary flow gas path through core engine <NUM>, in accordance with various embodiments. Core engine <NUM> may include engine static structure <NUM>, low-pressure compressor <NUM>, HPC <NUM>, combustor <NUM>, high-pressure turbine <NUM>, and low-pressure turbine C. Engine static structure <NUM> may be referred to as an engine case. Gas may flow into low-pressure compressor <NUM> along core flow-path C. Gas flowing through low-pressure compressor <NUM> along core flow-path C may be compressed, resulting in an increase in pressure and temperature relative to the pressure and temperature upon entering low-pressure compressor <NUM>. Gas may flow into HPC <NUM> along core flow-path C. Gas flowing through HPC <NUM> along core flow-path C may be compressed, resulting in an increase in pressure and temperature relative to the pressure and temperature upon entering HPC <NUM>. Uncombusted gas in core flow-path C leaving HPC <NUM> may be referred to as T3 gas <NUM>. T3 gas may have a varying temperature at different engine speeds. The temperature of T3 gas may be about <NUM>°F (<NUM>) when core engine <NUM> is at idle speeds and may reach about <NUM>,<NUM>°F (<NUM>) or higher as core engine <NUM> accelerates for takeoff, where the term "about" in this context only may refer to +/- <NUM>°F. Different engines may have higher temperatures or lower temperatures at each stage. T3 gas may be present at location <NUM> of core engine <NUM>. T3 gas leaving the HPC <NUM> may then flow into combustor <NUM> to supply combustor <NUM> with air for combustion.

With respect to <FIG>, elements with like element numbering as depicted in <FIG> and <FIG> are intended to be the same and will not necessarily be repeated for the sake of clarity.

With reference to <FIG>, an active core cooling system (hereinafter referred to as "cooling system") <NUM> is illustrated, in accordance with various embodiments. Cooling system <NUM> may include a valve system <NUM>. Valve system <NUM> may include a valve, such as butterfly valve <NUM> for example, and an actuation device, such as actuation device <NUM> for example. Actuation device <NUM> may include a muscle pressure line <NUM>. However, any suitable valve system having a valve and an actuation device for actuating the valve between an open and closed position is contemplated herein. Valve system <NUM> may be fluidly coupled to an air duct <NUM> via an inlet <NUM>. The air duct <NUM> comprises an engine starter air duct. Valve system <NUM> may be fluidly coupled to an engine case <NUM> via an outlet <NUM>. In various embodiments, engine case <NUM> comprises a high pressure compressor (HPC) case. Valve system <NUM> may be located radially outward (y-direction) from engine case <NUM>. Inlet <NUM> and outlet <NUM> may comprise a pipe, tube, or any other suitable duct. In various embodiments, inlet <NUM> and outlet <NUM> may comprise carbon steel, stainless steel, beryllium copper, an austenitic nickel-chromium-based super alloy, a titanium alloy, or any other material suitable for high temperatures.

In various embodiments, cooling air <NUM> may be supplied via an aircraft auxiliary power unit (APU). In various embodiments, cooling air <NUM> may be supplied via a ground cart air-supply. Cooling air <NUM> enters plenum <NUM> in response to valve system <NUM> actuating to an open position. In various embodiments, plenum <NUM> may comprise an HPC bleed air plenum. Cooling air <NUM> then enters core flow-path C. Cooling air <NUM> may travel in a forward direction (negative z-direction) in core flow-path C. Cooling air <NUM> may travel in an aft direction (positive z-direction) in core flow-path C. Cooling air <NUM> may also enter bore <NUM>. In this regard, cooling air <NUM> aids in evenly cooling HPC <NUM>. In this regard, cooling air <NUM> aids in evenly cooling components around core flow-path C. Components that may be cooled include blades, vanes (i.e., blade <NUM> or vane <NUM>) and rotor <NUM> of HPC <NUM>, as well as any other components located within HPC <NUM>. In this regard, inlet <NUM> is in fluid communication with air duct <NUM>. Furthermore, in this regard, outlet <NUM> is in fluid communication with engine case <NUM>, plenum <NUM>, and core flow-path C. Still furthermore, valve system <NUM> is in fluid communication with inlet <NUM> and outlet <NUM>.

In various embodiments, cooling system <NUM> may be for cooling HPC <NUM> in response to the gas turbine engine being powered down. For example, after a gas turbine engine operates, the HPC <NUM> may be hot, such as between about <NUM>°F (<NUM>) and about <NUM>,<NUM>°F (<NUM>), for example. Thus, after turning off the gas turbine engine, cooling system <NUM> may be operated in order to evenly cool down HPC <NUM> and core flow-path C.

In various embodiments, valve system <NUM> may comprise a butterfly valve <NUM>. In various embodiments, valve system <NUM> may include a muscle pressure line <NUM>. Air from inlet <NUM> may be communicated into valve system <NUM> via muscle pressure line <NUM> to actuate butterfly valve <NUM>. Thus, valve system <NUM> may be actuated using a pressure of the cooling air <NUM> located upstream from the valve. Air duct <NUM> may include a starter air valve <NUM>. Valve system <NUM> may be coupled to air duct <NUM> at a location upstream from starter air valve <NUM>.

In various embodiments, an electronic engine controller (EEC) <NUM> may be in electronic communication with valve system <NUM>. EEC <NUM> may be configured to monitor valve system <NUM>. Thus, EEC <NUM> may determine if valve system <NUM> is in an open position, a closed position, or any other position. EEC <NUM> may be in electronic communication with engine interface unit (EIU) <NUM>. In various embodiments, EIU <NUM> may control valve system <NUM>. In various embodiments, EIU may control valve system <NUM> via EEC <NUM>.

In various embodiments, a bleed air monitoring computer (BMC) <NUM> may be in electronic communication with valve system <NUM>. EIU <NUM> may be in electronic communication with BMC <NUM>. BMC <NUM> may be configured to control valve system <NUM> in response to EEC <NUM> being powered off. EIU <NUM> may receive valve position feedback from valve system <NUM>.

In general, aircraft EECs lose power upon engine shutdown. Thus, it may be necessary for valve system <NUM> to be controlled via a system other than the EEC. In this regard, EIU <NUM> may be in direct electronic communication with valve system <NUM> as illustrated, in accordance with various embodiments. Furthermore, EIU <NUM> may be in electronic communication with valve system <NUM> via BMC <NUM>, in accordance with various embodiments. Although illustrated as being in electronic communication with all three of EIU <NUM>, BMC <NUM>, and EEC <NUM>, it is contemplated herein that valve system <NUM> may be in electronic communication with any one of, or a combination of, EIU <NUM>, BMC <NUM>, and/or EEC <NUM>. Furthermore, it is contemplated herein that EEC <NUM> may be configured to remain powered on after an engine shutdown. Thus, EEC <NUM> may be used to control valve system <NUM> after an engine shutdown, in accordance with various embodiments.

In various embodiments, a temperature sensor <NUM> may be coupled to outlet <NUM>. Temperature sensor <NUM> may monitor the temperature of cooling air <NUM>. Temperature sensor <NUM> may monitor the temperature of outlet <NUM>. Although shown as being coupled to outlet <NUM>, it is contemplated herein that temperature sensor <NUM> may be coupled to engine case <NUM>. Thus, temperature sensor <NUM> may be configured to monitor the temperature of engine case <NUM> and/or plenum <NUM>. Temperature sensor <NUM> may send a signal to EEC <NUM> corresponding to the measured temperature. Cooling system <NUM> may use said signal from temperature sensor <NUM> to control cooling system <NUM>.

With reference to <FIG>, a method <NUM> for cooling a core flow-path of a compressor section of a gas turbine engine is provided. Method <NUM> may include coupling an inlet of a valve to an engine starter air duct (see step <NUM>). Method <NUM> includes coupling an outlet of the valve to a high pressure compressor (HPC) plenum (see step <NUM>).

In various embodiments, with additional reference to <FIG>, step <NUM> may include coupling inlet <NUM> of valve system <NUM> to air duct <NUM>, the inlet <NUM> being in fluid communication with air duct <NUM>. Step <NUM> may include coupling outlet <NUM> of valve system <NUM> to an HPC plenum <NUM>, the outlet <NUM> being in fluid communication with the plenum <NUM>. Core flow-path C may receive cooling air <NUM> from the air duct <NUM> via valve system <NUM> in response to the valve system <NUM> being actuated to an open position.

In various embodiments, the inlet <NUM> may be welded to air duct <NUM>.

In various embodiments, as previously mentioned, valve system <NUM> may comprise an actuation device <NUM> that acts to open and close valve system <NUM>. For example, actuation device <NUM> may actuate butterfly valve <NUM> via muscle pressure line <NUM> in response to a command received from EEC <NUM>, EIU <NUM>, and/or BMC <NUM>. In various embodiments, valve system <NUM> may include a spring which acts to close butterfly valve <NUM> in response to valve system <NUM> losing power.

With reference to <FIG>, cooling system <NUM> is illustrated, in accordance with various embodiments. Although <FIG> illustrates outlet <NUM> being coupled directly to engine case <NUM>, outlet <NUM> may be coupled to engine case <NUM> via environmental control system (ECS) duct <NUM>. ECS duct <NUM> may comprise a high pressure duct. Thus, outlet <NUM> may be coupled to ECS duct <NUM>. ECS duct <NUM> may comprise a duct for communicating compressor bleed air to other aircraft systems for environmental control of the aircraft. In various embodiments, outlet <NUM> may be welded to ECS duct <NUM>. In this regard, ECS duct <NUM> is in fluid communication with plenum <NUM>.

Connecting lines shown in the various figures contained herein are intended to represent exemplary functional relationships and/or physical couplings between the various elements. In the appended claims, a reference to an element in the singular is not intended to mean "one and only one" unless explicitly so stated, but rather "one or more.

Claim 1:
A gas turbine engine (<NUM>) comprising:
an engine case (<NUM>);
an engine starter air duct (<NUM>);
a high pressure compressor section (<NUM>) having a plenum (<NUM>);
a core flow-path (C); and
a cooling system (<NUM>) comprising a valve system (<NUM>) located radially outward from the engine case (<NUM>), and being coupled between the engine starter air duct (<NUM>) and the engine case (<NUM>), the valve system (<NUM>) having an actuation device (<NUM>) configured to open and/or close a valve in response to a command from an electronic engine controller (<NUM>), wherein cooling air (<NUM>) from the engine starter air duct (<NUM>) can pass through the valve system (<NUM>), enter the engine case (<NUM>) into the high pressure compressor plenum (<NUM>), and then enter the core flow-path (C), when the valve system (<NUM>) is in an open position, and wherein the valve (<NUM>) is actuated using a pressure of the cooling air (<NUM>) located upstream from the valve system (<NUM>).