Patent Description:
Multi-engine helicopters are often provided with two or more engines, typically gas turbine turboshaft engines, connected to a main rotor via a common gearbox. Each of the engines is sized to provide power greater than what is required for cruising using both/all engines. During normal cruise operating regimes, both engines typically operate at similar power output levels (e.g. each engine provides <NUM>% of the total power output provided to the aircraft).

Attempts have been made to operate the engines asymmetrically, that is, operating one engine at a higher power than the other. Doing so may provide improved better fuel efficiency overall. However, the engine operating at lower power needs to be able to rapidly speed back up when called upon, for example in the event of an emergency or if there is an urgent need for increased power. While existing systems may be suitable for their intended purposes, improvements remain desirable.

Documents <CIT> and <CIT> disclose prior art multi-engine systems for aircraft.

Documents <CIT> and <CIT> disclose prior art gas turbine engine systems for aircraft using pressurized air vessels for rapid relaunch of an engine operating at low power.

In one aspect, there is provided a method of operating a multi-engine system of an aircraft having a first engine and a second engine, the method comprising: accumulating compressed air in a pressure vessel external to the first and second engines; operating the first and second engines asymmetrically, by controlling the first engine to operate in an active operating condition providing sufficient power and/or rotor speed for demands of the aircraft, and controlling the second engine to operate in a standby operating condition wherein the second engine produces less power output than the first engine; and in response to a power demand request, accelerating the second engine by introducing the compressed air from the pressure vessel into the second engine at a location upstream of a combustor of the second engine.

The method as defined above and described herein may further include one or more of the following steps/elements, in whole or in part, and in any combination.

In certain embodiments, the method includes accumulating includes extracting bleed air from one or more of the first engine and the second engine, and feeding the bleed air into the pressure vessel.

In certain embodiments, the method includes extracting the bleed air includes extracting the bleed air from one or more of a high pressure compressor of the first engine and a high pressure compressor of the second engine.

In certain embodiments, the method includes extracting includes extracting the bleed air from the second engine prior to the second engine being controlled to produce less power output than the first engine.

In certain embodiments, the method includes accumulating includes gradually extracting the bleed air from both the one or more of the first and second engines from a beginning of a flight mission, after the one or more of the first and second engines is started or after the aircraft has taken-off.

In certain embodiments, the method includes accumulating includes using one or more valves to control the extracting of the bleed air.

In certain embodiments, the method includes accumulating includes extracting additional bleed air from the first engine after the second engine has been controlled to produce less power output than the first engine.

In certain embodiments, the method includes introducing the compressed air from the pressure vessel includes injecting the compressed air into the second engine upstream of a high pressure compressor of the second engine, to rapidly spin a high pressure spool of the second engine having the high pressure compressor.

In certain embodiments, both extracting bleed air to accumulate the compressed air in the pressure vessel and introducing the compressed air into the second engine is carried out via one or more common ports in the second engine.

In certain embodiments, the method includes controlling the second engine to operate in the standby operating condition includes operating the second engine to provide minimal or no propulsive power to the aircraft.

In certain embodiments, second engine is shut down in the standby operating condition, and the accelerating the second engine includes accelerating the second engine from shut down to an active operating condition corresponding to that of the first engine.

In certain embodiments, operating the second engine in the standby operating condition comprises, in flight, operating the second engine to produce substantially no motive power to the aircraft.

In certain embodiments, the method includes, prior to the introducing of the compressed air from the pressure vessel into the second engine at said location upstream of the combustor, the method includes blocking a main gas path of the second engine upstream of said location to prevent back-flow of the compressed air through the main gas path.

In another aspect, there is provided a multi-engine system for an aircraft comprising: a first engine and a second engine driving a common gearbox configured to drive a load, the first engine and the second engine respectively including a first compressor and a second compressor; a pressure vessel external to the first engine and the second engine, the pressure vessel in fluid flow communication with one or both of the first and second compressors via: an accumulation flow path along which compressed air produced by one or both of the first and second compressors flows to reach the pressure vessel; and an injection flow path along which compressed air accumulated within the pressure vessel flows to reach an injection port in at least the second engine; and a valve in fluid communication with the accumulation flow path and the injection flow path, the valve being operable to move between a closed position and an open position, wherein in the open position the compressed air is permitted to flow either along the accumulation flow path or along the injection flow path.

The multi-engine system as defined above and described herein may further include one or more of the following steps/elements, in whole or in part, and in any combination.

In certain embodiments, the system includes an engine controller configured to control the first engine to operate in an active operating condition and to control the second engine to operate in a standby operating condition, wherein in the active operating condition the first engine provides sufficient power and/or rotor speed for demands for the aircraft, and in the standby operating condition the second engine provides provide minimal or no propulsive power to the aircraft.

In certain embodiments, the engine controller is in communication with the valve to control operation thereof, the engine controller operable to position the valve at one or more intermediate positions between the closed position and the open position, to control injection and extraction of compressed air through the accumulation flow path and the injection flow path.

In certain embodiments, the pressure vessel is one of two or more pressure vessels, each of the two or more pressure vessels being in fluid flow communication with a respective one of the first and second engines.

In certain embodiments, the pressure vessel is inflatable.

In certain embodiments, the pressure vessel is thermally insulated, to retain heat of the compressed air accumulated within the pressure vessel.

In certain embodiments, the pressure vessel is cooled to increase a storage capacity thereof.

To maintain clarity of this description, some of the same reference numerals have been used in different embodiments to show features that may be common to the different embodiments.

<FIG> depicts an exemplary multi-engine aircraft <NUM>, which in this case is a helicopter. The aircraft <NUM> may however also be a fixed-wing aircraft. The aircraft <NUM> includes at least two aircraft engines <NUM> (or simply "engines"), labeled in <FIG> as "ENGINE <NUM>" and "ENGINE <NUM>". In a particular embodiment, these two engines are turboshaft gas turbine engines. However, it is to be understood that one or both of the engines may also and/or alternately be hybrid or other types suitable aircraft engines, and may therefore be at least partially electrically driven. The two engines <NUM> may be interconnected by a common gearbox <NUM> (see <FIG>), forming a multi-engine system <NUM> as shown in <FIG> and as will be described in further detail below.

<FIG> illustrates an exemplary multi-engine system <NUM> to be used as a power plant for the aircraft <NUM>, which can include but is not limited to a rotorcraft such as the helicopter (H) of <FIG>. The multi-engine system <NUM> includes multiple engines <NUM>, and in one embodiment includes two or more engines 10A, 10B. In the case of a helicopter application, these engines 10A, 10B may be turboshaft gas turbine engines. They may alternatively be other types of gas turbine engines, or any suitable aircraft engines such as hybrid and/or electrically powered 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 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.

When the aircraft conditions, such as cruise speed and altitude, are substantially stable - such as during a cruise flight segment of the aircraft - 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 "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 alternately referred to herein as an "asymmetric mode", an "asymmetric operating regime" or an "idle cruise regime" (ICR), 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). This operation management may therefore be referred to as a "asymmetric operation mode", or an "idle cruise regime" (ICR), wherein one of the two engines is operated in a low-power or "standby mode" (also referred to herein as a "standby operating condition") while the other engine is operated in a high-power or "active mode". In the standby mode, an engine provides significantly less propulsive power to the aircraft than does the other engine operating in the higher-power, active mode. In certain embodiments, the engine operating in the standby mode may even provide no or almost no propulsive power to the aircraft.

The multi-engine system <NUM> is used in an aircraft, such as but not limited to a helicopter, but also has applications in suitable marine and/or industrial applications or other ground operations, such applications not forming part of the claimed invention.

As shown in <FIG>, the multi-engine system <NUM> includes a first engine 10A and a second engine 10B configured to drive a common load <NUM>. In the depicted embodiment, the engines 10A, 10B are turboshaft gas turbine engines. In some embodiments, the common load <NUM> may comprise a rotary wing of a rotary-wing aircraft. For example, the common load <NUM> may be a main rotor of the helicopter. Depending on the type of the common load <NUM> and on the operating speed thereof, turboshaft engines 10A, 10B may be drivingly coupled to the common load <NUM> via a gearbox <NUM>, which may be any suitable type, such as a speed-changing (e.g., reducing) type. The gearbox <NUM> may have a plurality of transmission shafts <NUM> to receive mechanical energy from respective output shafts 40A, 40B of respective turboshaft engines 10A, 10B to direct at least some of the combined mechanical energy from the plurality of the turboshaft engines 10A, 10B to a common output shaft <NUM> for driving the common load <NUM> at a suitable operating (e.g., rotational) speed. The multi-engine system <NUM> may include a transmission <NUM> driven by the output shaft 40B and driving the rotatable transmission shaft <NUM>. The transmission <NUM> may be controlled to vary a ratio between the rotational speeds of the respective output shaft 40A / 40B and transmission shaft <NUM>.

The multi-engine system <NUM> may be configured, for example, to drive accessories of an associated aircraft in addition to the main rotor. The gearbox <NUM> may be configured to permit the common load <NUM> to be driven by either the first turboshaft engine 10A or the second turboshaft engine 10B, or, by a combination of both the first turboshaft engine 10A and the second turboshaft engine together 10B. A clutch <NUM> may be provided to permit each engine 10A, 10B to be engaged and disengaged with the transmission X, as desired. For example, an engine 10A, 10B running at low- or no-power conditions may be declutched from the transmission if desired. In some embodiments, a conventional clutch may be used.

Referring still to <FIG>, according to the present description the multi-engine system <NUM> driving a helicopter (H) or other aircraft <NUM> may be operated in such an asymmetric manner, in which a first one of the engines (say, 10A) is operated at high power in an active mode and the second one of the engines, for instance the engine 10B in this example, is capable of being operated in a low-power standby mode. In one example, the first 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> and thus the aircraft. The second engine 10B may be controlled by the controller(s) <NUM> to operate at low-power or no-output-power conditions to supply one of substantially little, 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. The term controller as used herein includes any one of: a single controller controlling the engines, and any suitable combination of multiple controllers controlling the engines, including one or more controllers for each engine, so long as the functionality described in this document is provided. 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 <CIT>. 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 in certain embodiments 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, 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 other possible embodiments, the standby engine may be completely shut down, such that no fuel flow is used by this engine when it is operating in the standby mode. In such a case, therefore, the fuel flow (which is zero) to the standby engine is thus <NUM>% less than the fuel flow to the active engine. It is therefore to be understood that the term "standby" mode as used herein is intended to include, in certain embodiments, a complete shut-down state of the standby engine, whereby only one of the two engines (i.e. the active engine) is in operation. A complete shut down of the second engine placed into the standby mode may be particularly interesting given that it completely eliminates fuel consumption by that engine, thereby reducing fuel consumption and thus fuel costs, while also reducing the total flight hours of the engine, thereby reducing related maintenance and operating costs.

In another embodiment, the controller <NUM> may operate one engine of the multiengine system <NUM>, for instance the engine 10B, 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 multi-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.

In use, the first turboshaft engine (say 10A) may operate in the active mode while the second turboshaft engine, such as the engine 10B, may operate in the standby mode, as described above. 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.

During such asymmetric operation, if the helicopter (H) needs a power increase (expected or otherwise), the second turboshaft engine 10B may be required to provide more power relative to the low power conditions of the standby mode, and possibly rapidly to a high-power or full-power condition. In such situations, the engine 10B operating previously in a low power condition must be able to quickly accelerate back up to cruise or full power output levels. This may be required, for example, in the event of an emergency or an urgent need for increased power (e.g. if the pilot requires additional power in order to perform a desired manoeuver). In certain conditions/applications, it may also be possible to completely shut down the second engine 10B engine. However, to do so would require the ability to rapidly re-start the second engine in the event of an emergency or sudden need for more power. Even absent an emergency, it will be desirable to repower the standby engine to exit the asymmetric mode, such that the two engines operate in a normal cruise operating regime whereby both engines operate at similar power output levels (e.g. each engine provides about <NUM>% of the total power output provided to the aircraft).

As will be described in further detail below with reference to <FIG> and <FIG>, the multiengine system <NUM> of the present disclosure includes an air accumulation system <NUM> which is operable to assist an engine operating in a low power condition to quickly accelerate back up to cruise or full power output levels. As will be seen, this is accomplished by introducing external compressed air from outside the engines 10A, 10B into the engine to be accelerated, in order to assist its rapid acceleration back up to cruise or full power. More particularly, compressed air is extracted (e.g. bled off) from one or both engines 10A, 10B of the multi-engine system <NUM> and accumulated in a tank or pressure vessel <NUM> that is external to both engines 10A, 10B. When rapid re-acceleration of an engine operating in a stand-by mode becomes required, for example in response to a power demand, then the compressed air that has been accumulated in the external pressure vessel <NUM> is introduced (or re-introduced, as the case may be) into the cold section of the engine in order to permit it to more rapidly accelerate to a higher power output. The cold section of the engine is understood to be located upstream of a combustor <NUM> of the engine, and more particularly is defined as extending from an air inlet of the engine <NUM> to, but not including, a combustion zone located within the combustion chamber liner of the combustor <NUM>.

Before additional details of the air accumulation system <NUM> and its method of operation are described, each of the engines <NUM>, 10A, 10B of the multi-engine system <NUM> will first be described in further detail, with reference to <FIG> and <FIG>.

As shown in <FIG> and <FIG>, each aircraft engine 10A, 10B (identified simply as engine <NUM> in <FIG>) of the multi-engine system <NUM> may, as in the depicted embodiment, be a turboshaft gas turbine engine generally comprising in serial flow communication a low pressure (LP) compressor section, which will be referred to herein as the LP compressor <NUM> and a high pressure (HP) compressor section, which will be referred below as the HP compressor <NUM> for pressurizing air received via an air inlet <NUM>. The air compressed by the LP compressor <NUM> and by the HP compressor <NUM> is fed to a combustor <NUM> in which the compressed air is mixed with a fuel flow, delivered to the combustor <NUM> via fuel nozzles <NUM> from a suitable fuel system, and ignited for generating a stream of hot combustion gases. A high pressure turbine section, which will referred to herein as the HP turbine <NUM>, extracts energy from the combustion gases. A low pressure turbine section, which will be referred to herein as the LP turbine <NUM> is located downstream of the HP turbine <NUM> for further extracting energy from the combustion gases and driving the LP compressor <NUM>. The combustion gases are then exhausted by an exhaust outlet <NUM>. The LP compressor <NUM> may include one or more compression stages, and the HP compressor <NUM> may include one or more compression stages.

In the embodiment shown, the turboshaft engine <NUM> includes a low-pressure spool, referred to below as LP spool <NUM>, and a high-pressure spool, referred to below as a HP spool <NUM>. The LP spool <NUM> includes a low-pressure shaft, referred to below as LP shaft <NUM>. The HP spool <NUM> includes a high-pressure shaft, referred to below as HP shaft <NUM>. The HP turbine <NUM> is drivingly engaged to the HP compressor <NUM> via the HP shaft <NUM>. The LP turbine <NUM> is drivingly engaged to the LP compressor <NUM> via the LP shaft <NUM>. The HP spool <NUM>, and the components mounted thereon, are configured to rotate independently from the LP spool <NUM> and from the components mounted thereon. These two spools may thus rotate at different speeds about an engine central axis <NUM>. The HP shaft <NUM> and the LP shaft <NUM> may be concentric. In the embodiment shown, the HP shaft <NUM> extends around the LP shaft <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.

Although the gas turbine engine <NUM> as shown in <FIG> is a multi-spool engine, having separate LP spool <NUM> and HP spool <NUM>, it is to be understood that in an alternate embodiment, one or more of the engines 10A, 10B of the multi-engine system <NUM> may have a single spool architecture, which is often the case for auxiliary power unit (APU) engines used in aircraft.

In the embodiment shown, the HP compressor <NUM> rotates at the same speed as the HP turbine <NUM>. And, the LP compressor <NUM> rotates at the same speed as the LP turbine <NUM>. However, this may not be the case if transmission(s) are provided on the LP spool <NUM> and HP spool <NUM> to create speed ratios between the interconnected compressors and turbines. This may increase or decrease rotational speeds of the compressors relative to that of the turbines. Any suitable transmissions may be used for this purpose.

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 be provided to vary a ratio between rotational speeds of the low pressure shaft <NUM> and the output shaft <NUM>. 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> of the engine <NUM> (and therefore of each of the engines 10A, 10B of <FIG> and <FIG>) may have a bleed valve <NUM> (shown schematically) configured to selectively bleed air from the LP compressor <NUM>, via an associated bleed port, according to a desired control regime of the engine <NUM>, for example to assist in control of compressor stability.

As mentioned, the HP compressor <NUM> is configured to independently rotate from the LP compressor <NUM> by virtue of their mounting on different engine spools. The HP compressor <NUM> may include one or more compression stages, such as a single stage, or two or more stages as shown in more detail in <FIG>. It is contemplated that the HP compressor <NUM> may include any suitable type and/or configuration of stages. The HP compressor <NUM> 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 <NUM>, via an associated bleed port, according to a desired control regime of the engine <NUM>, for example to assist in control of compressor stability.

One or both of the bleed valve <NUM> located within the LP compressor <NUM> and the HP compressor <NUM> may also serve as accumulator injection/bleed ports 86A, 86B, as shown in <FIG> and described further below as part of the air accumulation system <NUM>. Alternately, the engine <NUM> (and thus the two engines 10A, 10B) may each include regular bleed valve <NUM> and/or <NUM>, in addition to at least one accumulator injection/bleed ports 86A, 86B as described below. In one particular embodiment, the accumulator injection/bleed ports 86A, 86B may be located within the HP compressor <NUM>, for example just upstream of the combustor (at an engine station often referred to as "P3", where the static pressure of the compressed air produced by the compressor(s) of the engine is the highest). Regardless of the chosen configuration, both the standard compressor bleed valves <NUM>, <NUM> and the accumulator injection/bleed ports 86A, 86B are located within the "cold section" of the engine, that is upstream of the combustor <NUM> within the engine. Further details of the accumulator injection/bleed ports 86A, 86B will be provided below.

The expression "upstream of the combustor" as used herein, particularly with reference to the location at which compressed air from the pressure vessel <NUM> is introduced into the second engine, is therefore understood to mean anywhere within the cold section of the engine, between the air inlet <NUM> and the combustor <NUM> (and more precisely the combustion zone contained within the combustion chamber liner(s) of the combustor). This includes the air plenums, cavities or passages which may surround the combustor <NUM>, even if some or all of such plenums, cavities or passages are axially located at or forward of the combustion chamber liner itself), at which location(s) the pressure of the compressed air (e.g. P3 air) is the highest.

In use, suitable 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>, controls operation of the engine <NUM>. The FADEC(s) may be provided as for example conventional software and/or hardware, so long as the FADEC(s) is/are configured to perform the various control methods and sequences as described in this document. 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>.

Referring still to <FIG> and <FIG>, the turboshaft engine <NUM> may also include variable guide vanes (VGVs) <NUM>, 36A, 36B. As seen in <FIG>, a first set of VGVs 36A is located upstream of the LP compressor <NUM>, and a second set of VGVs 36B is located upstream of the HP compressor <NUM>. The VGVs <NUM> may be independently controlled by suitable one or more controllers <NUM>, as described above. The VGVs <NUM> may direct inlet air to the corresponding stage of the LP compressor <NUM> and of the HP compressor <NUM>. The VGVs <NUM> may be operated to modulate the inlet air flow to the compressors in a manner which may allow for improved control of the output power of the turboshaft engine <NUM>, as described in more detail below. The VGVs <NUM> may be provided with any suitable operating range. In some embodiments, VGVs <NUM> 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 air flow. In a more specific embodiment, the VGVs <NUM> 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 VGVs <NUM> may be configured for a similar range of positions, or other suitable position range.

In some embodiments, the first set of VGVs 36A upstream of the LP compressor <NUM> may be mechanically decoupled from the second set of VGVs 36B upstream of the HP compressor <NUM> and downstream of the LP compressor <NUM>, having no mechanical link between the two sets of VGVs to permit independent operation of the respective stages. The VGVs <NUM> may be operatively controlled by the controller(s) <NUM> described above, to be operated independently of each other. Indeed, the turboshaft engine <NUM> is 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 <NUM> 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 VGVs <NUM> 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 <NUM> may allow the spools <NUM>, <NUM> to run at speeds close to maximum power. In some embodiments, independent control of the VGVs <NUM> 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 turboshaft engine <NUM> is operated by the controller(s) <NUM> described above to introduce a fuel flow via the nozzles <NUM> to the combustor <NUM>. Combustion gases turn the HP turbine <NUM> and the LP turbine <NUM> which in turn drive the HP compressor <NUM> and the LP compressor <NUM>. The controller(s) <NUM> control(s) the angular position of VGVs <NUM> 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 turboshaft engine <NUM> to operate at a lower output speed than a higher fuel flow rate.

Referring now to <FIG>, the multi-engine system <NUM> also includes an air accumulation system <NUM> which includes an air tank or pressure vessel <NUM>, which is external to both the first and second engines 10A, 10B, and serves to receive and retain therein compressed air that is bled off from one or both of the engines 10A, 10B. As will be described in further detail, once the pressure vessel <NUM> is filled, either partially or fully, with compressed air, it is stored therein until such as time as it is needed for injection, or re-injection, into one of the two engines 10A, 10B that is operating in a standby mode as described above. Accordingly, in certain embodiments the standby engine can be completely shut down, once there is sufficient air pressure in the pressure vessel <NUM>. The external pressure vessel <NUM> will have an internal volume that is sufficient to guarantee engine start when the compressed air stored in the pressure vessel is injected back into the standby engine when required for an emergency re-start and/or rapid acceleration.

The pressure vessel <NUM> is external to both engines 10A, 10B, and may be physically located either within the overall multi-engine system package or may alternately be located elsewhere within the aircraft. While it will be appreciated that the pressure vessel <NUM> must be suitable to hold compressed air having a pressure corresponding to the air pressures generated by the engines, the exact construction of the pressure vessel <NUM> may be selected to be suitable for the purposes described herein. In certain embodiments, the pressure vessel <NUM> may be inflatable, such that it remains lightweight and when it is empty (and thus is deflated) it will take up relatively little space within the aircraft. Additionally, in certain embodiments, the external pressure vessel <NUM> may also be thermally insulated, so as to help to retain the heat in the extracted air that was generated when it was compressed. Alternately still, the pressure vessel <NUM> may also be cooled, either passively or actively using a suitable heat-exchanger for example. This may be useful so as to increase the storage capacity, in terms of total mass of the compressed air accumulated therein at a given pressure, thereby allowing - when the accumulated air is injected back into the re-acerbating engine - more fuel flow into the combustion chamber before reaching the hot section temperature limit.

Thus, the pressure vessel <NUM> is fluidly connected to each of the first engine 10A and the second engine 10B by one or more air conduits 82A, 82B which define each one or more airflow paths between an internal cavity of the pressure vessel <NUM> and the cold section of a respective one of the engines 10A, 10B. More particularly, each of the air conduits 82A, 82B may provide a first, accumulation, flow path used to transport compressed air from the cold section of the respective engines 10A, 10B (i.e. from the compressor(s), upstream of the combustors) to the pressure vessel <NUM>, and a second, injection, flow path used to transport compressed air from the pressure vessel <NUM> to the cold section of the respective engines 10A, 10B. It is to be understood, however, that a single conduit 82A, 82B can be used to direct air to and from each respective engine 10A, 10B, such that air flows out of the engines and into the engines through the same passage, line or conduit. Although the depicted embodiment shows only a single pressure vessel <NUM>, it is to be understood that two or more pressure vessels may also be used. For example, each engine 10A, 10B may have its own dedicated pressure vessel <NUM>, within which air is accumulated and stored for sub-sequent delivery to is respective engine.

Each of the first and second engines 10A, 10B includes a respective accumulator injection/bleed port 86A, 86B located within the cold section of the engine. These ports 86A, 86B may also be simply referred to herein as "bleed ports" even if they may also serve to inject compressed air flow into the engine in addition to or in stead of being used to bleed compressed air off from the engine for filling the pressure vessel <NUM>.

In the depicted embodiment, the accumulator bleed ports 86A, 86B are located within the HP compressor 14A, 14B, just upstream of their respective combustors 16A, 16B. Thus, the air bled off via the accumulator bleed ports 86A, 86B, when one or more associated control valves 84A, 84B are opened, and subsequently stored in the pressure vessel <NUM>, will have a high pressure. These control valves 84A, 84B may also be, or include, one-way valves such as to prevent flow of the compressed air in an unwanted direction.

The control valves 84A, 84B are thus operable to control the flow of air to and from the pressure vessel <NUM>, and thus to control the injection of the pressurized air contained within the pressure vessel <NUM> into one of the engines, when it becomes desirable to do so in the manner described herein. In the depicted embodiment, a first control valve 84A is disposed in flow communication with the first air conduit 82A and a second control valve 84B is disposed in flow communication with the second air conduit 82B, wherein the first and second control valves 84A, 84B can be used to either allow or prevent flow between the cold sections of the engines 10A, 10B and the pressure vessel <NUM>. In an alternate embodiment, a single, multi-port, valve may be able to be used in place of the two separate control valves 84A, 84B, provided that independent flow within each of the first and second air conduits 82A, 82B can be separately controlled.

As such, compressed air can be drawn or bled off from only one or both of the engines 10A, 10B via one or both of the accumulator bleed ports 86A, 86B located within the cold section of the engines, as controlled by the one or both control valves 84A, 84B prior to the shut-down (or reduction of power output) of one of the two engines into a low-power standby mode. The air extracted in this manner is accordingly directed through the respective one(s) of the first and second air conduits 82A, 82B (depending of course on which port is opened to allow bleed flow therethrough) to the pressure vessel <NUM>, where the compressed air is accumulated, and retained future use.

This accumulated of the compressed air within the pressure vessel <NUM> may, in one particular embodiment, be done gradually, while the engine from which the air is extracted is operating at standard cruise or high power (e.g. a high pressure ratios).

For example, in one operation scenario, high pressure air is drawn off from the HP compressor 14B of the second engine 10B, via its accumulator bleed port 86B, right from the beginning of a normal flight mission, as the engine 10B starts, then idles, then gets to take-off power. Because the pressure at this location of the engine will increase over a sufficiently long time interval, and the volume of compressed air extracted via accumulator bleed port 86B and fed to the pressure vessel <NUM> remains relative small (i.e. relative to the total volume of air flowing through the HP compressor 14B) throughout, this small extraction of flow from the HP compressor 14B will not be "perceptible" to the overall engine operability - in other words, this relatively small amount of compressed air that is extracted for feeding to the pressure vessel <NUM> will have little to no impact on the overall performance and operability of the engine.

However, if the compressed air used to fill the pressure vessel <NUM> is extracted from either engine while the engine or engines are already operating in a high power regime (e.g. following an emergency re-start or acceleration for instance), then it may be desirable to bleed this air off slowly, again via the accumulator bleed ports 86A, 86B, such as to prevent compressor surge and limit performance penalties. For that purpose, a servo valve (controlled via the FADEC for instance) could be used. Such a servo valve may form part of the control valve or valves 84A, 84B, or may be an additional servo valve(s) in other instances.

Further, when the external pressure vessel <NUM> has been "charged" (i.e. filled, either partially or fully, with the accumulated compressed air), and the second engine 10B has been shut down or placed into a low power operating condition, it is also possible to use the first engine 10A (e.g. the active engine operating a regular cruise or full power) to add small amounts of additional compressed air into the external pressure vessel <NUM> - e.g. either periodically or using a small but continuous trickle flow. In this manner, the compressed air within the pressure vessel <NUM> can be "topped up" using air extracted from the active engine. This may help to keep the pressure within the external pressure vessel <NUM> at a desired level and/or may be used to replenish any lost or used accumulated compressed air, which may be particularly useful if there is a cooling mechanism in place that could cause the pressure within the pressure vessel <NUM> to drop. In this manner, both the volume and pressure of the accumulated compressed air within the external pressure vessel can be maintained at desired levels, so as to ensure that if/when this accumulated air is needed it will be sufficient for injection into the standby engine for emergency start-up and/or rapid acceleration of the standby or shut-down engine.

In one possible scenario, at the end of aircraft take-off, compressed air from one engine (e.g. second engine 10B) may have been slowly bled off and fed to the external pressure vessel <NUM> such that this external pressure vessel <NUM> is fully "charged" with compressed air. The second engine 10B can then be shut down or placed in very low power standby mode, as described above, once the aircraft has reached its flight cruise phase, with the accumulated compressed air stored in the external pressure vessel <NUM>. When/if needed for an emergency re-start or power recovery of the second engine 10B, the accumulated compressed air retained in the pressure vessel <NUM> is then fed back into the compressor of the shut-down engine 10B to enable a rapid re-start thereof.

With specific respect to the introduction of the compressed air from the pressure vessel <NUM> into the cold section of the engine operating in a standby mode, when required for the purposes of executing an emergency re-start and/or a rapid acceleration of the low power standby engine, this injection of the accumulated compressed air is optimally done somewhere between the inlet 22A, 22B of the engine in question and the combustor 16A, 16B thereof (i.e. somewhere within the cold section of the engine. The precise location that this accumulated compressed air is injected into the cold section may vary, and is selected depending on the particular engine architecture (e.g. number of compressor stages, compressor inertia, etc.) and the available compressed air (volume, pressure) as well as the desirable engine power recovery time. However, in one particular embodiment, the injection of the compressed air from the pressure vessel <NUM> into the standby engine, when it needs to be rapidly re-accelerated and/or started, may be done at or just upstream of the HP compressor 14A, 14B of the engine so as to rapidly spin-up the HP spool <NUM> (see <FIG> and <FIG>). In such an embodiment, therefore, the injection of the accumulated compressed air stored in the external pressure vessel <NUM> may be done at a point upstream of the HP compressor <NUM>, 14A, 14B (e.g. impeller), which will cause the HP impeller <NUM>, 14A, 14B to rotate as the accumulated compressed air flows therethrough. This may help to facilitate the re-start. In an alternate embodiment, however, the accumulated compressed air from the pressure vessel <NUM> can be injected into the standby engine at a location that is just downstream end of the HP compressor <NUM>, 14A, 14B and immediately upstream of the combustor <NUM>, 16A, 16B. Additionally, in one particular embodiment, the air accumulated in the pressure vessel <NUM> can also be injected or re-injected back into the standby engine via the same accumulator bleed ports 86A, 86B described above, which may have been used to extract the compressed air in the first place.

Regardless of the specific location at which the accumulated compressed air from the pressure vessel <NUM> is injected into the engine that needs to be rapidly re-started or accelerated to full power, an active control system <NUM> may be provided to control this re-injection of the external compressed air from the pressure vessel <NUM>. The actively controlled reinjection system <NUM>, which is in communication with the engine controller <NUM> (also referred to herein simply as FADEC <NUM>) (see <FIG>) and is controlled thereby, may for example include at least one active valve (which may be, for example, one for each engine) and/or a pressure regulating valve which are collectively operable to control the injection of the compressed air. For example, the actively controlled reinjection system <NUM> may be configured to permit a more constant flow of the compressed air being injected back into the standby engine, and thus may avoid having too much compressed air injected back into the engine too early, thus avoiding all of the accumulated compressed air being used up and consequently ensuring that there is sufficient compressed air for use in a later phase of the start and re-acceleration sequence of the engine as it starts back up. The actively controlled reinjection system <NUM> may also include, or alternately form part of, the control valve or valves 84A, 84B and/or servo valves, all controlled via the FADEC, used for the extraction of the accumulated compressed air in the first place. Thus, in certain embodiments, a single flow control system (e.g. the system <NUM>) that is itself controlled by the FADEC, may be used to control both compressed air extraction from one or both engines, and the injection of the accumulated compressed from the external pressure vessel <NUM> into the standby engine when required for a rapid re-acceleration thereof.

When the accumulated compressed air is introduced back into the shut-down engine for the purposes of emergency restart or acceleration, the regular starter of the engine in question may still be used to re-start the engine. However, in certain situations, the flight conditions (e.g. altitude, temperature, Mach number, etc. and the operating status of each of the engines (e.g. out of usage, shut-down, sub-idle, idle, low or high power) may permit the engine to be re-started without requiring the use of the regular starter.

Additionally, a suitable flow blocking system 88A, 88B may also be provided within each engine 10A, 10B to prevent the accumulated high pressure compressed air within the pressure vessel <NUM>, when re-injected back into the shut-down engine, from flowing backwards through the engine (e.g. from the re-injection point within the cold section of the engine upstream through the main gas path of the engine core towards the air inlet of the engine, rather than downstream to the combustor). This may be achieved, for example, using one or more one-way valves or another backflow prevention system - e.g. valves or other devices to block the exits of HP compressor diffuser pipes, for example when the injection point is done past that component. Additionally, other engine components, such as variable guides vanes for instance, could also be used to block the compressor gas path when the compressor air is injected downstream that component and upstream to one or more compressor stage.

Solutions for improving operating fuel efficiencies of multi-engine systems, such as the present multi-engine system <NUM> as described herein, have been proposed in other disclosures. One such example is described in <CIT>, which describes a specific control logic for achieve fuel economy of a multi-engine system. In this document, a "breathing cycle" of an engine is described, wherein rotor inertia is used for energy accumulation and then re-used during a suitable point in the breathing cycle. It is of note that the compressed air accumulated in the external pressure vessel <NUM> of the present disclosure may be used in a similar manner, namely in order to extend the time of the "breathing phase" when fuel flow is low.

Referring now to <FIG>, a method of operating the multi-engine system <NUM> of an aircraft is shown at <NUM>. In accordance with the present description, there is therefore provided a method <NUM> of operating a multi-engine system <NUM> of an aircraft <NUM>, such as helicopter H, having a first engine 10A and a second engine 10B.

The method <NUM> includes: accumulating compressed air in a pressure vessel <NUM> that is external to the first and second engines 10A, 10B, at step <NUM>; operating the first and second engines asymmetrically, by controlling the first engine 10A to operate in an active, or high power, operating condition wherein it provides sufficient power and/or rotor speed to meet the demands of the aircraft, and controlling the second engine 10B to operate in a standby operating condition wherein the second engine 10B produces less power output than the first engine 10A, at <NUM>; and, in response to a power demand request, accelerating the second engine 10B by introducing the compressed air from the pressure vessel <NUM> into the second engine 10B at a location therein upstream of a combustor 16B of the second engine 10B, at <NUM>. Because this location at which the compressed air is introduces is within the cold section of the engine, it is necessarily downstream of an air inlet 22B of the second engine.

Step <NUM> of the method <NUM> may include extracting bleed air from one or more of the first engine 10A and the second engine 10B, and feeding the bleed air into the pressure vessel <NUM>. Additionally, extracting the bleed air may include extracting the bleed air from one or more of a high pressure compressor 14A of the first engine 10A and a high pressure compressor 14B of the second engine 10B. Extracting the bleed air may also include extracting the bleed air from the second engine 10B prior to the second engine being controlled to produce less power output than the first engine 10A.

Step <NUM> may also include gradually extracting the bleed air from both the one or more of the first and second engines from a beginning of a flight mission, after the one or more of first and second engines is started or after the aircraft has taken-off. Additionally, accumulating the compressed air may include using one or more valves, such as servo control valves operated by the engine controller or FADEC <NUM>, to control the extracting of the bleed air. Accumulating the compressed air may also include extracting additional bleed air from the first engine 10A, after the second engine 10B has been controlled to produce less power output than the first engine.

At step <NUM> of the method <NUM>, introducing the compressed air from the pressure vessel <NUM> may include injecting the compressed air into the second engine 10B upstream of the high pressure compressor 14B of the second engine 10B, such as to rapidly spin (e.g. to "spin-up") a high pressure spool <NUM> of the second engine 10B (the high pressure spool <NUM> including the high pressure compressor 14B mounted thereon).

At step <NUM>, both extracting the bleed air to accumulate the compressed air in the pressure vessel and introducing the compressed air into the second engine may be carried out via one or more common ports in the second engine 10B, such as the port 86B for example.

Step <NUM> may include operating the second engine 10B to provide minimal or no propulsive power to the aircraft.

Step <NUM> may include accelerating the second engine 10B from the standby operating condition to an active operating condition corresponding to that of the first engine.

Operating the first and second engines 10A, 10B asymmetrically, as in step <NUM>, may be performed during a cruise flight segment of the aircraft.

At step <NUM>, prior to the introducing of the compressed air from the pressure vessel <NUM> into the second engine 10B at said location upstream of the combustor 16B, there may also include blocking a main gas path of the second engine 10B upstream of said location to prevent back-flow of the compressed air through the main gas path of the second engine.

The method <NUM> may be used for example to operate the multi-engine engine system <NUM> during, in one example, a cruise flight segment which may be described as a continuous, steady-state flight segment which is typically at a relatively constant cruising speed and altitude. In a typical cruise mode for twin-engine helicopters, both engines provide -<NUM>% of the cruise power demand of the helicopter. This power level of each engine (~<NUM>% of total power required by the helicopter) may be referred to herein as a "cruise power level".

Step <NUM> of the method <NUM> can include using the engine controller <NUM>, such as a full authority digital control (FADEC) <NUM> to control the engines 10A, 10B to operate asymmetrically. The FADEC <NUM> may thus determine that the aircraft is in a suitable condition for entering asymmetric mode, for example during a cruise flight segment. The FADEC <NUM> may accelerate one engine (say 10A) of the multiengine system <NUM> from a cruise power level into an active engine mode, in which the first engine may provide a higher cruise power level and sufficient power to satisfy substantially all or all (e.g. <NUM>% or higher) of a helicopter power or rotor speed demand. The FADEC <NUM> may then decelerate another engine (say 10B) of the multi-engine system <NUM> to operate in a standby mode (as described herein) at a power substantially lower than cruise power level, and in some embodiments at zero output power (i.e. the standby engine is completely shut down) and in other embodiments less than <NUM>% output power relative to a reference power (provided at a reference fuel flow).

With reference to <FIG>, an example of a computing device <NUM> is illustrated. For simplicity only one computing device <NUM> is shown but the system may include more computing devices <NUM> operable to exchange data. The computing devices <NUM> may be the same or different types of devices. The controller <NUM> may be implemented with one or more computing devices <NUM>. Note that the controller <NUM> can be implemented as part of a full-authority digital engine controls (FADEC) or other similar device, including electronic engine control (EEC), engine control unit (ECU), electronic propeller control, propeller control unit, and the like. In some embodiments, the controller <NUM> is implemented as a Flight Data Acquisition Storage and Transmission system, such as a FASTTM system. The controller <NUM> may be implemented in part in the FASTTM system and in part in the EEC. Other embodiments may also apply.

The computing device <NUM> comprises a processing unit <NUM> and a memory <NUM> which has stored therein computer-executable instructions <NUM>, which serve to control the engine system in the manner described herein. More particularly, the processing unit <NUM> may comprise any suitable devices configured to implement the method <NUM> such that instructions <NUM>, when executed by the computing device <NUM> or other programmable apparatus, may cause the functions/acts/steps performed as part of the method <NUM> as described herein to be executed.

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, magnetooptical memory, erasable programmable read-only memory (EPROM), and electricallyerasable programmable read-only memory (EEPROM), Ferroelectric RAM (FRAM) or the like.

The methods and systems for operating the multi-engine system 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 the multi-engine system may be implemented in assembly or machine language. The language may be a compiled or interpreted language. Program code for implementing the methods and systems for operating the multi-engine system may be stored on a storage media or a device, for example a ROM, a magnetic disk, an optical disc, a flash drive, or any other suitable storage media or device. The program code may be readable by a general or special-purpose programmable computer for configuring and operating the computer when the storage media or device is read by the computer to perform the procedures described herein. Embodiments of the methods and systems for operating the multi-engine system 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 embodiments described in this document provide non-limiting examples of possible implementations of the present technology. Upon review of the present disclosure, a person of ordinary skill in the art will recognize that changes may be made to the embodiments described herein without departing from the scope of the present technology. Yet further modifications could be implemented by a person of ordinary skill in the art in view of the present disclosure, which modifications would be within the scope of the present technology.

Claim 1:
A method of operating a multi-engine system (<NUM>) of an aircraft (<NUM>) having a first engine (10A) and a second engine (10B), the method comprising:
accumulating compressed air in a pressure vessel (<NUM>) external to the first and second engines (10A, 10B);
operating the first and second engines (10A, 10B) asymmetrically, by controlling the first engine (10A) to operate in an active operating condition providing sufficient power and/or rotor speed for demands of the aircraft (<NUM>), and controlling the second engine (10B) to operate in a standby operating condition wherein the second engine (10B) produces less power output than the first engine (10A); and
in response to a power demand request, accelerating the second engine (10B) by introducing the compressed air from the pressure vessel (<NUM>) into the second engine (10B) at a location upstream of a combustor (16B) of the second engine (10B).