Patent Description:
Aircraft are often provided with two engines, typically gas turbine engines. In the case of helicopters, typically two turboshaft engines are connected to a main rotor via a common reduction gearbox, and each of the engines may be sized to such that the power of each engine is greater than what is required for cruising.

Bleed air produced by a gas turbine engine of a multi-engine aircraft is compressed air from a compressor stage and is used for various functions of that engine (such as cooling of turbines and to help seal bearing cavities, for example). Bleed air may also be used for aircraft functions (such as engine starting, cabin pressure, air systems, pressurizing liquid tanks, etc.). Engine bleed air can be derived from the high pressure or the low pressure compressor stage, depending on the air pressure requirements and the engine operating condition. <CIT>, <CIT>, <CIT>, <CIT> and <CIT> disclose air systems of the prior art.

In an aspect, there is provided a method of operating an aircraft according to claim <NUM>.

In another aspect, there is provided an aircraft according to claim <NUM>.

Optional embodiments are defined by the dependent claims.

In at least some multi-engine aircraft, such as helicopters, prior art bleed systems may not be capable of supplying an adequate flowrate and/or pressure of bleed air in some operating conditions, such as when a gas turbine engine providing the bleed air is operating in a low power, or standby mode, and not being used to provide substantive motive power to the aircraft.

For the purposes of this document, the term "active" used with respect to a given engine means that the given engine is providing motive power to the aircraft. For the purposes of this document, the terms "standby" and "sub-idle" are used with respect to a given engine to mean that the given engine is operating but is providing no motive power, or at least substantially no motive power, to the aircraft, with "sub-idle" operation being a particular type of standby operation according to the present technology as described in this document. It is however understood that when operating in a "standby" mode, as used herein, the engine provides little to no motive power to the aircraft, when the standby engine is running at, below, or above, idle speed.

For the purposes of the present description, the term "conduit" with respect to a fluid is used to describe an arrangement of one or more elements, such as one or more conventional hoses, connectors, filters, pumps and the like, as may be suitable for the described functionality of the conduit, and which together form the flow path(s) to provide the functionality described with respect to the conduit. For example, a given air conduit may be defined by any number and combination of air lines, filters, control actuators, and the like, selected to provide the particular functionality described with respect to the given air conduit. As another example, a given fuel conduit may be defined by any number and combination of hoses hydraulically interconnected in parallel and/or series, by or with one or more fuel filters, switches, pumps, and the like, selected to provide the particular functionality described with respect to the given fuel conduit.

<FIG> illustrates an aircraft engine <NUM> of a type preferably provided for use in subsonic flight, generally comprising a shaft <NUM> operatively connectable to a fan or other rotor, such as a helicopter rotor, and, in serial flow communication, a compressor section <NUM> for pressurizing ambient air, a combustor <NUM> in which the compressed air is mixed with fuel and ignited for generating an annular stream of hot combustion gases, and a turbine section <NUM> for extracting energy from the combustion gases. Components of the engine <NUM> are rotatable about a longitudinal center axis <NUM> of the engine <NUM>. In the present embodiment, the engine <NUM> is a turboshaft engine. It is contemplated that the engine <NUM> could be a different type of engine, such as a rotary engine, a turboprop, or a turbofan engine for example.

<FIG> schematically illustrates an aircraft <NUM>, in this example a helicopter, having a first engine <NUM>', and a second engine <NUM>". The engines <NUM>', <NUM>" are operable to provide motive power to the aircraft <NUM> via conventional transmission systems and controls. For simplicity, only the nonconventional aspects of the present technology are described in detail in this document. In this embodiment, each of the engines <NUM>', <NUM>" is substantially the same as engine <NUM> shown in <FIG> and described above. Therefore, only the first engine <NUM>' is described in further detail. Parts of the second engine <NUM>" that correspond to parts of the first engine <NUM>' are labeled with the same numerals.

The illustrated exemplary multi-engine system may be used as a power plant for the aircraft <NUM>, including but not limited to a rotorcraft such as a helicopter. The multi-engine system may include the two or more gas turbine engines <NUM>', <NUM>". In the case of the aircraft <NUM> being a helicopter, these gas turbine engines <NUM>', <NUM>" will be turboshaft engines. Control of the multi-engine system shown in <FIG> is effected by one or more controller(s) <NUM>' (shown in <FIG> only, to maintain clarity of the figures), which may be FADEC(s), electronic engine controller(s) (EEC(s)), or the like, that are programmed to manage, as described herein below, the operation of the engines <NUM>', <NUM>". In some embodiments and operating conditions, control sequences as described in the present application may reduce an overall fuel burn of the aircraft <NUM>, particularly during sustained cruise operating regimes, wherein the aircraft <NUM> is operated at a sustained (steady-state) cruising speed and altitude. The cruise operating regime is typically associated with the operation of prior art engines at equivalent part-power, such that each engine contributes approximately equally to the output power of the multi-engine system. Other phases of a typical helicopter mission would include transient phases like take-off, climb, stationary flight (hovering), approach and landing. Cruise may occur at higher altitudes and higher speeds, or at lower altitudes and speeds, such as during a search phase of a search-and-rescue mission.

In the present description, while the aircraft <NUM> conditions (cruise speed and altitude) are substantially stable, the engines <NUM>', <NUM>" of the multi-engine system may be operated asymmetrically, with one engine operated in a high-power "active" mode and the other engine operated in a lower-power "standby" mode. Doing so may provide fuel saving opportunities to the aircraft, however there may be other suitable reasons why the engines are desired to be operated asymmetrically. This operation management may therefore be referred to as an "asymmetric mode" or an "asymmetric operating regime", wherein one of the two engines is operated in a low-power "standby mode" while the other engine is operated in a high-power "active" mode. In such an asymmetric operation, which may be engaged during a cruise phase of flight (continuous, steady-state flight which is typically at a given commanded constant aircraft cruising speed and altitude). The multi-engine system may be used in an aircraft, such as a helicopter, but also has applications in suitable marine and/or industrial applications or other ground operations.

Referring still to <FIG>, according to the present description the multi-engine system driving a helicopter <NUM> may be operated in this asymmetric manner, in which one of the engines <NUM>', <NUM>" may be operated at high power in an active mode and another one of the engines <NUM>', <NUM>" may be operated in a low-power standby mode. In one example, the active engine may be controlled by the controller(s) <NUM>' to run at full (or near-full) power conditions in the active mode, to supply substantially all or all of a required power and/or speed demand of the aircraft <NUM>. The standby engine may be controlled by the controller(s) <NUM>' to operate at low-power or no-output-power conditions to supply substantially none or none of a required power and/or speed demand of the aircraft <NUM>. Optionally, a clutch may be provided to declutch the low-power engine. Controller(s) <NUM>' may control the engine's governing on power according to an appropriate schedule or control regime, for example as described in this document. The controller(s) <NUM>' may be one or multiple suitable controllers, such as for example a first controller for controlling the engine <NUM>' and a second controller for controlling the second engine <NUM>". The first controller and the second controller may be in communication with each other in order to implement the operations described herein. In some embodiments, and a single controller <NUM>' may be used for controlling the first engine <NUM>' and the second engine <NUM>". To this end, the term controller as used herein includes any one of: a single controller controlling the engines <NUM>', <NUM>", and multiple controllers controlling the engines <NUM>', <NUM>".

In another example, an asymmetric operating regime of the engines may be achieved through the one or more controller's differential control of fuel flow to the engines, as described in pending application <CIT>. Low fuel flow may also include zero fuel flow in some examples and/or times.

Although various differential control between the engines of the multi-engine engine system are possible and some such sequences are described in this document, in one particular embodiment the controller(s) <NUM>' may correspondingly control fuel flow rate to each engine <NUM>', <NUM>" as follows. In the case of the standby engine, a fuel flow (and/or a fuel flow rate) provided to the standby engine may be controlled to be between <NUM>% and <NUM>% less than the fuel flow (and/or the fuel flow rate) provided to the active engine. In the asymmetric mode, the standby engine may be maintained between <NUM>% and <NUM>% less than the fuel flow to the active engine. In some embodiments of the method <NUM>, the fuel flow rate difference between the active and standby engines may be controlled to be in a range of <NUM>% and <NUM>% of each other, with fuel flow to the standby engine being <NUM>% to <NUM>% less than the active engine. In some embodiments, the fuel flow rate difference may be controlled to be in a range of <NUM>% and <NUM>%, with fuel flow to the standby engine being <NUM>% to <NUM>% less than the active engine.

In another embodiment, the controller <NUM> may operate one engine of the multiengine system in a standby mode at a power substantially lower than a rated cruise power level of the engine, and in some embodiments at zero output power and in other embodiments less than <NUM>% output power relative to a reference power (provided at a reference fuel flow). Alternately still, in some embodiments, the controller(s) <NUM>' may control the standby engine to operate at a power in a range of <NUM>% to <NUM>% of a rated full-power of the standby engine (i.e. the power output of the second engine to the common gearbox remains between <NUM>% to <NUM>% of a rated full-power of the second engine when the second engine is operating in the standby mode).

In another example, the engine system of <FIG> may be operated in an asymmetric operating regime by control of the relative speed of the engines using controller(s) <NUM>', that is, the standby engine is controlled to a target low speed and the active engine is controlled to a target high speed. Such a low speed operation of the standby engine may include, for example, a rotational speed that is less than a typical ground idle speed of the engine (i.e. a "sub-idle" engine speed). Still other control regimes may be available for operating the engines in the asymmetric operating regime, such as control based on a target pressure ratio, or other suitable control parameters.

Although the examples described herein illustrate two engines, asymmetric mode is applicable to more than two engines, whereby at least one of the multiple engines is operated in a low-power standby mode while the remaining engines are operated in the active mode to supply all or substantially all of a required power and/or speed demand of a common load.

In use, the one of the engines <NUM>', <NUM>" may operate in the active mode while the other of the engines <NUM>', <NUM>" may operate in the standby mode, as described above. During this asymmetric operation, if the aircraft <NUM> needs a power increase (expected or otherwise), the active engine(s) may be required to provide more power relative to the low power conditions of the standby mode, and possibly return immediately to a high- or full-power condition. This may occur, for example, in an emergency condition of the multi-engine system powering the helicopter, wherein the "active" engine loses power the power recovery from the lower power to the high power may take some time. Even absent an emergency, it will be desirable to repower the standby engine to exit the asymmetric mode. The controller(s) <NUM>' may also be used to operate the various air valves described herein. To this end, any suitable operative connections and configurations of controls may be provided, so long as the functionality described herein is provided. Because such operative connections may be conventional, to maintain clarity, only one such operative connection is shown schematically in <FIG>. The rest of the operative connections may be similar, and hence are not shown.

As shown schematically in <FIG>, the first engine <NUM>' includes a bleed air system <NUM> that includes air conduits <NUM>, <NUM>, <NUM>, <NUM>, <NUM> and valves <NUM>', <NUM>', <NUM>", as will be seen. A first bleed air conduit <NUM> and a second bleed air conduit <NUM> are provided, both of which bleed compressed air from respective parts of the compressor section <NUM> of the first engine <NUM>'. While in this application the air sources are P2. <NUM> and P2. <NUM>, respectively, in other embodiments other locations in the compressor section <NUM> and/or other locations fluidly connected to the compressor section <NUM> may be used, for example to suit each particular embodiment and application of the engine(s) <NUM>', <NUM>". In the present embodiment, the first bleed air conduit <NUM> includes a check valve <NUM>' and branches off into supply bleed air conduits <NUM> downstream of the check valve. In this embodiment, the second bleed air conduit <NUM> includes a check valve <NUM>' and a check valve <NUM>". The second bleed air conduit <NUM> branches off into supply bleed air conduits <NUM> at one or more locations that are fluidly in between the check valves <NUM>', <NUM>". As shown, the check valves <NUM>', <NUM>" are pointing toward each other, for purposes explained below.

The supply bleed air conduits <NUM>, <NUM> deliver bleed air to various sealing and lubrication systems of the first engine <NUM>' and/or to various locations for various aircraft functions. The particular airflow destinations may be selected to suit and/or may depend on the particular embodiment and application of the engine(s) <NUM>', <NUM>". The particular number and configuration of the sealing systems may be any suitable number and configuration, and is therefore not described in detail. The supply bleed air conduits <NUM> and <NUM> may also provide bleed air for various other functions of the first engine <NUM>' and/or the aircraft. Examples of such functions include, but are not limited to, cooling of turbines, maintenance of cabin pressure, operation of air systems, and pressurizing liquid tanks. Any suitable air piping and controls arrangement may be used to provide for each particular combination of the functions provided for by the bleed air from the first and second bleed air conduits <NUM>, <NUM>.

Still referring to <FIG>, the first and second bleed air conduits <NUM>, <NUM> of the first engine <NUM>' fluidly converge / join into a common bleed air conduit <NUM>. The common bleed air conduit <NUM> fluidly connects to a control valve <NUM>. The control valve <NUM> may be any suitable one or more control valves so long as it provides for the functionality described in this document. The conduits <NUM>, <NUM>, <NUM>, <NUM>, <NUM> and valves <NUM>', <NUM>', <NUM>" of the first engine <NUM>' are part of the bleed air system <NUM> of the first engine <NUM>'. The rest of the bleed air system <NUM> may be conventional and is therefore not shown or described in detail herein.

As shown in <FIG>, in the present embodiment, the bleed air system <NUM> of the second engine <NUM>" is similar to the bleed air system <NUM> of the first engine <NUM>', described above. Therefore, to maintain simplicity of this description, the bleed air system <NUM> of the second engine <NUM>" is not described in detail. Suffice it to say that parts of the bleed air system <NUM> of the second engine <NUM>" that correspond to parts of the bleed air system <NUM> of the first engine <NUM>' are labeled with the same numerals.

As shown in <FIG>, the common bleed air conduit <NUM> of the second engine <NUM>", similar to the common bleed air conduit <NUM> of the first engine <NUM>', fluidly connects to a control valve <NUM>. The control valve <NUM> is operable by a controller of the aircraft <NUM> to selectively: i) fluidly connect the common bleed air conduit <NUM> of the first engine <NUM>' to the common bleed air conduit <NUM> of the second engine <NUM>", and ii) fluidly disconnect the common bleed air conduit <NUM> of the first engine <NUM>' from the common bleed air conduit <NUM> of the second engine <NUM>", as illustrated by the internal structure of the control valve <NUM> schematically shown in <FIG>. The control valve <NUM> may be actuated using any suitable actuator of the engines <NUM>', <NUM>" and/or of the aircraft <NUM>.

<FIG> shows a first in-flight, powered, mode of operation of the aircraft <NUM> during which both engines <NUM>', <NUM>" are "active" (a. operating in an active mode), and are therefore both providing motive power to the aircraft <NUM>. In this operating condition, the bleed air system <NUM> of the first engine <NUM>' and the bleed air system <NUM> of the second engine <NUM>" are both self-sufficient. For the purposes of this document, the term "self-sufficient" used with respect to a given bleed air system of a given engine means that the given bleed air system of the given engine provides all of its intended functions for the duration of the time during which it is called upon to provide for the functions. A given bleed air system of a given engine is not "self-sufficient" when one or more of the intended functions of the given bleed air system may be unavailable or unstable due to a lack of bleed air pressure and/or bleed air supply rate provided by the corresponding engine to the given bleed air system.

Reference is now made to <FIG>, which shows a second in-flight, powered, mode of operation of the aircraft <NUM> during which: i) the first engine <NUM>' is "active" and is therefore providing motive power to the aircraft <NUM>, and ii) the second engine <NUM>" is on "standby" (a. operating in a standby mode) and is therefore not providing any material amount of motive power to the aircraft <NUM>. In this operating condition (i.e. in the second in-flight mode of operation), the bleed air system <NUM> of the first engine <NUM>' is self-sufficient. On the other hand, depending on each particular embodiment of the engines <NUM>', <NUM>" and/or the aircraft <NUM> and/or on the characteristics of the particular "standby" operation of the second engine <NUM>", the bleed air system <NUM> of the second engine <NUM>" may or may not be self-sufficient in the standby mode.

For this reason, during the second in-flight mode of operation of the aircraft <NUM>, the control valve <NUM> may be actuated by a suitable controller of the aircraft <NUM> to fluidly connect the common bleed air conduit <NUM> of the first engine <NUM>' to the common bleed air conduit <NUM> of the second engine <NUM>", to provide for an additional supply of bleed air from the bleed air system <NUM> of the first engine <NUM>' to the bleed air system <NUM> of the second engine <NUM>". Self-sufficiency of the bleed air system <NUM> of the second engine <NUM>" may thereby be provided. After the second engine <NUM>" is brought into an "active" state while the first engine <NUM>' is in an "active" state, the control valve <NUM> may be actuated by a suitable controller of the aircraft <NUM> to fluidly disconnect the common bleed air conduit <NUM> of the first engine <NUM>' from the common bleed air conduit <NUM> of the second engine <NUM>", as shown in <FIG>. After the first engine <NUM>' is put into a standby mode or a sub-idle mode while the second engine <NUM>" is in an "active" mode, the control valve <NUM> may be actuated by a suitable controller of the aircraft <NUM> to fluidly connect the common bleed air conduit <NUM> of the first engine <NUM>' to the common bleed air conduit <NUM> of the second engine <NUM>". The bleed air system <NUM> of the second engine <NUM>" may thereby provide compressed air to the bleed air system <NUM> of the first engine <NUM>'. Self-sufficiency of the bleed air system <NUM> of the first engine <NUM>' may thereby be provided.

The bleed air systems <NUM>, <NUM> of the engines <NUM>', <NUM>" and the control valve <NUM> are part of an air system <NUM> of the aircraft <NUM>. As described above, the air system <NUM> of the aircraft <NUM> implemented according to the present technology may thereby provide for self-sufficient operation of at least one of the engines <NUM>', <NUM>" in at least some operating conditions of the aircraft <NUM> in which at least some prior art engine bleed air systems may not be self-sufficient.

Further, as shown in <FIG> and <FIG> for example, in the present embodiment, the check valves <NUM>' and <NUM>" are provided in the bleed air conduits <NUM>, downstream of the branching-out bleed air conduits <NUM>. In this embodiment, this the branching-out bleed air conduits <NUM> may supply compressed air to at least some subsystems of the respective engines <NUM>', <NUM>". Each of the check valves <NUM>' and <NUM>" ensures that when the engine <NUM>', <NUM>" having that check valve <NUM>', <NUM>" is providing compressed air from its bleed air system <NUM>, <NUM> to the bleed air system <NUM>, <NUM> of the other engine <NUM>', <NUM>", the compressed air is provided from the air source corresponding to the bleed air conduit <NUM> of that engine <NUM>', <NUM>". The check valves <NUM>' and <NUM>" therefore help ensure uncompromised self-sufficient operation of the subsystems of the one of the engines <NUM>', <NUM>" that may at a given time be providing compressed air to the other one of the engines <NUM>', <NUM>". In some embodiments, the check valve <NUM>' and/or the check valve <NUM>" may be omitted.

The rest of the air system <NUM> that is not shown in the figures of the present application may be conventional and is therefore not described in detail herein. Any suitable controls and any suitable control logic may be used to provide for the functionality of the air system <NUM>, and/or for various timings of actuation of the control valve <NUM> relative to switches between "active" and "standby" states that may occur with respect to each of the engines <NUM>', <NUM>" during in-flight or ground operations of the aircraft <NUM>.

Referring now to <FIG>, an air system <NUM> of the aircraft <NUM>, which is an alternative embodiment of the air system <NUM> is shown. The air system <NUM> is similar to the air system <NUM>, and therefore similar reference numerals have been used for the air system <NUM>. A difference of the air system <NUM> from the air system <NUM>, is that air system <NUM> includes a control valve <NUM>, a control valve <NUM>, and an external compressed air source <NUM> such as an auxiliary power unit (APU) and/or an air compressor for example. The external compressed air source <NUM> may be any conventional external compressed air source suitable for each particular embodiment of the engines <NUM>', <NUM>" and the aircraft <NUM>.

The control valve <NUM> selectively fluidly connects the external compressed air source <NUM> to the common bleed air conduit <NUM> of the first engine <NUM>', via any suitable corresponding air conduits. More particularly, when the first engine <NUM>' is "active", the control valve <NUM> may be actuated by a suitable controller of the aircraft <NUM> to fluidly disconnect the external compressed air source <NUM> from the common bleed air conduit <NUM> of the first engine <NUM>', and may thereby allow the bleed air system <NUM> of the first engine <NUM>' to run self-sufficiently.

When the first engine <NUM>' is on "standby", the control valve <NUM> may be actuated by a suitable controller of the aircraft <NUM> to fluidly connect the external compressed air source <NUM> to the common bleed air conduit <NUM> of the first engine <NUM>'. The control valve <NUM> may thereby provide "supplemental" compressed air to the bleed air system <NUM> of the first engine <NUM>' at a supply rate and pressure sufficient to allow the bleed air system <NUM> of the first engine <NUM>' to provide for all of its intended functions during the "standby" operation of the first engine <NUM>'. The control valve <NUM>, via corresponding air conduit(s), may selectively fluidly connect the external compressed air source <NUM> to a different part of the bleed air system <NUM> of the first engine <NUM>', so long as the functionality described above is provided.

The control valve <NUM> similarly fluidly connects the external compressed air source <NUM> to the common bleed air conduit <NUM> of the second engine <NUM>", and is actuated according to a similar control logic to allow the bleed air system <NUM> of the second engine <NUM>" to provide for all of its intended functions during the "standby" operation of the second engine <NUM>". As shown, the control valve <NUM> that fluidly connects the bleed air system <NUM> of the first engine <NUM>' to the bleed air system <NUM> of the second engine <NUM>" may be in a position in which it fluidly disconnects the first engine <NUM>' from the second engine <NUM>", to allow for the supplemental compressed air to be provided to either one, or to both, of the engines <NUM>', <NUM>" by the external compressed air source <NUM>. In some embodiments, the control valves <NUM>, <NUM>, <NUM> may be actuated correspondingly to switch between the various possible supply modes of air described above. For example, in some operating conditions, the bleed air system <NUM>, <NUM> of one of the engines <NUM>', <NUM>" may receive "supplemental" compressed air from one or both of: i) the bleed air system <NUM>, <NUM> of another one of the engines <NUM>', <NUM>", and ii) the external compressed air source <NUM>.

Referring now to <FIG>, an air system <NUM> of the aircraft <NUM>, which is yet another alternative embodiment of the air system <NUM> is shown. The air system <NUM> is similar to the air system <NUM>, and therefore similar reference numerals have been used for the air system <NUM>. A difference of the air system <NUM> from the air system <NUM>, is that air system <NUM> does not have a control valve <NUM> for fluidly connecting the bleed air system <NUM> of the first engine <NUM>' to the bleed air system <NUM> of the second engine <NUM>". Operation of the air system <NUM> is similar to operation of the air system <NUM> with respect to the external compressed air source <NUM>.

In at least some embodiments and applications, the air systems <NUM>, <NUM>, <NUM> may allow to provide "supplemental" compressed air to the bleed air system <NUM>, <NUM> of one of the engines <NUM>', <NUM>" in at least some cases where that bleed air system <NUM>, <NUM> is malfunctioning and/or leaking air for example. A person skilled in the art will appreciate that while a particular air conduit arrangement is shown in <FIG>, other air conduit arrangements may be used while providing for at least some of the functionality described in this document. While a single external compressed air source <NUM> is used in the embodiments of <FIG> and <FIG>, multiple different external compressed air sources may be used. Likewise, while the example aircraft <NUM> has two engines <NUM>', <NUM>", the present technology may be implemented with respect to more than two engines and/or with respect to other types of engines.

With the above systems in mind, the present technology provides a method <NUM> of using, in flight, a source of pressurized air external to an engine of an aircraft <NUM>. As seen above, in some embodiments and operating conditions, the source of pressurized air may be one of the engines <NUM>', <NUM>" of the aircraft <NUM>, and in some embodiments, an APU <NUM> or air compressor <NUM> of the aircraft <NUM>. In some embodiments, the method <NUM> includes a step <NUM> of operating a given engine <NUM>', <NUM>" of the aircraft <NUM> during flight. In some embodiments, the method <NUM> also includes a step <NUM> of directing pressurized air from the source of pressurized air external to the given engine <NUM>', <NUM>", to a bleed air system <NUM>, <NUM> of the given engine <NUM>', <NUM>".

In some embodiments, the given engine <NUM>', <NUM>" to which pressurized air is directed is a first engine <NUM>' of the aircraft <NUM>, the aircraft <NUM> includes a second engine <NUM>", and the source of pressurized air external to the first engine <NUM>' is a bleed air system <NUM> of the second engine <NUM>". As seen above, in some embodiments, the aircraft <NUM> is a multi-engine helicopter in which the engines <NUM>', <NUM>" are operatively connected to drive at least one main rotor of the helicopter to provide motive power to / propel the helicopter.

As seen above, in some embodiments, the directing pressurized air to the bleed air system <NUM> of the first engine <NUM>' is executed when the first engine <NUM>' is operating in a standby mode. In embodiments in which the source of the pressurized air is the bleed air system <NUM> of the second engine <NUM>", the second engine <NUM>" is active (i.e. providing motive power to the helicopter). Similarly, in some operating conditions during flight, the given engine <NUM>', <NUM>" to which pressurized air is directed is a second engine <NUM>" of the aircraft <NUM>. In some such cases, the source of pressurized air external to the second engine <NUM>" is a bleed air system <NUM> of the first engine <NUM>'. In some such cases, the second engine <NUM>" is in a standby mode while the first engine <NUM>' providing the compressed air is active (i.e. providing motive power to the helicopter).

As seen above, in some embodiments, the source of pressurized air is a first source of pressurized air (e.g. first engine <NUM>' or second engine <NUM>", depending on which of these engines is active and which is in a standby mode), the aircraft <NUM> includes a second source of pressurized air (e.g. APU / air compressor <NUM> of the aircraft <NUM>). In some such embodiments, the second source of pressurized air <NUM> is external to both the first engine <NUM>' and the second engine <NUM>". In some such embodiments and in some flight conditions, the method <NUM> comprises directing pressurized air from the second source of pressurized air <NUM> to the first engine <NUM>'. In some such embodiments and in some flight conditions, the method <NUM> comprises directing pressurized air from the second source of pressurized air <NUM> to the second engine <NUM>". Further in some such embodiments and in some flight conditions, the method <NUM> comprises directing pressurized air from the second source of pressurized air <NUM> to both the first engine <NUM>' and the second engine <NUM>".

Further with the structure of the aircraft <NUM> described above, the present technology also provides method <NUM> of operating a bleed air system <NUM> of a first gas turbine engine <NUM>' of a multi-engine aircraft <NUM> during flight. In some embodiments, the method <NUM> comprises a step <NUM> of operating the first gas turbine engine <NUM>' of the aircraft <NUM> during flight in a standby mode, such as an idle or a sub-idle mode that provides either no motive power or at least materially no motive power to the aircraft <NUM>. In some embodiments, the method <NUM> comprises a step <NUM> of operating a second gas turbine engine <NUM>" of the aircraft <NUM> during flight in an active mode (i.e. providing non-substantially-zero motive power to the aircraft <NUM>).

In some cases, the steps <NUM> and <NUM> are executed simultaneously. In some such cases, the method <NUM> comprises directing pressurized air from a bleed air system <NUM> of the second gas turbine engine <NUM>" to a bleed air system <NUM> of the first gas turbine engine <NUM>'.

In some cases, the method <NUM> further includes a step <NUM> of operating a source of pressurized air (E. APU / air compressor <NUM>, and the like) of the aircraft <NUM> external to both the first gas turbine engine <NUM>' and the second gas turbine engine <NUM>", and a step of directing pressurized air from the source of pressurized air <NUM> to at least one of the first gas turbine engine <NUM>' and the second gas turbine engine <NUM>".

In some cases, the directing pressurized air from one of the bleed air systems <NUM>, <NUM> to the other one of the bleed air systems <NUM>, <NUM> (depending on which one of the bleed air systems <NUM>, <NUM> requires supplemental compressed air) may be executed simultaneously with directing pressurized air from a second source of pressurized air to the one of the bleed air systems <NUM>, <NUM> that is receiving the supplemental compressed air. In some embodiments, the second source of pressurized air <NUM> includes, or is, at least one of: an APU <NUM> of the aircraft <NUM>, and an air compressor <NUM> of the aircraft <NUM>.

In some such cases, the air pressure in the one of the bleed air systems <NUM>, <NUM> receiving supplemental compressed air may be lower than the pressure of the supplemental compressed air. It is contemplated that any suitable controls and control arrangements may be used to provide for this pressure differential, where required. While two engines <NUM>', <NUM>" of an aircraft <NUM> are described, it is contemplated that the present technology could be implemented with regard to a larger number of engines of an aircraft to provide supplemental compressed air from one or more of the engines or other compressed air source(s), to one or more other ones of the engines.

Claim 1:
A method (<NUM>) of operating (<NUM>) an aircraft (<NUM>) having a first engine (<NUM>') and a second engine (<NUM>"), the first engine (<NUM>, <NUM>') having a bleed air system (<NUM>) and the second engine (<NUM>") having a bleed air system (<NUM>), the method characterised by
in flight, and with the first engine (<NUM>, <NUM>') operating in a standby mode and the second engine (<NUM>") operating in an active mode, directing (<NUM>) pressurized air from a source of pressurized air external to the first engine (<NUM>, <NUM>') to the bleed air system (<NUM>) of the first engine (<NUM>, <NUM>'), the source of pressurized air including the bleed air system (<NUM>) of the second engine (<NUM>"),
in flight, and with the first engine (<NUM>, <NUM>') operating in the standby mode, operating the second engine in the active mode to supply all or substantially all of a required power and/or speed demand of a common load.