Patent Description:
When operating aircraft with multiple engines, there may be certain portions of a mission that do not require both engines to be operating at full regime. In cruising conditions, operating a single engine at a relatively high regime, instead of both engines at lower regimes, may allow for better fuel efficiency.

Improvements are needed for managing the various engine operating regimes.

<CIT> discloses a prior art method for operating an aircraft according to the preamble of claim <NUM>.

<CIT> discloses a prior art multi-engine aircraft power plant with heat recuperation.

<CIT> also discloses a prior art control system for an in-flight engine restart system of a rotorcraft.

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

In an embodiment, the method is performed by a Full Authority Digital Engine Control (FADEC).

In an embodiment according to any of the previous embodiments, the method further comprises monitoring the set of engine parameters and outputting the engine availability confirmation when the engine operating conditions are met.

In an embodiment according to any of the previous embodiments, the set of engine conditions are monitored and the engine availability confirmation is output by a Full Authority Digital Engine Control (FADEC).

In an embodiment according to any of the previous embodiments, the method further comprises monitoring the set of aircraft parameters and outputting the aircraft availability confirmation when the aircraft operating conditions are met.

In an embodiment according to any of the previous embodiments, the set of aircraft parameters are monitored and the aircraft availability confirmation is output by a Full Authority Digital Engine Control (FADEC).

In an embodiment according to any of the previous embodiments, the aircraft availability confirmation is received from aircraft avionics.

In an embodiment according to any of the previous embodiments, the engine operating conditions comprise at least one of:.

In an embodiment according to any of the previous embodiments, the aircraft operating conditions comprise at least one of:.

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

In one embodiment, the system is a Full Authority Digital Engine Control (FADEC).

In an embodiment according to any of the previous embodiments, the program code is further executable for monitoring the set of engine parameters and outputting the engine availability confirmation when the engine operating conditions are met.

In an embodiment according to any of the previous embodiments, the program code is further executable for monitoring the set of aircraft parameters and outputting the aircraft availability confirmation when the aircraft operating conditions are met.

<FIG> illustrates a gas turbine engine <NUM>. In this example, the gas turbine <NUM> is a turboshaft engine generally 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 a fuel flow, delivered to the combustor <NUM> via fuel nozzles <NUM> from fuel system (not depicted), and ignited for generating a stream of hot combustion gases, a high pressure turbine section <NUM> for extracting energy from the combustion gases and driving the high pressure compressor section <NUM> via a high pressure shaft <NUM>, and a low pressure turbine section <NUM> for further extracting energy from the combustion gases and driving the low pressure compressor section <NUM> via a low pressure shaft <NUM>.

The turboshaft engine <NUM> may include a transmission <NUM> driven by the low pressure shaft <NUM> and driving a rotatable output shaft <NUM>. The transmission <NUM> may optionally be provided to vary a ratio between rotational speeds of the low pressure shaft <NUM> and the output shaft <NUM>.

The compressors and turbines are arranged in low and high pressures spools <NUM>, <NUM>, respectively. In use, one or more controllers <NUM>, such as one or more full authority digital controllers (FADEC) providing full authority digital control of the various relevant parts of the engine <NUM>, control operation of the engine <NUM>. The controller <NUM> may also be an engine control unit (ECU) or electronic engine control (EEC), forming part of the FADEC. Each controller <NUM> may be used to control one or more engines <NUM> of an aircraft (H). Additionally, in some embodiments the controller(s) <NUM> may be configured for controlling operation of other elements of the aircraft (H), for instance the main rotor <NUM>.

The low pressure compressor section <NUM> is configured to independently rotate from the high pressure compressor section <NUM> by virtues of their mounting on different engine spools. The low pressure compressor section <NUM> may include one or more compression stages, and the high pressure compressor section <NUM> may include one or more compression stages. In the embodiment shown in <FIG>, the low pressure (LP) compressor section <NUM> includes a single compressor stage 12A, which includes a single mixed flow rotor (MFR), for example such as described in <CIT>, entitled "MIXED FLOW AND CENTRIFUGAL COMPRESSOR FOR GAS TURBINE ENGINE", the contents of which are hereby expressly incorporated herein by reference in its entirety.

The LP compressor <NUM> and the HP compressor <NUM> are configured to deliver desired respective pressure ratios in use, as will be described further below. The LP compressor <NUM> may have a bleed valve <NUM> (shown schematically) which may be configured to selectively bleed air from the LP compressor <NUM> according to a desired control regime of the engine <NUM>, for example to assist in control of compressor stability. The design of such valve <NUM> is well known and not described herein in further detail. Any suitable bleed valve arrangement may be used.

As mentioned, the HP compressor section <NUM> is configured to independently rotate from the LP compressor section <NUM> by virtue of their mounting on different engine spools. The HP compressor section <NUM> may include one or more compression stages, such as a single stage, or two or more stages 14A as shown in more detail in <FIG>. It is contemplated that the HP compressor section <NUM> may include any suitable type and/or configuration of stages. The HP compressor is configured to deliver a desired pressure ratio in use, as will be described further below. The HP compressor <NUM> may have a bleed valve <NUM> (shown schematically) which may be configured to selectively bleed air from the HP compressor section <NUM> according to a desired control regime of the engine <NUM>, for example to assist in control of compressor stability. The design of such valve <NUM> is well known and not described herein in further detail. Any suitable bleed valve arrangement may be used.

The engine <NUM> has two or more compression stages <NUM>, <NUM> to pressurize the air received through an air inlet <NUM>, and corresponding turbine stages <NUM>, <NUM> which extract energy from the combustion gases before they exit via an exhaust outlet <NUM>. In the illustrated embodiment, the turboshaft engine <NUM> includes a low pressure spool <NUM> and a high pressure spool <NUM> mounted for rotation about an engine axis <NUM>. The low pressure and high pressure spools <NUM>, <NUM> are independently rotatable relative to each other about the axis <NUM>. The term "spool" is herein intended to broadly refer to drivingly connected turbine and compressor rotors, and need not mean the simple shaft arrangements depicted.

The low pressure spool <NUM> may include a low pressure shaft <NUM> interconnecting the low pressure turbine section <NUM> with the low pressure compressor section <NUM> to drive rotors of the low pressure compressor section <NUM>. The low pressure compressor section <NUM> may include at least one low pressure compressor rotor directly drivingly engaged to the low pressure shaft <NUM>, and the low pressure turbine section <NUM> may include at least one low pressure turbine rotor directly drivingly engaged to the low pressure shaft <NUM> so as to rotate the low pressure compressor section <NUM> at a same speed as the low pressure turbine section <NUM>. In other embodiments (not depicted), the low pressure compressor section <NUM> may be connected via a suitable transmission (not depicted) to run faster or slower (as desired) than the low pressure turbine section <NUM>.

The high pressure spool <NUM> includes a high pressure shaft <NUM> interconnecting the high pressure turbine section <NUM> with the high pressure compressor section <NUM> to drive rotor(s) of the high pressure compressor section <NUM>. The high pressure compressor section <NUM> may include at least one high pressure compressor rotor (in this example, two rotors are provided, a MFR compressor 14A and a centrifugal compressor 14B) directly drivingly engaged to the high pressure shaft <NUM>. The high pressure turbine section <NUM> may include at least one high pressure turbine rotor (in this example there is one HP turbine 18A) directly drivingly engaged to the high pressure shaft <NUM> so as to drive the high pressure compressor section <NUM> at a same speed as the high pressure turbine section <NUM>. In some embodiments, the high pressure shaft <NUM> and the low pressure shaft <NUM> are concentric, though any suitable shaft and spool arrangement may be employed.

The turboshaft engine <NUM> may include a set of variable guide vanes (VGVs) <NUM> upstream of the LP compressor section <NUM>, and may include a set of variable guide vanes (VGVs) <NUM> upstream of the HP compressor section <NUM>. The first set of variable guide vanes 36A may be provided upstream of the low pressure compressor section <NUM>. A set of variable guide vanes 36B may be provided upstream of the high pressure compressor section <NUM>. The variable guide vanes 36A, 36B may be independently controlled by suitable one or more controllers <NUM>, as described above. The variable guide vanes 36A, 36B may direct inlet air to the corresponding stage of compressor sections <NUM>, <NUM>. The set of variable guide vanes 36A, 36B may be operated to modulate the inlet airflow to the compressors in a manner which allows for improved control of the output power of the turboshaft engines <NUM>, as described in more detail below. The VGVs may be provided with any suitable operating range. In some embodiments, VGV vanes 36B may be configured to be positioned and/or modulated between about +<NUM> degrees and about -<NUM> degrees, with <NUM> degrees being defined as aligned with the inlet airflow, as depicted schematically in <FIG>. In a more specific embodiment, the VGV vanes 36A and/or 36B may rotate in a range from +<NUM> degrees to -<NUM> degrees, or from +<NUM> degrees to -<NUM> degrees, and more particularly still from <NUM> degrees to -<NUM> degrees. The two set of VGV vanes <NUM> may be configured for a similar range of positions, or other suitable position range.

In some embodiments, the set of variable guide vanes 36A upstream of the low pressure compressor section <NUM> may be mechanically decoupled from the set of variable guide vanes 36B upstream of the high pressure compressor section <NUM>, having no mechanical link between variable guide vanes 36A, 36B to permit independent operation of the respective stages. The VGV vanes 36A, 36B may be operatively controlled by the controller(s) <NUM> described above, to be operated independently of each other. Indeed, the engines 10A, 10B are also controlled using controller(s) <NUM> described above, to carry out the methods described in this document. For the purposes of this document, the term "independently" in respects of the VGVs <NUM> means that the position of one set of the VGV vanes (e.g. 36A) may be set without effecting any change to a position of the other set of the VGV vanes (e.g. 36B), and vice versa.

Independent control of the VGVs 36A, 36B may allow the spools <NUM>, <NUM> to be operated to reduce or eliminate or reduce aerodynamic coupling between the spools <NUM>, <NUM>. This may permit the spools <NUM>, <NUM> to be operated at a wider range of speeds than may otherwise be possible. The independent control of the VGV vanes 36A, 36B may allow the spools <NUM>, <NUM> to be operated at constant speed over a wider operating range, such as from a "standby" speed to a "cruise" power speed, or a higher speed. In some embodiments, independent control of the VGVs 36A, 36B may allow the spools <NUM>, <NUM> to run at speeds close to maximum power. In some embodiments, independent control of the VGVs 36A, 36B may also allow one of the spools <NUM>, <NUM> to run at high speed while the other one run at low speed.

In use, the engine <NUM> is operated by the controller(s) <NUM> described above to introduce a fuel flow via nozzles <NUM> to the combustor <NUM>. Combustion gases turn turbine sections <NUM>, <NUM> which in turn drive the compressor sections <NUM>, <NUM>. The controller(s) <NUM> control(s) the angular position of VGVs 36A, 36B in accordance with a desired control regime, as will be described further below. The speed of the engine <NUM> is controlled, at least in part, by the delivery of a desired fuel flow rate to the engine, with a lower fuel flow rate causing the engine <NUM> to operate at a lower output speed than a higher fuel flow rate.

Such control strategies may allow for a faster "power recovery" of the engine <NUM> (when an engine is accelerated from a low output speed to a high output speed), possibly because the spools <NUM>, <NUM> can be affected relatively less by their inherent inertia through the described use of spool <NUM>,<NUM> speed control using VGVs <NUM>, as will be further described below. In some embodiments, using the vanes VGV 36A, 36B as described herein, in combination with the use of MFR-based low pressure compressor section <NUM> and/or MFR-based high pressure compressor section <NUM> may provide relatively more air and/or flow control authority and range through the core of the engine <NUM>, and/or quicker power recovery.

Where MFR compressors <NUM> and/or <NUM> of the engines 10A, 10B are provided as described herein, the control of the VGVs 36A and/or VGV 36B provides for improved stability of engine operation. This may be so even where the VGV is operated at an extreme end of its range, such as in the "closed down" position (e.g. at a position of +<NUM> degrees in one embodiment described herein). This control of the VGVs facilitates the ability of the engine to operate at a very low power setting, such as may be associated with a "standby" mode as described further below herein, wherein the compressor of an engine operating in standby mode is operating in a very low flow and/or low pressure ratio regime.

Turning now to <FIG>, illustrated is an exemplary multi-engine system <NUM> that may be used as a power plant for an aircraft (H), including but not limited to a rotorcraft such as a helicopter. The multi-engine system <NUM> may include two or more gas turbine engines 10A, 10B. In the case of a helicopter application, these gas turbine engines 10A, 10B will be turboshaft engines. Control of the multi-engine system <NUM> 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 10A, 10B to reduce an overall fuel burn, particularly during sustained cruise operating regimes, wherein the aircraft 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 system <NUM>. Other phases of a typical helicopter mission 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 conditions (cruise speed and altitude) are substantially stable, the engines 10A, 10B of the system <NUM> may be operated asymmetrically, with one engine operated in a high-power "active" mode and the other engine operated in a lower-power (which could be no power, in some cases) "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 lower-power (which could be no power, in some cases) "standby mode" while the other engine is operated in a high-power "active" mode. Such an asymmetric operation is engaged for a cruise phase of flight (continuous, steady-state flight which is typically at a given commanded constant aircraft cruising speed and altitude). The multi-engine system <NUM> 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 <NUM> is driving in this example a helicopter (H) which may be operated in this asymmetric regime, in which a first of the turboshaft engines (say, 10A) may be operated at high power in an active mode and the second of the turboshaft engines (10B in this example) may be operated in a lower-power (which could be no power, in some cases) standby mode. In one example, the first turboshaft engine 10A 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 common load <NUM>. The second turboshaft engine 10B may be controlled by the controller(s) <NUM> to operate at lower-power or no-output-power conditions to supply substantially none or none of a required power and/or speed demand of the common load <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. The controller(s) <NUM> may comprise a first controller for controlling the first engine 10A and a second controller for controlling the second engine 10B. 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, a single controller <NUM> may be used for controlling the first engine 10A and the second engine 10B.

In another example, an asymmetric operating regime of the engines may be achieved through the one or more controller's <NUM> differential control of fuel flow to the engines, as described in pending application <NUM>/<NUM>,<NUM>, the entire contents of which are incorporated herein by reference. Low fuel flow may also include zero fuel flow in some examples.

Although various differential control between the engines of the engine system <NUM> are possible, in one particular embodiment the controller(s)<NUM> may correspondingly control fuel flow rate to each engine 10A, 10B accordingly. 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 operating regime, the standby engine may be maintained between <NUM>% and <NUM>% less than the fuel flow to the active engine. In some embodiments, 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>% to <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 (say 10B) of the multiengine system <NUM> in a standby mode at a power substantially lower than a rated cruise power level of the engine, and in some embodiments at substantially zero output power and in other embodiments at 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 <NUM> 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 first engine (say 10A) may operate in the active mode while the other engine (say 10B) may operate in the standby mode, as described above. During this operation in the asymmetric regime, if the helicopter (H) needs a power increase (expected or otherwise), the second engine 10B may be required to provide more power relative to the low power conditions of the standby mode, and possibly return immediately to a high- or full-power condition. This may occur, for example, in an emergency condition of the multi-engine system <NUM> powering the helicopter, wherein the "active" engine loses power and 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 operating regime.

Referring to <FIG>, there is illustrated an aircraft H, comprising two engines 10A, 10B. More than two engines 10A, 10B may be present on a same aircraft H. An AOR system <NUM> is configured for operating the engines 10A, 10B of the aircraft H in an asymmetric operating regime.

In some embodiments, the AOR system <NUM> forms part or all of the controller <NUM>, which may be a FADEC, ECU, EEC, or the like. In some embodiments, the AOR system <NUM> is a separate computing device that communicates with a FADEC, an ECU, an EEC, and/or any related accessories.

In order to enter the asymmetric operating regime, both engine and aircraft parameters must meet certain operating conditions associated with the asymmetric operating regime. When one or more of these parameters no longer meet the operating conditions, the asymmetric operating regime may be exited. One or more first sensors 204A are operatively coupled to engine 10A, and one or more second sensors 204B are operatively coupled to engine 10B. The sensors 204A, 204B may be any type of sensor used to measure engine parameters, such as but not limited to speed sensors, pressure sensors, temperature sensors, and the like.

In some embodiments, sensor measurements are transmitted to a monitoring device <NUM> for monitoring the engine parameters and determining whether the engine operating conditions are met or no longer met. Note that not all engine parameters necessarily come from the sensors 204A, 204B. In some embodiments, some of the engine parameters monitored by the monitoring device <NUM> are received from one or more other source, such as but not limited to a FADEC, an ECU, an EEC, or any related accessories that control any aspect of engine performance. In some embodiments, measurements obtained from the sensors 204A, 204B are used to calculate or determine other related engine parameters.

Aircraft parameters are also monitored to determine whether certain aircraft operating conditions for the asymmetric operating regime are met or no longer met. In some embodiments, the aircraft parameters are obtained by the monitoring device <NUM> from aircraft avionics <NUM>. The aircraft avionics <NUM> may include any and all systems related to control and management of the aircraft, such as but not limited to communications, navigation, display, monitoring, flight-control systems, collision-avoidance systems, flight recorders, weather systems, and aircraft management systems. In some embodiments, the aircraft avionics <NUM> perform all monitoring of the aircraft parameters and communicate with the AOR system <NUM> and/or the monitoring device <NUM> when the aircraft operating conditions for the asymmetric operating regime are met or no longer met.

In the embodiment of <FIG>, the monitoring device <NUM> is shown to form part of the AOR system <NUM>. Alternatively, the monitoring device <NUM> is separate therefrom and communicates with the AOR system <NUM> when the engine operating parameters are met and the aircraft operating parameters are met. Alternatively or in combination therewith, monitoring of some or all of the parameters is performed externally to the AOR system <NUM> and involves a pilot monitoring some or all of the parameters.

In some embodiments, the AOR system <NUM> monitors engine and/or aircraft conditions required to enter and exit the asymmetric operating regime. Monitoring may be done continuously or by periodical queries. If at any time the conditions are not respected, the asymmetric operating regime is either exited/aborted or disabled (i.e. cannot be entered).

In some embodiments, the AOR system <NUM> receives an engine availability confirmation when the engine parameters meet the engine operating conditions for the asymmetric operating regime, for example from the monitoring device <NUM> or the cockpit <NUM>. In some embodiments, the AOR system <NUM> receives an aircraft availability confirmation when the aircraft parameters meet the aircraft operating conditions for the asymmetric operating regime, for example from the aircraft avionics <NUM>, the monitoring device <NUM>, or the cockpit <NUM>.

In response to the engine and aircraft availability confirmations, the AOR system <NUM> outputs an availability message to a cockpit <NUM> of the aircraft H. The availability message may be a visual and/or audible message to the cockpit of the aircraft, to provide awareness to the pilot of the possibility of entering the asymmetric operating regime. An audible message may consist in a chime, ring, buzzer, or other suitable sound. Alternatively, or in addition, a visible message may consist in a coloured light, a particular flashing pattern, a dialog box on a screen of a cockpit computer, or any other suitable visual marker. Other approaches are also considered.

Once the pilot is advised, he or she can command entry to the asymmetric operating regime. This may be done using any interface in the cockpit, for example discrete (or other) inputs from a button press or a long hold for added protection against inadvertent selection. The AOR system <NUM> receives the pilot-initiated request to operate the engines 10A, 10B in the asymmetric operating regime and in response, commands the engines 10A, 10B accordingly.

Referring now to <FIG>, there is illustrated a flowchart of an example method <NUM> for operating a multi-engine aircraft. At step <NUM>, an engine availability confirmation for entering the asymmetric operating regime is received. In some embodiments, the method <NUM> comprises a step <NUM> of monitoring the engine parameters and determining whether the operating conditions for the engine to enter the asymmetric operating regime are met. Once the engine operating conditions are met, the engine availability confirmation is issued.

Some example engine operating conditions for entering the asymmetric operating regime are as follows:.

Other engine operating conditions may also be used, alone or in combination with any of the engine operating conditions listed above.

At step <NUM>, an aircraft availability confirmation for entering the asymmetric operating regime is received. The aircraft availability confirmation may be received, for example, from the aircraft avionics <NUM> or from the FADEC. In some embodiments, the method <NUM> comprises a step <NUM> of monitoring the aircraft parameters and determining whether the operating conditions for the aircraft to enter the asymmetric operating regime are met. Once the aircraft operating conditions are met, the aircraft availability confirmation is issued.

Some example aircraft operating conditions for entering the asymmetric operating regime are as follows:.

Other aircraft operating conditions may also be used, alone or in combination with any of the aircraft operating conditions listed above.

At step <NUM>, an availability message is output to a cockpit of the aircraft, such as cockpit <NUM> of aircraft H, in response to receiving the engine availability confirmation and the aircraft availability confirmation. At step <NUM>, a pilot-initiated request to operate the engines in the asymmetric operating regime is received. At step <NUM>, the engines are commanded to operate in the asymmetric operating regime in response to the pilot-initiated request.

In some embodiments, the method <NUM> is performed by the FADEC of the aircraft H. In some embodiments, a portion of the method <NUM> is performed by the FADEC. For example, the set of engine parameters are monitored and the engine availability confirmation is output by the FADEC.

With reference to <FIG>, the method <NUM> may be implemented by a computing device <NUM> as an embodiment of the AOR system <NUM>. The processing unit <NUM> may comprise any suitable devices configured to implement the functionality of the AOR system <NUM> such that instructions <NUM>, when executed by the computing device <NUM> or other programmable apparatus, may cause the functions/acts/steps performed by the system <NUM> as described herein to be executed. The processing unit <NUM> may comprise, for example, any type of general-purpose microprocessor or microcontroller, a digital signal processing (DSP) processor, a central processing unit (CPU), an integrated circuit, a field programmable gate array (FPGA), a reconfigurable processor, other suitably programmed or programmable logic circuits, custom-designed analog and/or digital circuits, or any combination thereof.

The memory <NUM> may include a suitable combination of any type of computer memory that is located either internally or externally to device, for example random-access memory (RAM), read-only memory (ROM), compact disc read-only memory (CDROM), electro-optical memory, magneto-optical memory, erasable programmable read-only memory (EPROM), and electrically-erasable programmable read-only memory (EEPROM), Ferroelectric RAM (FRAM) or the like.

The methods and systems for operating engines of an aircraft in an asymmetric operating regime as described herein may be implemented in a high level procedural or object oriented programming or scripting language, or a combination thereof, to communicate with or assist in the operation of a computer system, for example the computing device <NUM>. Alternatively, the methods and systems for operating engines of an aircraft in an asymmetric operating regime may be implemented in assembly or machine language. The language may be a compiled or interpreted language.

Embodiments of the methods and systems for operating engines of an aircraft in an asymmetric operating regime may also be considered to be implemented by way of a non-transitory computer-readable storage medium having a computer program stored thereon. The computer program may comprise computer-readable instructions which cause a computer, or more specifically the processing unit <NUM> of the computing device <NUM>, to operate in a specific and predefined manner to perform the functions described herein, for example those described in the method <NUM>.

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 present disclosure. Still other modifications which fall within the scope of the present disclosure will be apparent to those skilled in the art, in light of a review of this disclosure.

Claim 1:
A method for operating an aircraft having two or more engines (10A, 10B), the method comprising:
receiving a pilot-initiated request to operate the engines (10A, 10B) in an asymmetric operating regime; and
commanding the engines (10A, 10B) to operate in the asymmetric operating regime in response to the pilot-initiated request,
characterised in that the method further comprises:
receiving an engine availability confirmation when a set of engine parameters meet engine operating conditions for the asymmetric operating regime, wherein a first of the engines (10A, 10B) is in an active mode to provide motive power to the aircraft and a second of the engines (10A, 10B) is receiving fuel and operating in a standby mode to provide substantially no motive power to the aircraft;
receiving an aircraft availability confirmation when a set of aircraft parameters meet aircraft operating conditions for the asymmetric operating regime; and
outputting an availability message to a cockpit (<NUM>) of the aircraft in response to receiving the engine availability confirmation and the aircraft availability confirmation; and
delivering fuel to both the first engine and the second engine when the first engine and the second engine are operating in the asymmetric operating regime.