Patent Publication Number: US-2021172334-A1

Title: Gas turbine engine

Description:
CROSS-REFERENCE TO RELATED APPLICATIONS 
     This specification is based upon and claims the benefit of priority from United Kingdom patent application number 1917911.8, filed on 6 Dec. 2019, the entire contents of which are incorporated herein by reference. 
     TECHNICAL FIELD 
     This disclosure relates to gas turbine engines. 
     BACKGROUND 
     In aerospace, more electric engine (MEE) and more electric aircraft (MEA) concepts are increasingly being adopted due to the reductions in both fuel consumption and complexity that they facilitate. For example, one known aircraft&#39;s engines include electric machines operable as both motors and generators to facilitate both generation of electrical power during flight and starting of the engine. This permits removal of the air-turbine starter. One engine configuration for this known aircraft 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. 
     Analysis has shown that further reduction in fuel consumption may be achieved by generating electrical power using the low-pressure spool rather than using the high-pressure spool. Further, it has also been shown that transfer of power from the low- and high-pressure spools may improve fuel efficiency at a number of phases in the operational envelope. In an example, approximately 1 megawatt of electrical power may be produced from the low-pressure shaft, with 400 kilowatts being transferred to the high-pressure spool and the remainder supplied to the airframe. 
     It is also contemplated that future airframe designs may incorporate fuselage boundary layer ingestion systems to reduce wake drag and thus reduce fuel consumption. Most practical proposals are based on tube-and-wing twinjets, in which the underwing turbofan engines also operate as turboelectric generators for supplying power to a boundary layer ingestion fan at the tail of the aircraft. Such aft-mounted fans may command in excess of 2 megawatts to operate effectively. Thus, each turbofan engine may be required to produce a further megawatt of electrical power for the airframe. 
     In such scenarios, therefore, medium voltage (as defined by IEC 60038:2009, i.e. 1 kilovolt ac or greater) electrical systems will be required to maintain acceptable current ratings, as too high a current has a detrimental effect in terms of Joule losses and conductor weight. However, it is desirable to reduce the number of systems operating at medium voltage due to the risk of arcing and corona at altitude, thus it is prudent to retain lower voltage ratings for the lower powered systems. 
     Thus, it is an object of the invention to provide an electrical system for a gas turbine engine that facilitates transfer of power from one electrical network to another operating at a different voltage. 
     SUMMARY 
     In an aspect, there is provided a gas turbine engine having a spool, further comprising: 
     a first rotary electric machine mechanically coupled with the spool and having connected therewith a first power converter to interface the electric machine with a first dc network operating at a first voltage; 
     a second rotary electric machine mechanically coupled with the spool and having connected therewith a second power converter to interface the electric machine with a second dc network operating at a second voltage greater than said first voltage; 
     wherein the first electric machine and the second electric machine are operable as a motor-generator set for transfer of electrical power to the first dc network from the second dc network. 
     In an embodiment, the first electric machine has a lower power rating than the second electric machine. 
     In an embodiment, the first power converter is a unidirectional rectifier. In an embodiment, the second power converter is a unidirectional inverter. 
     In an embodiment, the first power converter and the second power converter are bidirectional converters, whereby the first electric machine and the second electric machine are operable as a motor-generator set for transfer of electrical power between the first dc network to the second dc network. 
     In an embodiment, the spool is a low-pressure spool, and the gas turbine engine further comprises a high-pressure spool. 
     In an embodiment, the gas turbine engine further comprises a third rotary electric machine mechanically coupled with the high-pressure spool and having connected therewith a third power converter to interface the third electric machine with the first dc network. 
     In an embodiment, the third power converter is a unidirectional inverter. 
     In an embodiment, the third power converter is a bidirectional converter. 
     In an embodiment, the electric machines comprise permanent magnet rotors. 
     In an embodiment, the first electric machine and the second electric machine are configured as a dynamotor. 
     In an embodiment, the second dc network includes an engine electronic controller. 
     In an embodiment, the first dc network includes one or more electrical accessory drives connected with respective engine accessories. 
     In an embodiment, the engine accessories are one or more engine accessories selected from: 
     a fuel pump; 
     an oil pump; 
     a hydraulic pump; 
     a cabin blower. 
     In an embodiment, the second dc network includes an energy storage device, whereby the motor-generator set is operable to supply power to the first dc network from the energy storage device. 
     In another aspect, there is provided a method comprising transferring electrical power to a first dc network operating at a first voltage from a second dc network operating at a second voltage greater than the first voltage, the method further comprising: 
     providing a gas turbine engine of the aforesaid type; 
     providing electrical power at the second voltage to the second power converter from the second dc network; 
     causing rotation of the second electric machine by the second power converter and the first electric machine thereby; 
     generating electrical power by the first electric machine; 
     providing electrical power to the first dc network at the first voltage by the first power converter. 
    
    
     
       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  shows an electrical system for the engine of  FIG. 1 ; 
         FIG. 3  shows a first embodiment of a motor-generator set for the engine of  FIG. 1 ; and 
         FIG. 4  shows a second embodiment of a motor-generator set for the engine of  FIG. 1 . 
     
    
    
     DETAILED DESCRIPTION 
     
       FIG. 1 
     
     A general arrangement of an engine  101  for an aircraft is shown in  FIG. 1 . In the present embodiment, the engine  101  is of turbofan configuration, and thus comprises a ducted fan  102  that receives intake air A and generates two pressurised airflows: a bypass flow B which passes axially through a bypass duct  103  and a core flow C which enters a core gas turbine. It will be appreciated by those skilled in the art that the principles of the present invention may be applied to other engine configurations, such as aft-fan, prop-fan, counter-rotating ducted fan, etc. or even to non-aircraft engines such as land- or sea-based gas turbine engines. 
     Referring again to the Figure, 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 operation, 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. The low-pressure turbine  108  drives the low-pressure compressor  104  via another interconnecting shaft. Together, the high-pressure compressor  105 , high-pressure turbine  107 , and associated interconnecting shaft form part of a high-pressure spool of the engine  101 . Similarly, the low-pressure compressor  104 , low-pressure turbine  108 , and associated interconnecting shaft form part of a low-pressure spool of the engine  101 . Such nomenclature will be familiar to those skilled in the art. 
     In the present embodiment, the fan  102  is driven by the low-pressure turbine  108  via a reduction gearbox in the form of a planetary-configuration epicyclic gearbox  109 . Thus in this configuration, the low-pressure turbine  108  is connected with a sun gear of the gearbox  109 . The sun gear is meshed with a plurality of planet gears located in a rotating carrier, which planet gears are in turn are meshed with a static ring gear. The rotating carrier drives the fan  102  via a fan shaft  110 . 
     It will be appreciated that in alternative embodiments a star-configuration epicyclic gearbox (in which the planet carrier is static, and the ring gear rotates and provides the output) may be used instead. In other alternative embodiments, the fan  102  may be driven directly by the low-pressure turbine  108 , thus forming part of the low-pressure spool. Further alternative arrangements may involve provision of a further turbine to specifically for driving the fan, known as a three-shaft or triple-spool configuration. 
     In the present embodiment, a first rotary electric machine  111  is mechanically coupled with the low-pressure spool. In the present embodiment, the first electric machine  111  is mounted in the tail cone  112  of the engine  101  coaxially with the turbomachinery and is coupled to the low-pressure turbine  108 . In alternative embodiments, the first electric machine  111  may be located axially in line with low-pressure compressor  104 , which may adopt a bladed disc or bladed drum configuration to provide space for the first rotary electric machine  111 . ac 
     As set out in the introduction, higher levels of electrical power may be required for certain features on advanced aircraft, such as electrically-powered propulsors. To this end, the engine  101  further comprises a second rotary electric machine  113  which is also mechanically coupled to the low-pressure turbine  108 . In the present embodiment, the second electric machine  113  is mounted in the tail cone  112  of the engine  101  coaxially with the turbomachinery and the first electric machine  111 . ac 
     In the present embodiment, an additional, third rotary electric machine  114  is coupled with the high-pressure spool. In the present embodiment, the third electric machine  114  is coupled to the high-pressure spool via a high-pressure spool driven, core-mounted accessory gearbox  115  of conventional drive configuration. 
     It will of course be appreciated by those skilled in the art that any other suitable location for the first, second and third electric machines may be adopted. 
     For example, in alternative embodiments, the third electric machine  114  may be mounted coaxially with the turbomachinery in the engine  101 . The third electric machine  114  may be mounted axially in line with the duct between the low- and high-pressure compressors  104  and  105 , or between the low-pressure compressor  104  and the gearbox  109 . In alternative embodiments, the first electric machine  111  may be mounted axially in line with the duct between the low- and high-pressure compressors  104  and  105 . 
     In the present embodiment, each of the electric machines  111 ,  113  and  114  are connected with a power electronics module (PEM)  116 . In the present embodiment, the PEM  116  is mounted on the fan case  117  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. Further, different parts of the PEM  116  may be distributed between different locations. For example, some components may be mounted on the engine  101  and some may be mounted in the in the vehicle to which the engine  101  is attached. 
     Control of the PEM  116  and thus of the first, second and third electric machines  111 ,  112  and  113  is in the present example performed by an electronic engine controller (EEC)  118 . In the present embodiment the EEC  118  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 first, second and third electric machines  111 ,  112  and  113 . In this way, the EEC  118  may holistically respond to both thrust demand and electrical power demand. 
     Operation of the PEM  116  by the EEC  118  will be described further with reference to  FIG. 2 . 
     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, 2.6 metres, 2.7 metres, 2.8 metres, 2.9 metres, 3 metres, 3.1 metres, 3.2 metres, 3.3 metres, 3.4 metres, 3.5 metres, 3.6 metres, 3.7 metres, 3.8 metres or 3.9 metres. 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 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. 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 kelvin or 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 kelvin or 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 those skilled in the art. 
     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 which 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 to 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.  2   
     An electrical system  201  for connecting the first electric machine  111  and the second electric machine  113  on the low-pressure spool is shown in  FIG. 2 . In this embodiment, the electrical system  201  also connects with the third electric machine  114  on the high-pressure spool. 
     The electrical system  201  is shown in the form of a single line diagram, the conventions of which will be familiar to those skilled in the art. Thus for alternating current (ac) a single line replaces a plurality of polyphase lines, and for direct current (dc) a single line replaces the +V and −V lines. 
     The PEM  116  comprises a first power converter  201  having an ac side connected with the first electric machine  111 , and a second power converter  202  having an ac side connected with the second electric machine  113 . 
     A dc side of the first power converter  201  is connected with a first dc network  203  operating at a first dc voltage. In an embodiment the first dc voltage is up to 1 kilovolt. In an embodiment, the first dc voltage is up to ±500 volts. In an embodiment, the first dc voltage is 540 volts. In an embodiment, the first dc voltage is ±270 volts. 
     A dc side of the second power converter  202  is connected with a second dc network  204  operating at a second dc voltage. The second dc voltage is greater than the first dc voltage. In an embodiment the second dc voltage is 1 kilovolt or greater. In an embodiment, the second dc voltage is 3 kilovolts. In an embodiment, the first dc voltage is ±1.5 kilovolts. 
     In an embodiment, the first electric machine  111  and the second electric machine  113  are both permanent-magnet rotor electric machines. Alternatively, one or both of the first electric machine  111  and the second electric machine  113  could instead be a different electric machine type, such as induction, wound-field, switched-reluctance, etc. 
     In an embodiment, the first electric machine  111  is configured with a lower rated power than the second electric machine  113 . For example, in an embodiment the first electric machine  111  is configured with a rated power of 500 kilowatts, and the second electric machine  113  is configured with a rated power of 2 megawatts. 
     In operation, the first electric machine  111  and the second electric machine  113  are configured to cooperate to transfer power from the second dc network  204  to the first dc network  203 . This is achieved by operating the first electric machine  111  and the second electric machine  113  as a motor-generator set. Thus, the second electric machine  113  is operated as a motor, and the first electric machine  111  is operated as a generator. In this way, electrical power available on the second dc bus  204 , for example from another generator or a battery or any other source of electrical power, may be transferred to the first dc network  203 . 
     Configuring the first electric machine  111  and the second electric machine  113  to operate in this way provides a number of benefits compared with use of a dc-dc converter between the dc networks  202  and  203 . 
     In particular, it allows electrical power sources on the second dc network  204  to be used to power devices on the first dc network  203 . For example, in an embodiment, the EEC  118  is powered via the first dc network  203 , and thus can be initialised using a power source such as a battery on the second dc network  204  prior to engine start. Other engine accessories  205  driven by electrical accessory drives may also be powered in this way, such as for example, a fuel pump, an oil pump, a hydraulic pump, and/or a cabin blower. 
     Further, the configuration of the electric machines as a motor-generator set provides harmonic cancellation, frequency conversion, and line isolation. This allows different power quality standards to be adopted on the two dc networks  202  and  203 . The combination of the electric machines also provides resistance to electrostatic discharge. Further, the motor-generator set formed by the electric machines  111  and  113  can handle large short-term overloads far better than semiconductor devices of the same average load rating due to the much larger mass of the thermally-limited components (windings and stator, compared with solid-state switches and heatsinks). 
     In an embodiment, the first power converter  201  is a unidirectional rectifier and the second power converter  202  is a unidirectional inverter. Thus the first power converter  201  is configured to convert ac generated by the first electric machine  111  to dc for the first dc network  203 , and the second power converter  202  is configured to convert ac generated by the second electric machine  113  to dc for the second dc network  204 . 
     In a different embodiment, the first power converter  201  and the second power converter  202  are bidirectional converters capable of converting from ac to dc and vice versa. In this way, electrical power from the first dc network  202  may also be transferred to the second dc network  202  by operating the first electric machine  111  as a motor and the second electric machine  113  as a generator. 
     In an embodiment, the third electric machine  114  is connected with an ac side of a third power converter  206 . A dc side of the third power converter  206  is connected with the first dc network  202 . 
     In a specific embodiment, the third power converter  206  is a unidirectional inverter. Thus, in operation, engine start may be achieved by driving the second electric machine  113  as a motor using electrical power from a source thereof on the second dc network  204 , generating electrical power for the first dc network  203  by the first electric machine  111  acting as a generator, and driving the high-pressure spool by the third electric machine  114  acting as a motor using electrical power from the first dc network  203 . 
     In another embodiment, the third power converter  206  is a bidirectional power converter. Thus, in operation, power may be generated by the third electric machine  114  and transferred to the second dc network  204  via the first and second electric machines  111  and  113  operating as a motor-generator set. 
     Providing bidirectional electrical power transfer between the spools in the engine  101  allows the turbomachinery to be designed to exploit the attendant advantages conferred by power transfer. For example, transfer of power from the low-pressure spool to the high-pressure spool during the approach phase of an aircraft&#39;s mission reduces the effective thrust of the engine  101  whilst maintaining sufficient high-pressure spool rotational speed to safely initiate a go-around manoeuvre. Further, in engine  101 , transfer of power from the high-pressure spool to the low-pressure spool during a deceleration manoeuvre reduces the risk of weak extinction, therefore enabling a more optimal combustor design. 
     FIG.  3   
     One possible arrangement of the first electric machine  111  and the second electric machine  113  is shown in  FIG. 3 . In this embodiment, the two electric machines are separate electric machines connected by a shaft. 
     FIG.  4   
     Another possible arrangement of the first electric machine  111  and the second electric machine  113  is shown in  FIG. 4 . In this embodiment, the two electric machines form a dynamotor, having their own winding sets but a shared rotor. 
     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.