METHOD OF OPERATING A GAS TURBINE ENGINE, COMPRESSED AIR DELIVERY SYSTEM AND AIRCRAFT INCORPORATING SAME

There is described a method of operating a gas turbine engine having a rotor rotatable about a rotation axis, a core gas path defined annularly around the rotation axis, and a casing defining a wall to the core gas path. The method generally has: generating compressed air from a compressed air source, the compressed air source being external relative the gas turbine engine; and guiding the compressed air in sequence from the compressed air source, radially inwardly relative the rotation axis, through a bleed port and into the core gas path.

TECHNICAL FIELD

The disclosure generally relates to the operation of a gas turbine engine, and more particularly relates to the delivery of compressed air to the gas turbine engine.

BACKGROUND OF THE ART

Rotor bow, sometimes referred to as thermal bow, is a phenomena associated with engine shafts where different engine parts are each exposed to different thermal loading. For example, when an aircraft engine enters a shut down phase after landing, the engine/combustor continues to reject heat to its adjacent components. The rejected heat rises against gravity, leaving the upper sectors hotter than the lower sectors. This uneven thermal loading results in a differential thermal expansion which can ultimately result in slender components such as shafts to deflect, resulting in a bowed rotor condition. Under a bowed rotor condition, a subsequent start could lead to undesired occurrences of component rubbing.

Although existing rotor bow prevention techniques such as continuing to operate the engine shaft to rotate the fans are satisfactory to a certain degree, there remains room for improvement.

SUMMARY

In one aspect, there is provided a method of operating a gas turbine engine, the gas turbine engine having a rotor rotatable about a rotation axis, a core gas path defined annularly around the rotation axis, and a casing defining a wall to the core gas path, the method comprising: generating compressed air from a compressed air source, the compressed air source being external relative the gas turbine engine; and guiding the compressed air in sequence from the compressed air source, radially inwardly relative the rotation axis, through a bleed port and into the core gas path.

In another aspect, there is provided a compressed air delivery system for an aircraft having a gas turbine engine, an aircraft system, and a fluid conduit network, the gas turbine engine having a rotor rotatable about a rotation axis, a core gas path defined annularly around the rotation axis, a casing defining a wall to the core gas path, a bleed port in the wall, open to the core gas path, the fluid conduit network configured for fluidly connecting the bleed port to the aircraft system sequentially via an engine conduit extending radially outwardly relative the rotation axis and an aircraft conduit extending externally relative the gas turbine engine, the compressed air delivery system comprising: at least one valve and an external source conduit fluidly integrated to the fluid conduit network, the at least one valve selectively operable to a first configuration fluidly connecting the engine conduit to the aircraft conduit and partitioning the external source conduit, and a second configuration fluidly connecting the engine conduit to the external source conduit and partitioning the aircraft conduit.

In a further aspect, there is provided an aircraft comprising: an aircraft system; a gas turbine engine having a rotor rotatable about a rotation axis, a core gas path defined annularly around the rotation axis, a casing defining a wall to the core gas path, a bleed port in the casing open to the core gas path; a fluid conduit network configured for fluidly connecting the bleed port to the aircraft system sequentially via an engine conduit extending radially outwardly relative the rotation axis and an aircraft conduit extending externally relative the gas turbine engine, the fluid conduit network having a valve and an external source conduit fluidly integrated thereto; and a compressed air source configured for generating compressed air and guiding the compressed air into the external source conduit; the valve being selectively operable to a first configuration fluidly connecting the engine conduit to the aircraft conduit, and to a second configuration fluidly connecting the engine conduit to the external source conduit and partitioning the aircraft conduit.

Embodiments may include combinations of the above features.

DETAILED DESCRIPTION

The fan12is drivingly interconnected, directly or indirectly, to low pressure rotor(s) of the turbine section18through a low pressure shaft20, and the high pressure rotor(s) of the compressor section14is/are drivingly connected to high pressure rotor(s) of the turbine section18through a high pressure shaft22concentrically surrounding the low pressure shaft20. The gas turbine engine10has a rotation axis11about which rotatable components of the gas turbine engine10, such as the low and high pressure rotor(s), rotates during use.

The gas turbine engine10may include an accessory drive assembly24which includes an accessory gearbox (AGB)26. Although not shown, the accessory drive assembly24can also include a pump assembly and/or an air starter generator. The accessory drive assembly24may be driven by the high pressure shaft22via an accessory shaft28which drivingly interconnects the high pressure shaft22and the accessory gearbox26. Bearings30may be used to rotatably support different components of the engine10, such as the low and high pressure shafts20and22, and components of the accessory gearbox26.

The gas turbine engine10has a casing32which encloses the turbo machinery of the engine, and an outer casing or nacelle34disposed radially outwardly of the casing32. The casing32and the nacelle34collectively define two compressed gas paths36in this embodiment, including a core gas path36aand a bypass gas path36B. The air propelled by the fan12is split into a first portion which flows through the core of the engine via the core gas path36a, which is circumscribed by a radially outer wall37of the casing32and allows the flow to circulate through the multistage compressor14, combustor16and turbine section18as described above. The remainder portion of the propelled air flows around the casing32within the bypass gas path36B.

As depicted, the nacelle34is generally suspended to an aircraft wing40(or fuselage) via a pylon42. A fluid conduit network43is provided. As shown, the fluid conduit network43is configured for fluidly connecting the core gas path36ato an aircraft system45sequentially via engine conduit44extending radially outwardly relative the rotation axis11and an aircraft conduit47extending externally relative the gas turbine engine10. Examples of such engine conduits include compressed air conduits44aand44bwhich in this example fluidly connect portions of the core gas path36ato the aircraft conduit47which itself leads to an environmental control system (ECS)46or any other aircraft system. As shown in this embodiment, the compressed air conduits44aand44bcan convey compressed air bleeding from the core gas path36avia corresponding bleed ports, such as low and high pressure bleed ports48aand48b. The compressed air conduits may44aand44brun through strut(s)45, the pylon42and/or the aircraft wing40depending on the embodiment. As compressed air originating from the low pressure bleed port48amay be of lower pressure and lower temperature, compressed air originating from the high pressure bleed port48bmay be of higher pressure and higher temperature. As shown in this embodiment, the low pressure bleed port48ais located at a lower pressure portion of the core gas path36awhereas the high pressure bleed port48bis located at a higher pressure portion of the core gas path36a. The locations of the bleed ports are carefully optimized depending upon the flight envelope of the aircraft and the engine power settings, and can thereby differ from one embodiment to another. Bleed port(s) may be located in the nacelle34and open to the bypass gas path36b. In some embodiments, the gas turbine engine may have one, two, three or more bleed ports located at different portions of the core gas path36a. During normal working conditions of the gas turbine engine10, the pressurized air originating from the low and high pressure bleed ports48aand48bcan be mixed to one another so as to be used by the ECS46to control cabin pressure. Other aircraft systems can be configured for de-icing the nacelle front lip and wings, and start the other engines, for instance. For instance, the pressurized air originating from the low and high pressure bleed ports48aand48b, and perhaps from bleed ports open to the bypass gas path36b, can be mixed to one another using a mixer to obtain pressurized air of a given pressure and temperature where desired.

It is known that when the gas turbine engine10enters a shut down phase after landing, the hot components progressively cool down and by doing so dissipate heat therearound, which can heat up adjacent components unevenly. The rejected heat can rise against gravity, leaving the hardware in upper sectors hotter than the hardware in the lower sectors. This uneven thermal loading results in a differential thermal expansion of long and slender components like the low and high pressure shafts20and22causing them to deflect, resulting in a bowed rotor condition. Under a bowed rotor condition, a subsequent start of the gas turbine engine10can lead to unexpected component rubs. It is thus desirable to deliver compressed air to either one or both the low and high pressure shafts20and22, or other components of the gas turbine engine10, for a given period of time after the shutting down of the gas turbine engine10has been initiated, for instance. Delivering compressed air to the components surrounding the core gas path36may also be used in embodiments where the gas turbine engine10is to be pre-heated prior to ignition.

FIG. 2shows an example of a compressed air delivery system100for an aircraft having a gas turbine engine10, an aircraft system45and a fluid conduit network43. The compressed air delivery system100can be used with any type of gas turbine engine including, but not limited to, a turbofan engine, a turboshaft engine and the like. The operation of the compressed air delivery system100can be manually or automatically controlled depending on the embodiment. As shown, the gas turbine engine10having a rotor rotatable about a rotation axis11, a core gas path36adefined annularly around the rotation axis11, a casing32defining a wall37to the core gas path36a, a bleed port148in the casing32open to the core gas path36a. The fluid conduit network43fluidly connects the bleed port148to the aircraft system45sequentially via engine conduit(s)144extending radially outwardly relative the rotation axis11and an aircraft conduit47extending externally relative the gas turbine engine10.

As depicted, the compressed air delivery system100has one or more valve152and an external source conduit149fluidly integrated to the fluid conduit network43. The valve(s)152is selectively operable to a first configuration fluidly connecting the engine conduit(s)144to the aircraft conduit47, and to a second configuration fluidly connecting the engine conduit(s)144to the external source conduit149and partitioning the aircraft conduit47. The valve152can be any suitable type of valve including, but not limited to, butterfly valve, knife-gate valve, ball-type valve, bi-directional valve, and the like. For instance, in the first configuration the valve152can prevent the compressed air to flow through the bleed port(s)148. In the second configuration the valve152to can allow the compressed air to flow radially inwardly through the bleed port(s)148and into the core gas path36awhile preventing the compressed air from being guided in the aircraft conduit47.

As depicted, the gas turbine engine10includes one or more bleed ports148. Examples of the bleed port(s)148can include, but are not limited to, the low and high bleed ports48aand48bdescribed above. As shown, the bleed port(s)148are in fluid communication with the core gas path36aof the gas turbine engine10. More specifically, the bleed port(s)148can be axially or circumferentially spaced from one another along casing32.

The fluid conduit network43includes one or more engine conduit(s)144. Examples of the engine conduit(s)144can include, but not limited to, the engine conduits44aand44bthat are in communication with the low and high pressure bleed ports48aand48b. As shown, the engine conduit(s)144are in fluid communication with the bleed port(s)148. The engine conduits144may be used for bleeding compressed air out of the core gas path36avia the bleed port(s)148and towards the aircraft conduit47leading to the environmental control system46during normal working conditions of the gas turbine engine10. Although each bleed port148has its own dedicated engine conduit144in the illustrated embodiment, a single engine conduit144can be fluidly connected to more than one of the bleed ports148. Also, a single bleed port148can be fluidly connected to more than one engine conduit144in some other embodiments. The bleed ports148are part of the compressor section14or combustor16of the gas turbine engine10. As depicted, the corresponding engine conduits44band44crun from the bleed ports158, through the casing32, the strut45, the nacelle34and pylon42towards the aircraft.

A compressed air source150external to the gas turbine engine10can also be provided. The compressed air source150is in fluid communication with the external source conduit149. As shown, in a configuration of the valve152, the compressed air source150is configured for guiding compressed air in sequentially in the external source conduit149and in the engine conduit(s)144, through the bleed port(s)148and into the core gas path38aof the gas turbine engine10. Still referring toFIG. 2, the flow of the compressed air is depicted by dashed arrows A. By doing so, the compressed air can flow against the low and high pressure shafts20and22, and the casing32, towards an exhaust portion of the turbine section18. By way of this flow, the low and high pressure shafts20and22, and other components of the gas turbine engine10, may be heated up or cooled down depending on whether the compressed gas delivery system100is operated before ignition of the gas turbine engine10or after the shut down of the gas turbine engine10has been initiated. In the latter case, the cooling down of the low and high pressure shafts20and22may reduce the risks of rotor bending. In some embodiments, the compressed air source150is an auxiliary power unit (APU) which is part of the aircraft but external to the gas turbine engine10. In some other embodiments, the compressed air source150can be part of a ground power unit. For instance, the compressed air source150can be a ground-based compressed air source such as a blower motor. In these embodiments, ends of the compressed air conduit(s)144may be exposed thereby facilitating the connections with the ground-based compressed air source. To this end, external compressed air conduits such as external conduit149may be used.

In some embodiments, the flow of compressed air incoming from the compressed air source150may be sufficient to induce rotation of the low and/or high pressure shafts20and22up to a given rotation speed. In these embodiments, the rotation of the low and/or high pressure shafts20and22may help in dissipating heat during the pre-heating and/or cooling periods, depending on when the compressed air delivery system100is operated. In some embodiments, the rotation speed of the low and/or high pressure shafts20and22can be monitored over time. In these embodiments, the flow of the compressed air source150can be maintained and/or increased for a given period of time until the monitored rotation speed reaches a given rotation speed threshold. Either one or both of the given period of time and the given rotation speed threshold can be depending on surrounding environmental conditions. For instance, in colder weather conditions, the period of time during which compressed air is flowed radially inwardly through the bleed port(s)148and into the core gas path36amay be shorter than in warmer weather conditions as the surrounding environmental conditions may contribute to the cooling of the gas turbine engine10. The rotational speed threshold can also depend on the surrounding environmental conditions. In embodiments where cooling of the engine10is sought, the rotation speed threshold can be higher in colder weather conditions than in warmer weather conditions, and vice versa.

The construction of the compressed air delivery system100can differ from one embodiment to another. For instance, the compressed air delivery system can use existing valve and conduits of the aircraft, such as described with reference toFIG. 3, or use dedicated valve and conduits, such as described with reference toFIG. 4.

FIG. 3shows a schematic view of an aircraft302having a gas turbine engine310, an ECS346and an APU350. The gas turbine engine310has a core gas path336a, a casing332through which the core gas path336aextends, and a rotor rotatable about a rotation axis311inside the casing332. The casing332has a low pressure bleed port348alocated at a low pressure portion of the core gas path36a, and a high pressure bleed port348blocated at a high pressure portion of the core gas path336a. During normal working conditions of the gas turbine engine310, compressed air bleeding from the low and high pressure bleed ports348aand348bcan be mixed using a mixer360and then directed towards the ECS346. Using external source conduit(s)349, the APU350can convey compressed air towards an air starter362of the gas turbine engine310during the starting of the gas turbine engine310as well as to the ECS346during normal working conditions of the gas turbine engine310. The compressed air is flowed between the APU350, the low and high pressure bleed ports348aand348band the ECS346using a fluid conduit network343incorporating a number of engine conduit(s)344, aircraft conduit(s)347and external source conduit(s)349. In some embodiments, the APU350supplies compressed air not only to the low pressure and high pressure bleed ports348aand348bbut also to the air starter362to energize it at very low speeds. In some embodiments, the APU350has an auxiliary rotor casing332′ defining an auxiliary gas path (not shown), an auxiliary bleed port348′ open to the auxiliary gas path and in fluid communication with the valve via the external source conduit349. As such, compressed air originating from the auxiliary bleed port348′ of the APU350can be guided towards the low and high pressure bleed ports348aand348bof the gas turbine engine310

In this example, a compressed air delivery system300is shown. As depicted, the compressed air delivery system300is implemented using the fluidic circuit of the aircraft302. The compressed air delivery system300has the low and high pressure bleed ports348aand348bthat are open to the core gas path36in the casing332. The compressed air delivery system300also includes a compressed gas source external, in this case provided in the form of the aircraft-based or onboard APU350, to the gas turbine engine310and in fluid communication with the low and high pressure bleed ports348aand348via corresponding engine conduits344. The fluid conduit network343incorporates a valve352in fluid communication with one or more of the engine conduit(s)344. As shown, the valve352is movable from a first configuration preventing the compressed air to flow radially inwardly across the low and high pressure bleed ports348aand348bto a second configuration allowing the compressed air to flow radially inwardly across the low and high pressure bleed ports348aand348band into the core gas path336a. In the second configuration the valve352may prevent the compressed air from reaching the aircraft conduit347. In the first position, the valve352can guide the compressed air drawn from the low and high pressure bleed ports348aand348bto the ECS346via a sequence of the engine conduits344and the aircraft conduit347.

FIG. 4shows the aircraft302equipped with another example of a compressed air delivery system400. As shown, the compressed air delivery system400has dedicated external source conduits449, i.e., separate from the compressed air conduits344described above, to convey compressed air from the APU350to the low and high pressure bleed ports bleed ports348aand348b. As shown, in this example, the first valve352is used to determine whether compressed air originates from the APU350or from the low and high pressure bleed ports348aand348b. However, the compressed air delivery system400has a dedicated, second valve452which can direct compressed air either towards the first valve352, and ultimately the ECS346, or towards the dedicated external source conduits449and associated bleed ports. As shown in this example, the low and high pressure bleed ports348aand348bhave corresponding low pressure and high pressure valves452aand452bthat are controllable to allow or restrict compressed air flow. Each of the valves352,452,452aand452bare controllable independently from one another, depending on the embodiment. The low pressure and high pressure valves452aand452btypically regulate the flow of compressed air supplied to the ECS346during normal working conditions of the gas turbine engine310. In view of the above, the low pressure and high pressure valves452aand452bcan also regulate the flow of compressed air to the core gas path338during shutdown of the gas turbine engine310. The compressed air guided through the low pressure bleed port348acan provide a pressure resistance (back-pressure) to the compressed air guided through the high pressure bleed port348bsuch that its diffusion towards the forward, cold-end of the gas turbine engine is restricted. In some embodiments, the external source conduit449in fluid communication with the low pressure bleed port can be omitted.

As shown in this embodiment, there is provided a controller470communicatively coupled to the ECS346, the APU350, the valves352,452,452aand452b, and/or other aircraft or engine sensor(s). In this embodiment, the controller470has a processor and a computer-readably memory having stored thereon instructions that when executed by the processor can control the operation of the compressed air delivery system400. For instance, upon receiving a command to pre-heat the gas turbine engine310, the controller470may move the second valve452into a configuration guiding compressed air from the APU350in the dedicated external source conduits449and radially inwardly across the low and high pressure ports348aand348b. Once it has been determined that the gas turbine engine310has been satisfactorily pre-heated, the controller470may move the valve452into another configuration favoring compressed air to be bled out of the casing332. In some other embodiments, the controller470may detect that a shutdown sequence of the gas turbine engine310has been initiated, for instance by detecting that the engine rotors rotate at a rotation speed below corresponding idle rotation speeds and/or by detecting that fuel supply has been stopped. In these embodiments, the controller470may the initiate a shutdown of the gas turbine engine310, the controller470may move the second valve452into a configuration guiding compressed air from the APU350in the dedicated external source conduits449and radially inwardly across the low and high pressure ports348aand348band into the core gas path336a.

The compressed air delivery systems described above may be implemented in any type of gas turbine engine. However, it was found useful to implement such systems in gas turbine engines which are more prone to rotor bow.FIGS. 5A and 5Bshow exemplary gas turbine engines having their rotors supported by bearings which are spaced-apart by first and second axial distances d1and d2, respectively. As can be appreciated, the gas turbine engine ofFIG. 5Bmay be more prone to rotor bow than the gas turbine engine ofFIG. 5Aas the length of unsupported rotor is greater, i.e. d2>d1. Accordingly, the benefits of implementing a compressed air delivery system such as the one described herein may be more apparent for the gas turbine engine ofFIG. 5Athan for that ofFIG. 5B, although there are also benefits in implementing the compressed air delivery system to the gas turbine engine ofFIG. 5Ain some circumstances. One aspect of importance in the rotor bow prevention is the spacing of the high pressure shaft bearings and the portion of unsupported shaft that lies within a “hot” zone as depicted by the dashed line inFIGS. 5A and 5B. As typical engines have a “hot bearing,” i.e., a bearing usually underneath the combustion chamber, this shortens the amount of unsupported shaft in the hot zone. Typical engines also use an impeller, versus an all axial high pressure compressor. This shortens the overall length of the high pressure shaft; hence, shortens the unsupported section in the hot zone, and the overall impact of rotor bow. As some gas turbine engines have <40 or 30% of their unsupported length in the hot zone, some other gas turbine engines may have >50% of their unsupported length in the hot zone. Using a compressed air delivery system such as the ones described herein may have significant benefits when used with the latter engines.

FIG. 6shows a flow chart of a method600of operating a gas turbine engine. As discussed, the gas turbine engine has a rotor rotatable about a rotation axis, a core gas path defined annularly around the rotation axis, and a casing defining a radially outer wall to the core gas path.

The method has a step602of generating compressed air from a compressed air source, the compressed air source being external relative the gas turbine engine, and a step604of guiding the compressed air in sequence from the source, radially inwardly relative the rotation axis, through a bleed port fluidly connected to the core gas path.

In some embodiments, the compressed air source is an aircraft-based APU. In these embodiments, the step604of guiding the compressed air can include a step of moving a valve from a first configuration preventing the compressed air to flow through the bleed port to a second configuration allowing the compressed air to flow radially inwardly through the bleed port. In some embodiments, the compressed air source is a ground-based compressed air source. In these embodiments, a step of moving a valve from the first configuration to the second configuration may be performed as well. Additionally, in these embodiments, the step604of guiding the compressed air includes a step of bringing an external compressed air conduit in fluid communication between the compressed air source and the bleed port. In this case, the compressed air conduit may be an external compressed air conduit having end connectors connectable to corresponding connectors of the aircraft and/or gas turbine engine. Using both the aircraft-based APU and the ground-based compressed air source may be envisaged in some embodiments. In some embodiments, the aircraft-based APU can be substituted by an auxiliary air-pump that is powered by alternate means.

In embodiments where the gas turbine is already hot, the step602of guiding the compressed air includes a step of cooling the core gas path, the rotor, the radially outer wall of the casing, and other engine components. In embodiments where the gas turbine is cold, the step602of guiding the compressed air includes a step of warming the core gas path, the rotor, the radially outer wall of the casing, and other engine components. In either case, the compressed air guided through the bleed port can be used for forced-convective cooling and/or warming of the gas turbine engine using the compressed air from a non-engine related compressed air source such as the APU or a stand-alone blower motor.

In some embodiments, the step604of guiding the compressed air induces rotation of the rotor up to a given rotation speed. In these embodiments, the rotation of the rotor may not be induced by rotation of the air starter but rather by compressed air acting on blades (e.g., aerofoils) of the rotor. The rotation speed is generally kept well below an idle rotation speed of the corresponding rotor. For example, the idle rotation speed for a high pressure rotor can be 17,000 RPM whereas the idle rotation speed for a low pressure rotor can be 1,500 RPM. As such, the compressed air guided through the bleed port can cause the high pressure rotor to rotate at a speed of about 300-400 RPM or even lower in some embodiments. The compressed air can be used to rotate the low pressure and/or high pressure rotors at 2% of the corresponding idle rotation speed, and preferably 10% of the corresponding idle rotation speed of the rotation speed. The flow of compressed air radially inwardly through the bleed port(s) may be maintained and/or increased for a given period of time up until the rotation speed of the rotor reaches a given rotation speed threshold. In these embodiments, either one or both of the given period of time and the given rotation speed threshold can be dependent upon surrounding environmental conditions.

When the gas turbine engine has been left inoperative for several hours in cold climate, it can be difficult to start. In such conditions, it may be desirable to pre-heat the engine, and more specifically the components surrounding the combustor including the combustor itself, prior to ignition. In such embodiments, the method600, and more specifically the step604, can be performed until the core gas path and/or the rotor(s) have reached a given temperature. When it is determined that the gas turbine engine has been satisfactorily pre-heated, a step606of pre-heating the gas turbine engine until reaching suitable metal temperature. Then, in some embodiments, the method can include a step of igniting the gas turbine engine. In some embodiments, the flow of compressed air radially inwardly through the bleed port(s) and into the core gas path is maintained throughout pre-heating and ignition and perhaps until a given period of time has elapsed. In some other embodiments, the flow of compressed air radially inwardly through the bleed port(s) and into the core gas path is stopped immediately prior to the step606of pre-heating the gas turbine engine.

In some embodiments, especially those in which the gas turbine engine has been running for a while and in which shutting down of the gas turbine engine is sought, the step604can be performed in order to cool the core gas path to prevent rotor bow such as introduced above. In these embodiments, the step604may be performed only when some conditions indicating of the shutting down of the gas turbine engine have been met. For instance, the method600has a step608of reducing a rotation speed of the engine's rotor is below idle rotation speed. For example, the idle rotation speed for a high pressure rotor can be 17,000 RPM whereas the idle rotation speed for a low pressure rotor can be 1,500 RPM. When that such speed reduction occurs, step604may be performed. In some embodiments, the method600has a step610of stopping fuel supply to the gas turbine engine. In these embodiments, when the fuel supply is stopped, step604is performed thereafter. In some embodiments, step604may be performed only upon determining that both the steps608and610have been performed.

In some embodiments, the method600can pre-heat the gas turbine engine prior to ignition in order to assist in cold-starts. When using a ground-based compressed air source, the fuel consumption may also be reduced at the aircraft level. The method600can allow the gas turbine engine to be completely shut-down sooner after landing thanks to the cooling thereby reducing noise pollution. Moreover, using the APU for engine cooling can help reduce the wait times from arrival at the gate until the deplaning.

The embodiments described in this document provide non-limiting examples of possible implementations of the present technology. Upon review of the present disclosure, a person of ordinary skill in the art will recognize that changes may be made to the embodiments described herein without departing from the scope of the present technology. For example, the valve between the external compressed air source and the bleed port(s) can be omitted. External to the gas turbine engine can refer to external to the casing, or external to the nacelle, for instance. It is intended that the method described herein can be used to assist in cold-starts by pre-heating the engine hardware to avoid frame-quenching. In some embodiments, compressed air may be flowed radially inwardly across bleed port(s) open in the bypass gas path to pre-heat and/or cool the core engine casing and/or the nacelle, in some embodiments. Yet further modifications could be implemented by a person of ordinary skill in the art in view of the present disclosure, which modifications would be within the scope of the present technology.