Patent Publication Number: US-11384696-B2

Title: Reducing low flight mach number fuel consumption

Description:
CROSS-REFERENCE TO RELATED APPLICATIONS 
     This specification is based upon and claims the benefit of priority from UK Patent Application Number 1908373.2 filed on 12 Jun. 2019, the entire contents of which are incorporated herein by reference. 
     TECHNICAL FIELD 
     This disclosure relates to gas turbine engines. 
     BACKGROUND 
     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. 
     SUMMARY 
     In an aspect, there is provided a gas turbine engine for an aircraft, comprising: 
     a high-pressure (HP) spool comprising an HP compressor and a first electric machine driven by an HP turbine, the first electric machine having a first maximum output power; 
     a low-pressure (LP) spool comprising an LP compressor and a second electric machine driven by an LP turbine, the second electric machine having a second maximum output power; and 
     an engine controller configured to identify a condition to the effect that the LP turbine is operating in an unchoked regime, and, in response to an electrical power demand being between zero and the first maximum output power, only extracting electrical power from the first electric machine to meet the electrical power demand. 
     In another aspect, there is provided a method of reducing fuel consumption at low inlet Mach numbers in a gas turbine engine of the type having a high-pressure (HP) spool comprising an HP compressor and a first electric machine driven by an HP turbine, the first electric machine having a first maximum output power, and a low-pressure (LP) spool comprising an LP compressor and a second electric machine driven by an LP turbine, the second electric machine having a second maximum output power, the method comprising: 
     identifying a condition to the effect that the LP turbine is operating in an unchoked regime; and 
     in response to an electrical power demand being between zero and the first maximum output power, only extracting electrical power from the first electric machine to meet the electrical power demand 
    
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
       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: 
         FIG. 1  shows a general arrangement of an engine for an aircraft; 
         FIG. 2  is a block diagram of the engine of  FIG. 1 ; 
         FIG. 3  is a block diagram of the interface of the electronic engine controller and other systems on the engine of  FIG. 1 ; 
         FIG. 4  is block diagram of the functional modules of the power controller in the electronic engine controller of  FIG. 3 ; 
         FIG. 5  is block diagram of the functional modules of the classifier in the power controller of  FIG. 4 ; 
         FIG. 6  shows a procedure to optimise operation of the engine of  FIG. 1  during maximum climb or maximum take-off conditions; 
         FIG. 7  shows a procedure to optimise operation of the engine of  FIG. 1  in conditions in which the low pressure turbine is unchoked; 
         FIG. 8  shows a procedure to optimise operation of the engine of  FIG. 1  during an approach idle condition; 
         FIG. 9  shows a characteristic for an axial flow compressor; 
         FIGS. 10A and 10B  show transient working lines for, respectively, a high-pressure compressor and a low-pressure compressor; 
         FIGS. 11A and 11B  show transient working lines for, respectively, a high-pressure compressor and a low-pressure compressor where control of power offtake and/or input is implemented; 
         FIG. 12  shows a procedure to optimise operation of the engine of  FIG. 1  during an acceleration event; 
         FIG. 13  shows a plot of fuel-air ratio against mass flow in a combustor for different operating conditions; 
         FIG. 14  shows a procedure to optimise operation of the engine of  FIG. 1  during a deceleration event; 
         FIGS. 15A and 15B  show, respectively, an increase in electrical power demand, and the corresponding transient working line for the compressor on the same spool as the motor-generator meeting said power demand; 
         FIGS. 16A and 16B  show, respectively, an increase in electrical power demand, and the corresponding transient working line for the compressor on the same spool as the motor-generator meeting said power demand with assistance from an energy storage system; 
         FIG. 17  shows a procedure to optimise operation of the engine of  FIG. 1  in the event of an increase in electrical power demand; 
         FIGS. 18A and 18B  show characteristics for, respectively, a high-pressure compressor and a low-pressure compressor and the movement of operating point when shaft power is transferred from the low-pressure spool to the high-pressure spool; 
         FIG. 19  shows a procedure to optimise power offtake and/or transfer to increase surge margin; 
         FIGS. 20A and 20B  show characteristics for, respectively, a high-pressure compressor and a low-pressure compressor and the movement of operating point when shaft power is transferred from the high-pressure spool to the low-pressure spool; 
         FIG. 21  shows a procedure to optimise power offtake and/or transfer to increase compression efficiency; and 
         FIG. 22  shows a procedure to implement a speed limiter function. 
     
    
    
     DETAILED DESCRIPTION 
     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. 
     
       FIGS. 1 &amp; 2 
     
     A general arrangement of an engine  101  for an aircraft is shown in  FIG. 1 , with an equivalent block diagram of the main components thereof being presented in  FIG. 2 . 
     In the present embodiment, the engine  101  is a turbofan, and thus comprises a ducted fan  102  that receives intake air A and generates two airflows: a bypass flow B which passes axially through a bypass duct  103  and a core flow C which enters a core gas turbine. 
     The core gas turbine comprises, in axial flow series, a low-pressure compressor  104 , a high-pressure compressor  105 , a combustor  106 , a high-pressure turbine  107 , and a low-pressure turbine  108 . 
     In use, the core flow C is compressed by the low-pressure compressor  104  and is then directed into the high-pressure compressor  105  where further compression takes place. The compressed air exhausted from the high-pressure compressor  105  is directed into the combustor  106  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  107  and in turn the low-pressure turbine  108  before being exhausted to provide a small proportion of the overall thrust. 
     The high-pressure turbine  107  drives the high-pressure compressor  105  via an interconnecting shaft  109 . The low-pressure turbine  108  drives the low-pressure compressor  104  via an interconnecting shaft  110 . Together, the high-pressure compressor  105 , interconnecting shaft  109  and high-pressure turbine  107  form part of a high-pressure spool of the engine  101 . Similarly, the low-pressure compressor  104 , interconnecting shaft  110  and low-pressure turbine  108  form part of a low-pressure spool of the engine  101 . 
     The fan  102  is driven by the low-pressure turbine  101  via a reduction gearbox in the form of a planetary-configuration epicyclic gearbox  111 . Thus in addition to the low-pressure compressor  104 , the interconnecting shaft  110  is also connected with a sun gear  112  of the gearbox  111 . The sun gear  112  is meshed with a plurality of planet gears  113  located in a rotating carrier  114 , which planet gears  113  are in turn are meshed with a static ring gear  115 . The rotating carrier  114  is connected with the fan  102  via a fan shaft  116 . 
     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  101 , a first electric machine  117  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  109  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  117  may be mounted coaxially with the turbomachinery in the engine  101 . For example, the HP motor-generator  117  may be mounted axially in line with the duct  118  between the low- and high-pressure compressors. 
     Similarly, a second electric machine  119  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  110  to receive drive therefrom. In the present embodiment, the LP motor-generator  119  is mounted in the tailcone  120  of the engine  101  coaxially with the turbomachinery. In alternative embodiments, the LP motor-generator  119  may be located axially in line with low-pressure compressor  104 , which may adopt a bladed disc or drum configuration to provide space for the LP motor-generator  119 . 
     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)  121 . In the present embodiment, the PEM  121  is mounted on the fancase  122  of the engine  101 , 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  101  is attached, for example. 
     Control of the PEM  121  and of the HP and LP motor-generator is in the present example performed by an electronic engine controller (EEC)  123 . In the present embodiment the EEC  123  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  101 , i.e. both of the core gas turbine and the motor-generators  117  and  119 . In this way, the EEC  123  may holistically respond to both thrust demand and electrical power demand. 
     An embodiment of the overall system will be described with reference to  FIG. 3 , and the control software architecture will be described with reference to  FIGS. 4 and 5 . The various control strategies implemented in response to various engine operational phenomena will be described with reference to  FIGS. 6 to 22 . 
     Various embodiments of the engine  101  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  101  may instead be a turboprop comprising a propeller for producing thrust. 
     The low- and high-pressure compressors  104  and  105  may comprise any number of stages, for example multiple stages. Each stage may comprise a row of rotor blades and a row of stator vanes, which may be variable stator vanes (in that their angle of incidence may be variable). In addition to, or in place of, axial stages, the low-or high-pressure compressors  104  and  105  may comprise centrifugal compression stages. 
     The low- and high-pressure turbines  107  and  108  may also comprise any number of stages. 
     The fan  102  may have any desired number of fan blades, for example 16, 18, 20, or 22 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 0 percent span position, to a tip at a 100 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: 0.4, 0.39, 0.38 0.37, 0.36, 0.35, 0.34, 0.33, 0.32, 0.31, 0.3, 0.29, 0.28, 0.27, 0.26, or 0.25. 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  102  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: 2.5 metres (around 100 inches), 2.6 metres, 2.7 metres (around 105 inches), 2.8 metres (around 110 inches), 2.9 metres (around 115 inches), 3 metres (around 120 inches), 3.1 metres (around 122 inches), 3.2 metres (around 125 inches), 3.3 metres (around 130 inches), 340 cm (around 135 inches), 3.5 metres (around 138 inches), 3.6 metres (around 140 inches), 3.7 metres (around 145 inches), 3.8 metres (around 150 inches) or 3.9 metres (around 155 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  102  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 2500 rpm, for example less than 2300 rpm. Purely by way of further non-limitative example, the rotational speed of the fan  102  at cruise conditions for an engine having a fan diameter in the range of from 2.5 metres to 3 metres (for example 2.5 metres to 2.8 metres) may be in the range of from 1700 rpm to 2500 rpm, for example in the range of from 1800 rpm to 2300 rpm, or, for example in the range of from 1900 rpm to 2100 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 3.2 metres to 3.8 metres may be in the range of from 1200 rpm to 2000 rpm, for example in the range of from 1300 rpm to 1800 rpm, for example in the range of from 1400 rpm to 1600 rpm. 
     In use of the engine  101 , the fan  102  (with its associated fan blades) rotates about a rotational axis. This rotation results in the tip of the fan blade moving with a velocity U tip . 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/U tip   2 , where dH is the enthalpy rise (for example the one dimensional average enthalpy rise) across the fan and U tip  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: 0.3, 0.31, 0.32, 0.33, 0.34, 0.35, 0.36, 0.37, 0.38, 0.39 or 0.4 (all units in this paragraph being Jkg −1 K −1 /(ms −1 ) 2 ). 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  101  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: 10, 10.5, 11, 11.5, 12, 12.5, 13, 13.5, 14, 14.5, 15, 15.5, 16, 16.5, or 17. 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  103 . 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  101  may be defined as the ratio of the stagnation pressure upstream of the fan  102  to the stagnation pressure at the exit of the high-pressure compressor  105  (before entry into the combustor). By way of non-limitative example, the overall pressure ratio of the engine  101  at cruise may be greater than (or on the order of) any of the following: 35, 40, 45, 50, 55, 60, 65, 70, 75. 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  101  may be defined as the net thrust of the engine divided by the total mass flow through the engine  101 . At cruise conditions, the specific thrust of the engine  101  may be less than (or on the order of) any of the following: 110 Nkg −1 s, 105 Nkg −1 s, 100 Nkg −1 s, 95 Nkg −1 s, 90 Nkg −1 s, 85 Nkg −1 s, or 80 Nkg −1 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  101  may have any desired maximum thrust. For example, the engine  101  may be capable of producing a maximum thrust of at least (or on the order of) any of the following: 160 kilonewtons, 170 kilonewtons, 180 kilonewtons, 190 kilonewtons, 200 kilonewtons, 250 kilonewtons, 300 kilonewtons, 350 kilonewtons, 400 kilonewtons, 450 kilonewtons, 500 kilonewtons, or 550 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 15 degrees Celsius (ambient pressure 101.3 kilopascals, temperature 30 degrees Celsius), with the engine  101  being static. 
     In use, the temperature of the flow at the entry to the high-pressure turbine  107  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  106 , 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: 1400 kelvin, 1450 kelvin, 1500 kelvin, 1550 kelvin, 1600 kelvinor 1650 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  101  may be, for example, at least (or on the order of) any of the following: 1700 kelvin, 1750 kelvin, 1800 kelvin, 1850 kelvin, 1900 kelvin, 1950 kelvinor 2000 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  102  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  101  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 50 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 2533 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 10000 to 15000 metres, such as from 10000 to 12000 metres, or from 10400 to 11600 metres (around 38000 feet), or from 10500 to 11500 metres, or from 10600 to 11400 metres, or from 10700 metres (around 35000 feet) to 11300 metres, or from 10800 to 11200 metres, or from 10900 to 11100 metres, or 11000 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 0.7 to 0.9, for example one of Mach 0.75 to 0.85, Mach 0.76 to 0.84, Mach 0.77 to 0.83, Mach 0.78 to 0.82, Mach 0.79 to 0.81, Mach 0.8, Mach 0.85, or in the range of from Mach 0.8 to 0.85. 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 0.7 or above Mach 0.9. 
     Thus, for example, the cruise conditions may correspond specifically to a pressure of 23 kilopascals, a temperature of minus 55 degrees Celsius, and a forward Mach number of 0.8. 
     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. 
     
       FIG. 3 
     
     A block diagram of the interface of the EEC  123  and other engine systems is shown in  FIG. 3 . 
     As described previously, in the present embodiment, the EEC  123  is coupled with the PEM  121  to control the HP motor-generator  117  and the LP motor-generator  119 . In this way, power may either be extracted from or added to each of shafts  109  and  110 . 
     In the present embodiment, the PEM  121  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  121  comprises a first bidirectional power converter  301  connected with the HP motor-generator  117 , and a second bidirectional power converter  302  connected with the LP motor-generator  119 . The dc sides of the power converters  301  and  302  are in the present example connected with each other to facilitate bidirectional power transfer between the motor-generators  117  and  119 . 
     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 250 kilowatts to 2 megawatts. In a more specific embodiment, they are rated from 300 kilowatts to 1 megawatt. In a more specific embodiment, they are rated at 350 kilowatts. 
     In other embodiments, the HP motor-generator  117  and the first bidirectional power converter  301  are rated at a different continuous power than the LP motor-generator  119  and the second bidirectional power converter  302 . In a specific embodiment, they are rated at from 250 kilowatts to 2 megawatts. In a more specific embodiment, they are rated from 350 kilowatts to 1 megawatt. In a more specific embodiment, they are rated at 350 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  301  and  302  are also connected with a dc bus  303 . 
     In the present example the dc bus  303  has connected to it various loads, which may be either located on the engine  101  or on the vehicle instead. Some such as anti-icing systems  304  may be part of the engine, such as electric nacelle anti-icing systems, or part of the aircraft on which the engine  101  is installed, such as electric wing anti-icing systems. 
     Other loads may be connected with the dc bus  303  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  305 . In the present example, control of the charge/discharge state of the battery  305  is facilitated by a dc-dc converter  306 . Other energy storage devices may be connected to the dc bus  303  as well as or in place of the battery  305 , such as a capacitor. 
     Other loads, indicated at  307 , may be connected with the dc bus  303  such as cabin environmental control systems, electric actuation systems, auxiliary power units, etc. 
     In operation, the EEC  123  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 (P D ). In addition, the EEC receives a plurality of sets of measured parameters, namely a set concerning flight parameters of the vehicle, Σ AIRCRAFT , and a set concerning operational parameters of the engine, Σ ENGINE . As will be described further with reference to  FIGS. 4 and 5 , 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 W F  to be metered by a fuel metering unit  308  on the engine  101 , or variable-stator vane angle, handling bleed settings, etc. in the present embodiment the EEC  123  comprises a power controller module  309  to generate a control signal P H  for the first bidirectional power converter  301  and a second control signal P L  for the second bidirectional power converter  302 . The control signals P H  and P L  control the operation of the power converters in terms of both direction and magnitude of electrical power. In this way, the EEC  123  may meet the demanded power P D  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 P H  and P L , in the present example the power controller  309  is configured to derive a control signal P BAT  for the dc-dc converter  306  to facilitate charge or discharge of the battery  305 . 
     In a specific embodiment, the power controller  309  is configured to, in normal operation, set P BAT  to zero, and only change its status as set out in the optimisation routines described herein, for example the routines described with reference to  FIGS. 8 and 12 . 
     In another specific embodiment, the power controller  309  includes battery optimisation functionality and modifies the power demand parameter P D  by adding or subtracting a value P BAT  depending upon whether it is more optimal to charge, discharge, or maintain the charge of the battery  305 . 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  309  modifies the power demand parameter P D  by performing an addition-assignment operation P D +=P BAT . The sign convention used herein is such that a positive P BAT  means that the battery  305  is to be charged, and thus additional generation is required from the HP and LP motor-generators, whilst a negative P BAT  means that the battery  305  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 P D  will not be modified at all. 
     In the present example another control signal P AI  is generated to activate the anti-icing systems  304 . In the present example only the nacelle anti-ice system of the engine  101  is controlled, however it is envisaged that the EEC  123  may in alternative implementations have authority over wing anti-ice in certain circumstances. In a similar way to the modification of P D  by the addition-assignment of P BAT , the power controller  309  is configured to perform the addition assignment P D +=P AI  so as to account for the power required for the anti-icing systems  304 . Again, it will be appreciated that in embodiments lacking electric anti-ice systems, this signal will not be generated and no modification of P D  will be performed in this manner. 
     
       FIG. 4 
     
     In practice, the EEC  123  houses microprocessors for executing program modules to control the engine  101 . A block diagram of the functional modules of the power controller  309  is shown in  FIG. 4 . 
     Input parameters previously described with reference to  FIG. 3  are initially received by a classifier module  401  to output an optimiser setting mode for an optimiser module  402 . The operation of the classifier module  401  will be described further with reference to  FIG. 5 , and the various modes of the optimiser module will be described with reference to  FIGS. 7 to 22 . 
     The input parameters are also supplied to a filter  403  prior to updating an engine model module  404 . The filter  403  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  404  runs a real time model of the engine  101  so as to facilitate prediction of changes is operational parameters, such as W F , P H , and P L  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  402  finds the optimal set of parameters for operation of the engine  101  given the current operational state of aircraft on which the engine is installed and the engine itself. 
     
       FIG. 5 
     
     The classifier module  401  is detailed in  FIG. 5 . 
     In the present embodiment, the classifier module  401  comprises a comparator  501  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 5000 feet, whilst the Mach number threshold is 0.3. 
     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 30000 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. 6 . 
     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 30000 feet, whilst the Mach number range is from 0.8 to 0.9. 
     In such circumstances, the optimisation strategy that can be employed is to minimise fuel flow W F  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  FIGS. 19 and 21 , 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  108  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  108  and the rest of the engine  101 . 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 0.2. 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. 7 . 
     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 0 to 5000 feet above ground level, and the Mach number range is from 0.2 to 0.3. 
     The optimisation strategy for the approach idle condition will be described further with reference to  FIG. 8 . 
     The classifier module  401  further comprises a differentiator  502  which is configured to monitor the PLA and P D  parameters and identify a transient type. 
     In response to the change in power lever angle being positive, the differentiator  502  is configured to identify the initiation of an acceleration event. The optimisation strategy for this manoeuvre will be described further with reference to  FIG. 12 . 
     In response to the change in power lever angle being negative, the differentiator  502  is configured to identify the initiation of a deceleration event. The optimisation strategy for this manoeuvre will be described further with reference to  FIG. 14 . 
     In response to a change in electrical power demand P D , the differentiator  502  is configured to cause the optimiser to invoke the optimisation strategy described with reference to  FIG. 17 . 
     The classifier module  401  further comprises a limiter  503  which is configured to monitor the high-pressure and low-pressure spool speeds, N H  and N L . 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. 22  may be invoked. 
     In the present example, the outputs of the comparator  501 , differentiator  502  and limiter  503  are compared by a prioritiser  504 . 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  503  are priorities over outputs from the differentiator  502 , which are in turn prioritised over outputs from the comparator  501 . 
     
       FIG. 6 
     
     Following the identification of a maximum take-off or maximum climb condition by the classifier module  401 , the optimiser  402  enters the corresponding optimisation routine at step  601 . At step  602 , a question is asked as to whether the power demand P D  is less than the maximum power rating of the LP motor-generator  119 , P Lmax . If so, then control proceeds to step  603  where the optimiser  402  maximises the power generation by the LP motor-generator  119 , P L . 
     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 P D , this results in a higher high-pressure spool rotational speed N H . 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 350 kilowatts, it is possible to reduce the stator outlet temperature by 2 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 P D  after accounting for battery charge/discharge and/or anti-icing system operation is greater than the maximum power rating of the LP motor-generator  119 , P Lmax , then the question asked at step  602  will be answered in the negative. In this case, control proceeds to step  604  where the optimiser  402  maximises the power generation by the LP motor-generator  119 , P L , and minimises the power generation by the HP motor-generator  117 , P H . Thus, the LP motor-generator  119  is directed to generate P Lmax  and the HP motor-generator  117  is directed to generate the remainder, P D -P Lmax . 
     In an embodiment, should spare capacity be available from the LP motor-generator  119 , then the optimiser  402  may elect to divert P Lmax -P D  to the HP motor-generator  117  which may further increase core flow and reduce stator outlet temperature. 
     
       FIG. 7 
     
     Following the identification of an unchoked regime for the low-pressure turbine  108  by the classifier module  401 , the optimiser  402  enters the corresponding optimisation routine at step  701 . As described previously, this may be triggered by a low Mach number idle operating condition, for example the ground idle operating point. 
     At step  702 , a question is asked as to whether the power demand P D  is less than the maximum power rating of the HP motor-generator  117 , P Hmax . If so, then control proceeds to step  703  where the optimiser  402  maximises the power generation by the HP motor-generator  119 , P H . 
     As the low-pressure turbine  108  is unchoked, there is a large impact when it is required to provide electrical power via the LP motor-generator  119 . The reduction in the speed of the low-pressure spool, N L , is such that there is a steep drop in the efficiency of the low-pressure compressor  104 , and thus an increase in fuel burn. In practice, there may also be a drop in efficiency of the LP motor-generator  119  due to the lower rotational speed. Thus, at the low flight Mach numbers that cause the low-pressure turbine  108  to unchoke, it is possible to improve fuel consumption by this optimisation routine. 
     In a similar way to the situation of  FIG. 6 , should the power demand P D  be greater than the maximum power rating of the HP motor-generator  117 , P Hmax , then the question asked at step  702  will be answered in the negative. In this case, control proceeds to step  704  where the optimiser  402  maximises the power generation by the HP motor-generator  117 , P H , and minimises the power generation by the LP motor-generator  119 , P H . Thus, the HP motor-generator  117  is directed to generate P Hmax  and the LP motor-generator  119  is directed to generate the remainder, P D -P Hmax . 
     In an embodiment, should spare capacity be available from the HP motor-generator  117 , then the optimiser  402  may elect to divert P Hmax -P D  to the LP motor-generator  119  which may further reduce fuel consumption. 
     
       FIG. 8 
     
     Following the identification of an approach idle condition by the classifier module  401 , the optimiser  402  enters the corresponding optimisation routine at step  801 . At step  802 , a question is asked as to whether the power demand P D  is less than the maximum power rating of the LP motor-generator  119 , P Lmax . If so, then control proceeds to step  603  where the optimiser  402  maximises the power generation by the LP motor-generator  119 , P L . The excess capacity of the LP motor-generator  119 , P Lmax -P D , is transferred to the HP motor-generator  117 . 
     This has two effects. First, extraction of the maximum power possible from the low-pressure spool significantly reduces the low-pressure spool rotational speed, N L . Recalling that in the engine  101  the fan  102 , which is primary thrust generating element, rotates with the low-pressure spool, it will be clear that a reduction in N L  reduces the thrust generated by the engine  101 . 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 N H  has a powerful effect on the response time during an engine acceleration, so maximising N H  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 N H  and thus a shorter response time, along with a reduced idle thrust due to the lower N L . It has been demonstrated that in an engine of the type described herein, 105 kilowatts of power transfer achieves a sufficiently high N H  and a 75 percent reduction in idle thrust. 
     If the power demand P D  is greater than the maximum power rating of the LP motor-generator  119 , P Lmax , then the question asked at step  802  will be answered in the negative. In this case, control proceeds to step  804  where the optimiser  402  maximises the power generation by the LP motor-generator  119 , P L , and minimises the power generation by the HP motor-generator  117 , P H . Thus, the LP motor-generator  119  is directed to generate P Lmax  and the HP motor-generator  117  is directed to generate the remainder, P D -P Lmax . 
     In an embodiment, the optimiser  402  is further configured to identify that maintaining the requisite N H  will cause unsafe operation of the low-pressure compressor  104 . This may be caused by the operating point of the LP low-pressure compressor  104  becoming too close to surge or to choke. In response to the onset of such a condition, the fuel flow W F  may be increased. 
     Alternatively, in order to reduce fuel burn on approach, the optimiser  402  may be configured to supplement the HP motor-generator using an energy storage device, for example the battery  305  via control of the P BAT  parameter, or another energy source such as the auxiliary power unit on the aircraft. In this way, community emissions on approach may be reduced. 
     
       FIG. 9 
     
     One of the primary considerations for control of a gas turbine engine is the prevention of surge in the compression stages. A characteristic  901  for an axial flow compressor, such as low-pressure compressor  104  or high-pressure compressor  105 , is shown in  FIG. 9 . 
     The characteristic  901  plots pressure ratio against flow function, which in this case is non-dimensional flow (W√T/P). The characteristic  901  shows a plurality of non-dimensional speed lines  902 ,  903 ,  904 ,  905 , along with the compressor steady state working line  906 , which is the locus of operating points for various steady state throttle settings at different non-dimensional speeds. In addition, the surge line  907  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  906  and the value of R on the surge line  907  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 15 percent surge margin, and high-pressure compressors to have around 20 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  FIGS. 10A, 10B, 11A, and 11B . 
     
       FIGS. 10A &amp; 10B 
     
     Characteristics for the high-pressure compressor  105  and the low-pressure compressor  106  showing transient phenomena during acceleration events are plotted in  FIGS. 10A and 10B  respectively. 
     In  FIG. 10A , the high-pressure compressor steady-state working line  1001  is shown along with lines of constant corrected speed  1002 ,  1003 ,  1004 , and  1005 , and the surge line  1006 . During an acceleration manoeuvre, the high-pressure compressor  105  moves from an initial operating point  1007  to a final operating point  1008  via a transient working line  1009  above the steady-state working line  1001 . 
     Similarly, in  FIG. 10B , the low-pressure compressor steady-state working line  1011  is shown along with lines of constant corrected speed  1012 ,  1013 ,  1014 , and  1015 , and the surge line  1016 . During an acceleration manoeuvre, the low-pressure compressor  106  moves from an initial operating point  1017  to a final operating point  1018  via a transient working line  1019  which crosses the steady-state working line  1011 . 
     During the acceleration manoeuvre, the high-pressure compressor  105  initially moves towards surge due to the flow compatibility requirement with the high-pressure turbine  107 . The flow function of the high-pressure turbine  107  (W 405 √T 405 /P 405 ) is substantially fixed during most operating conditions of the engine  101 , 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  308  (W F ) is increased. This leads to an increase in turbine entry temperature, T 405 . Normally, the high-pressure spool speed N H  is prevented from changing instantaneously due to its inertia. Consequently, the operating point of the high-pressure compressor  105  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  1009  parallel to the surge line  1006 . As the acceleration finishes, the high-pressure compressor  105  adopts its final operating point  1008  on the steady-state working line  1001 . 
     Referring now to  FIG. 10B , at the initiation of the acceleration manoeuvre, the operating point of the low-pressure compressor  104  moves a little towards surge, and then crosses the steady-state working line  1011 . The initial move towards surge is due to the reduction in flow in the high-pressure compressor  105  due to the high-pressure spool inertia as described above. As the speed of the high-pressure compressor  105  increases it can accept more flow. However, due to the greater inertia of the low-pressure spool (recalling that it drives the fan  102  via the gearbox  111 ), it cannot accelerate at the same rate and so the operating points of the low-pressure compressor  104  during the acceleration manoeuvre fall below the steady-state working line  1011 . 
     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. 
     
       FIGS. 11A &amp; 11B 
     
     By contrast, in the engine  101  it is possible to utilise the HP motor-generator  117  and optionally the LP motor-generator  119  to reduce the transient excursion. 
       FIG. 11A  shows the characteristic for the high-pressure compressor  105  when the HP motor-generator  117  is used to overcome the high-pressure spool inertia. It can be seen that for the same degree of overfuelling, the transient working line  1101  is much closer to the steady-state working line  1001  and further from the surge line  1006 . 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. 11B  shows the characteristic for the low-pressure compressor  104  during application of the same technique on the low-pressure spool using the LP motor-generator  119 . It may be seen that the transient working line  1102  is again much closer to the steady-state working line  1011  due to the reduction in effective low-pressure spool inertia by the LP motor-generator  119 . 
     
       FIG. 12 
     
     Steps carried out by the optimiser  402  to achieve the advantages described previously for an acceleration event are set out in  FIG. 12 . 
     Following the identification of an acceleration condition by the classifier module  401 , the optimiser  402  enters the corresponding optimisation routine at step  1201 . In the present embodiment, with an energy storage system such as the battery  305  available on the dc bus  303 , a question is asked at step  1202  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 20 percent, below which the battery may be damaged. 
     If so, a further question is asked at step  1203  as to whether the unmodified total aircraft power demand P D  (i.e. prior to any modification thereof to account for battery optimisation) is less than the maximum power available from the battery  305 , P BATmax . If so, then control proceeds to step  1204  where the optimiser  402  overrides any concurrent battery optimisation processes, and fully supplies the power demand Po using the battery  305 , with any excess being supplied to the HP motor-generator  117  to overcome the HP spool inertia. Optionally, any further excess may be supplied to the LP motor-generator  119 . 
     Should either of the questions asked at steps  1202  or  1203  be answered in the negative, i.e. the battery  305  has a minimal state of charge, or the unmodified total aircraft power demand P D  is greater than the maximum power of the battery  305 , P BATmax  (or if indeed the particular embodiment of the engine  101  does not include a battery), then control proceeds to step  1205  in which power generation by the LP motor-generator  119 , P L , is maximised and power generation by the HP motor-generator  117 , P H , is minimised. 
     Following optimisation of the power generation strategy to satisfy the power demand P D  in the preceding steps, the fuel flow W F  metered by the fuel metering unit  308  is increased at step  1206 . 
     
       FIG. 13 
     
     When a deceleration event is initiated, fuel flow by the fuel metering unit  308  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  107  are choked and the flow function remains constant. As the high-pressure spool speed N H  is prevented from changing instantaneously due to its inertia, the reduction in turbine entry temperature T 405  causes the operating point of the high-pressure compressor  105  to move down a line of constant corrected speed on its characteristic. This manifests as an increase in mass flow W 31  and a decrease in pressure P 31  at the exit of the high-pressure compressor  105 . 
     The result of this for the combustor  106  is that not only is the amount of fuel delivered lower, but the mass flow W 31  therethrough has increased. This means that the combustor  106  operates at a lower fuel-air ratio (FAR) than normal, which risks weak extinction (also known as lean blowout). 
     Referring to  FIG. 13 , which is a plot of FAR against flow function, the weak extinction boundary  1301  for the combustor  105  is illustrated. Corrected flow through the combustor  105  to the right of the weak extinction boundary  1301  results in extinction of the flame and is an unacceptable operating condition. A steady-state FAR for a particular mass flow through the combustor  106 , is shown at point  1302 . 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 1  due to the proximity of the fuel-air ratio to the weak extinction boundary  1301  during the deceleration, shown at point  1303 . 
     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  107 , the operating point of the high-pressure compressor  105  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 405 . As the speed of the high-pressure compressor  105  substantially instantaneously begins to decelerate, the mass flow through the combustor  106  also reduces substantially instantaneously. As shown at point  1304  in  FIG. 13 , the additional margin M 2  allows a greater degree of underfuelling. 
     It will be understood that this approach allows the design of the combustor  106  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  105 . Using the engine model  404 , for example, the onset of weak extinction may be identified by evaluating the current fuel-air ratio in the combustor  105 . This may be achieved, for example, by utilising the flight Mach number, altitude and temperature to determine the mass flow in the engine  101 , the characteristic of the fan  102  to determine the mass flow C into the core gas turbine, and the characteristics of the compressors  104  and  105  to determine the mass flow into the combustor  105 . This may be combined with the commanded fuel flow W F  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  1301 , the EEC  123  may use the power controller  309  to extract mechanical shaft power from the high-pressure spool using the HP motor-generator  117  to prevent a further drop in fuel-air ratio in the combustor  105 . 
     
       FIG. 14 
     
     Steps carried out by the optimiser  402  to achieve the advantages described previously for a deceleration event are set out in  FIG. 14 . 
     Following the identification of a deceleration condition by the classifier module  401 , the optimiser  402  enters the corresponding optimisation routine at step  1401 . A question is asked at step  1402  as to whether the power demand P D  is less than the maximum power generation capability of the HP motor-generator  117 , P Hmax . If so, then control proceeds to step  1403  whereupon the power generation of the HP motor-generator  117 , P H , is maximised to satisfy P D . 
     In the present embodiment, the excess capacity P Hmax -P D  is transferred to other loads. In an embodiment, the excess capacity is directed to an energy storage system, such as the battery  305 . 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  304 , which may be the nacelle anti-ice system of the engine  101 . Alternatively, it may be the wing anti-ice system of the vehicle on which the engine  101  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  1403  solely maximises P H  up to P D  to assist in the reduction of the high-pressure spool speed. 
     If the question asked at step  1402  is answered in the negative, to the effect that the power demand P D  is greater than the maximum power generation capability of the HP motor-generator  117 , P Hmax , then control proceeds to step  1404  where first the power generation of the HP motor-generator  117 , P H , is maximised, then the power generation of the LP motor-generator  119 , P L , is maximised to supply remainder of P D . 
     Following optimisation of the power generation strategy to satisfy the power demand P D  in the preceding steps, the fuel flow W F  metered by the fuel metering unit  308  is reduced at step  1206 . 
     In an alternative, embodiment the excess capacity P Hmax -P D  may be directed to the LP motor-generator  119 . This may be possible due to this excess power representing a small proportion of the power generated by the low-pressure turbine  108 , therefore leading to a very small change in thrust generated by the fan  102 . 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. 
     
       FIGS. 15A &amp; 15B 
     
     When an increase in power demand P D  occurs, the power controller  309  must respond by in turn demanding an increase in power output by the gas turbine engine. 
       FIG. 15A  illustrates an exemplary increase in power demand P D  of magnitude dP D  within a timeframe dt.  FIG. 15B  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 dP D  the specific work of the turbine must increase. The steady-state working line is shown at  1501 , with the surge line shown at  1502 . 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. 15A , 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. 15B . Line  1503  shows the movement of the operating point at constant corrected speed. Line  1504  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  1502 . 
     Thus it may be seen that at low engine throttle settings in particular, such an increase in power demand P D  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. 
     
       FIGS. 16A &amp; 16B 
     
     In the present embodiment, however, the approach is taken to utilise the energy storage system to mitigate the possibility of surge. Thus, as illustrated in  FIG. 16A , the same increase in power demand P D  of magnitude dP D  within a timeframe dt is demanded. Instead of this solely being met by one or both of the HP motor-generator  117  and the LP motor-generator  119 , it is met during the manoeuvre by the battery  305 . Thus, as shown in the Figure, initially the power demand is met by the battery  305 , as shown by the shaded region  1601 . As the engine  101  accelerates, the proportion provided by the motor-generator(s) increases gradually until the new power demand is fully met by the engine  101 . 
       FIG. 16B  shows the transient working line  1602  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  FIGS. 15A and 15B . 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  305  may provide all of the higher power demand whilst the engine  101  accelerates to a higher corrected speed, at which point provision of the power demand Po is switched from the battery to the motor-generator(s) in the engine  101 . 
     
       FIG. 17 
     
     Steps carried out by the optimiser  402  to achieve the functionality described previously for an increase in power demand P D  are set out in  FIG. 17 . 
     Following the identification of an increase in power demand within a given timeframe dt by the differentiator  502  in the classifier module  401 , the optimiser  402  enters the corresponding optimisation routine at step  1701 . At step  1702 , the optimiser  401  evaluates the operating points of the low-pressure compressor  104  and the high-pressure compressor  105  for the demanded P D . In the present embodiment, this may be achieved using the engine model  404  and knowledge of the current power lever angle setting etc. Alternatively, a look-up table or similar may be used instead. 
     At step  1702 , the current surge margin in the low-pressure compressor  104 , dR L /R L , and the current surge margin in the high-pressure compressor  105 , dR H /R H  are evaluated, again using the engine model  404  in the present embodiment, or suitable alternatives if required. 
     At step  1703 , 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  1706  where a decision is taken to utilise the battery  305  (or other energy storage unit such as a capacitor) to satisfy the demanded P D . 
     Then, or if the question asked at step  1705  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  308 . As described previously, at this point the new power demand may then be fully met by one or more of the HP motor-generator  117  and the LP motor-generator  119 . The transition may be gradual, or the battery  305  may solely supply the additional power demand dP D  until the new operating points are achieved. 
     
       FIGS. 18A &amp; 18B 
     
     The effect of power transfer from the LP motor-generator  119  to the HP motor-generator  117  on the operating point of the high-pressure compressor  105  is shown in  FIG. 18A  on the compressors&#39; characteristic. The effect on the operating point of the low-pressure compressor  104  is shown on its characteristic in  FIG. 18B . 
     As power is added to the high-pressure spool, the pressure ratio and flow function increase, as shown in  FIG. 18A  by the transition from an initial operating point  1801  to a final operating point  1802  at a higher non-dimensional speed on the compressor&#39;s working line  1803 . 
     Extraction of power from the low-pressure spool lowers the working line of the low-pressure compressor  104 . Recalling that the low-pressure compressor rotational speed is fixed relative to the fan  102 , at constant thrust the low-pressure compressor operating point may only move on a constant non-dimensional speed line, in this case speed line  1804 . Due to the increase in flow function in the high-pressure compressor  105 , the low-pressure compressor  104  is unthrottled and so also sees a raise in flow function. Thus the operating point moves from an initial operating point  1805  to a final operating point  1806  on speed line  1804  away from the surge line  1807 . 
     It will therefore be understood that controlling the degree of electrical power generated by one or both of the HP motor-generator  117  and the LP motor-generator  119  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  101  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  101  to be varied. This has particular advantages in terms of optimising the jet velocity of the engine  101  for particular airspeeds. 
     In an embodiment, power transfer may be used to further vary the bypass ratio by operating the LP motor-generator  119  as a generator and operating the HP motor-generator  117  as a motor. 
     In this way, it will be understood that the engine  101  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 P D  be greater than or equal to the capability of the LP motor-generator  119 , then an increase in surge margin is still achieved by satisfying the power demand P D  by maximising low-pressure spool offtake. This is because a greater enthalpy drop is required across the low-pressure turbine  108 , which requires a greater mass flow. The greater mass flow through the high-pressure compressor  105 , whilst not as high as with power transfer, still unthrottles the low-pressure compressor  104  and increases its surge margin. Thus it will be understood that this strategy provides a suitable means for increasing surge margin in the engine  101 . 
     
       FIG. 19 
     
     Steps carried out by the optimiser  402  to increase surge margin are therefore set out in  FIG. 19 . 
     Following the identification of an operating condition in which surge margin needs to be increased by the classifier module  401 , the optimiser  402  enters the corresponding optimisation routine at step  1901 . 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  1902 , a question is asked as to whether the current power demand P D  is less than the maximum power rating of the LP motor-generator  119 , P Lmax . If so, then control proceeds to step  1903  where the optimiser  402  maximises the power generation by the LP motor-generator  119 , P L  to increase surge margin in the low-pressure compressor  104 , and transfers any excess electrical power P Lmax -P D  to the HP motor-generator  117  to raise its operating point up its working line, also increasing surge margin. 
     If the question asked at step  1902  is answered in the negative, to the effect that the LP motor-generator  119  is not solely capable of satisfying the power demand P D , then control proceeds to step  1904  where the optimiser  402  maximises the power generation by the LP motor-generator  119 , P L  to increase surge margin in the low-pressure compressor  104 . Recall that power extraction from the high-pressure spool normally moves the operating point of the high-pressure compressor  105  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  1904  the optimiser  402  minimises power generation by the HP motor-generator  117 , P H  which substantially maintain its operating point at around its steady state value, or slightly higher on its working line. 
     
       FIGS. 20A &amp; 20B 
     
     Whilst surge margin may be increased by the method of  FIG. 19 , 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  117  to the LP motor-generator  119 .  FIG. 20A  shows the movement of the operating point of the high-pressure compressor  105  in this scenario.  FIG. 20B  shows the movement of the operating point of the low-pressure compressor  104  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  101  by moving the operating points of the compressors into regions of high efficiency. 
     
       FIG. 21 
     
     Steps carried out by the optimiser  402  to increase compression efficiency are therefore set out in  FIG. 21 . 
     Following the identification of an operating condition in which compression efficiency may be increased by the classifier module  401 , the optimiser  402  enters the corresponding optimisation routine at step  2101 . As described previously, operating conditions such as sufficiently steady inlet flow may permit entry into this routine. 
     At step  2102 , a question is asked as to whether the current power demand P D  is less than the maximum power rating of the HP motor-generator  117 , P Hmax . If so, then control proceeds to step  2103  where the optimiser  402  maximises the power generation by the HP motor-generator  117 , P H  to increase compression efficiency in the high-pressure compressor  104 , and transfers any excess electrical power P Hmax -P D  to the LP motor-generator  119  to lower its operating point on its working line, also increasing compression efficiency in this example. 
     If the question asked at step  2102  is answered in the negative, to the effect that the HP motor-generator  117  is not solely capable of satisfying the power demand P D , then control proceeds to step  2104  where the optimiser  402  maximises the power generation by the HP motor-generator  117 , P H  to increase compression efficiency in the high-pressure compressor  105 . Furthermore, the optimiser  402  minimises power generation by the LP motor-generator  119 , P L  to keep the low-pressure compressor  104  in as high a region of compression efficiency as possible. 
     
       FIG. 22 
     
     As described previously, it is also possible to utilise the HP motor-generator  117  and LP motor-generator  119  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  503  monitors the shaft speeds N H  and N L . 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  401 , the optimiser  402  enters the corresponding optimisation routine at step  2201 . At step  2202 , a question is asked as to whether the limit is either the low-pressure shaft speed, N L  (either mechanical or aerodynamic), or the high-pressure shaft speed, N H  (either mechanical or aerodynamic). 
     If the trigger was low-pressure shaft speed, N L , then control proceeds to step  2203  in which the optimiser  302  maximises the power generation by the LP motor-generator  119 , P L  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  305 , or alternatively it may be diverted to other systems such as anti-icing systems or potentially the HP motor-generator  117 . 
     If the trigger was high-pressure shaft speed, N H , then control instead proceeds to step  2204  in which the optimiser  302  maximises the power generation by the HP motor-generator  117 , P H  to decrease the high-pressure shaft speed. Again, the power generated thereby may be diverted to storage or to loads. 
     Various examples have been described, each of which feature various combinations of features. It will be appreciated by those skilled in the art that, except where clearly mutually exclusive, any of the features may be employed separately or in combination with any other features and the invention extends to and includes all combinations and sub-combinations of one or more features described herein.