Patent Description:
Gas turbine engines featuring electric machines operable as both motors and generators are known, such as those used for more electric aircraft. Whilst such engines may include a plurality of such electric machines for redundancy, they are only coupled to one of the spools. For example, one known configuration includes such electric machines coupled to the high-pressure spool of a twin-spool turbofan. Another includes such electric machines coupled to the intermediate-pressure spool of a triple-spool turbofan.

An issue with such a configuration is that for a given electrical power demand, there is no choice but to supply it from the single spool in the engine. Thus the design of the turbomachinery must be capable of accommodating all possible electrical power demands throughout the operational envelope, which inevitably leads to compromise.

It has therefore been proposed to include an electric machine on two or more shafts of a multi-spool engine. Whilst numerous documents put forward candidates for the optimal physical implementations of such an architecture, few make reference to the optimal control strategy to operate such configurations.

European Patent Application Publication <CIT> describes an electrical generation system for a jet engine. The system includes first and second generators connected to low- and high-pressure shafts of the engine. A power control device controls first and second power regulation devices to ensure the engine has a specified or greater amount of surge margin.

European Patent Application Publication <CIT> describes an integrated gas turbine engine, electrical power subsystem and thermal management subsystem for aerospace application. First and second motor/generators are coupled to low and high pressure shafts of the engine. An engine controller receives an engine thrust demand and electrical power demand, and determines total system power demand and the amount of power sharing between the first and second motor/generators to meet the thrust and electrical power demand.

United States Patent Application Publication <CIT> relates to systems and method for monitoring surge conditions. A method includes detecting a surge condition through vibration signals measured at at least one location in a turbomachine. Detecting a surge condition includes determining a ratio-metric indicator by comparing a blade frequency band to a reference frequency band.

The scope of protection is defined in the appended claims, to which reference should now be made.

In an aspect, there is provided a gas turbine engine for an aircraft, comprising:.

In another aspect, there is provided a method of controlling a gas turbine engine for an aircraft of the type having a high-pressure (HP) spool comprising an HP compressor and a first electric machine driven by an HP turbine, a low-pressure (LP) spool comprising an LP compressor and a second electric machine driven by an LP turbine, a combustor, and an electrical energy storage unit, the method comprising:.

Embodiments will now be described by way of example only with reference to the accompanying drawings, which are purely schematic and not to scale, and in which the claimed invention in described in <FIG>, in and which:.

The present invention is described in the context of two-spool, geared turbofan engine architectures. However, it will be apparent to those skilled in the art that the principles of the present invention may be applied to other engine types including gas turbines with two or more spools, such as direct-drive turbofans, turboprops, or open rotor engines.

A general arrangement of an engine <NUM> for an aircraft is shown in <FIG>, with an equivalent block diagram of the main components thereof being presented in <FIG>.

In the present embodiment, the engine <NUM> is a turbofan, and thus comprises a ducted fan <NUM> that receives intake air A and generates two airflows: a bypass flow B which passes axially through a bypass duct <NUM> and a core flow C which enters a core gas turbine.

The core gas turbine comprises, in axial flow series, a low-pressure compressor <NUM>, a high-pressure compressor <NUM>, a combustor <NUM>, a high-pressure turbine <NUM>, and a low-pressure turbine <NUM>.

In use, the core flow C is compressed by the low-pressure compressor <NUM> and is then directed into the high-pressure compressor <NUM> where further compression takes place. The compressed air exhausted from the high-pressure compressor <NUM> is directed into the combustor <NUM> where it is mixed with fuel and the mixture is combusted. The resultant hot combustion products then expand through, and thereby drive, the high-pressure turbine <NUM> and in turn the low-pressure turbine <NUM> before being exhausted to provide a small proportion of the overall thrust.

The high-pressure turbine <NUM> drives the high-pressure compressor <NUM> via an interconnecting shaft <NUM>. The low-pressure turbine <NUM> drives the low-pressure compressor <NUM> via an interconnecting shaft <NUM>. Together, the high-pressure compressor <NUM>, interconnecting shaft <NUM> and high-pressure turbine <NUM> form part of a high-pressure spool of the engine <NUM>. Similarly, the low-pressure compressor <NUM>, interconnecting shaft <NUM> and low-pressure turbine <NUM> form part of a low-pressure spool of the engine <NUM>.

The fan <NUM> is driven by the low-pressure turbine <NUM> via a reduction gearbox in the form of a planetary-configuration epicyclic gearbox <NUM>. Thus in addition to the low-pressure compressor <NUM>, the interconnecting shaft <NUM> is also connected with a sun gear <NUM> of the gearbox <NUM>. The sun gear <NUM> is meshed with a plurality of planet gears <NUM> located in a rotating carrier <NUM>, which planet gears <NUM> are in turn are meshed with a static ring gear <NUM>. The rotating carrier <NUM> is connected with the fan <NUM> via a fan shaft <NUM>.

It will be appreciated however that a different number of planet gears may be provided, for example three planet gears, or six, or any other suitable number. Further, it will be appreciated that in alternative embodiments a star-configuration epicyclic gearbox may be used instead.

In order to facilitate electrical generation by the engine <NUM>, a first electric machine <NUM> capable of operating both as a motor and generator (hereinafter, "HP motor-generator") forms part of the high-pressure spool and is thus connected with the interconnecting shaft <NUM> to receive drive therefrom. In the present embodiment, this is implemented using a tower-shaft arrangement of the known type. In alternative embodiments, the HP motor-generator <NUM> may be mounted coaxially with the turbomachinery in the engine <NUM>. For example, the HP motor-generator <NUM> may be mounted axially in line with the duct <NUM> between the low- and high-pressure compressors.

Similarly, a second electric machine <NUM> capable of operating both as a motor and generator (hereinafter, "LP motor-generator") forms part of the low-pressure spool and is thus connected with the interconnecting shaft <NUM> to receive drive therefrom. In the present embodiment, the LP motor-generator <NUM> is mounted in the tailcone <NUM> of the engine <NUM> coaxially with the turbomachinery. In alternative embodiments, the LP motor-generator <NUM> may be located axially in line with low-pressure compressor <NUM>, which may adopt a bladed disc or drum configuration to provide space for the LP motor-generator <NUM>.

It will of course be appreciated by those skilled in the art that any suitable location for the HP and LP motor-generators may be adopted.

In the present embodiment, the HP and LP motor-generators are permanent-magnet type motor-generators. Thus, the rotors of the machines comprise permanent-magnets for generation of magnetic fields for interaction with the stator windings. Extraction of power from, or application of power to the windings is performed by a power electronics module (PEM) <NUM>. In the present embodiment, the PEM <NUM> is mounted on the fancase <NUM> of the engine <NUM>, but it will be appreciated that it may be mounted elsewhere such as on the core gas turbine, or in the vehicle to which the engine <NUM> is attached, for example.

Control of the PEM <NUM> and of the HP and LP motor-generator is in the present example performed by an electronic engine controller (EEC) <NUM>. In the present embodiment the EEC <NUM> is a full-authority digital engine controller (FADEC), the configuration of which will be known and understood by those skilled in the art. It therefore controls all aspects of the engine <NUM>, i.e. both of the core gas turbine and the motor-generators <NUM> and <NUM>. In this way, the EEC <NUM> may holistically respond to both thrust demand and electrical power demand.

An embodiment of the overall system will be described with reference to <FIG>, and the control software architecture will be described with reference to <FIG> and <FIG>. The various control strategies implemented in response to various engine operational phenomena will be described with reference to <FIG>.

Various embodiments of the engine <NUM> may include one or more of the following features.

It will be appreciated that instead of being a turbofan having a ducted fan arrangement, the engine <NUM> may instead be a turboprop comprising a propeller for producing thrust.

The low- and high-pressure compressors <NUM> and <NUM> may comprise any number of stages, for example multiple stages. In addition to, or in place of, axial stages, the low-or high-pressure compressors <NUM> and <NUM> may comprise centrifugal compression stages.

The low- and high-pressure turbines <NUM> and <NUM> may also comprise any number of stages.

The fan <NUM> may have any desired number of fan blades, for example <NUM>, <NUM>, <NUM>, or <NUM> fan blades.

Each fan blade may be defined as having a radial span extending from a root (or hub) at a radially inner gas-washed location, or <NUM> percent span position, to a tip at a <NUM> percent span position. The ratio of the radius of the fan blade at the hub to the radius of the fan blade at the tip - the hub-tip ratio - may be less than (or on the order of) any of: <NUM>, <NUM>, <NUM><NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, or <NUM>. The hub-tip ratio may be in an inclusive range bounded by any two of the aforesaid values (i.e. the values may form upper or lower bounds). The hub-tip ratio may both be measured at the leading edge (or axially forwardmost) part of the blade. The hub-tip ratio refers, of course, to the gas-washed portion of the fan blade, i.e. the portion radially outside any platform.

The radius of the fan <NUM> may be measured between the engine centreline and the tip of a fan blade at its leading edge. The fan diameter may be greater than (or on the order of) any of: <NUM> metres (around <NUM> inches), <NUM> metres, <NUM> metres (around <NUM> inches), <NUM> metres (around <NUM> inches), <NUM> metres (around <NUM> inches), <NUM> metres (around <NUM> inches), <NUM> metres (around <NUM> inches), <NUM> metres (around <NUM> inches), <NUM> metres (around <NUM> inches), <NUM> (around <NUM> inches), <NUM> metres (around <NUM> inches), <NUM> metres (around <NUM> inches), <NUM> metres (around <NUM> inches), <NUM> metres (around <NUM> inches) or <NUM> metres (around <NUM> inches). The fan diameter may be in an inclusive range bounded by any two of the aforesaid values (i.e. the values may form upper or lower bounds).

The rotational speed of the fan <NUM> may vary in use. Generally, the rotational speed is lower for fans with a higher diameter. Purely by way of non-limitative example, the rotational speed of the fan at cruise conditions may be less than <NUM> rpm, for example less than <NUM> rpm. Purely by way of further non-limitative example, the rotational speed of the fan <NUM> at cruise conditions for an engine having a fan diameter in the range of from <NUM> metres to <NUM> metres (for example <NUM> metres to <NUM> metres) may be in the range of from <NUM> rpm to <NUM> rpm, for example in the range of from <NUM> rpm to <NUM> rpm, or, for example in the range of from <NUM> rpm to <NUM> rpm. Purely by way of further non-limitative example, the rotational speed of the fan at cruise conditions for an engine having a fan diameter in the range of from <NUM> metres to <NUM> metres may be in the range of from <NUM> rpm to <NUM> rpm, for example in the range of from <NUM> rpm to <NUM> rpm, for example in the range of from <NUM> rpm to <NUM> rpm.

In use of the engine <NUM>, the fan <NUM> (with its associated fan blades) rotates about a rotational axis. This rotation results in the tip of the fan blade moving with a velocity Utip. The work done by the fan blades on the flow results in an enthalpy rise dH of the flow. A fan tip loading may be defined as dH/Utip<NUM>, where dH is the enthalpy rise (for example the one dimensional average enthalpy rise) across the fan and Utip is the (translational) velocity of the fan tip, for example at the leading edge of the tip (which may be defined as fan tip radius at leading edge multiplied by angular speed). The fan tip loading at cruise conditions may be greater than (or on the order of) any of: <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM> or <NUM> (all units in this paragraph being Jkg-<NUM>K-<NUM>/(ms-<NUM>)<NUM>). The fan tip loading may be in an inclusive range bounded by any two of the values in the previous sentence (i.e. the values may form upper or lower bounds).

The engine <NUM> may have any desired bypass ratio, where the bypass ratio is defined as the ratio of the mass flow rate of the flow B through the bypass duct to the mass flow rate of the flow C through the core at cruise conditions. Depending upon the selected configuration, the bypass ratio may be greater than (or on the order of) any of the following: <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, or <NUM>. The bypass ratio may be in an inclusive range bounded by any two of the aforesaid values (i.e. the values may form upper or lower bounds). The bypass duct may be substantially annular. The bypass duct may be radially outside the core engine <NUM>. The radially outer surface of the bypass duct may be defined by a nacelle and/or a fan case.

The overall pressure ratio of the engine <NUM> may be defined as the ratio of the stagnation pressure upstream of the fan <NUM> to the stagnation pressure at the exit of the high-pressure compressor <NUM> (before entry into the combustor). By way of non-limitative example, the overall pressure ratio of the engine <NUM> at cruise may be greater than (or on the order of) any of the following: <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>. The overall pressure ratio may be in an inclusive range bounded by any two of the aforesaid values (i.e. the values may form upper or lower bounds).

Specific thrust of the engine <NUM> may be defined as the net thrust of the engine divided by the total mass flow through the engine <NUM>. At cruise conditions, the specific thrust of the engine <NUM> may be less than (or on the order of) any of the following: <NUM> Nkg-<NUM>s, <NUM> Nkg-<NUM>s, <NUM> Nkg-<NUM>s, <NUM> Nkg-<NUM>s, <NUM> Nkg-<NUM>s, <NUM> Nkg-<NUM>s, or <NUM> Nkg-<NUM>s. The specific thrust may be in an inclusive range bounded by any two of the values in the previous sentence (i.e. the values may form upper or lower bounds). Such engines may be particularly efficient in comparison with conventional gas turbine engines.

The engine <NUM> may have any desired maximum thrust. For example, the engine <NUM> may be capable of producing a maximum thrust of at least (or on the order of) any of the following: <NUM> kilonewtons, <NUM> kilonewtons, <NUM> kilonewtons, <NUM> kilonewtons, <NUM> kilonewtons, <NUM> kilonewtons, <NUM> kilonewtons, <NUM> kilonewtons, <NUM> kilonewtons, <NUM> kilonewtons, <NUM> kilonewtons, or <NUM> kilonewtons. The maximum thrust may be in an inclusive range bounded by any two of the aforesaid values (i.e. the values may form upper or lower bounds). The thrust referred to above may be the maximum net thrust at standard atmospheric conditions at sea level plus <NUM> degrees Celsius (ambient pressure <NUM> kilopascals, temperature <NUM> degrees Celsius), with the engine <NUM> being static.

In use, the temperature of the flow at the entry to the high-pressure turbine <NUM> may be particularly high. This temperature, which may be referred to as turbine entry temperature or TET, may be measured at the exit to the combustor <NUM>, for example immediately upstream of the first turbine vane, which itself may be referred to as a nozzle guide vane. At cruise, the TET may be at least (or on the order of) any of the following: <NUM> kelvin, <NUM> kelvin, <NUM> kelvin, <NUM> kelvin, <NUM> kelvin or <NUM> kelvin. The TET at cruise may be in an inclusive range bounded by any two of the aforesaid values (i.e. the values may form upper or lower bounds). The maximum TET in use of the engine <NUM> may be, for example, at least (or on the order of) any of the following: <NUM> kelvin, <NUM> kelvin, <NUM> kelvin, <NUM> kelvin, <NUM> kelvin, <NUM> kelvin or <NUM> kelvin. The maximum TET may be in an inclusive range bounded by any two of the aforesaid values (i.e. the values may form upper or lower bounds). The maximum TET may occur, for example, at a high thrust condition, for example at a maximum take-off (MTO) condition.

A fan blade and/or aerofoil portion of a fan blade described and/or claimed herein may be manufactured from any suitable material or combination of materials. For example at least a part of the fan blade and/or aerofoil may be manufactured at least in part from a composite, for example a metal matrix composite and/or an organic matrix composite, such as carbon fibre. By way of further example at least a part of the fan blade and/or aerofoil may be manufactured at least in part from a metal, such as a titanium based metal or an aluminium based material (such as an aluminium-lithium alloy) or a steel based material. The fan blade may comprise at least two regions manufactured using different materials. For example, the fan blade may have a protective leading edge, which may be manufactured using a material that is better able to resist impact (for example from birds, ice or other material) than the rest of the blade. Such a leading edge may, for example, be manufactured using titanium or a titanium-based alloy. Thus, purely by way of example, the fan blade may have a carbon-fibre or aluminium-based body with a titanium leading edge.

The fan <NUM> may comprise a central hub portion, from which the fan blades may extend, for example in a radial direction. The fan blades may be attached to the central portion in any desired manner. For example, each fan blade may comprise a fixture which may engage a corresponding slot in the hub. Purely by way of example, such a fixture may be in the form of a dovetail that may slot into and/or engage a corresponding slot in the hub/disc in order to fix the fan blade to the hub. By way of further example, the fan blades maybe formed integrally with a central hub portion. Such an arrangement may be a bladed disc or a bladed ring. Any suitable method may be used to manufacture such a bladed disc or bladed ring. For example, at least a part of the fan blades may be machined from a billet and/or at least part of the fan blades may be attached to the hub/disc by welding, such as linear friction welding.

The engine <NUM> may be provided with a variable area nozzle (VAN). Such a variable area nozzle may allow the exit area of the bypass duct to be varied in use. The general principles of the present disclosure may apply to engines with or without a VAN.

As used herein, cruise conditions have the conventional meaning and would be readily understood by the skilled person. Thus, for a given gas turbine engine for an aircraft, the skilled person would immediately recognise cruise conditions to mean the operating point of the engine at mid-cruise of a given mission (which may be referred to in the art as the "economic mission") may mean cruise conditions of an aircraft to which the gas turbine engine is designed to be attached. In this regard, mid-cruise is the point in an aircraft flight cycle at which <NUM> percent of the total fuel that is burned between top of climb and start of descent has been burned (which may be approximated by the midpoint - Such cruise conditions may be conventionally defined as the conditions at mid-cruise, for example the conditions experienced by the aircraft and/or engine at the midpoint (in terms of time and/or distance-) between top of climb and start of descent. Cruise conditions thus define an operating point of, the gas turbine engine that provides a thrust that would ensure steady state operation (i.e. maintaining a constant altitude and constant Mach number) at mid-cruise of an aircraft to which it is designed to be attached, taking into account the number of engines provided to that aircraft. For example where an engine is designed to be attached to an aircraft that has two engines of the same type, at cruise conditions the engine provides half of the total thrust that would be required for steady state operation of that aircraft at mid-cruise.

In other words, for a given gas turbine engine for an aircraft, cruise conditions are defined as the operating point of the engine that provides a specified thrust (required to provide - in combination with any other engines on the aircraft - steady state operation of the aircraft to which it is designed to be attached at a given mid-cruise Mach Number) at the mid-cruise atmospheric conditions (defined by the International Standard Atmosphere according to ISO <NUM> at the mid-cruise altitude). For any given gas turbine engine for an aircraft, the mid-cruise thrust, atmospheric conditions and Mach number are known, and thus the operating point of the engine at cruise conditions is clearly defined.

The cruise conditions may correspond to ISA standard atmospheric conditions at an altitude that is in the range of from <NUM> to <NUM> metres, such as from <NUM> to <NUM> metres, or from <NUM> to <NUM> metres (around <NUM> feet), or from <NUM> to <NUM> metres, or from <NUM> to <NUM> metres, or from <NUM> metres (around <NUM> feet) to <NUM> metres, or from <NUM> to <NUM> metres, or from <NUM> to <NUM> metres, or <NUM> metres. The cruise conditions may correspond to standard atmospheric conditions at any given altitude in these ranges.

The forward speed at the cruise condition may be any point in the range of from Mach <NUM> to <NUM>, for example one of Mach <NUM> to <NUM>, Mach <NUM> to <NUM>, Mach <NUM> to <NUM>, Mach <NUM> to <NUM>, Mach <NUM> to <NUM>, Mach <NUM>, Mach <NUM>, or in the range of from Mach <NUM> to <NUM>. Any single speed within these ranges may be the cruise condition. For some aircraft, the cruise conditions may be outside these ranges, for example below Mach <NUM> or above Mach <NUM>.

Thus, for example, the cruise conditions may correspond specifically to a pressure of <NUM> kilopascals, a temperature of minus <NUM> degrees Celsius, and a forward Mach number of <NUM>.

It will of course be appreciated, however, that the principles of the invention claimed herein may still be applied to engines having suitable design features falling outside of the aforesaid parameter ranges, while still falling within the scope of the appended claims.

A block diagram of the interface of the EEC <NUM> and other engine systems is shown in <FIG>.

As described previously, in the present embodiment, the EEC <NUM> is coupled with the PEM <NUM> to control the HP motor-generator <NUM> and the LP motor-generator <NUM>. In this way, power may either be extracted from or added to each of shafts <NUM> and <NUM>.

In the present embodiment, the PEM <NUM> facilitates conversion of alternating current to and from direct current. This is achieved in the present embodiment by employing a respective bidirectional power converter for conversion of ac to and from dc. Thus, as shown in the Figure, the PEM <NUM> comprises a first bidirectional power converter <NUM> connected with the HP motor-generator <NUM>, and a second bidirectional power converter <NUM> connected with the LP motor-generator <NUM>. The dc sides of the power converters <NUM> and <NUM> are in the present example connected with each other to facilitate bidirectional power transfer between the motor-generators <NUM> and <NUM>.

In an embodiment, both motor-generators and associated bidirectional power converters are rated at the same continuous power. In a specific embodiment, they are rated at from <NUM> kilowatts to <NUM> megawatts. In a more specific embodiment, they are rated from <NUM> kilowatts to <NUM> megawatt. In a more specific embodiment, they are rated at <NUM> kilowatts.

In other embodiments, the HP motor-generator <NUM> and the first bidirectional power converter <NUM> are rated at a different continuous power than the LP motor-generator <NUM> and the second bidirectional power converter <NUM>. In a specific embodiment, they are rated at from <NUM> kilowatts to <NUM> megawatts. In a more specific embodiment, they are rated from <NUM> kilowatts to <NUM> megawatt. In a more specific embodiment, they are rated at <NUM> kilowatts.

Those skilled in the art will be familiar with the term "continuous power," i.e. a maximum sustainable power output that does not damage components due to over-current, over-voltage, or over-temperature for example.

In the present example, the dc sides of the power converters <NUM> and <NUM> are also connected with a dc bus <NUM>.

In the present example the dc bus <NUM> has connected to it various loads, which may be either located on the engine <NUM> or on the vehicle instead. Some such as anti-icing systems <NUM> may be part of the engine, such as electric nacelle anti-icing systems, or part of the aircraft on which the engine <NUM> is installed, such as electric wing anti-icing systems.

Other loads may be connected with the dc bus <NUM> and be able to draw power from and supply power to the bus, such as an energy storage device in the form of a battery <NUM>. In the present example, control of the charge/discharge state of the battery <NUM> is facilitated by a dc-dc converter <NUM>. Other energy storage devices may be connected to the dc bus <NUM> as well as or in place of the battery <NUM>, such as a capacitor.

Other loads, indicated at <NUM>, may be connected with the dc bus <NUM> such as cabin environmental control systems, electric actuation systems, auxiliary power units, etc..

In operation, the EEC <NUM> receives a plurality of demand signals, namely a thrust demand in the form of a power lever angle (PLA) signal which may be manually set or by an autothrottle system, and an electrical power demand (PD). In addition, the EEC receives a plurality of sets of measured parameters, namely a set concerning flight parameters of the vehicle, ΣAIRCRAFT, and a set concerning operational parameters of the engine, ΣENGINE. As will be described further with reference to <FIG> and <FIG>, these demands and parameters facilitate the derivation of a set of output parameters to control the core gas turbine and the motor-generators.

Thus in addition to the routine set of output parameters generated by a FADEC, such as the fuel flow WF to be metered by a fuel metering unit <NUM> on the engine <NUM>, or variable-stator vane angle, handling bleed settings, etc. in the present embodiment the EEC <NUM> comprises a power controller module <NUM> to generate a control signal PH for the first bidirectional power converter <NUM> and a second control signal PL for the second bidirectional power converter <NUM>. The control signals PH and PL control the operation of the power converters in terms of both direction and magnitude of electrical power. In this way, the EEC <NUM> may meet the demanded power PD using a suitable balance of electrical power from the HP and LP motor-generators. As will be described with reference to the later Figures, the optimum way to do this varies throughout a mission.

In addition to the control signals PH and PL, in the present example the power controller <NUM> is configured to derive a control signal PBAT for the dc-dc converter <NUM> to facilitate charge or discharge of the battery <NUM>.

In a specific embodiment, the power controller <NUM> is configured to, in normal operation, set PBAT to zero, and only change its status as set out in the optimisation routines described herein, for example the routines described with reference to <FIG> and <FIG>.

In another specific embodiment, the power controller <NUM> includes battery optimisation functionality and modifies the power demand parameter PD by adding or subtracting a value PBAT depending upon whether it is more optimal to charge, discharge, or maintain the charge of the battery <NUM>. Those skilled in the art will be familiar with such types of battery state-of-charge optimisation routines.

Thus, in such an embodiment, the power controller <NUM> modifies the power demand parameter PD by performing an addition-assignment operation PD += PBAT. The sign convention used herein is such that a positive PBAT means that the battery <NUM> is to be charged, and thus additional generation is required from the HP and LP motor-generators, whilst a negative PBAT means that the battery <NUM> is to be discharged.

It will be appreciated that in embodiments lacking an energy storage device, this signal will not be generated and thus the power demand parameter PD will not be modified at all.

In the present example another control signal PAl is generated to activate the anti-icing systems <NUM>. In the present example only the nacelle anti-ice system of the engine <NUM> is controlled, however it is envisaged that the EEC <NUM> may in alternative implementations have authority over wing anti-ice in certain circumstances. In a similar way to the modification of PD by the addition-assignment of PBAT, the power controller <NUM> is configured to perform the addition assignment PD += PAl so as to account for the power required for the anti-icing systems <NUM>. Again, it will be appreciated that in embodiments lacking electric anti-ice systems, this signal will not be generated and no modification of PD will be performed in this manner.

In practice, the EEC <NUM> houses microprocessors for executing program modules to control the engine <NUM>. A block diagram of the functional modules of the power controller <NUM> is shown in <FIG>.

Input parameters previously described with reference to <FIG> are initially received by a classifier module <NUM> to output an optimiser setting mode for an optimiser module <NUM>. The operation of the classifier module <NUM> will be described further with reference to <FIG>, and the various modes of the optimiser module will be described with reference to <FIG>.

The input parameters are also supplied to a filter <NUM> prior to updating an engine model module <NUM>. The filter <NUM> in the present example is an integrator to smooth out short-term transients so as to not cause large variations in the engine model. The engine model module <NUM> runs a real time model of the engine <NUM> so as to facilitate prediction of changes is operational parameters, such as WF, PH, and PL given a thrust demand. Such models and their integration within an EEC will be familiar to those skilled in the art.

Following entry into a given optimisation setting, the optimiser module <NUM> finds the optimal set of parameters for operation of the engine <NUM> given the current operational state of aircraft on which the engine is installed and the engine itself.

The classifier module <NUM> is detailed in <FIG>.

In the present embodiment, the classifier module <NUM> comprises a comparator <NUM> which compares the present altitude (ALT), the Mach number (Mn) and the power lever angle to determine the flight regime. It will be appreciated that other inputs may be utilised to increase the fidelity of the comparison process, such as engine conditions, ambient temperature, etc..

In the present embodiment, the comparator is configured to identify if the engine is operating in a maximum take-off condition if the altitude is less than a threshold, the Mach number is less than a threshold, and the power lever angle is at a maximum. In an embodiment, the altitude threshold is <NUM> feet, whilst the Mach number threshold is <NUM>.

In the present embodiment, the comparator is configured to identify if the engine is operating in a maximum climb condition if the altitude is above a threshold, and the power lever angle is at a maximum. In an embodiment, the altitude threshold is <NUM> feet.

The optimisation strategy for the maximum take-off condition and the maximum climb mode of operation will be described further with reference to <FIG>.

In the present embodiment, the comparator is configured to identify if the engine is operating in a cruise if the altitude is above a threshold, the Mach number is in a cruise range, and the power lever angle is at a cruise setting. In an embodiment, the altitude threshold is <NUM> feet, whilst the Mach number range is from <NUM> to <NUM>.

In such circumstances, the optimisation strategy that can be employed is to minimise fuel flow WF at constant thrust. Alternatively, the optimiser may be set to optimise surge margin in the engine, or to optimise compression efficiency depending on the engine and aircraft parameters. Such strategies will be described further with reference to <FIG> and <FIG>, respectively.

Alternatively, the optimiser may be set to optimise the bypass ratio by varying the core flow, implementing a variable cycle.

In the present embodiment, the comparator is configured to identify if the engine is operating in a regime in which the low-pressure turbine <NUM> is operating in an unchoked condition. This typically manifests at low Mach number idle conditions, although the unchoked condition may exist at other operating points depending upon the specific design of the low-pressure turbine <NUM> and the rest of the engine <NUM>. In the present embodiment, the comparator is configured to identify this condition if the Mach number is below a threshold, and the power lever angle is at an idle setting. In an embodiment, the Mach number threshold is <NUM>. In other embodiments, other inputs such as the corrected speed of the low pressure spool may be utilised to identify the unchoked condition.

The optimisation strategy for the low Mach number idle condition will be described further with reference to <FIG>.

In the present embodiment, the comparator is configured to identify if the engine is operating in an approach Mach number idle condition if the altitude is within a range, the Mach number is in a range, and the power lever angle is at an idle setting. In an embodiment, the altitude range is from <NUM> to <NUM> feet above ground level, and the Mach number range is from <NUM> to <NUM>.

The optimisation strategy for the approach idle condition will be described further with reference to <FIG>.

The classifier module <NUM> further comprises a differentiator <NUM> which is configured to monitor the PLA and PD parameters and identify a transient type.

In response to the change in power lever angle being positive, the differentiator <NUM> is configured to identify the initiation of an acceleration event. The optimisation strategy for this manoeuvre will be described further with reference to <FIG>.

In response to the change in power lever angle being negative, the differentiator <NUM> is configured to identify the initiation of a deceleration event. The optimisation strategy for this manoeuvre will be described further with reference to <FIG>.

In response to a change in electrical power demand PD, the differentiator <NUM> is configured to cause the optimiser to invoke the optimisation strategy described with reference to <FIG>.

The classifier module <NUM> further comprises a limiter <NUM> which is configured to monitor the high-pressure and low-pressure spool speeds, NH and NL. Should either spool speed approach a limit, which may be a maximum limit or a keep-out zone, the optimisation strategy described with reference to <FIG> may be invoked.

In the present example, the outputs of the comparator <NUM>, differentiator <NUM> and limiter <NUM> are compared by a prioritiser <NUM>. It will be appreciated that there may be concurrent outputs from each initial stage of the comparator module, and thus in the present embodiment the comparator is configured to filter to only one output optimiser setting. In the present embodiment, outputs from the limiter <NUM> are priorities over outputs from the differentiator <NUM>, which are in turn prioritised over outputs from the comparator <NUM>.

Following the identification of a maximum take-off or maximum climb condition by the classifier module <NUM>, the optimiser <NUM> enters the corresponding optimisation routine at step <NUM>. At step <NUM>, a question is asked as to whether the power demand PD is less than the maximum power rating of the LP motor-generator <NUM>, PLmax. If so, then control proceeds to step <NUM> where the optimiser <NUM> maximises the power generation by the LP motor-generator <NUM>, PL.

Preferring electrical offtake from the low-pressure spool at temperature-limited conditions such as maximum take-off and maximum climb conditions reduces the load on the high-pressure spool. For a given power demand PD, this results in a higher high-pressure spool rotational speed NH. This increases the core flow C through the core gas turbine, and so reduces the stator outlet temperature required to deliver the low-pressure turbine power for a given thrust.

It has been found that for motor-generators rated at <NUM> kilowatts, it is possible to reduce the stator outlet temperature by <NUM> kelvin. It will be appreciated that the higher the rating of the motor-generators, the greater the reduction that may be achieved.

Clearly, if the power demand PD after accounting for battery charge/discharge and/or anti-icing system operation is greater than the maximum power rating of the LP motor-generator <NUM>, PLmax, then the question asked at step <NUM> will be answered in the negative. In this case, control proceeds to step <NUM> where the optimiser <NUM> maximises the power generation by the LP motor-generator <NUM>, PL, and minimises the power generation by the HP motor-generator <NUM>, PH. Thus, the LP motor-generator <NUM> is directed to generate PLmax and the HP motor-generator <NUM> is directed to generate the remainder, PD - PLmax.

In an embodiment, should spare capacity be available from the LP motor-generator <NUM>, then the optimiser <NUM> may elect to divert PLmax - PD to the HP motor-generator <NUM> which may further increase core flow and reduce stator outlet temperature.

Following the identification of an unchoked regime for the low-pressure turbine <NUM> by the classifier module <NUM>, the optimiser <NUM> enters the corresponding optimisation routine at step <NUM>. As described previously, this may be triggered by a low Mach number idle operating condition, for example the ground idle operating point.

At step <NUM>, a question is asked as to whether the power demand PD is less than the maximum power rating of the HP motor-generator <NUM>, PHmax. If so, then control proceeds to step <NUM> where the optimiser <NUM> maximises the power generation by the HP motor-generator <NUM>, PH.

As the low-pressure turbine <NUM> is unchoked, there is a large impact when it is required to provide electrical power via the LP motor-generator <NUM>. The reduction in the speed of the low-pressure spool, NL, is such that there is a steep drop in the efficiency of the low-pressure compressor <NUM>, and thus an increase in fuel burn. In practice, there may also be a drop in efficiency of the LP motor-generator <NUM> due to the lower rotational speed. Thus, at the low flight Mach numbers that cause the low-pressure turbine <NUM> to unchoke, it is possible to improve fuel consumption by this optimisation routine.

In a similar way to the situation of <FIG>, should the power demand PD be greater than the maximum power rating of the HP motor-generator <NUM>, PHmax, then the question asked at step <NUM> will be answered in the negative. In this case, control proceeds to step <NUM> where the optimiser <NUM> maximises the power generation by the HP motor-generator <NUM>, PH, and minimises the power generation by the LP motor-generator <NUM>, PH. Thus, the HP motor-generator <NUM> is directed to generate PHmax and the LP motor-generator <NUM> is directed to generate the remainder, PD - PHmax.

In an embodiment, should spare capacity be available from the HP motor-generator <NUM>, then the optimiser <NUM> may elect to divert PHmax - PD to the LP motor-generator <NUM> which may further reduce fuel consumption.

Following the identification of an approach idle condition by the classifier module <NUM>, the optimiser <NUM> enters the corresponding optimisation routine at step <NUM>. At step <NUM>, a question is asked as to whether the power demand PD is less than the maximum power rating of the LP motor-generator <NUM>, PLmax. If so, then control proceeds to step <NUM> where the optimiser <NUM> maximises the power generation by the LP motor-generator <NUM>, PL. The excess capacity of the LP motor-generator <NUM>, PLmax - PD, is transferred to the HP motor-generator <NUM>.

This has two effects. First, extraction of the maximum power possible from the low-pressure spool significantly reduces the low-pressure spool rotational speed, NL. Recalling that in the engine <NUM> the fan <NUM>, which is primary thrust generating element, rotates with the low-pressure spool, it will be clear that a reduction in NL reduces the thrust generated by the engine <NUM>. This relaxes the requirement to use high drag devices on the airframe to achieve a required descent rate. This reduces noise and reduces fuel consumption.

Second, on approach, the engine idle setting is often constrained by the requirement for the engine to respond to a throttle transient in a timely manner - in the event of a go-around, the engine must deliver full rated thrust as quickly as possible. Initial high-pressure spool speed NH has a powerful effect on the response time during an engine acceleration, so maximising NH at idle significantly reduces the engine acceleration time. However, this is usually at the expense of a low idle thrust.

The inventor has found that by transferring power from the low-pressure spool to the high-pressure spool allows a higher NH and thus a shorter response time, along with a reduced idle thrust due to the lower NL. It has been demonstrated that in an engine of the type described herein, <NUM> kilowatts of power transfer achieves a sufficiently high NH and a <NUM> percent reduction in idle thrust.

If the power demand PD is greater than the maximum power rating of the LP motor-generator <NUM>, PLmax, then the question asked at step <NUM> will be answered in the negative. In this case, control proceeds to step <NUM> where the optimiser <NUM> maximises the power generation by the LP motor-generator <NUM>, PL, and minimises the power generation by the HP motor-generator <NUM>, PH. Thus, the LP motor-generator <NUM> is directed to generate PLmax and the HP motor-generator <NUM> is directed to generate the remainder, PD - PLmax.

In an embodiment, the optimiser <NUM> is further configured to identify that maintaining the requisite NH will cause unsafe operation of the low-pressure compressor <NUM>. This may be caused by the operating point of the LP low-pressure compressor <NUM> becoming too close to surge or to choke. In response to the onset of such a condition, the fuel flow WF may be increased.

Alternatively, in order to reduce fuel burn on approach, the optimiser <NUM> may be configured to supplement the HP motor-generator using an energy storage device, for example the battery <NUM> via control of the PBAT parameter, or another energy source such as the auxiliary power unit on the aircraft. In this way, community emissions on approach may be reduced.

One of the primary considerations for control of a gas turbine engine is the prevention of surge in the compression stages. A characteristic <NUM> for an axial flow compressor, such as low-pressure compressor <NUM> or high-pressure compressor <NUM>, is shown in <FIG>.

The characteristic <NUM> plots pressure ratio against flow function, which in this case is non-dimensional flow (W√T/P). The characteristic <NUM> shows a plurality of non-dimensional speed lines <NUM>, <NUM>, <NUM>, <NUM>, along with the compressor steady state working line <NUM>, which is the locus of operating points for various steady state throttle settings at different non-dimensional speeds. In addition, the surge line <NUM> is shown, which is the locus of points at which the compressor enters surge at the various non-dimensional speeds. For a given value of the flow function, the pressure ratio at which surge is encountered is denoted R. The difference in pressure ratio on the working line <NUM> and the value of R on the surge line <NUM> for a given value of the flow function is denoted dR. Therefore, the surge margin for a given compressor operating point may be defined as dR/R.

It is important to maintain a degree of surge margin at all points in the operational envelope. This is to mitigate random threats, such as inlet flow instabilities due to crosswinds or turbulence, for example. To a first order, it is often recommended for low-pressure compressors to have around <NUM> percent surge margin, and high-pressure compressors to have around <NUM> percent surge margin. A significant proportion of the surge margin, typically up to half, is to make allowance for transient excursions of the working line during acceleration and deceleration manoeuvres. Such transient phenomena will be described further with reference to <FIG>, <FIG>.

Characteristics for the high-pressure compressor <NUM> and the low-pressure compressor <NUM> showing transient phenomena during acceleration events are plotted in <FIG> respectively.

In <FIG>, the high-pressure compressor steady-state working line <NUM> is shown along with lines of constant corrected speed <NUM>, <NUM>, <NUM>, and <NUM>, and the surge line <NUM>. During an acceleration manoeuvre, the high-pressure compressor <NUM> moves from an initial operating point <NUM> to a final operating point <NUM> via a transient working line <NUM> above the steady-state working line <NUM>.

Similarly, in <FIG>, the low-pressure compressor steady-state working line <NUM> is shown along with lines of constant corrected speed <NUM>, <NUM>, <NUM>, and <NUM>, and the surge line <NUM>. During an acceleration manoeuvre, the low-pressure compressor <NUM> moves from an initial operating point <NUM> to a final operating point <NUM> via a transient working line <NUM> which crosses the steady-state working line <NUM>.

During the acceleration manoeuvre, the high-pressure compressor <NUM> initially moves towards surge due to the flow compatibility requirement with the high-pressure turbine <NUM>. The flow function of the high-pressure turbine <NUM> (W<NUM>√T<NUM>/P<NUM>) is substantially fixed during most operating conditions of the engine <NUM>, due to the nozzle guide vanes therein being choked. In order to initiate the acceleration manoeuvre, for example due to an increase in power lever angle setting, the amount of fuel metered by the fuel metering unit <NUM> (WF) is increased. This leads to an increase in turbine entry temperature, T<NUM>. Normally, the high-pressure spool speed NH is prevented from changing instantaneously due to its inertia. Consequently, the operating point of the high-pressure compressor <NUM> moves up a line of constant corrected speed. As the overfuelling continues and the high-pressure spool inertia is overcome, the operating point moves along the transient working line <NUM> parallel to the surge line <NUM>. As the acceleration finishes, the high-pressure compressor <NUM> adopts its final operating point <NUM> on the steady-state working line <NUM>.

Referring now to <FIG>, at the initiation of the acceleration manoeuvre, the operating point of the low-pressure compressor <NUM> moves a little towards surge, and then crosses the steady-state working line <NUM>. The initial move towards surge is due to the reduction in flow in the high-pressure compressor <NUM> due to the high-pressure spool inertia as described above. As the speed of the high-pressure compressor <NUM> increases it can accept more flow. However, due to the greater inertia of the low-pressure spool (recalling that it drives the fan <NUM> via the gearbox <NUM>), it cannot accelerate at the same rate and so the operating points of the low-pressure compressor <NUM> during the acceleration manoeuvre fall below the steady-state working line <NUM>.

It will be understood that handling during acceleration manoeuvres may significantly affect the design of the compressor turbomachinery and impose requirements for systems to manage the transients by either modifying the transient working line or the surge line, such as variable guide vanes and bleed valves.

By contrast, in the engine <NUM> it is possible to utilise the HP motor-generator <NUM> and optionally the LP motor-generator <NUM> to reduce the transient excursion.

<FIG> shows the characteristic for the high-pressure compressor <NUM> when the HP motor-generator <NUM> is used to overcome the high-pressure spool inertia. It can be seen that for the same degree of overfuelling, the transient working line <NUM> is much closer to the steady-state working line <NUM> and further from the surge line <NUM>. Thus for a given compressor configuration, this technique may either be used to improve surge margin during an acceleration manoeuvre, or facilitate a greater degree of overfuelling (up to the stator outlet temperature limit) and thus a faster acceleration time.

<FIG> shows the characteristic for the low-pressure compressor <NUM> during application of the same technique on the low-pressure spool using the LP motor-generator <NUM>. It may be seen that the transient working line <NUM> is again much closer to the steady-state working line <NUM> due to the reduction in effective low-pressure spool inertia by the LP motor-generator <NUM>.

Steps carried out by the optimiser <NUM> to achieve the advantages described previously for an acceleration event are set out in <FIG>.

Following the identification of an acceleration condition by the classifier module <NUM>, the optimiser <NUM> enters the corresponding optimisation routine at step <NUM>. In the present embodiment, with an energy storage system such as the battery <NUM> available on the dc bus <NUM>, a question is asked at step <NUM> as to whether the battery state of charge is greater than a minimum value. As will be appreciated by those skilled in the art, this may not be absolutely zero but instead will be a minimum state of charge, for example <NUM> percent, below which the battery may be damaged.

If so, a further question is asked at step <NUM> as to whether the unmodified total aircraft power demand PD (i.e. prior to any modification thereof to account for battery optimisation) is less than the maximum power available from the battery <NUM>, PBATmax. If so, then control proceeds to step <NUM> where the optimiser <NUM> overrides any concurrent battery optimisation processes, and fully supplies the power demand PD using the battery <NUM>, with any excess being supplied to the HP motor-generator <NUM> to overcome the HP spool inertia. Optionally, any further excess may be supplied to the LP motor-generator <NUM>.

Should either of the questions asked at steps <NUM> or <NUM> be answered in the negative, i.e. the battery <NUM> has a minimal state of charge, or the unmodified total aircraft power demand PD is greater than the maximum power of the battery <NUM>, PBATmax (or if indeed the particular embodiment of the engine <NUM> does not include a battery), then control proceeds to step <NUM> in which power generation by the LP motor-generator <NUM>, PL, is maximised and power generation by the HP motor-generator <NUM>, PH, is minimised.

Following optimisation of the power generation strategy to satisfy the power demand PD in the preceding steps, the fuel flow WF metered by the fuel metering unit <NUM> is increased at step <NUM>.

When a deceleration event is initiated, fuel flow by the fuel metering unit <NUM> is reduced. In the opposite sense to the scenario described above, as fuel flow is decreased, the turbine entry temperature decreases instantaneously. This is because, as described previously, during the majority of operating scenarios, the nozzle guide vanes in the high-pressure turbine <NUM> are choked and the flow function remains constant. As the high-pressure spool speed NH is prevented from changing instantaneously due to its inertia, the reduction in turbine entry temperature T<NUM> causes the operating point of the high-pressure compressor <NUM> to move down a line of constant corrected speed on its characteristic. This manifests as an increase in mass flow W<NUM> and a decrease in pressure P<NUM> at the exit of the high-pressure compressor <NUM>.

The result of this for the combustor <NUM> is that not only is the amount of fuel delivered lower, but the mass flow W<NUM> therethrough has increased. This means that the combustor <NUM> operates at a lower fuel-air ratio (FAR) than normal, which risks weak extinction (also known as lean blowout).

Referring to <FIG>, which is a plot of FAR against flow function, the weak extinction boundary <NUM> for the combustor <NUM> is illustrated. Corrected flow through the combustor <NUM> to the right of the weak extinction boundary <NUM> results in extinction of the flame and is an unacceptable operating condition. A steady-state FAR for a particular mass flow through the combustor <NUM>, is shown at point <NUM>. The constraint on how aggressive a deceleration manoeuver may be is dictated by the allowable underfuelling margin. In prior art engines, in which the flow function increases slightly at the outset of the deceleration manoeuver, the underfuelling margin is limited to M<NUM> due to the proximity of the fuel-air ratio to the weak extinction boundary <NUM> during the deceleration, shown at point <NUM>.

However, by actively reducing the speed of the high-pressure spool at the initiation of and during the deceleration event by the HP motor-generator <NUM>, the operating point of the high-pressure compressor <NUM> is no longer forced to move down a constant speed line at the outset of the manoeuvre. Instead, there is only a slight deviation from the steady state working line due to the reduction in turbine entry temperature, T<NUM>. As the speed of the high-pressure compressor <NUM> substantially instantaneously begins to decelerate, the mass flow through the combustor <NUM> also reduces substantially instantaneously. As shown at point <NUM> in <FIG>, the additional margin M<NUM> allows a greater degree of underfuelling.

It will be understood that this approach allows the design of the combustor <NUM> to be optimised due to the greater weak extinction margin, and also allows the vehicle design to be optimised as a greater thrust reduction is achievable by the engine alone without resort to high drag devices to reduce forward airspeed in, for example, a slam decel manoeuvre.

It will also be appreciated that the approach provides a method of controlling weak extinction in the combustor <NUM>. Using the engine model <NUM>, for example, the onset of weak extinction may be identified by evaluating the current fuel-air ratio in the combustor <NUM>. This may be achieved, for example, by utilising the flight Mach number, altitude and temperature to determine the mass flow in the engine <NUM>, the characteristic of the fan <NUM> to determine the mass flow C into the core gas turbine, and the characteristics of the compressors <NUM> and <NUM> to determine the mass flow into the combustor <NUM>. This may be combined with the commanded fuel flow WF along with a model of the combustion process to determine the fuel-air ratio.

In response to identifying that the fuel-air ratio is approaching the weak extinction boundary <NUM>, the EEC <NUM> may use the power controller <NUM> to extract mechanical shaft power from the high-pressure spool using the HP motor-generator <NUM> to prevent a further drop in fuel-air ratio in the combustor <NUM>.

Steps carried out by the optimiser <NUM> to achieve the advantages described previously for a deceleration event are set out in <FIG>.

Following the identification of a deceleration condition by the classifier module <NUM>, the optimiser <NUM> enters the corresponding optimisation routine at step <NUM>. A question is asked at step <NUM> as to whether the power demand PD is less than the maximum power generation capability of the HP motor-generator <NUM>, PHmax. If so, then control proceeds to step <NUM> whereupon the power generation of the HP motor-generator <NUM>, PH, is maximised to satisfy PD.

In the present embodiment, the excess capacity PHmax - PD is transferred to other loads. In an embodiment, the excess capacity is directed to an energy storage system, such as the battery <NUM>. As described previously, the energy storage system may additionally or alternatively comprise a capacitor. Additionally or alternatively, the excess capacity may be directed to an electrical consumer such as the anti-icing system <NUM>, which may be the nacelle anti-ice system of the engine <NUM>. Alternatively, it may be the wing anti-ice system of the vehicle on which the engine <NUM> is installed.

If it is inappropriate to direct excess capacity anywhere, for example if further heating of using anti-ice systems may cause damage given the atmospheric conditions, or the energy storage system is full, then in an embodiment step <NUM> solely maximises PH up to PD to assist in the reduction of the high-pressure spool speed.

If the question asked at step <NUM> is answered in the negative, to the effect that the power demand PD is greater than the maximum power generation capability of the HP motor-generator <NUM>, PHmax, then control proceeds to step <NUM> where first the power generation of the HP motor-generator <NUM>, PH, is maximised, then the power generation of the LP motor-generator <NUM>, PL, is maximised to supply remainder of PD.

Following optimisation of the power generation strategy to satisfy the power demand PD in the preceding steps, the fuel flow WF metered by the fuel metering unit <NUM> is reduced at step <NUM>.

In an alternative, embodiment the excess capacity PHmax - PD may be directed to the LP motor-generator <NUM>. This may be possible due to this excess power representing a small proportion of the power generated by the low-pressure turbine <NUM>, therefore leading to a very small change in thrust generated by the fan <NUM>. Whilst the change in thrust may be small, the effect on the high-pressure spool is large in terms preventing an increase in mass flow at the point of reduction of fuel flow, and thereby on the ability to prevent weak extinction.

When an increase in power demand PD occurs, the power controller <NUM> must respond by in turn demanding an increase in power output by the gas turbine engine.

<FIG> illustrates an exemplary increase in power demand PD of magnitude dPD within a timeframe dt. <FIG> shows a characteristic for an exemplary axial flow compressor, forming part of a single gas turbine spool coupled to a generator. In order to satisfy the increase in power demand dPD the specific work of the turbine must increase. The steady-state working line is shown at <NUM>, with the surge line shown at <NUM>. To increase the work by the engine, an increase in fuel flow is required.

For the situation in which the generator load follows the step of <FIG>, the spool may be held at constant corrected speed, or allowed to accelerate to a higher non-dimensional speed. The movement of the operating point of the exemplary compressor for each option is shown on the characteristic of <FIG>. Line <NUM> shows the movement of the operating point at constant corrected speed. Line <NUM> shows the movement of the operating point to a higher corrected speed. It may be seen that responding in this manner would mean that as the generator load increases, the compressor non-dimensional speed exhibits a slight initial reduction as a greater proportion of the turbine work is used to drive the generator rather than the compressor. As fuel flow increases, the compressor operating point moves towards and in both examples exceeds the surge line <NUM>.

Thus it may be seen that at low engine throttle settings in particular, such an increase in power demand PD may unchecked cause a compressor to enter surge, requiring additional handling systems prevent this and guarantee adequate surge margin. In practice, this situation may for example occur in an aircraft engine during descent when anti-ice systems need to be enabled but the engines are at an idle setting.

According to the claimed invention, however, the approach is taken to utilise the energy storage system to mitigate the possibility of surge. Thus, as illustrated in <FIG>, the same increase in power demand PD of magnitude dPD within a timeframe dt is demanded. Instead of this solely being met by one or both of the HP motor-generator <NUM> and the LP motor-generator <NUM>, it is met during the manoeuvre by the battery <NUM>. Thus, as shown in the Figure, initially the power demand is met by the battery <NUM>, as shown by the shaded region <NUM>. As the engine <NUM> accelerates, the proportion provided by the motor-generator(s) increases gradually until the new power demand is fully met by the engine <NUM>.

<FIG> shows the transient working line <NUM> on a compressor characteristic when this approach is adopted. As the initial increase in power demand is fulfilled by a different energy source to the core gas turbine engine, there is no attendant drop in compressor non-dimensional speed. In addition, the increase in fuel flow may be tempered, so that the raise in working line during the transient manoeuvre is not as great as in the example of <FIG>. In this way, adequate surge margin is maintained, potentially allowing a more optimum compressor design and/or removal of handling systems.

In an alternative embodiment, the battery <NUM> may provide all of the higher power demand whilst the engine <NUM> accelerates to a higher corrected speed, at which point provision of the power demand PD is switched from the battery to the motor-generator(s) in the engine <NUM>.

Steps carried out by the optimiser <NUM> to achieve the functionality described previously for an increase in power demand PD are set out in <FIG>.

Following the identification of an increase in power demand within a given timeframe dt by the differentiator <NUM> in the classifier module <NUM>, the optimiser <NUM> enters the corresponding optimisation routine at step <NUM>. At step <NUM>, the optimiser <NUM> evaluates the operating points of the low-pressure compressor <NUM> and the high-pressure compressor <NUM> for the demanded PD. In the present embodiment, this may be achieved using the engine model <NUM> and knowledge of the current power lever angle setting etc. Alternatively, a look-up table or similar may be used instead.

At step <NUM>, the current surge margin in the low-pressure compressor <NUM>, dRL/RL, and the current surge margin in the high-pressure compressor <NUM>, dRH/RH are evaluated, again using the engine model <NUM> in the present embodiment, or suitable alternatives if required.

At step <NUM>, the maximum allowable rate of acceleration for each spool is evaluated given the requirement to maintain adequate surge margin during the manoeuvre. In the present embodiment, this may be achieved by referring to the respective acceleration schedules for the spools.

A question is then asked as to whether acceleration of the high-pressure spool and low-pressure spool only will meet the required power demand within the demanded timeframe. If not, for example if the new power demand is very high or is required in a very short amount of time, then control proceeds to step <NUM> where a decision is taken to utilise the battery <NUM> (or other energy storage unit such as a capacitor) to satisfy the demanded PD.

Then, or if the question asked at step <NUM> was answered in the negative, the high-pressure and low-pressure spools are accelerated to their new operating points by increasing the fuel flow metered by the fuel metering unit <NUM>. As described previously, at this point the new power demand may then be fully met by one or more of the HP motor-generator <NUM> and the LP motor-generator <NUM>. The transition may is gradual.

The effect of power transfer from the LP motor-generator <NUM> to the HP motor-generator <NUM> on the operating point of the high-pressure compressor <NUM> is shown in <FIG> on the compressors' characteristic. The effect on the operating point of the low-pressure compressor <NUM> is shown on its characteristic in <FIG>.

As power is added to the high-pressure spool, the pressure ratio and flow function increase, as shown in <FIG> by the transition from an initial operating point <NUM> to a final operating point <NUM> at a higher non-dimensional speed on the compressor's working line <NUM>.

Extraction of power from the low-pressure spool lowers the working line of the low-pressure compressor <NUM>. Recalling that the low-pressure compressor rotational speed is fixed relative to the fan <NUM>, at constant thrust the low-pressure compressor operating point may only move on a constant non-dimensional speed line, in this case speed line <NUM>. Due to the increase in flow function in the high-pressure compressor <NUM>, the low-pressure compressor <NUM> is unthrottled and so also sees a raise in flow function. Thus the operating point moves from an initial operating point <NUM> to a final operating point <NUM> on speed line <NUM> away from the surge line <NUM>.

It will therefore be understood that controlling the degree of electrical power generated by one or both of the HP motor-generator <NUM> and the LP motor-generator <NUM> allows the mass flow rate of the core flow C to be varied even at fixed thrust settings. Recalling that the bypass ratio of the engine <NUM> is defined as the ratio of the mass flow rate of the flow B through the bypass duct to the mass flow rate of the flow C through the core gas turbine, this allows the bypass ratio of the engine <NUM> to be varied. This has particular advantages in terms of optimising the jet velocity of the engine <NUM> for particular airspeeds.

In an embodiment, power transfer may be used to further vary the bypass ratio by operating the LP motor-generator <NUM> as a generator and operating the HP motor-generator <NUM> as a motor.

In this way, it will be understood that the engine <NUM> may operate as a variable-cycle engine.

It will also be seen that transfer of power from the low-pressure spool to the high-pressure spool is an effective way of increasing surge margin in both compressors. It should also be noted that these fundamental effects occur even in the absence of active power transfer: should the power demand PD be greater than or equal to the capability of the LP motor-generator <NUM>, then an increase in surge margin is still achieved by satisfying the power demand PD by maximising low-pressure spool offtake. This is because a greater enthalpy drop is required across the low-pressure turbine <NUM>, which requires a greater mass flow. The greater mass flow through the high-pressure compressor <NUM>, whilst not as high as with power transfer, still unthrottles the low-pressure compressor <NUM> and increases its surge margin. Thus it will be understood that this strategy provides a suitable means for increasing surge margin in the engine <NUM>.

Steps carried out by the optimiser <NUM> to increase surge margin are therefore set out in <FIG>.

Following the identification of an operating condition in which surge margin needs to be increased by the classifier module <NUM>, the optimiser <NUM> enters the corresponding optimisation routine at step <NUM>. As described previously, operating conditions such as in high cross winds or other unsteady inlet flow phenomena may trigger entry into this routine.

At step <NUM>, a question is asked as to whether the current power demand PD is less than the maximum power rating of the LP motor-generator <NUM>, PLmax. If so, then control proceeds to step <NUM> where the optimiser <NUM> maximises the power generation by the LP motor-generator <NUM>, PL to increase surge margin in the low-pressure compressor <NUM>, and transfers any excess electrical power PLmax - PD to the HP motor-generator <NUM> to raise its operating point up its working line, also increasing surge margin.

If the question asked at step <NUM> is answered in the negative, to the effect that the LP motor-generator <NUM> is not solely capable of satisfying the power demand PD, then control proceeds to step <NUM> where the optimiser <NUM> maximises the power generation by the LP motor-generator <NUM>, PL to increase surge margin in the low-pressure compressor <NUM>. Recall that power extraction from the high-pressure spool normally moves the operating point of the high-pressure compressor <NUM> down its normal working line, but that the extraction of power from the low-pressure spool normally moves the operating point up its working line. Thus in step <NUM> the optimiser <NUM> minimises power generation by the HP motor-generator <NUM>, PH which substantially maintain its operating point at around its steady state value, or slightly higher on its working line.

Whilst surge margin may be increased by the method of <FIG>, it may in some cases be beneficial to reverse the direction of the power transfer such that power is transferred from the HP motor-generator <NUM> to the LP motor-generator <NUM>. <FIG> shows the movement of the operating point of the high-pressure compressor <NUM> in this scenario. <FIG> shows the movement of the operating point of the low-pressure compressor <NUM> in this scenario. The characteristics include lines of constant isentropic efficiency for the compressors. It can be seen that the transfer of power from the high-pressure spool to the low-pressure spool may enable an increase in compression efficiency in the engine <NUM> by moving the operating points of the compressors into regions of high efficiency.

Steps carried out by the optimiser <NUM> to increase compression efficiency are therefore set out in <FIG>.

Following the identification of an operating condition in which compression efficiency may be increased by the classifier module <NUM>, the optimiser <NUM> enters the corresponding optimisation routine at step <NUM>. As described previously, operating conditions such as sufficiently steady inlet flow may permit entry into this routine.

At step <NUM>, a question is asked as to whether the current power demand PD is less than the maximum power rating of the HP motor-generator <NUM>, PHmax. If so, then control proceeds to step <NUM> where the optimiser <NUM> maximises the power generation by the HP motor-generator <NUM>, PH to increase compression efficiency in the high-pressure compressor <NUM>, and transfers any excess electrical power PHmax - PD to the LP motor-generator <NUM> to lower its operating point on its working line, also increasing compression efficiency in this example.

If the question asked at step <NUM> is answered in the negative, to the effect that the HP motor-generator <NUM> is not solely capable of satisfying the power demand PD, then control proceeds to step <NUM> where the optimiser <NUM> maximises the power generation by the HP motor-generator <NUM>, PH to increase compression efficiency in the high-pressure compressor <NUM>. Furthermore, the optimiser <NUM> minimises power generation by the LP motor-generator <NUM>, PL to keep the low-pressure compressor <NUM> in as high a region of compression efficiency as possible.

As described previously, it is also possible to utilise the HP motor-generator <NUM> and LP motor-generator <NUM> to implement speed limiting. This may provide advantages in terms of safety, by preventing overspeed conditions or by managing operation around keep-out zones for example speed ranges where vibration levels are high.

The limiter <NUM> monitors the shaft speeds NH and NL. In an embodiment, the limiter triggers if a mechanical limit is exceed, i.e. purely on the basis of revolutions per minute. Alternatively, the limiter triggers on the basis of an aerodynamic limit, i.e. a corrected speed, and so takes temperatures into account. In this way, the breakdown of flow in the compressors may be prevented.

Thus, following the identification of a limit condition by the classifier module <NUM>, the optimiser <NUM> enters the corresponding optimisation routine at step <NUM>. At step <NUM>, a question is asked as to whether the limit is either the low-pressure shaft speed, NL (either mechanical or aerodynamic), or the high-pressure shaft speed, NH (either mechanical or aerodynamic).

If the trigger was low-pressure shaft speed, NL, then control proceeds to step <NUM> in which the optimiser <NUM> maximises the power generation by the LP motor-generator <NUM>, PL to decrease the low-pressure shaft speed. As described previously with respect to other optimisation routines, the electrical energy generated may be stored if capacity is available in an energy storage system such as battery <NUM>, or alternatively it may be diverted to other systems such as anti-icing systems or potentially the HP motor-generator <NUM>.

Claim 1:
A gas turbine engine (<NUM>) for an aircraft, comprising:
a high-pressure (HP) spool comprising an HP compressor (<NUM>) and a first electric machine (<NUM>) driven by an HP turbine (<NUM>);
a low-pressure (LP) spool comprising an LP compressor (<NUM>) and a second electric machine (<NUM>) driven by an LP turbine (<NUM>);
a combustor (<NUM>); and
an engine controller (<NUM>);
characterized in that the gas turbine engine (<NUM>) further comprises an electrical energy storage unit (<NUM>) and in that the engine controller (<NUM>) is configured to:
in response to receipt of a request for an increased electrical power supply by the engine (dPD) within a requested timeframe (dt), evaluate a current HP compressor surge margin (dRH/RH) and a current LP compressor surge margin (dRL/RL) based on the operating points thereof;
in response to evaluating that one or more of the HP compressor (<NUM>) and the LP compressor (<NUM>) has insufficient surge margin to facilitate a change of electrical power supply at a requested rate (dPD/dt), meet the requested electrical power demand using the electrical energy storage unit (<NUM>) whilst increasing fuel flow to the combustor (<NUM>) to accelerate the engine (<NUM>), and gradually increase the amount of electrical power supplied from one or more of the first (<NUM>) and second (<NUM>) electric machine whilst the engine (<NUM>) accelerates.