Patent Description:
Prior art fuel delivery systems for aircraft may be suitable for their intended purposes. However, improvements in aviation are always desirable.

<CIT> discloses a prior art fuel delivery system according to the preamble of claim <NUM>. <CIT> and <CIT> disclose other prior art systems.

In one aspect, there is provided a fuel delivery system for an aircraft engine according to claim <NUM>.

The fuel delivery system as described herein may also include one or more of the following features.

In some embodiments, the cross-flow fuel conduit fluidly connects to the first fuel manifold at a location in a fuel nozzle of the first fuel manifold.

In some embodiments, the cross-flow fuel conduit is defined in part by a flow modulating device operable to modulate flow through the cross-flow fuel conduit.

In some embodiments, the fuel delivery system further comprises a fuel control valve upstream of the first and second fuel manifolds, the fuel control valve being operable to supply fuel from a fuel tank to: i) both the first and second fuel manifolds, ii) the first fuel manifold while blocking fuel supply to the second fuel manifold, and iii) the second fuel manifold while blocking fuel supply to the first fuel manifold; and iv) to supply a trickle flow of fuel from a fuel conduit upstream of the first fuel manifold to a fuel conduit upstream of the second fuel manifold while blocking fuel supply to the second fuel manifold.

In some such embodiments, the valve in the second fuel manifold is a check valve oriented to allow fuel flow into the combustor out of the fuel nozzle of the second fuel manifold.

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

The aircraft gas turbine engine as described herein may also include one or more of the following features.

In some such embodiments, the cross-flow fuel conduit is a plurality of cross-flow fuel conduits, and each fuel nozzle of the first plurality of fuel nozzles is fluidly connected to one fuel nozzle of the second plurality of fuel nozzles via a cross-flow fuel conduit of the plurality of cross-flow fuel conduits at least when the gas turbine engine operates in a standby mode.

In some suck embodiments, each fuel nozzle of the second plurality of fuel nozzles comprises a valve operable to block fuel flow out of that fuel nozzle, and each cross-flow fuel conduit of the plurality of cross-flow fuel conduits that fluidly connects to that fuel nozzle, fluidly connects to that fuel nozzle at a location downstream of the valve of that fuel nozzle.

In some such embodiments, the aircraft gas turbine engine comprises at least one flow modulating device operatively connected to at least one of the plurality of cross-flow fuel conduits to modulate flow through the at least one of the plurality of cross-flow fuel conduits.

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

The method as described herein may also include one or more of the following features.

In some embodiments, the method further comprises modulating the trickle flow.

In some embodiments, the method further comprises providing a trickle flow of fuel into the combustor out of the second fuel manifold.

In some embodiments, the providing the trickle flow includes bypassing at least one closed valve in at least one fuel nozzle of the second fuel manifold.

In some embodiments, the providing the trickle flow includes providing a trickle flow of fuel from a fuel conduit upstream of the first fuel manifold to a fuel conduit upstream of the second fuel manifold.

In some embodiments, the providing the trickle flow includes controlling a flow control valve fluidly connected to both the fuel conduit upstream of the first fuel manifold and the fuel conduit upstream of the second fuel manifold.

In some embodiments, the maintaining combustion is part of operating the aircraft engine in a standby mode, and the method further includes switching the aircraft engine from the standby mode to an active mode, the switching including unblocking fuel flow out of the second fuel manifold.

In some embodiments, the unblocking fuel flow includes opening at least one valve in at least one fuel nozzle of the second fuel manifold.

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 example of a gas turbine engine <NUM>. In this example, the gas turbine <NUM> is a turboshaft engine <NUM> comprising in serial flow communication a low pressure (LP) compressor section <NUM> and a high pressure (HP) compressor section <NUM> for pressurizing air, a combustor <NUM> in which the compressed air is mixed with fuel and ignited for generating an annular stream of hot combustion gases, a high pressure (HP) turbine section <NUM>, and a lower pressure (LP) turbine section <NUM>. The respective pairs of the compressor and turbine sections <NUM>, <NUM>, <NUM>, <NUM> are interconnected via respective independently rotatable low pressure (LP) and high pressure (HP) spools, or shafts. This arrangement enables, inter alia, the core flow and processing of air through the engine <NUM>, which is received through an air inlet <NUM> and exhausted via an exhaust outlet <NUM> of the turboshaft engine <NUM>.

In the present embodiment, the turboshaft engine <NUM> further includes a set of variable guide vanes <NUM>, <NUM> at an inlet of one or both of the compressor sections <NUM>, <NUM>. In other words, relative to a direction of airflow through the core of the turboshaft engine <NUM>, a set of variable guide vanes <NUM> may be provided upstream of the LP compressor section <NUM> to modulate airflow into the LP compressor section <NUM>. A set of variable guide vanes <NUM> may be provided upstream of the HP compressor section <NUM> to modulate airflow into the HP compressor section <NUM> and to modulate a power output of the turboshaft engine <NUM>.

The turboshaft engine <NUM> may include a transmission <NUM>' driven by the LP turbine section <NUM> via the low pressure shaft and driving a rotatable output shaft <NUM>". In some embodiments, the transmission <NUM>' may vary a ratio between rotational speeds of the low pressure shaft and the output shaft <NUM>".

<FIG> schematically illustrates an aircraft <NUM>, in this non-limiting example a helicopter, having a first engine <NUM>', and a second engine <NUM>". The technology as described herein may be implemented with respect to a prior art multi-engine helicopter, and therefore a particular helicopter is not shown or described in detail. For simplicity, only the non-conventional aspects of the present technology are described in detail in this document.

The engines <NUM>', <NUM>" are operable to provide motive power to the aircraft <NUM> via, for example, one or more conventional transmission systems, which include the transmission <NUM>' shown in <FIG>, and conventional controls. 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 multiengine 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>', 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 multiengine 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. 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 multiengine 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.

As shown schematically in <FIG>, the first engine <NUM>' includes a first bleed air conduit <NUM> and a second bleed air conduit <NUM>, both of which bleed compressed air from respective parts of the LP and HP compressor sections <NUM>, <NUM> of the first engine <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>'. 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 a bleed air system <NUM> of the first engine <NUM>'.

As noted above, in this embodiment, the bleed air system <NUM> bleeds compressed air, via conduits <NUM> and <NUM>, from the LP compressor section <NUM> and the HP compressor section <NUM> of the first engine <NUM>', and supplies it to various parts of the first engine <NUM>' including bearing assemblies for sealing and intershaft for lubrication. It is contemplated that the bleed air system <NUM> may have a different combination of functions and/or other functions. The rest of the bleed air system <NUM> may be conventional and is therefore not shown or described in detail herein. Details of the bleed air system <NUM> that are not shown or described herein may be conventional, and are omitted to maintain clarity of this description.

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. Each of the bleed air systems <NUM>, <NUM> of the aircraft <NUM> is sized and designed to provide all of its functions at least when the engine <NUM>', <NUM>" that has the bleed air system <NUM>, <NUM> is in an "active" mode (i.e. providing motive power to the aircraft <NUM>). However, as described in more detail later in this document, each of the engines <NUM>', <NUM>" in this embodiment is also configured to operate in a "sub-idle" mode while at least another one of the engines <NUM>', <NUM>" is in an active mode.

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 with which it is used. 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 which it is used, with the "sub-idle" operation being a particular type of standby operation according to the present technology as described in this document.

In the "sub-idle" mode, the engine <NUM>', <NUM>" operates at a power level at which the engine <NUM>', <NUM>" provides no motive power, or substantially motive power, to the aircraft <NUM>. In at least some operating conditions, while in the sub-idle mode, and but-for the selective air interconnection between the bleed air systems <NUM>, <NUM> described below, a given engine <NUM>', <NUM>" may not provide sufficient pressure and/or supply rate of bleed air to its bleed air system <NUM>, <NUM> in order to enable that bleed air system <NUM>, <NUM> to provide all of its intended functions.

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 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. The selective air interconnection between the bleed air systems <NUM>, <NUM> provides for "self-sufficient" of each of the bleed air systems <NUM>, <NUM> when the engine <NUM>', <NUM>" having that bleed air system <NUM>, <NUM> operates in a sub-idle mode. A non-limiting embodiment of the selective air interconnection according to the present technology is described next, in detail.

As shown in <FIG> and <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 <NUM>' of the aircraft <NUM>, such as one or more full authority digital controllers (FADEC) (<FIG>, which may be part of the aircraft <NUM> and/or the engine(s) <NUM> for example), 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, cruise, mode of operation of the aircraft <NUM> during which both engines <NUM>', <NUM>" are 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 without a need to use the control valve <NUM>. Reference is now made to <FIG>, which shows a second in-flight, cruise, 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 operating in a sub-idle 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 sub-idle 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 sub-idle 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>". The common bleed air conduit <NUM>, the throughput of the control valve <NUM>, and the size of the bleed air system <NUM> may be selected so as to provide enough of a flow and pressure of the additional supply of bleed air to the bleed air system <NUM> so as to enable self-sufficient operation of the bleed air system <NUM> simultaneously with self-sufficient operation of the bleed air system <NUM>, with the second engine <NUM>" being in sub-idle mode. Conventional engineering principles may be used to provide for such sizing, to suit each particular embodiment and/or application of the aircraft <NUM>.

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>'. Similarly, the common bleed air conduit <NUM>, the throughput of the control valve <NUM>, and the size of the bleed air system <NUM> may be selected so as to provide enough of a flow and pressure of the additional supply of bleed air to the bleed air system <NUM> so as to enable self-sufficient operation of the bleed air system <NUM> simultaneously with self-sufficient operation of the bleed air system <NUM>, with the first engine <NUM>' being in sub-idle mode. Self-sufficiency of both of the bleed air systems <NUM>, <NUM> of the aircraft <NUM> during all modes of operation of the engines <NUM>', <NUM>" may thereby be provided.

Further, the sub-idle mode of operation as described herein has been developed as a way to improve upon prior art methods of idle operation of one or more aircraft engines, and is therefore not part of the prior art as of the time of writing this description. However, the selective air interconnection of two or more engines of an aircraft as described herein may be implemented in multi-engine aircraft, such as at least some helicopters, in which one or more of the engines are operable in a prior art idle mode. 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 therefore provide for self-sufficient operation of at least one of the engines <NUM>', <NUM>" and/or the engines' <NUM>', <NUM>" bleed air system(s) <NUM>, <NUM> in at least some operating conditions of the aircraft <NUM> in which at least some prior art engines and/or engine bleed air systems may not be self-sufficient.

Further according to the present technology, 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> to suit the various different 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 in a sub-idle mode according to the present technology (further, simply "in a sub-idle mode"), or on "standby" according to prior art methods, 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 that additional / "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 sub-idle or 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 sub-idle or 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 of the air system <NUM> difference 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 sub-idle mode on or standby. 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 operating in a sub-idle mode or on standby 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 sub-idle operation or on standby), 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 sub-idle or 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 nonsubstantially-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.

In at least some cases and in at least some embodiments, the technology described above may be implemented with, and may help provide stable sub-idle operation of one or more engines of a multi-engine aircraft. Operating one or more of an aircraft's multiple engines in a sub-idle mode according to the present technology is described in detail next.

Referring to <FIG>, the present technology provides a sub-idle mode <NUM> of operation of an aircraft engine <NUM>, <NUM>', <NUM>". The present embodiment of the sub-idle mode <NUM> is illustrated with respect to the engine <NUM> of <FIG>, but may be executed with regard to a different aircraft engine as well, such as one of the engines <NUM>' and <NUM>" of the aircraft <NUM> described above.

Operating the engine <NUM> in sub-idle mode <NUM> according to the present technology uses the rotor(s), such as the shafts / compressors / turbines <NUM>, <NUM>, <NUM>, <NUM>, of the engine <NUM> as energy accumulators, by running the engine <NUM> at lower speeds at least during repeating predetermined time intervals, in comparison with speeds at which a similar engine may be operated according to at least some prior art stand-by methods of operation. It has been found by the developers of the present technology that the sub-idle mode <NUM> provides a lower overall fuel consumption over a given operating time period in comparison with at least some prior art stand-by methods applied to similar sized/powered engines in at least some similar applications and/or operating conditions.

As shown in <FIG>, in the present embodiment, the sub-idle mode <NUM> includes a breathing-in phase <NUM> and a breathing-out phase <NUM>, executed in a repeating sequence. In the present embodiment, the breathing-in phase <NUM> includes opening, or open, the variable guide vanes <NUM>, <NUM> at the air inlet <NUM> of the engine <NUM>, and supplying fuel to the combustor <NUM> of the engine <NUM> while combustion is occurring therein or while initiating combustion, until the rotors <NUM>, <NUM>, <NUM>, <NUM> have reached an upper pre-determined level of kinetic energy. In some embodiments, the upper pre-determined level of kinetic energy corresponds to a speed of the engine <NUM> at or above an idle speed of the engine <NUM>.

Once the upper pre-determined level of kinetic energy is reached, a breathing-out phase <NUM> may be executed. As shown, the present embodiment, the breathing-out phase <NUM> includes closing the variable guide vanes <NUM>, <NUM> at the air inlet <NUM> of the engine <NUM> and reducing fuel flow to the combustor <NUM> of the engine <NUM> to at least one level that is below an idle speed of the engine <NUM>. In some embodiments, the fuel flow is terminated during the breathing-out phase <NUM>.

In an aspect, the closing of the variable guide vanes <NUM>, <NUM> at the air inlet <NUM> during each breathing-out phase <NUM> limits entry of air into the engine <NUM> and thereby reduces drag and other losses at the rotors <NUM>, <NUM>, <NUM>, <NUM> of the engine <NUM>. This helps conserve, for as long as possible, the kinetic energy stored in the rotors <NUM>, <NUM>, <NUM>, <NUM> as a result of a sequentially preceding breathing-in phase <NUM>. Rotating the rotors <NUM>, <NUM>, <NUM>, <NUM> of the engine <NUM> with the variable guide vanes <NUM>, <NUM> at the air inlet <NUM> being closed to restrict airflow through the air inlet <NUM> is referred to herein as a lower drag mode.

According to the present embodiment, the breathing-out phase <NUM> is executed until the kinetic energy in the rotors <NUM>, <NUM>, <NUM>, <NUM> drops to a lower pre-determined level of kinetic energy, at which point a sequentially next breathing-in phase <NUM> is executed to restore the kinetic energy to the upper pre-determined level for a sequentially next breathing-out phase <NUM>. The breathing-in phases <NUM> and the breathing-out phases <NUM> are executed sequentially one after the other to provide for the sub-idle mode <NUM> of operation of the engine <NUM>. In an aspect, in at least some embodiments and applications of the engine <NUM>, the sub-idle mode <NUM> allows to reduce fuel consumed by the engine <NUM> over a given time period, in comparison with prior-art idle operation methods that, for example, may run the engine <NUM> at a constant idle speed. In another aspect, in at least some embodiments and applications of the engine <NUM>, the sub-idle mode <NUM> allows the engine <NUM> to respond quickly to a sudden power demand stemming from the application in which the engine <NUM> is used.

Now referring to <FIG>, the breathing-in phases <NUM> and the breathing-out phases <NUM> of the sub-idle mode <NUM> of the present embodiment are shown and described in more detail. As shown, the breathing-in phases <NUM> all have one and the same VGV and fuelflow profile, and the breathing-out phases <NUM> all have one and the same VGV and fuelflow profile. However, it is contemplated that in other embodiments the breathing-in phases <NUM> and/or the breathing-out phases <NUM> could each include more than one profile. More particularly, the breathing-in phases <NUM> and the breathing-out phases <NUM> of the present technology may be executed with respect to an engine <NUM>, <NUM>', <NUM>" in a sequential series of cycles <NUM>, referred to herein as "breathing cycles <NUM>", and may thereby provide for sub-idle operation of the engine <NUM>, <NUM>', <NUM>" which maintains rotation of the engine's rotors <NUM>, <NUM>, <NUM>, <NUM> and allows the engine <NUM>, <NUM>', <NUM>" to be ready to quickly respond to a sudden demand for motive power from that engine <NUM>, <NUM>', <NUM>".

More particularly, with the above structure in mind and now also referring to <FIG>, there is provided a method <NUM> of operating an engine <NUM>, <NUM>', <NUM>" of a multi-engine aircraft <NUM>. In some embodiments, the method <NUM> includes operating the engine <NUM>, <NUM>', <NUM>" in a sub-idle mode, such as the sub-idle mode shown in <FIG>. As shown in <FIG>, in some embodiments, the sub-idle mode includes executing a step <NUM> of opening a set of variable guide vanes, such as the set of variable guide vanes <NUM> and/or <NUM> shown in <FIG>, upstream an air compressor section, such as the LP compressor section <NUM> and/or the HP compressor section <NUM>, of the engine <NUM>, <NUM>', <NUM>".

Also as shown in <FIG>, in some embodiments, the sub-idle mode includes executing a step <NUM> of increasing a supply rate (shown as the slope of the fuel flow graph labeled "SR" in <FIG>) of a fuel to the combustor <NUM> of the engine <NUM>, <NUM>', <NUM>" to an upper supply rate "UR". As shown, in some such embodiments, the upper supply rate "UR" is lower than a minimum fuel supply rate "MRM" required for the engine <NUM>, <NUM>', <NUM>" to provide a material amount of motive power (or simply, "to provide motive power") to the aircraft <NUM>, and greater than a minimum constant fuel supply rate "CIR" required to maintain rotation of the rotors <NUM>, <NUM>, <NUM>, <NUM> at a substantially constant minimum idle rotation speed "MIRS" of the engine <NUM>, <NUM>', <NUM>". Each of the minimum fuel supply rate "MRM", the minimum constant fuel supply rate "CIR", and the substantially constant minimum idle rotation speed "MIRS" may be a function of each particular embodiment and type of engine <NUM>, <NUM>', <NUM>" and/or the aircraft <NUM> and/or the application with which the present technology is used, and may be different and in at least some cases may be specified by the manufacturer(s) for each particular embodiment and type of engine <NUM>, <NUM>', <NUM>" / aircraft <NUM> / application.

In some embodiments, the sub-idle mode includes executing a step <NUM> of closing the set(s) of variable guide vanes <NUM> and/or <NUM>, and a step <NUM> of decreasing the supply rate "SR" of the fuel to a lower supply rate "LR" that is lower than the upper supply rate "UR", to maintain rotation of the rotors <NUM>, <NUM>, <NUM>, <NUM> of the engine <NUM>, <NUM>', <NUM>". As shown, in some embodiments, the lower supply rate "LR" is a zero supply rate, meaning that the flow of fuel to the combustor <NUM> is shut off. However, in other embodiments and depending on the particular embodiment of the engine <NUM>, <NUM>', <NUM>" for example, the lower supply rate "LR" is a non-zero supply rate, but is in at least some cases lower than the minimum constant fuel supply rate "CIR" required to maintain rotation of the rotors <NUM>, <NUM>, <NUM>, <NUM> at the substantially constant minimum idle rotation speed "MIRS" of the engine <NUM>, <NUM>', <NUM>".

Also as shown in <FIG>, in each given breathing cycle <NUM>, the step <NUM> of increasing the supply rate "SR" to the upper supply rate "UR" is followed by the step <NUM> of decreasing the supply rate "SR" to the lower supply rate "LR", and the step <NUM> of opening the set(s) of variable guide vanes <NUM>, <NUM> is followed by the step <NUM> of closing the set(s) of variable guide vanes <NUM>, <NUM>. Yet further as shown in <FIG>, in some embodiments in each given breathing cycle <NUM>, the decreasing <NUM> the supply rate "SR" is started substantially immediately after an end of the increasing <NUM> the supply rate "SR", and closing <NUM> the set(s) of variable guide vanes <NUM>, <NUM> is started substantially immediately after an end of the opening <NUM> the set(s) of variable guide vanes <NUM>, <NUM>.

Yet further as shown in <FIG>, in some embodiments in each given breathing cycle <NUM>, the opening <NUM> the set(s) of variable guide vanes <NUM>, <NUM> is simultaneous with at least part of the increasing <NUM> the supply rate "SR", and the closing <NUM> the set(s) of variable guide vanes <NUM>, <NUM> is simultaneous with at least part of the decreasing <NUM> the supply rate "SR". Yet further as shown in <FIG>, in some embodiments in each given breathing cycle <NUM>, the breathing-out phase <NUM> includes maintaining the set(s) of variable guide vanes <NUM>, <NUM> closed, and maintaining the supply rate "SR" at the lower supply rate "LR".

As shown, in some embodiments, in the sequentially-next breathing out phase <NUM>, the set(s) of variable guide vanes <NUM>, <NUM> are further maintained closed for a predetermined time period after the start of the sequentially-next breathing out phase <NUM>. In some embodiments, this time delay is omitted, for example to suit a particular embodiment and/or application of the engine <NUM>, <NUM>', <NUM>". In such embodiments, during the breathing-in phase <NUM> of a sequentially next breathing cycle <NUM>, the opening <NUM> the set(s) of variable guide vanes <NUM>, <NUM> starts at a substantially same time as a start of increasing <NUM> the supply rate "SR" of fuel to the combustor <NUM>.

In some embodiments, the method <NUM> further includes monitoring, for example via the controller(s) <NUM>', such as FADEC(s) <NUM>', and corresponding sensor(s), a rotor speed "RS" (e.g. a relative rotational speed of the rotors <NUM>, <NUM>, <NUM>, <NUM> in one non-limiting embodiment) of the engine <NUM>, <NUM>', <NUM>", and in response to the rotor speed "RS" decreasing to a pre-determined sub-idle threshold "LT" during the breathing-out phase <NUM> of a given one of the breathing cycles <NUM>, terminating the breathing-out phase <NUM> of the given cycle <NUM> and starting the breathing-in phase <NUM> of a sequentially next one of the breathing cycles <NUM>. As shown, in some embodiments, during the breathing-in phase <NUM> of the sequentially next breathing cycle <NUM>, the increasing <NUM> the supply rate "SR" of fuel to the upper supply rate "UR" starts substantially immediately after the rotor speed "RS" reaches the pre-determined sub-idle threshold "LR", and thereby ensures that the rotor speed "RS" does not materially drop below the pre-determined sub-idle threshold "LR" speed.

In another aspect and now referring to <FIG>, the present technology further provides a method <NUM> of operating an engine <NUM>, <NUM>', <NUM>" of a multi-engine aircraft <NUM>, which includes operating the engine <NUM>, <NUM>', <NUM>" in a sequential plurality of breathing cycles <NUM>, <NUM>'. 86n, with each of the breathing cycles <NUM>, <NUM>'. 86n including a breathing-in phase <NUM> followed by a breathing-out phase <NUM> as described above. For clarity, only the breathing-in phase <NUM> and the breathing-out phase <NUM> of one of the breathing cycles <NUM>, <NUM>'. 86n is shown in <FIG> in detail.

In some embodiments of the method <NUM>, a given breathing-in phase <NUM> may include a step <NUM> of: i) in response to a speed of a rotor <NUM>, <NUM>, <NUM>, <NUM> of the engine <NUM>, <NUM>', <NUM>" being at least approximately at a pre-determined sub-idle threshold "LT", opening a set variable guide vanes <NUM> and/or <NUM> disposed upstream of an air compressor section <NUM>, <NUM> of the engine <NUM>, <NUM>', <NUM>" and injecting <NUM> a fuel into a combustor <NUM> of the engine <NUM>, <NUM>', <NUM>" to increase the speed to at least approximately a pre-determined upper threshold "UT", followed by a step <NUM> of ii) in response to the speed reaching at least approximately the pre-determined upper threshold "UT", at least reducing a supply rate "SR" of the fuel into the combustor <NUM> and closing the set of variable guide vanes <NUM> and/or <NUM>.

In some such embodiments, a given breathing-out phase <NUM> may include a step <NUM> of maintaining the set of variable guide vanes <NUM> and/or <NUM> closed at least until the speed of the rotor <NUM>, <NUM>, <NUM>, <NUM> drops from the pre-determined upper threshold "UT" to at least approximately the pre-determined sub-idle threshold "LT". In some such embodiments, in the breathing-in phase <NUM> of at least one repeating breathing cycle <NUM> of the breathing cycles <NUM>, the at least reducing the supply rate "SR" starts before a start of the closing the set of variable guide vanes <NUM> and/or <NUM>. In some such embodiments, in the breathing-in phase <NUM> of at least one repeating breathing cycle <NUM> of the breathing cycles <NUM>, the opening <NUM> the set of variable guide vanes <NUM> and/or <NUM> starts at least approximately simultaneously with the injecting the fuel into the combustor <NUM>, and the injecting the fuel includes rapidly increasing, and more particularly spiking, the supply rate of the fuel into the combustor <NUM>. In some such embodiments, the at least reducing <NUM> the supply rate includes reducing the supply rate to a zero supply rate.

It is contemplated that particular timings, including starts and stops, of the steps, relative to each other, of the methods <NUM>, <NUM> described above for each given engine <NUM>, <NUM>', <NUM>" may be determined based on and/or dictated by each particular embodiment of that engine <NUM>, <NUM>', <NUM>" and/or that engine's <NUM>, <NUM>', <NUM>" application, using for example conventional engineering and design methods.

Thus, now referring back to any one of <FIG>, the present technology provides a multi-engine aircraft <NUM>, such as a multi-engine helicopter <NUM>, that includes a first engine <NUM>' operable to provide motive power to the aircraft <NUM>, a second engine <NUM>" operable to provide motive power to the aircraft <NUM>, and at least one controller <NUM>' operatively connected to the first and second engines <NUM>', <NUM>".

In some such embodiments, the controller <NUM>', which may be one or more suitable controllers of a control system of the aircraft <NUM> and/or the engine(s) <NUM>', <NUM>" for example, is configured to operate the first engine <NUM>' in a sub-idle mode while operating the second engine <NUM>" in an active mode (further, "first sub-idle configuration"). In some such embodiments, the controller <NUM>' is configured to operate the second engine <NUM>" in a sub-idle mode while operating the first engine <NUM>' in an active mode (further, "second sub-idle configuration"), either in addition to or instead of being configured to operate in the first sub-idle configuration. Since the first and second sub-idle configurations may be similar, only the first sub-idle configuration is described in detail herein next.

Referring also to <FIG>, operating the first engine <NUM>' in the sub-idle mode according to the present technology may include sequentially executing, by the controller(s) <NUM>', a plurality of breathing cycles <NUM>, with each cycle <NUM> of the plurality of cycles <NUM> including a breathing-in phase <NUM> followed by a breathing-out phase <NUM>. In some embodiments, the breathing-in phase <NUM> may include: i) modulating a set of variable guide vanes <NUM> and/or <NUM> in <FIG> upstream an air compressor section <NUM>, <NUM> of the first engine <NUM>' to an open position, and a fuel supply to a combustor <NUM> of the first engine <NUM>' to an upper supply rate "UR", followed by ii) modulating the set of variable guide vanes <NUM> and/or <NUM> to at least a substantially closed position, and the fuel supply to a lower supply rate "LR" that is lower than the upper supply rate "UR". In some embodiments, the breathing-out phase <NUM> may include maintaining the set of variable guide vanes <NUM> and/or <NUM> at least substantially closed, at least until a sequentially next breathing-in phase <NUM> for example.

In some embodiments, the modulating the set of variable guide vanes <NUM> and/or <NUM> to the open position includes modulating the set of variable guide vanes <NUM> and/or <NUM> to at least a substantially open position. In some such embodiments, the modulating the set of variable guide vanes <NUM> and/or <NUM> includes modulating the set of variable guide vanes <NUM> and/or <NUM> to a completely open position. In some embodiments, the closing the modulating the set of variable guide vanes <NUM> and/or <NUM> includes completely closing modulating the set of variable guide vanes <NUM> and/or <NUM>. In some such embodiments, the maintaining the set of variable guide vanes <NUM> and/or <NUM> at least substantially closed includes maintaining the set of variable guide vanes <NUM> and/or <NUM> completely closed.

As shown in <FIG>, in some embodiments the at least one controller <NUM>' is configured to start the modulating the fuel supply to the lower supply rate "LR" substantially immediately after terminating the modulating the fuel supply to the upper supply rate "UR", and to start the modulating the set of variable guide vanes <NUM> and/or <NUM> to the closed position substantially immediately after terminating the modulating the set of variable guide vanes <NUM> and/or <NUM> to the open position.

This control logic may be said to provide for a "spiking" of the opening and closing the set of variable guide vanes <NUM> and/or <NUM>, and for at least a partially simultaneous "spiking" of the fuel supply rate. In some such embodiments, the at least one controller <NUM>' is configured to start the spiking the fuel supply rate at least substantially simultaneously with starting the spiking of the set of variable guide vanes <NUM> and/or <NUM>, and to terminate the spiking the fuel supply rate at least substantially simultaneously with terminating the spiking of the set of variable guide vanes <NUM> and/or <NUM>. In some embodiments and applications, the spiking of the fuel supply rate according to the present technology may reduce a rate of and/or a likelihood of fuel coking in fuel manifold sections of the engine <NUM>, <NUM>', <NUM>" being operated in a sub-idle mode of the present technology, in comparison with at least some prior art engine idling methods for example.

In some embodiments and such as where the aircraft <NUM> is a helicopter for example, the at least one controller <NUM>' is configured to switch operation of the first engine <NUM>' from the sub-idle mode into an active mode of the first engine <NUM>' at any point in time during operation of the first engine <NUM>' in the sub-idle mode. Similarly, in some embodiments, the at least one controller <NUM>' is configured to switch operation of the second engine <NUM>" from the sub-idle mode into an active mode of the second engine <NUM>" at any point in time during operation of the second engine <NUM>" in the sub-idle mode. The at least one controller <NUM>' may be therefore selectively operable between the first and second sub-idle configurations described above.

The standby mode, including the various embodiments of the sub-idle mode, and active modes described above with respect to each of the first engine <NUM>' and the second engine <NUM>" may be implemented with, for example, a fuel delivery system <NUM>, a non-limiting embodiment of which is shown in <FIG>. While illustrated with respect to the abovementioned methods of operation, the fuel delivery system <NUM> may also be used to execute other types of engine operations, such as other types of standby modes for example. It should be noted that while the fuel delivery system <NUM> is described in detail with respect to the turboshaft gas turbine engines <NUM>, <NUM>', and <NUM>", the fuel delivery system <NUM> may also be used with other types of aircraft engines. To maintain clarity of the description, the fuel delivery system <NUM> is described herein next with respect to the engine <NUM> of <FIG>. Since in some embodiments as described above, each of the first engine <NUM>' and the second engine <NUM>" may be similar to the engine <NUM> with respect to which the fuel delivery system <NUM> is described, one or both of the first engine <NUM>' and the second engine <NUM>" may have and/or may be operated with a fuel delivery system <NUM>.

Referring to <FIG> and <FIG>, the fuel delivery system <NUM> includes a first fuel manifold <NUM> positioned in the engine <NUM> and operable to deliver fuel to the combustor <NUM> of the engine <NUM>. More particularly, as shown in <FIG>, in the present embodiment the first fuel manifold <NUM> is annular and includes fuel nozzles <NUM>' that are distributed circumferentially around the first fuel manifold <NUM>. To maintain clarity, only some of the fuel nozzles <NUM>' of the first fuel manifold <NUM> are shown in <FIG>. The first fuel manifold <NUM> is fed with fuel from a fuel tank <NUM>.

The fuel tank <NUM> may be a part of the aircraft with which the engine <NUM> may be used, may be a conventional fuel tank, and is therefore not described in detail herein. In some embodiments, the fuel tank <NUM> may include and/or be connected to an ecology tank. In this embodiment, the first fuel manifold <NUM> fluidly connects to the fuel tank <NUM> via a fuel conduit <NUM> and a fuel control valve <NUM> to receive fuel from the fuel tank <NUM>. The fuel control valve <NUM> may be for example a conventional fuel control valve selected to provide for the functionality of the various embodiments of the fuel delivery system <NUM> described herein.

The fuel delivery system <NUM> further includes a second fuel manifold <NUM> positioned in the engine <NUM> and operable to deliver fuel to the combustor <NUM> of the engine <NUM>. More particularly, as shown in <FIG>, in the present embodiment the second fuel manifold <NUM> is annular and includes fuel nozzles <NUM>' that are distributed circumferentially around the second fuel manifold <NUM>. To maintain clarity, only some of the fuel nozzles <NUM>' of the second fuel manifold <NUM> are shown in <FIG>. As shown in <FIG>, each of the fuel nozzles <NUM>' includes a valve <NUM>" therein.

In this embodiment, the valve <NUM>" is a check valve oriented to allow fuel flow into the combustor <NUM> out of the fuel nozzle <NUM>'. The check valve <NUM>" prevents fuel flow through the fuel nozzle <NUM>' in a direction from the fuel nozzle <NUM>' toward the fuel tank <NUM>, and more particularly in this embodiment from the combustor <NUM> into the fuel nozzle <NUM>'. In some embodiments, the valve <NUM>" is a control valve <NUM>", such as a conventional control valve for example. In some such embodiments, the control valve <NUM>" may be operatively connected to a suitable actuator, such as a conventional actuator, and may be controlled via a suitable controller, such as a conventional controller.

In some such embodiments, the control valve <NUM>" may be selectively operated to block fuel flow to, and out of / from, the fuel nozzle <NUM>' into the combustor <NUM>, to allow fuel flow out of the fuel nozzle <NUM>' into the combustor <NUM>, and in some embodiments to modulate this flow. As used herein, the terms "block" and "blocking" with respect to fuel flow, except for where explicitly stated otherwise, mean that the flow is either completely blocked, or that the flow is restricted so as to provide a trickle flow, as the term "trickle flow" is defined in this document.

In some embodiments, such as the present embodiment for example, the control valve <NUM>" includes the check valve functionality described above, for example by including a check valve therein. In such embodiments, the control valve <NUM>" may be said to be a check valve. In some embodiments, the valve <NUM>" may be multiple valves, such as a combination of a check valve and a control valve for example, which may be selected to provide the functionality described herein.

In some embodiments, the fuel nozzle <NUM>' may be configured to provide a trickle flow of fuel to the combustor <NUM> at least when the engine <NUM> is operated in the standby mode. More particularly, in this embodiment, the control valve <NUM>" is operable to provide a trickle flow of fuel out of the fuel nozzle <NUM>' to the combustor <NUM> at least when the engine <NUM> is operated in a standby mode, such as one of the standby modes described herein above for example. As used herein, the term "trickle flow" through a given element, such as a fuel nozzle or a fuel conduit, refers to a flow that is a fraction of a design flow rate through the given element, the range being selected so as to provide substantially no motive power to the engine <NUM> when the trickle flow is combusted in the combustor <NUM>. Stated otherwise, the trickle flow may be defined as a fuel flow rate selected to prevent flame-out (i.e. unintended loss) of combustion in the combustor <NUM> while providing one of: substantially no motive power to the engine <NUM>, and no motive power to the engine <NUM>, via the combustion of the trickle flow of fuel.

As an example, in some embodiments where the engine <NUM> is one of multiple similar engines powering for example a helicopter via a common conventional gearbox of the helicopter, the trickle flow of fuel to the one of the engines <NUM> may be selected to lower an output of the one engine <NUM> sufficiently low so as to decouple the one engine <NUM> from the common gearbox while maintaining combustion in the one engine <NUM>. In such cases, the engine <NUM> may be said to provide no motive power, since whatever output the engine <NUM> may maintain does not get transferred into the common gearbox and hence does not get transferred into powering the aircraft / helicopter. According to other possible control sequences in a similar helicopter example, the trickle flow of fuel to the one of the engines <NUM> may be selected to lower an output of the one engine <NUM> sufficiently low so as to maintain the one engine <NUM> coupled to the common gearbox just above the point (e.g. rotations per minute and/or power) below which the one engine <NUM> will decouple from the common gearbox. In such scenarios, the trickle fuel flow may be said to be controlled so as to provide substantially no motive power to the aircraft / helicopter.

For example, the range may equal to between <NUM>% and <NUM>% of a full throttle flowrate that the given fuel nozzle <NUM>' may be configured to provide. In some embodiments, the range may equal to between <NUM>% and <NUM>% of a full throttle flowrate that the given fuel nozzle <NUM>' may be configured to provide. In some embodiments, the range may equal to between <NUM>% and <NUM>% of a full throttle flowrate that the given fuel nozzle <NUM>' may be configured to provide. In some embodiments, the range may equal to between <NUM>% and <NUM>% of a full throttle flowrate that the given fuel nozzle <NUM>' may be configured to provide. In some embodiments, the range may equal to between <NUM>% and <NUM>% of a full throttle flowrate that the given fuel nozzle <NUM>' may be configured to provide. In some such embodiments, the ranges exclude the <NUM>% so as to provide for at least a marginal flow. In some embodiments, the range may equal to between <NUM>% and <NUM>% of a full throttle flowrate that the given fuel nozzle <NUM>' may be configured to provide. The range may be different in other embodiments, and more particularly may be defined to suit each particular embodiment of the engine <NUM>. In some embodiments, a lower limit of a range of trickle fuel flow may be defined as a minimum fuel flow rate required for stable engine operation within a range of flight conditions for which the given aircraft may be designed. In some such embodiments, the range of fuel flow may be subject to hot end durability requirements. In some such embodiments, an upper limit of the range of trickle fuel flow may be defined as a maximum flow rate may be determined by normal engine idle fuel flow rate at those flight conditions.

In the present embodiment, the second fuel manifold <NUM> is fed with fuel from the fuel tank <NUM>. As shown, the fuel tank <NUM> may be the same fuel tank <NUM> that feeds the first fuel manifold <NUM>. In other embodiments, this may not be the case. In some embodiments, the fuel tank <NUM> may be multiple fuel tanks. In this embodiment, the second fuel manifold <NUM> fluidly connects to the fuel tank <NUM> via a fuel conduit <NUM> and the same fuel control valve <NUM> as feeds the first fuel manifold <NUM>, to receive fuel from the fuel tank <NUM>. In other embodiments, the fuel control valve <NUM> may be for example multiple conventional fuel control valves selected to provide for the functionality of the various embodiments of the fuel delivery system <NUM> described herein. For example, in some embodiments each of the fuel conduits <NUM>, <NUM> may be defined in part by one or more dedicated fuel control valves <NUM> selected to provide for the functionality of the various embodiments of the fuel delivery system <NUM> described herein.

As shown schematically in <FIG> with reference numeral <NUM>', in this embodiment of the fuel delivery system <NUM>, there is no cross-flow between the fuel conduits <NUM>, <NUM> at the fuel control valve <NUM> or the fuel tank <NUM>. In this embodiment, the fuel control valve <NUM> is operable to do any one of: i) supply fuel from the fuel tank <NUM> to both the first and second fuel manifolds <NUM>, <NUM>, ii) supply fuel from the fuel tank <NUM> to the first fuel manifold <NUM> while blocking fuel supply to the second fuel manifold <NUM>, and iii) supply fuel from the fuel tank <NUM> to the second fuel manifold <NUM> while blocking fuel supply to the first fuel manifold <NUM>. While in the present embodiment the fuel delivery system <NUM> includes two fuel manifolds <NUM>, <NUM> that are selectively fed with fuel from the fuel tank <NUM>, in other embodiments, the fuel delivery system <NUM> may include additional one or more fuel manifolds that may be fed from the one or more fuel tanks <NUM> via additional corresponding fuel conduit(s). In some such embodiments, at least some of the additional fuel manifolds may be similar to the second fuel manifold <NUM>.

Now referring to <FIG>, a fuel delivery system <NUM> is shown. The fuel delivery system <NUM> is similar to the fuel delivery system <NUM>. Therefore, elements of the fuel delivery system <NUM> that correspond to elements of the fuel delivery system <NUM> have been labeled with the same reference numerals.

A difference between the fuel delivery system <NUM> and the fuel delivery system <NUM> is that the fuel delivery system <NUM> includes cross-flow conduits <NUM> that fluidly connect fuel nozzles <NUM>' of the first fuel manifold <NUM> to the fuel nozzles <NUM>' of the second fuel manifold <NUM>. More particularly, in this embodiment and as shown in <FIG> with respect to one corresponding pair of fuel nozzles <NUM>', <NUM>' to maintain clarity of the figure, each fuel nozzle <NUM>' of the first fuel manifold <NUM> fluidly connects to one respective fuel nozzle <NUM>' of the second fuel manifold <NUM> via a cross-flow conduit <NUM> at a location that is downstream of the valve <NUM>" of the respective fuel nozzle <NUM>'. Since in this embodiment the cross-flow connection is similar for all corresponding pair of fuel nozzles <NUM>', <NUM>', to maintain clarity of the figure, the cross-flow connection is shown in <FIG> and described next with respect to only one of the corresponding pairs of fuel nozzles <NUM>', <NUM>'.

Also in this embodiment, and as shown in <FIG>, the first fuel manifold <NUM> is fluidly connected to a fuel tank <NUM> via a fuel conduit <NUM> that connects into the first fuel manifold <NUM> at a point that is upstream of all of the fuel nozzles <NUM>' of the first fuel manifold <NUM>. Similarly in this embodiment, and as shown in <FIG>, the second fuel manifold <NUM> is fluidly connected to a fuel tank <NUM>, such as but not necessarily the same fuel tank <NUM>, via a fuel conduit <NUM> that connects into the second fuel manifold <NUM> at a point that is upstream of all of the fuel nozzles <NUM>' of the second fuel manifold <NUM>. Accordingly, each of the cross-flow conduits <NUM> may be said to be fluidly connected to the first fuel manifold <NUM> at a point that is downstream of the point of connection of the fuel conduit <NUM> into the first fuel manifold <NUM>, and to the second fuel manifold <NUM> at a point that is downstream of the point of connection of the fuel conduit <NUM> into the second fuel manifold <NUM>.

In the present embodiment, the cross-flow conduit <NUM> is configured, such as via one or more suitable flow control devices for example, to provide a trickle flow from the fuel nozzle <NUM>' of the first fuel manifold <NUM> to the fuel nozzle <NUM>' of the second fuel manifold <NUM> at least when the engine <NUM> is operated in a standby mode, such as one of the standby modes described herein above. In some embodiments, the flow limiter may be a conventional control valve for example. As another example, in some embodiments, the flow control device may be a flow control aperture and/or a flow control valve and/or a flow restrictor, selected to provide the trickle flow through the cross-flow conduit <NUM>. It is contemplated that any suitable conventional one or more flow control devices may be used to provide the trickle flow.

In some embodiments, the trickle flow may be in a range of <NUM>% and <NUM>% of a full throttle flowrate that the fuel nozzle <NUM>' may be configured to provide. In some embodiments, the range may equal to between <NUM>% and <NUM>% of a full throttle flowrate that the fuel nozzle <NUM>' may be configured to provide. In some embodiments, the range may equal to between <NUM>% and <NUM>% of a full throttle flowrate that the fuel nozzle <NUM>' may be configured to provide. In some embodiments, the range may equal to between <NUM>% and <NUM>% of a full throttle flowrate that the fuel nozzle <NUM>' may be configured to provide. In some embodiments, the range may equal to between <NUM>% and <NUM>% of a full throttle flowrate that the fuel nozzle <NUM>' may be configured to provide. In some such embodiments, a given range may exclude the <NUM>% so as to provide for at least a marginal flow. In some embodiments, the range may equal to between <NUM>% and <NUM>% of a full throttle flowrate that the fuel nozzle <NUM>' may be configured to provide. This range may be different in other embodiments, and more particularly may be defined to suit each particular embodiment of the engine <NUM>. In some embodiments, a range of trickle fuel flow may be defined as described above.

In other embodiments, a different number and/or configuration of the cross-flow conduits <NUM> may be used to provide for the functionality described herein. For example, in some embodiments, a cross-flow conduit <NUM> may fluidly connect to multiple fuel nozzles <NUM>' of the second fuel manifold <NUM> downstream of the valves <NUM>" of those fuel nozzles <NUM>'.

Still referring to <FIG>, another difference between the fuel delivery system <NUM> and the fuel delivery system <NUM> is that in the fuel delivery system <NUM>, as shown with reference numeral <NUM>', the fuel control valve <NUM> is operable to provide a trickle flow of fuel from the first fuel conduit <NUM> to the second fuel conduit <NUM>. More particularly, the fuel control valve <NUM> may be selected to be operable to do any one of: i) supply fuel from the fuel tank <NUM> to both the first and second fuel manifolds <NUM>, <NUM>, ii) supply fuel from the fuel tank <NUM> to the first fuel manifold <NUM> while blocking fuel supply to the second fuel manifold <NUM>, iii) supply fuel from the fuel tank <NUM> to the second fuel manifold <NUM> while blocking fuel supply to the first fuel manifold <NUM>, and iv) supply a trickle flow of fuel from the first fuel conduit <NUM> to the second fuel conduit <NUM> while blocking fuel supply to the second fuel manifold <NUM>.

In some embodiments, the trickle flow may be in a range between <NUM>% and <NUM>% of a full throttle flowrate that the second fuel manifold <NUM> may be configured to provide. In some embodiments, the range may equal to between <NUM>% and <NUM>% of a full throttle flowrate that the second fuel manifold <NUM> may be configured to provide. In some embodiments, the range may equal to between <NUM>% and <NUM>% of a full throttle flowrate that second fuel manifold <NUM> may be configured to provide. In some embodiments, the range may equal to between <NUM>% and <NUM>% of a full throttle flowrate that the second fuel manifold <NUM> may be configured to provide. In some embodiments, the range may equal to between <NUM>% and <NUM>% of a full throttle flowrate that the second fuel manifold <NUM> may be configured to provide. In some such embodiments, the ranges exclude the <NUM>% so as to provide for at least a marginal flow. In some embodiments, the range may equal to between <NUM>% and <NUM>% of a full throttle flowrate that the second fuel manifold <NUM> may be configured to provide. This range may be different in other embodiments, and more particularly may be defined to suit each particular embodiment of the engine <NUM>. For example, in some embodiments, a range of trickle fuel flow may be defined as described above.

As shown, in some embodiments, the fuel control valve <NUM> may include a cross-flow conduit <NUM>" therein for providing the trickle flow. In other embodiments, a cross-flow conduit <NUM> may be provided separate from the valve <NUM>. For example, the trickle flow of fuel may be provided via a cross-flow conduit <NUM> that may extend for example from the first fuel conduit <NUM> to the second fuel conduit <NUM> at points upstream of the first and second fuel manifolds <NUM>, <NUM> but downstream of the valve <NUM>. In some such embodiments, this trickle flow may be provided while blocking fuel supply to the second fuel manifold <NUM> using the valve <NUM>, and/or while blocking fuel supply out of the second fuel manifold <NUM> into the combustor <NUM> using each of the valves <NUM>" in the second fuel manifold <NUM>. In some embodiments, the cross-flow conduit <NUM> may include flow control devices <NUM> therein, such as one or more conventional flow modulating valves and/or flow restrictors and/or flow control apertures for example, to provide for and/or modulate the trickle flow to be within a given range of trickle flows, as described above.

Now referring to <FIG>, a fuel delivery system <NUM> is shown. The fuel delivery system <NUM> is similar to the fuel delivery system <NUM>. Therefore, elements of the fuel delivery system <NUM> that correspond to elements of the fuel delivery system <NUM> have been labeled with the same reference numerals. Similar to the description of the fuel delivery systems <NUM> and <NUM> above, the fuel delivery system <NUM> is described with respect to one corresponding pair of fuel nozzles <NUM>' and <NUM>'. In this embodiment, the other corresponding pairs of fuel nozzles <NUM>' and <NUM>' are similar to the corresponding pair of fuel nozzles <NUM>' and <NUM>' that is shown in detail in <FIG> and described in detail herein below. In some embodiments, the fuel delivery system <NUM> may have other types of fuel nozzles as well.

As shown in <FIG>, a difference between the fuel delivery system <NUM> and the fuel delivery system <NUM> is that each of the fuel nozzles <NUM>' of the second fuel manifold <NUM> of the fuel delivery system <NUM> includes a by-pass <NUM> around the valve <NUM>" of that fuel nozzle <NUM>'. The by-pass <NUM> in this embodiment is provided via a suitable fuel conduit connecting a point upstream of the valve <NUM>" to a point downstream of the valve <NUM>". However, any other suitable by-pass may be used to provide for the functionality described herein.

For example, in some embodiments, the valve <NUM>" may include the by-pass <NUM> therein. In some embodiments, the by-pass <NUM> may include one or more flow control devices therein, such as conventional flow modulating valves and/or flow restrictors for example, to provide for and/or modulate the trickle by-pass flow to be within a given range of trickle by-pass flows, as described above. In some embodiments, the range may equal to between <NUM>% and <NUM>% of a full throttle flowrate that the fuel nozzle <NUM>' may be configured to provide. In some embodiments, the range may equal to between <NUM>% and <NUM>% of a full throttle flowrate that the fuel nozzle <NUM>' may be configured to provide. In some embodiments, the range may equal to between <NUM>% and <NUM>% of a full throttle flowrate that the fuel nozzle <NUM>' may be configured to provide. In some embodiments, the range may equal to between <NUM>% and <NUM>% of a full throttle flowrate that the fuel nozzle <NUM>' may be configured to provide. In some embodiments, the range may equal to between <NUM>% and <NUM>% of a full throttle flowrate that the fuel nozzle <NUM>' may be configured to provide. In some such embodiments, a given range may exclude the <NUM>% so as to provide for at least a marginal flow. In some embodiments, the range may equal to between <NUM>% and <NUM>% of a full throttle flowrate that the fuel nozzle <NUM>' may be configured to provide. This range may be different in other embodiments, and more particularly may be defined to suit each particular embodiment of the engine <NUM>. For example, in some embodiments, a range of trickle fuel flow may be defined as described above.

The various embodiments of fuel delivery system structures described above permit the second fuel manifold(s) <NUM> to be maintained at least substantially wet (i.e. filled with fuel) while the second fuel manifold(s) <NUM> is/are blocked to keep the engine <NUM> in a standby mode. In an aspect, the placement of the valves <NUM>" in the fuel nozzles <NUM>' of the second fuel manifold(s) <NUM> allows to increase a percentage of the second fuel manifold(s) <NUM> that is maintained wet in such operating conditions.

Accordingly, in an aspect, the fuel delivery systems <NUM>, <NUM>, <NUM> described herein help reduce a response time associated with a switch of the engine <NUM> from a standby mode to an active mode. In a further aspect, the provision of the one or more trickle flows by the fuel delivery systems <NUM>, <NUM>, <NUM> as described herein above helps reduce fuel stagnation and/or coking in the fuel nozzles <NUM>' that are kept "idle" to keep the engine <NUM> in a standby mode. In a further aspect, the provision of the one or more trickle flows by the fuel delivery systems <NUM>, <NUM>, <NUM> helps improve thermal distribution within the fuel nozzles <NUM>'.

With the various embodiments of the fuel delivery systems <NUM>, <NUM>, <NUM> described above in mind, the present technology provides methods of operating engines and fuel delivery systems. These methods are described in detail next.

Referring to <FIG>, the present technology provides a method <NUM> of operating an aircraft engine <NUM>. In some embodiments, the method <NUM> includes a step <NUM> of maintaining combustion in a combustor <NUM> of the aircraft engine <NUM>. The maintaining combustion includes supplying fuel to the combustor <NUM> via a first fuel manifold <NUM> while executing a step <NUM> of blocking fuel flow out of a second fuel manifold <NUM> into the combustor <NUM> at a location within the second fuel manifold <NUM>. As seen above, in some embodiments, the step <NUM> of substantially blocking fuel flow may include at least substantially closing at least one valve, such as the valve <NUM>' for example, in the second fuel manifold <NUM>. Also as seen above, in some embodiments, the step <NUM> of substantially blocking fuel flow may include at least substantially closing, or at least moving into a corresponding position, at least one valve, such as the fuel control valve <NUM> for example, upstream of the second fuel manifold <NUM>.

In some embodiments, the step <NUM> of blocking fuel flow may include providing a trickle flow of fuel into the combustor <NUM> out of the second fuel manifold <NUM>. In some embodiments, the providing the trickle flow may include incompletely closing the at least one valve <NUM>" in the second fuel manifold <NUM>. In some such embodiments, the providing the trickle flow may include incompletely closing each of the valves <NUM>" in the second fuel manifold <NUM>, and more particularly in each of the fuel nozzles <NUM>', <NUM>' thereof. In some such embodiments, the providing the trickle flow may include modulating the trickle flow provided by each of the valves <NUM>", by modulating each of the valves <NUM>". In some embodiments, the modulating may be executed using a suitable control sequence, such as a conventional control sequence, to maintain the trickle flow provided by each of the valves <NUM>" within a given range of flows, such as described herein above for example.

In some embodiments, the providing the trickle flow may include providing a by-pass <NUM>" across at least one valve <NUM>" in the second fuel manifold <NUM>. In some such embodiments, the providing the trickle flow may include providing a by-pass <NUM>" across each of the valves <NUM>" in the second fuel manifold <NUM>, and more particularly in each of the fuel nozzles <NUM>', <NUM>' thereof. Stated otherwise, the providing the trickle flow may include bypassing at least one closed valve <NUM>" in at least one fuel nozzle <NUM>' of the second fuel manifold <NUM>, and in some embodiments, bypassing the closed valve <NUM>" in each fuel nozzle <NUM>' of the second fuel manifold <NUM>.

In some embodiments, the providing the trickle flow may include providing a trickle flow of fuel from a fuel conduit <NUM> upstream of the first fuel manifold <NUM> to a fuel conduit <NUM> upstream of the second fuel manifold <NUM>. In some embodiments, this trickle flow is provided via a cross-flow conduit <NUM> disposed upstream of the first and second fuel manifolds <NUM>, <NUM> and interconnecting the fuel conduits <NUM>, <NUM> that supply fuel to the first and second fuel manifolds <NUM>, <NUM>. In some embodiments, the providing the trickle flow includes controlling a flow control valve <NUM> fluidly connected to both the fuel conduit <NUM> upstream of the first fuel manifold <NUM> and the fuel conduit <NUM> upstream of the second fuel manifold <NUM>. In some such embodiments, the cross-flow conduit <NUM> is disposed downstream of the flow control valve <NUM>.

In some embodiments, the step <NUM> of blocking fuel flow may include providing a trickle flow of fuel from the first fuel manifold <NUM> to the second fuel manifold <NUM>. In some embodiments, the providing the trickle flow of fuel from the first fuel manifold <NUM> to the second fuel manifold <NUM> includes providing the trickle flow to a location downstream of at least one valve <NUM>" in the second fuel manifold <NUM>, and more particularly in at least one fuel nozzle <NUM>' thereof. In some such embodiments, the providing the trickle flow of fuel from the first fuel manifold <NUM> to the second fuel manifold <NUM> includes providing the trickle flow from fuel nozzles <NUM>' of the first fuel manifold <NUM> to the fuel nozzles <NUM>' of the second fuel manifold <NUM>.

More particularly, in some embodiments, the trickle flow may be provided from each fuel nozzle <NUM>' of the first fuel manifold <NUM> to one fuel nozzle <NUM>' of the second fuel manifold <NUM>, at a location downstream of the valve <NUM>" of that fuel nozzle <NUM>' of the second fuel manifold <NUM>. In some embodiments, the providing the trickle flow may include modulating the trickle flow, for example to maintain the trickle flow within a given range of flows, such as described herein above for example. In some embodiments, the providing the trickle flow is into the combustor <NUM> out of the second fuel manifold <NUM>, and more particularly out of each fuel nozzle <NUM>' of the second fuel manifold <NUM>.

In some embodiments, the step <NUM> of maintaining combustion is part of operating the aircraft engine <NUM> in a standby mode, such as one of the standby modes described above for example, and the method <NUM> further includes switching the aircraft engine <NUM> from the standby mode to an active mode, as the active mode is described above, the switching including unblocking fuel flow out of the second fuel manifold <NUM> into the combustor <NUM>. In some embodiments, the unblocking fuel flow includes opening the at least one valve <NUM>" in the second fuel manifold <NUM>, and in some cases opening all of the valves <NUM>" in the second fuel manifold <NUM>.

In at least some embodiments and applications, the various embodiments of the method <NUM> described herein may help reduce a response time associated with a switch of the engine <NUM> from a standby mode to an active mode. In at least some of the embodiments of the method <NUM> described above, the method <NUM> includes maintaining the second fuel manifold <NUM> and its associated fuel supply conduit <NUM> wet (i.e. at least substantially filled with fuel) while fuel supply out of the second fuel manifold <NUM> into the combustor <NUM> is blocked.

While the second fuel manifold <NUM> is blocked, providing at least one of the various trickle flows associated with the second fuel manifold <NUM> as described herein above, in at least some applications, helps reduce fuel stagnation and/or coking in the fuel nozzles <NUM>' of the second fuel manifold <NUM>. In another aspect, provision of at least one of the various trickle flows associated with the second fuel manifold <NUM> as described herein above, in at least some applications, helps improve thermal management of the fuel nozzles <NUM>' of the second fuel manifold <NUM>.

Now referring to <FIG>, the present technology also provides a method <NUM> of operating an engine fuel delivery system <NUM>, <NUM>, <NUM> of an aircraft engine <NUM>. The method <NUM> includes a step <NUM> of maintaining combustion in a combustor <NUM> of the aircraft engine <NUM> by supplying fuel to the combustor <NUM> via a first fuel manifold <NUM>. In some embodiments, the method <NUM> may include a step <NUM>, which may be executed while step <NUM> is being executed. Step <NUM> may include providing a trickle flow of fuel out of a second fuel manifold <NUM> into the combustor <NUM>. In some embodiments and applications, the method <NUM> may be implemented with respect to fuel delivery systems <NUM>, <NUM>, <NUM> having more than two fuel manifolds <NUM>, <NUM>. In some such embodiments, the step <NUM> may include providing a trickle flow of fuel out of the second fuel manifold <NUM> and the one or more additional fuel manifolds. The one or more additional fuel manifolds may be similar to the second fuel manifold <NUM>, and are therefore not described in detail herein.

In some embodiments, the providing the trickle flow includes providing a trickle flow of fuel out of the second fuel manifold via a by-pass <NUM>" across a valve <NUM>" in the second fuel manifold <NUM>, and in some embodiments across each valve <NUM>" in the second fuel manifold <NUM>. In some embodiments, the providing the trickle flow includes providing a trickle flow of fuel to a location downstream of a valve <NUM>" in each fuel nozzle <NUM>' of the second fuel manifold <NUM> from at least one fuel nozzle <NUM>' of the first fuel manifold <NUM>. In some embodiments, the providing the trickle flow includes incompletely closing a valve <NUM>" in each fuel nozzle <NUM>' of the second fuel manifold <NUM>. Stated otherwise, the providing the trickle flow of fuel out of the second fuel manifold <NUM> may include positioning a valve <NUM>" in each fuel nozzle <NUM>' in the second fuel manifold <NUM> to a substantially closed position.

In some embodiments, the method <NUM> further includes providing a trickle flow of fuel from a fuel conduit <NUM> upstream of the first fuel manifold <NUM> to a fuel conduit <NUM> upstream of the second fuel manifold <NUM> while providing the trickle flow of fuel out of the second fuel manifold <NUM> via at least one of the ways as described above. In some embodiments, the method <NUM> further includes modulating at least one of the trickle flow out of the second fuel manifold <NUM> and the trickle flow provided into the fuel conduit <NUM> upstream of the second fuel manifold <NUM>.

In some embodiments, the step <NUM> of maintaining combustion is part of operating the aircraft engine <NUM> in a standby mode, such as one of the standby modes described above for example, and the method <NUM> further includes switching the aircraft engine <NUM> from the standby mode to an active mode, as the active mode is described above, the switching including unblocking fuel flow out of the second fuel manifold <NUM> into the combustor <NUM>. In some embodiments, the unblocking fuel flow includes opening at least one valve <NUM>" in the second fuel manifold <NUM>, and in some cases opening a valve <NUM>" in each fuel nozzle <NUM>' of the second fuel manifold <NUM>.

The particulars of how some of the functions described above are not described in detail to maintain clarity of this description, because those particulars may depend on each given embodiment of the aircraft <NUM> and the controller(s) <NUM>' with which the present technology is implemented, and because those particulars may be implemented using suitable corresponding conventional components of the aircraft <NUM> and using suitable conventional control methods.

The above description is meant to be exemplary only, and one skilled in the art will recognize that changes may be made to the embodiments described without departing from the scope of the disclosed technology. For example, in some embodiments, the subidling methods and technology described herein may be implemented by using dedicated actuated panel(s) <NUM>' (<FIG>) disposed at or in the air inlet(s) <NUM> of an engine <NUM> upstream of the compressor section(s) <NUM>, <NUM>, instead of or in addition to using the set(s) of variable guide vanes <NUM>, <NUM>. In some such embodiments, the panel(s) <NUM>' may be configured to substantially completely shut off airflow to the respective air inlet(s) <NUM> when actuated to a closed position, and to substantially fully open the air inlet(s) <NUM> when actuated to an open position.

Similar to the variable guide vanes <NUM>, <NUM>, actuation and control of the panel(s) <NUM>' may be executed using any suitable actuator(s), and using any suitable controller(s) of the engine <NUM> and/or the aircraft <NUM>. As yet another example, as shown in <FIG> for example, in some embodiments, one or more of the air systems <NUM>, <NUM>, <NUM> may include one or more pressure wave dampers <NUM>.

Claim 1:
A fuel delivery system (<NUM>, <NUM>, <NUM>) for an aircraft engine (<NUM>, <NUM>', <NUM>"), comprising:
a first fuel manifold (<NUM>) operable to deliver fuel to a combustor (<NUM>) of the aircraft engine (<NUM>, <NUM>', <NUM>") via a first set of fuel nozzles (<NUM>'),
a first fuel conduit (<NUM>) fluidly connected to the first fuel manifold (<NUM>) at a first point, the first fuel conduit (<NUM>) being fluidly connectable to a fuel source (<NUM>),
a second fuel manifold (<NUM>) operable to deliver fuel to the combustor (<NUM>) of the aircraft engine (<NUM>, <NUM>', <NUM>") via a second set of fuel nozzles (<NUM>', <NUM>'), and
a second fuel conduit (<NUM>) fluidly connected to the second fuel manifold (<NUM>) at a second point, the second fuel conduit (<NUM>) being fluidly connectable to the fuel source (<NUM>);
wherein the fuel delivery system (<NUM>, <NUM>, <NUM>) further comprises a cross-flow fuel conduit (<NUM>, <NUM>) fluidly connected to the first fuel manifold (<NUM>) at a point in the first fuel manifold (<NUM>) that is downstream of the first point, and to the second fuel manifold (<NUM>) at a point in the second fuel manifold (<NUM>) that is downstream of the second point, and
the second fuel manifold (<NUM>) includes a fuel nozzle (<NUM>') comprising a valve (<NUM>") operable to block fuel flow from the second fuel manifold (<NUM>) to the combustor (<NUM>), and the cross-flow fuel conduit (<NUM>) fluidly connects to the fuel nozzle (<NUM>') at a location that is fluidly downstream of the valve (<NUM>")
characterised in that:
the cross-flow fuel conduit (<NUM>, <NUM>) is configured to deliver trickle flow of fuel from the first fuel manifold (<NUM>) to the second fuel manifold (<NUM>) while the valve (<NUM>") blocks fuel flow from the second fuel manifold (<NUM>) to the combustor (<NUM>).