Method of controlling a gas turbine engine

Abstract

There is provided a method (820, figure 8) of controlling a gas turbine engine 10 comprising a blower system 400 including a rotor 410 drivingly coupled to a spool 440 of the gas turbine engine. The m

Description

TECHNICAL FIELD The present disclosure relates to a method of controlling a gas turbine engine comprising a blower system. The disclosure also concerns a gas turbine engine 5 comprising a blower system and a controller configured to perform a method of controlling the gas turbine engine. BACKGROUND It is known to pressurise and ventilate a cabin of an aircraft using engine bleed air which is bled from a compressor section of the core of a gas turbine engine. 10 Bleeding high pressure air from the gas turbine engine reduces its efficiency and thereby increases its fuel consumption. Blower systems which make use of air which is bled from a lower pressure source of a gas turbine engine (such as a bypass duct) and which subsequently compress the engine bleed air prior to delivering it to the cabin are also known, as 15 described in EP3517436 B1, EP3517437 B1 and EP3517438 B1. Such systems may 1 extract rotational power from a spool of a gas turbine engine for compressing the engine bleed air prior to delivery to the cabin. 1— C\J SUMMARY According to a first aspect there is provided a method of controlling a gas 20 turbine engine comprising a blower system including a rotor drivingly coupled to a spool of the gas turbine engine via a variable transmission, the variable transmission including an electrical machine operable as a motor and a differential gearbox mechanically coupled to the electrical machine, the spool of the gas turbine engine and the rotor of the blower system, the method comprising: operating the blower 25 system to extract rotational power from the spool to drive the rotor to rotate and compress air received from the gas turbine engine; receiving a signal associated with an operating parameter of the gas turbine engine; evaluating a surge criterion relating to engine compressor surge based on the received signal; and controlling the electrical machine to cause a reduction of the rotational power extracted from the 30 spool by the blower system in response to an output of the surge criterion evaluation. The method may be computer-implemented. The operating parameter may be one or more operating parameters of the gas turbine engine and there may be one or CO CM more associated signals. The surge criterion may be evaluated based on one or more received signals associated with one or more operating parameters. The method may be performed during a transient operation of the gas turbine engine in which a rotational speed of the spool changes overtime. 5 The method may comprise controlling the electrical machine to reverse a torque applied to the differential gearbox by the electrical machine and thereby cause the reduction of the rotational power extracted from the spool by the blower system. In addition, the method may comprise controlling a valve in fluid communication with the rotor to cause the reduction of the rotational power extracted 10 from the spool by the blower system. The operating parameter may be selected from a group consisting of: a thrust demand for the gas turbine engine; a rotational speed of a compressor, a turbine or a fan of the gas turbine engine; a fuel flow to a combustor of the gas turbine engine; a pressure ratio across a compressor or a turbine of the gas turbine engine; and an 15 acceleration of the gas turbine engine. The surge criterion may be defined with reference to an assessment parameter selected from a group consisting of: a rate-of-change in a thrust demand for the gas turbine engine; a rate-of-change in a rotational speed of a compressor, a turbine or a fan of the gas turbine engine; a rate-of-change in a fuel flow to a 20 combustor of the gas turbine engine; a pressure ratio across a compressor or a turbine of the gas turbine engine; and an acceleration of the gas turbine engine. Evaluating the surge criterion relating to engine compressor surge may include determining whether the assessment parameter exceeds a predetermined threshold value. It may be that the predetermined threshold value corresponds to a value of the 25 assessment parameter above which a compressor of the gas turbine engine is at risk of exhibiting compressor surge. It may be that, subsequent to the reduction of the rotational power, the method comprises: receiving a signal associated with the operating parameter of the gas turbine engine; evaluating a surge criterion relating to engine compressor surge 30 based on the received signal; and causing an increase of the rotational power extracted from the spool by the blower system in response to an output of the surge criterion evaluation. The method may comprise controlling the variable transmission to cause the increase of the rotational power extracted from the spool by the blower system. CO CM The method may comprise controlling the electrical machine to cause the increase of the rotational power extracted from the spool by the blower system. Further, the method may comprise controlling the electrical machine to reverse a torque applied to the differential gearbox by the electrical machine and thereby cause 5 the increase of the rotational power extracted from the spool by the blower system. In addition or instead, the method may comprise controlling a valve in fluid communication with the rotor to cause the increase of the rotational power extracted from the spool by the blower system. According to a second aspect there is provided a controller configured to 10 perform the method of the first aspect. According to a third aspect there is provided a non-transitory machine readable medium comprising instructions which, when executed by the controller of the second aspect, cause the controller to carry out the method of the first aspect. According to a fourth aspect there is provided a gas turbine engine comprising 15 the controller of the second aspect and a blower system including a rotor drivingly coupled to a spool of the gas turbine engine via a variable transmission, the variable transmission including an electrical machine operable as a motor and a differential gearbox mechanically coupled to the electrical machine, the spool of the gas turbine engine and the rotor of the blower system. 20 It may be that the gas turbine engine comprises an engine core comprising a turbine, a compressor, and a core shaft connecting the turbine to the compressor; a fan located upstream of the engine core, the fan comprising a plurality of fan blades; and a gearbox that receives an input from the core shaft and outputs drive to the fan so as to drive the fan at a lower rotational speed than the core shaft. 25 It may also be that the turbine is a first turbine, the compressor is a first compressor, and the core shaft is a first core shaft; the engine core further comprises a second turbine, a second compressor, and a second core shaft connecting the second turbine to the second compressor; and the second turbine, second compressor, and second core shaft are arranged to rotate at a higher rotational speed 30 than the first core shaft. According to a fifth aspect there is provided an aircraft comprising the gas turbine engine of the fourth aspect. As noted elsewhere herein, the present disclosure may relate to a gas turbine engine. Such a gas turbine engine may comprise an engine core comprising a turbine, 35 a combustor, a compressor, and a core shaft connecting the turbine to the CO CM compressor. Such a gas turbine engine may comprise a fan (having fan blades) located upstream of the engine core. Arrangements of the present disclosure may be particularly, although not exclusively, beneficial for fans that are driven via a gearbox. Accordingly, the gas 5 turbine engine may comprise a gearbox that receives an input from the core shaft and outputs drive to the fan so as to drive the fan at a lower rotational speed than the core shaft. The input to the gearbox may be directly from the core shaft, or indirectly from the core shaft, for example via a spur shaft and / or gear. The core shaft may rigidly connect the turbine and the compressor, such that the turbine and compressor rotate 10 at the same speed (with the fan rotating at a lower speed). The gas turbine engine as described and / or claimed herein may have any suitable general architecture. For example, the gas turbine engine may have any desired number of shafts that connect turbines and compressors, for example one, two or three shafts. Purely by way of example, the turbine connected to the core shaft 15 may be a first turbine, the compressor connected to the core shaft may be a first compressor, and the core shaft may be a first core shaft. The engine core may further comprise a second turbine, a second compressor, and a second core shaft connecting the second turbine to the second compressor. The second turbine, second compressor, and second core shaft may be arranged to rotate at a higher rotational 20 speed than the first core shaft. In such an arrangement, the second compressor may be positioned axially downstream of the first compressor. The second compressor may be arranged to receive (for example directly receive, for example via a generally annular duct) flow from the first compressor. 25 The gearbox may be arranged to be driven by the core shaft that is configured to rotate (for example in use) at the lowest rotational speed (for example the first core shaft in the example above). For example, the gearbox may be arranged to be driven only by the core shaft that is configured to rotate (for example in use) at the lowest rotational speed (for example only by the first core shaft, and not the second core 30 shaft, in the example above). Alternatively, the gearbox may be arranged to be driven by any one or more shafts, for example the first and / or second shafts in the example above. The gearbox may be a reduction gearbox (in that the output to the fan is a lower rotational rate than the input from the core shaft). Any type of gearbox may be 35 used. For example, the gearbox may be a “planetary” or “star” gearbox, as described CM in more detail elsewhere herein. The gearbox may have any desired reduction ratio (defined as the rotational speed of the input shaft divided by the rotational speed of the output shaft), for example greater than 2.5, for example in the range of from 3 to 4.2, or 3.2 to 3.8, for example on the order of or at least 3, 3.1, 3.2, 3.3, 3.4, 3.5, 3.6, 5 3.7, 3.8, 3.9, 4, 4.1 or 4.2. The gear ratio may be, for example, between any two of the values in the previous sentence. Purely by way of example, the gearbox may be a “star” gearbox having a ratio in the range of from 3.1 or 3.2 to 3.8. In some arrangements, the gear ratio may be outside these ranges. In any gas turbine engine as described and / or claimed herein, a combustor 10 may be provided axially downstream of the fan and compressor(s). For example, the combustor may be directly downstream of (for example at the exit of) the second compressor, where a second compressor is provided. By way of further example, the flow at the exit to the combustor may be provided to the inlet of the second turbine, where a second turbine is provided. The combustor may be provided upstream of the 15 turbine(s). The or each compressor (for example the first compressor and second compressor as described above) 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 20 variable). The row of rotor blades and the row of stator vanes may be axially offset from each other. The or each turbine (for example the first turbine and second turbine as described above) 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. The row of 25 rotor blades and the row of stator vanes may be axially offset from each other. 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% span position, to a tip at a 100% span position. The ratio of the radius of the fan blade at the hub to the radius of the fan blade at the tip may be less than (or on the order of) any of: 0.4, 0.39, 0.38, 0.37, 30 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 ratio of the radius of the fan blade at the hub to the radius of the fan blade at the tip 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), for example in the range of from 0.28 to 0.32. These ratios may commonly be referred to as the hub-to-tip ratio. The radius at 35 the hub and the radius at the tip may both be measured at the leading edge (or axially forwardmost) part of the blade. The hub-to-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 may be measured between the engine centreline and the tip of a fan blade at its leading edge. The fan diameter (which may simply be twice 5 the radius of the fan) may be greater than (or on the order of) any of: 210 cm, 220 cm, 230 cm, 240 cm, 250 cm (around 100 inches), 260 cm, 270 cm (around 105 inches), 280 cm (around 110 inches), 290 cm (around 115 inches), 300 cm (around 120 inches), 310 cm, 320 cm (around 125 inches), 330 cm (around 130 inches), 340 cm (around 135 inches), 350 cm, 360 cm (around 140 inches), 370 cm (around 145 10 inches), 380 cm (around 150 inches), 390 cm (around 155 inches), 400 cm, 410 cm (around 160 inches) or 420 cm (around 165 inches). The fan diameter 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), for example in the range of from 210 cm to 240 cm, or 250 cm to 280 cm, or 320 cm to 380 cm. 15 The rotational speed of the fan 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 2600 rpm, for example less than 2500 rpm, or less than 2300 rpm. Purely by way of further non-limitative example, the rotational speed of the fan at cruise conditions for an engine 20 having a fan diameter in the range of from 210 cm to 300 cm (for example 240 cm to 280 cm or 250 cm to 270 cm) may be in the range of from 1700 rpm to 2600 rpm, for example in the range of from 1800 rpm to 2300 rpm, 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 25 of from 330 cm to 380 cm 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 1800 rpm. In use of the gas turbine engine, the fan (with associated fan blades) rotates about a rotational axis. This rotation results in the tip of the fan blade moving with a 30 velocity UtiP. The work done by the fan blades 13 on the flow results in an enthalpy rise dH of the flow. A fan tip loading may be defined as dH / Utip2, where dH is the enthalpy rise (for example the 1-D average enthalpy rise) across the fan and UtiP is the (translational) velocity of the fan tip, for example at the leading edge of the tip (which may be defined as fan tip radius at leading edge multiplied by angular speed). 35 The fan tip loading at cruise conditions may be greater than (or on the order of) any of: 0.28, 0.29, 0.30, 0.31, 0.32, 0.33, 0.34, 0.35, 0.36, 0.37, 0.38, 0.39 or 0.4 (all values being dimensionless ). 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), for example in the range of from 0.28 to 0.31, or 0.29 to 0.3. 5 Gas turbine engines in accordance with the present disclosure may have any desired bypass ratio, where the bypass ratio is defined as the ratio of the mass flow rate of the flow through the bypass duct to the mass flow rate of the flow through the core at cruise conditions. In some arrangements 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, 10 14.5, 15, 15.5, 16, 16.5, 17, 17.5, 18, 18.5, 19, 19.5 or 20. The bypass ratio 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), for example in the range of form 12 to 16, 13 to 15, or 13 to 14. The bypass duct may be substantially annular. The bypass duct may be radially outside the core engine. The radially outer surface of the bypass 15 duct may be defined by a nacelle and / or a fan case. The overall pressure ratio of a gas turbine engine as described and / or claimed herein may be defined as the ratio of the stagnation pressure upstream of the fan to the stagnation pressure at the exit of the highest pressure compressor (before entry into the combustor). By way of non-limitative example, the overall pressure ratio of a 20 gas turbine engine as described and / or claimed herein 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 values in the previous sentence (i.e., the values may form upper or lower bounds), for example in the range of from 50 to 70. 25 Specific thrust of an engine may be defined as the net thrust of the engine divided by the total mass flow through the engine. At cruise conditions, the specific thrust of an engine described and / or claimed herein may be less than (or on the order of) any of the following: 110 Nkg~1s, 105 Nkg’1s, 100 Nkg~1s, 95 Nkg’1s, 90 Nkg~1s, 85 Nkg-1s or 80 Nkg-1s. The specific thrust may be in an inclusive range bounded by any 30 two of the values in the previous sentence (i.e., the values may form upper or lower bounds), for example in the range of from 80 Nkg-1s to 100 Nkg-1s, or 85 Nkg-1s to 95 Nkg-1s. Such engines may be particularly efficient in comparison with conventional gas turbine engines. A gas turbine engine as described and / or claimed herein may have any 35 desired maximum thrust. Purely by way of non-limitative example, a gas turbine as CO CM described and / or claimed herein may be capable of producing a maximum thrust of at least (or on the order of) any of the following: 130kN, 135kN, 140kN, 145kN, 150kN, 155kN, 160kN, 170kN, 180kN, 190kN, 200kN, 250kN, 300kN, 350kN, 400kN, 450kN, 500kN, or 550kN. The maximum thrust may be in an inclusive range bounded by any 5 two of the values in the previous sentence (i.e., the values may form upper or lower bounds). Purely by way of example, a gas turbine as described and / or claimed herein may be capable of producing a maximum thrust in the range of from 155kN to 170kN, 330kN to 420 kN, or 350kN to 400kN. The thrust referred to above may be the maximum net thrust at standard atmospheric conditions at sea level plus 15 degrees 10 C (ambient pressure 101.3kPa, temperature 30 degrees C), with the engine static. In use, the temperature of the flow at the entry to the high pressure turbine may be particularly high. This temperature, which may be referred to as TET, may be measured at the exit to the combustor, for example immediately upstream of the first turbine vane, which itself may be referred to as a nozzle guide vane. At cruise, the 15 TET may be at least (or on the order of) any of the following: 1400K, 1450K, 1500K, 1520K, 1530K, 1540K, 1550K, 1600K or 1650K. The TET at cruise 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), for example 1530Kto 1600K. The maximum TET in use of the engine may be, for example, at least (or on the order of) any of the 20 following: 1700K, 1750K, 1800K, 1850K, 1900K, 1950Kor2000K. The maximum TET 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), for example in the range of from 1800K to 1950K, or 1900K to 2000K. The maximum TET may occur, for example, at a high thrust condition, for example at a maximum take-off (MTO) 25 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 30 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 35 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 (such as an aluminium lithium alloy) 5 with a titanium leading edge. A fan as described and / or claimed herein may comprise a central 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 (or 10 disc). 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 / disc. By way of further example, the fan blades maybe formed integrally with a central portion. Such an arrangement may be referred to as a bladed disc or a bladed ring. Any suitable method may be used to manufacture such a bladed 15 disc or bladed ring. For example, at least a part of the fan blades may be machined from a block and / or at least part of the fan blades may be attached to the hub / disc by welding, such as linear friction welding. The gas turbine engines described and / or claimed herein may or may not be provided with a variable area nozzle (VAN). Such a variable area nozzle may allow 20 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. The fan of a gas turbine as described and / or claimed herein may have any desired number of fan blades, for example 14, 16, 18, 20, 22, 24 or 26 fan blades. As used herein, cruise conditions have the conventional meaning and would 25 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 industry as the “economic mission”) of an aircraft to which the gas turbine engine is designed to be attached. In this regard, mid-cruise is the point in an 30 aircraft flight cycle at which 50% 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 -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 35 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. 5 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 midcruise Mach Number) at the mid-cruise atmospheric conditions (defined by the 10 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. Purely by way of example, the forward speed at the cruise condition may be 15 any point in the range of from Mach 0.7 to 0.9, for example 0.75 to 0.85, for example 0.76 to 0.84, for example 0.77 to 0.83, for example 0.78 to 0.82, for example 0.79 to 0.81, for example on the order of Mach 0.8, on the order of Mach 0.85 or in the range of from 0.8 to 0.85. Any single speed within these ranges may be part of the cruise condition. For some aircraft, the cruise conditions may be outside these ranges, for 20 example below Mach 0.7 or above Mach 0.9. Purely by way of example, the cruise conditions may correspond to standard atmospheric conditions (according to the International Standard Atmosphere, ISA) at an altitude that is in the range of from 10000m to 15000m, for example in the range of from 10000m to 12000m, for example in the range of from 10400m to 11600m 25 (around 38000 ft), for example in the range of from 10500m to 11500m, for example in the range of from 10600m to 11400m, for example in the range of from 10700m (around 35000 ft) to 11300m, for example in the range of from 10800m to 11200m, for example in the range of from 10900m to 11100m, for example on the order of 11000m. The cruise conditions may correspond to standard atmospheric conditions 30 at any given altitude in these ranges. Purely by way of example, the cruise conditions may correspond to a forward Mach number of 0.8 and standard atmospheric conditions (according to the International Standard Atmosphere) at an altitude of 35000ft (10668m). At such cruise conditions, the engine may provide a known required net thrust level. The known required net thrust level is, of course, dependent on the engine and its intended application and may be, for example, a value in the range of from 20kN to 40kN. Purely by way of further example, the cruise conditions may correspond to a forward Mach number of 0.85 and standard atmospheric conditions (according to the 5 International Standard Atmosphere) at an altitude of 38000ft (11582m). At such cruise conditions, the engine may provide a known required net thrust level. The known required net thrust level is, of course, dependent on the engine and its intended application and may be, for example, a value in the range of from 35kN to 65kN. In use, a gas turbine engine described and / or claimed herein may operate at 10 the cruise conditions defined elsewhere herein. Such cruise conditions may be determined by the cruise conditions (for example the mid-cruise conditions) of an aircraft to which at least one (for example 2 or 4) gas turbine engine may be mounted in order to provide propulsive thrust. According to an aspect, there is provided an aircraft comprising a gas turbine 15 engine as described and / or claimed herein. The aircraft according to this aspect is the aircraft for which the gas turbine engine has been designed to be attached. Accordingly, the cruise conditions according to this aspect correspond to the midcruise of the aircraft, as defined elsewhere herein. According to an aspect, there is provided a method of operating a gas turbine 20 engine as described and / or claimed herein. The operation may be at the cruise conditions as defined elsewhere herein (for example in terms of the thrust, atmospheric conditions and Mach Number). According to an aspect, there is provided a method of operating an aircraft comprising a gas turbine engine as described and / or claimed herein. The operation 25 according to this aspect may include (or may be) operation at the mid-cruise of the aircraft, as defined elsewhere herein. The skilled person will appreciate that except where mutually exclusive, a feature or parameter described in relation to any one of the above aspects may be applied to any other aspect. Furthermore, except where mutually exclusive, any 30 feature or parameter described herein may be applied to any aspect and / or combined with any other feature or parameter described herein. 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 35 which: Figure 1 is a sectional side view of a gas turbine engine; Figure 2 is a close up sectional side view of an upstream portion of a gas turbine engine; Figure 3 is a partially cut-away view of a gearbox for a gas turbine engine; 5 Figure 4 is a diagram which shows an example blower system for providing air to a cabin of an aircraft; Figure 5 is a diagram which shows an example configuration of a variable transmission suitable for use in the blower system of Figure 4; Figure 6 is a schematic view of an aircraft comprising a gas turbine engine; 10 Figure 7 shows, schematically, a non-transitory machine readable medium comprising instructions which, when executed by a controller, cause the controller to carry out a method of operating the gas turbine engine of Figure 1 or Figure 2; and Figure 8 is a flowchart which shows an example method of controlling a gas turbine engine comprising a blower system. 15 DETAILED DESCRIPTION Figure 1 illustrates a gas turbine engine 10 having a principal rotational axis 9. The engine 10 comprises an air intake 12 and a propulsive fan 23 that generates two airflows: a core airflow A and a bypass airflow B. The gas turbine engine 10 comprises a core 11 that receives the core airflow A. The engine core 11 comprises, 20 in axial flow series, a low pressure compressor 14, a high-pressure compressor 15, combustion equipment 16, a high-pressure turbine 17, a low pressure turbine 19 and a core exhaust nozzle 20. A nacelle 21 surrounds the gas turbine engine 10 and defines a bypass duct 22 and a bypass exhaust nozzle 18. The bypass airflow B flows through the bypass duct 22. The fan 23 is attached to and driven by the low pressure 25 turbine 19 via a shaft 26 and an epicyclic gearbox 30. In use, the core airflow A is accelerated and compressed by the low pressure compressor 14 and directed into the high pressure compressor 15 where further compression takes place. The compressed air exhausted from the high pressure compressor 15 is directed into the combustion equipment 16 where it is mixed with 30 fuel and the mixture is combusted. The resultant hot combustion products then expand through, and thereby drive, the high pressure and low pressure turbines 17, 19 before being exhausted through the nozzle 20 to provide some propulsive thrust. The high pressure turbine 17 drives the high pressure compressor 15 by a suitable interconnecting shaft 27. The fan 23 generally provides the majority of the propulsive 35 thrust. The epicyclic gearbox 30 is a reduction gearbox. CO An exemplary arrangement for a geared fan gas turbine engine 10 is shown in Figure 2. The low pressure turbine 19 (see Figure 1) drives the shaft 26, which is coupled to a sun wheel, or sun gear, 28 of the epicyclic gear arrangement 30. Radially outwardly of the sun gear 28 and intermeshing therewith is a plurality of planet gears 5 32 that are coupled together by a planet carrier 34. The planet carrier 34 constrains the planet gears 32 to precess around the sun gear 28 in synchronicity whilst enabling each planet gear 32 to rotate about its own axis. The planet carrier 34 is coupled via linkages 36 to the fan 23 in order to drive its rotation about the engine axis 9. Radially outwardly of the planet gears 32 and intermeshing therewith is an annulus or ring gear 10 38 that is coupled, via linkages 40, to a stationary supporting structure 24. Note that the terms “low pressure turbine” and “low pressure compressor” as used herein may be taken to mean the lowest pressure turbine stages and lowest pressure compressor stages (i.e., not including the fan 23) respectively and / or the turbine and compressor stages that are connected together by the interconnecting 15 shaft 26 with the lowest rotational speed in the engine (i.e., not including the gearbox output shaft that drives the fan 23). In some literature, the “low pressure turbine” and “low pressure compressor” referred to herein may alternatively be known as the “intermediate pressure turbine” and “intermediate pressure compressor”. Where such alternative nomenclature is used, the fan 23 may be referred to as a first, or lowest 20 pressure, compression stage. The epicyclic gearbox 30 is shown by way of example in greater detail in Figure 3. Each of the sun gear 28, planet gears 32 and ring gear 38 comprise teeth about their periphery to intermesh with the other gears. However, for clarity only exemplary portions of the teeth are illustrated in Figure 3. There are four planet gears 25 32 illustrated, although it will be apparent to the skilled reader that more or fewer planet gears 32 may be provided within the scope of the claimed invention. Practical applications of a planetary epicyclic gearbox 30 generally comprise at least three planet gears 32. The epicyclic gearbox 30 illustrated by way of example in Figures 2 and 3 is 30 of the planetary type, in that the planet carrier 34 is coupled to an output shaft via linkages 36, with the ring gear 38 fixed. However, any other suitable type of epicyclic gearbox 30 may be used. By way of further example, the epicyclic gearbox 30 may be a star arrangement, in which the planet carrier 34 is held fixed, with the ring (or annulus) gear 38 allowed to rotate. In such an arrangement the fan 23 is driven by 35 the ring gear 38. By way of further alternative example, the gearbox 30 may be a differential gearbox in which the ring gear 38 and the planet carrier 34 are both allowed to rotate. It will be appreciated that the arrangement shown in Figures 2 and 3 is by way of example only, and various alternatives are within the scope of the present 5 disclosure. Purely by way of example, any suitable arrangement may be used for locating the gearbox 30 in the engine 10 and / or for connecting the gearbox 30 to the engine 10. By way of further example, the connections (such as the linkages 36, 40 in the Figure 2 example) between the gearbox 30 and other parts of the engine 10 (such as the input shaft 26, the output shaft and the fixed structure 24) may have any 10 desired degree of stiffness or flexibility. By way of further example, any suitable arrangement of the bearings between rotating and stationary parts of the engine (for example between the input and output shafts from the gearbox and the fixed structures, such as the gearbox casing) may be used, and the disclosure is not limited to the exemplary arrangement of Figure 2. For example, where the gearbox 30 has a 15 star arrangement (described above), the skilled person would readily understand that the arrangement of output and support linkages and bearing locations would typically be different to that shown by way of example in Figure 2. Accordingly, the present disclosure extends to a gas turbine engine having any arrangement of gearbox styles (for example star or planetary), support structures, 20 input and output shaft arrangement, and bearing locations. Optionally, the gearbox may drive additional and / or alternative components (e.g., the intermediate pressure compressor and / or a booster compressor). Other gas turbine engines to which the present disclosure may be applied may have alternative configurations. For example, such engines may have an alternative 25 number of compressors and / or turbines and / or an alternative number of interconnecting shafts. By way of further example, the gas turbine engine shown in Figure 1 has a split flow nozzle 18, 20 meaning that the flow through the bypass duct 22 has its own nozzle 18 that is separate to and radially outside the core engine nozzle 20. However, this is not limiting, and any aspect of the present disclosure may also 30 apply to engines in which the flow through the bypass duct 22 and the flow through the core 11 are mixed, or combined, before (or upstream of) a single nozzle, which may be referred to as a mixed flow nozzle. One or both nozzles (whether mixed or split flow) may have a fixed or variable area. Whilst the described example relates to a turbofan engine, the disclosure may apply, for example, to any type of gas turbine 35 engine, such as an open rotor (in which the fan stage is not surrounded by a nacelle) CO CM or turboprop engine, for example. In some arrangements, the gas turbine engine 10 may not comprise a gearbox 30. The geometry of the gas turbine engine 10, and components thereof, is defined by a conventional axis system, comprising an axial direction (which is aligned 5 with the rotational axis 9), a radial direction (in the bottom-to-top direction in Figure 1), and a circumferential direction (perpendicular to the page in the Figure 1 view). The axial, radial and circumferential directions are mutually perpendicular. A diagram of an example blower system 400 for providing air to a cabin of an aircraft is shown schematically in Figure 4. The blower system 400 comprises a rotor 10 410 which is drivingly coupled to a spool 440 of a gas turbine engine. The spool may, for example, be the high-pressure (HP) spool of a two- or three- shaft gas turbine or an intermediate pressure (IP) spool of a three-spool engine, though any one or more spools of a gas turbine engine may be coupled to the rotor 410. The rotor 410 is disposed within a rotor housing 420. In the example of Figure 4, the blower system 15 400 comprises a variable transmission 430 which drivingly couples the rotor 410 to the spool 440. In this embodiment, the blower system 400 is moveable between a compressor configuration and a turbine configuration by relative movement between the rotor 410 and a flow modifier. Therefore, the example blower system 400 is operable in both the compressor configuration (for supplying compressed air to the 20 airframe) and the turbine configuration (for applying power to the spool to, e.g., start the engine). In other embodiments, the blower system 400 may only be operable in a compressor configuration. The rotor 410 is configured to be driven to rotate by the spool 440 in the compressor configuration, wherein the rotor 410 compresses air it receives from the 25 gas turbine engine. The compressed air is discharged to a cabin discharge nozzle 426 for supply to an aircraft cabin 450. In the example of Figure 4, the rotor 410 is configured to receive air from the gas turbine engine via an engine bleed nozzle 422. The engine bleed nozzle 422 is in fluid communication with an air pathway (shown schematically at 460) of the gas turbine engine. Accordingly, in the compressor 30 configuration, the blower system 400 is configured to draw air from the air pathway 460 of the gas turbine engine and supply air to the aircraft cabin 450 to pressurise and / or ventilate the aircraft cabin 450. The blower system 400 is configured to function as a compressor in the compressor configuration, such that air supplied to the airframe system 450 is at a 35 higher pressure than air drawn from the air pathway 460 of the gas turbine engine. As a result, the blower system 400 is not required to draw air from a relatively high pressure region of the gas turbine engine in order to supply pressurised air to the airframe system 450. Instead, the blower system 400 may draw air via the engine bleed nozzle 422 from a relatively low pressure region of the gas turbine engine, such 5 as from a bypass duct 22 of the gas turbine engine as shown in Figure 1. If the blower system 400 were alternatively required to draw air from a relatively high pressure region of the gas turbine engine (e.g., the high pressure compressor), an efficiency of the gas turbine engine may be reduced. Therefore, the blower system 400 provides a more efficient airframe system pressurisation and ventilation system when 10 incorporated into an aircraft. In addition, this approach reduces a scope for contamination of the air supply to the airframe system 450. The rotor 410 is driven to rotate in the compressor configuration by drive output received from a first external terminal 430A of the variable transmission 430, which itself receives drive input from the spool 440 at a second external terminal 430B 15 of the variable transmission 430. Accordingly, the blower system 400 is configured to extract rotational power from the spool 440 to drive the rotor 410 to rotate and compress air received from the gas turbine engine when operated in the compressor configuration. The speed of rotation of the spool 440 depends on the operating point of the gas turbine engine, which dictates a speed of the spool 440. The variable 20 transmission 430 allows a rotational speed of the rotor 410 in the compressor configuration to be decoupled from a rotational speed of the spool 440, so that a compression performance of the blower system 400 in the compressor configuration is not solely governed by the operating point of the gas turbine engine (e.g., it can be controlled to operate at a target speed independent of the rotational speed of the 25 spool, and / or at a variable speed ratio relative to the rotational speed of the spool). Inclusion of a variable transmission 430 within the blower system 400 therefore provides more versatile and adaptable means for supplying pressurised air to an airframe system. An example configuration for the variable transmission 430, in the form of an electric variator, is described below with reference to Figure 5. Other 30 suitable types of variable transmission will occur to those skilled in the art. In some embodiments, the blower system 400 is also configured to be able to receive compressed air from an external air source 470 to drive the spool 440 to rotate for starting the gas turbine engine in the turbine configuration. In the example of Figure 4, the blower system is configured to receive compressed air from the external air 35 source 470 via the airframe discharge nozzle 426. In addition, the blower system 400 further comprises a start control and isolation valve 455 which is configured to isolate the airframe discharge nozzle 426 from the external air source 470 in the compressor configuration, and to isolate the airframe discharge nozzle 426 from the airframe system 450 in the turbine configuration. The start control and isolation valve 455 may 5 be further configured to control a mass flow and a pressure of an air flow from the external air source 470 to the airframe discharge nozzle 426 in the turbine configuration. However, it will be appreciated that the blower system 400 may otherwise be configured to receive compressed air from the external air source 470, such as via an external air nozzle, for example. 10 The external air source 470 may be derived from, for example, an auxiliary power unit (APU) of the aircraft or ground starting equipment (GSE). In the example of Figure 4, the blower system 400 is configured to discharge air to the engine bleed nozzle 422 in the turbine configuration. However, the blower system 400 may otherwise discharge air in the turbine configuration, such as to a dedicated auxiliary 15 nozzle, for example. Air discharged from the blower system 400 via a dedicated auxiliary nozzle may be used for cooling other systems and / or components of the gas turbine engine and / or the aircraft in the turbine configuration. The blower system 400 is configured to function as a turbine in the turbine configuration, such that the spool 440 may be driven to rotate by the rotor 410. 20 Generally, the blower system 400 can drive rotation of the spool 440 to a rotational speed which is sufficient to enable the gas turbine engine to successfully execute an ignition process. Consequently, the blower system 400 dispenses with a need to provide a dedicated air turbine starting system or an electric starting system to the gas turbine engine, each of which are associated with additional weight and system 25 complexity. Additionally or alternatively, the blower system 400 may be able to drive the spool 440 to rotate at a lower speed, for example to prevent a bowed engine rotor condition following engine shutdown or to reduce a bowed engine rotor condition prior to engine start. To this end, the start control and isolation valve 455 may be configured to control the mass flow and pressure of the air flow to a somewhat lower level than 30 that required for engine starting. The use of a two-configuration blower system 400 allows for a system in which the rotor 410 rotates in the same rotation direction (i.e., clockwise or anti-clockwise) in both the compressor configuration and the turbine configuration. In this way, in the turbine configuration of the blower system 400 the rotor 410 will drive the spool 440 35 to rotate in direction that the spool 440 rotates when it drives the rotor 410 in the compressor configuration. This allows for the omission of a separate reversing mechanism to permit the spool 440 to be driven to rotate in its starting direction, which will be the same as the direction it rotates during when driving the rotor 410 in the compressor configuration. A separate reversing mechanism would result in additional 5 mechanical efficiency losses in, and increased weight of and / or a reduced reliability of, the blower system 400. Figure 4 also schematically shows a gas turbine engine 10 comprising the example blower system 400. The gas turbine engine 10 may be in accordance with the gas turbine engine 10 described above with respect to Figure 1 and / or Figure 2. 10 Although it has been described that the example blower system 400 is operable in both the compressor configuration and the turbine configuration, the methods described herein may be used in the context of controlling a gas turbine engine comprising a blower system which is only operable in a compressor configuration. 15 The gas turbine engine 10 also comprises a controller 490 which is configured to carry out a method of controlling the gas turbine engine 10, as described in further detail below with reference to Figure 8. The blower system 400 is provided with apparatus for varying a rotational power extracted from the spool 440 of the gas turbine engine 10 by the blower system 400. 20 The apparatus for varying the rotational power extracted from the spool 440 by the blower system 400 may comprise a valve in fluid communication with the rotor 410. The valve may be, for instance, a recirculation valve provided to a recirculation pathway for recirculating air from the rotor 410. Otherwise the valve may be, for example, an overboard bleed valve provided to an overboard bleed pathway for 25 discharging air from the rotor 410 to an external environment such as ambient air. The valve is generally operable to reduce a compressive load on the rotor 410 (e.g., a torque applied to the rotor 410 due to the work required to compress air provided thereto) when operating in the compressor configuration, and is therefore adapted to vary the rotational power extracted from the spool by the rotor 410 of the blower 30 system 400. A position of the valve may determine a bleed state of the rotor 410. In addition or instead, the apparatus for varying the rotational power extracted from the spool 440 by the blower system 400 may be the variable transmission 430. The variable transmission 430 is controllable by the controller 490 to vary the rotational power extracted from the spool 440 by the blower system 400. Various arrangements for the variable transmission 430 suitable for achieving this effect will occur to those skilled in the art. Figure 5 is a diagram which shows an example configuration of a variable transmission 430 suitable for use in the blower system 400 of Figure 4. In this 5 example, the variable transmission 430 is in the form of an electric variator. The variable transmission 430 comprises a differential gearbox 512, an accessory gearbox 514, a first electrical machine 522 and a second electrical machine 524. Each gearbox 512, 514 is provided with internal terminals 512A-512C, 514A-514C for mechanical couplings. In particular, each gearbox 512, 514 may be an 10 epicyclic gearbox having respective sun gears, planet gear arrangements and ring gears. In the differential gearbox 512, each of the sun gear, planet gear arrangement and ring gear may be free to rotate. If so, a mechanical coupling at each internal terminal 512A-512C may be provided to one of the sun gear, the planet gear arrangement and the ring gear of the differential gearbox 512 such that the differential 15 gearbox 512 functions as a power splitting differential (PSD) gearbox 512. The accessory gearbox 514 is mechanically coupled to the spool 440 of the gas turbine engine 10 through the first internal terminal 514A of the accessory gearbox 514 and the second external terminal 430B, whereas the differential gearbox 512 is mechanically coupled to the rotor 410 via a first internal terminal 512A of the 20 differential gearbox 512 and the first external terminal 430A. The differential gearbox 512 and the accessory gearbox 514 are mechanically coupled to one another through respective second terminals 512B, 514B such that, in use, a driving output from the second terminal 514B of the accessory gearbox 514 may be provided as a driving input to the second terminal 512B of the differential gearbox 512 and vice versa. 25 Therefore, the differential gearbox 512 is mechanically coupled to the spool 440 via the accessory gearbox. The differential gearbox 512 is also mechanically coupled to the first electrical machine 522 through a third internal terminal 512C of the differential gearbox 512. Further, the accessory gearbox 514 is mechanically coupled to second electrical machine 524 through the third internal terminal 514C of the accessory 30 gearbox 514. The first electrical machine 522 is electrically coupled to a first bidirectional power converter 532, whereas the second electrical machine 524 is electrically coupled to a second bidirectional power converter 534. Each electrical machine 522, 524 is operable as a motor or a generator, and so may be referred to as motor-35 generators. When operating as a motor, the first electrical machine 522 receives polyphase (e.g., three-phase) power from the first bidirectional power converter 532 and provides torque to the differential gearbox 512 through the third internal terminal 512C of the differential gearbox 512. Likewise, when operating as a motor, the second electrical machine 524 receives polyphase power from the second bidirectional power 5 converter 534 and provides torque to the accessory gearbox 514 through the third internal terminal 514C of the accessory gearbox 514. Conversely, when operating as generators, each electrical machine 522, 524 receives torque from the respective gearbox 512, 514 and provides polyphase power to the respective bidirectional power converter 532, 534. A DC bus 540 electrically connects the bidirectional power 10 converters to one another and to an energy storage system 550. Depending on whether the corresponding electrical machine 522, 524 is operating as a motor or a generator, each bidirectional power converter 532, 534 may draw DC power from or provide DC power to the DC bus 540. A total net converter power draw from the DC bus 540 is defined as a sum of: the DC power drawn from the DC bus 540 by the first 15 bidirectional power converter 532; and the DC power drawn from the DC bus 540 by second bidirectional power converter 534, where the DC power drawn from the DC bus 540 by each bidirectional power converter is negative if the respective bidirectional power converter is providing DC power to the DC bus 540. If the total net converter power draw from the DC bus 540 is positive, the 20 energy storage system 550 supplies DC power to the DC bus 540 so that the DC bus 540 is maintained at a nominal voltage. Conversely, if the total net converter power draw from the DC bus 540 is negative, the energy storage system 550 receives and stores DC power from the DC bus 540 such that the DC bus 540 is maintained at the nominal voltage. The nominal voltage of the DC bus generally corresponds to an 25 operating or rated voltage of the energy storage system 550. In other words, the transfer (e.g., supply and receipt) of DC power from the DC bus 540 into the energy storage system 550 is suitable for maintaining the DC bus 540 at the nominal voltage. The energy storage system 550 may comprise, for example, a battery and / or a fuel cell. The energy storage system 550 may be disposed inside a housing of the variable 30 transmission 430 or the energy storage system 550 may be disposed outside the housing of the variable transmission 430 as shown in Figure 5, such as within an aircraft to which the gas turbine engine 10 is provided. In typical use, rotational power extracted from the spool 440 is provided to the accessory gearbox 514 at the first internal terminal 514A of the accessory gearbox 35 514. If the blower system 400 is operating in the compressor configuration, at least a CM fraction of the rotational power extracted from the spool 440 is then transmitted to the differential gearbox 512 for subsequent supply to the rotor 410. If the second electrical machine 524 is operating as a generator, a remaining fraction of the rotational power extracted from the spool 440 is provided to the second electrical machine 524. 5 However, if the second electrical machine 524 is operating as a motor, all of the rotational power extracted from the spool 440 is transmitted to the differential gearbox 512 together with any additional rotational power provided to the accessory gearbox 514 by the second electrical machine 524 at the third internal terminal 514C of the accessory gearbox 514. At least a fraction of the rotational power supplied to the 10 differential gearbox 512 is then supplied to the rotor 410. If the first electrical machine 522 is operating as a generator, a remaining fraction of the rotational power supplied to the differential gearbox 512 is provided to the first electrical machine 522. However, if the first electrical machine 522 is operating as a motor, all of the rotational power transmitted to the differential gearbox 512 is supplied to the rotor 410 together with 15 any additional rotational power provided to the differential gearbox 512 by the first electrical machine 522 at the third internal terminal 512C of the differential gearbox 512. In use, the variable transmission 430 may be controlled to vary the rotational power extracted from the spool 440 by controlling the first electrical machine 522 20 and / or the second electrical machine 524 so as to vary the rotational power extracted from the spool 440 by the blower system 400. By way of example, if the first electrical machine 522 is initially operating as a motor, the first electrical machine 522 may be controlled by the controller 490 to increase the provision of additional rotational power to the differential gearbox 512 such that the rotational power extracted from the spool 25 440 by the blower system 400 is reduced. By way of another example, if the first electrical machine 522 is initially operating as a generator, the first electrical machine 522 may be controlled by the controller 490 to decrease the remaining fraction of rotational power supplied to the differential gearbox 512 which is provided to the first electrical machine 522 such that 30 the rotational power extracted from the spool 440 by the blower system 400 is reduced. By way of a further example, if the first electrical machine 522 is initially operating as a generator, the first electrical machine 522 may be switched into operating as a motor. This results in a direction of the torque applied to the differential 35 gearbox 512 by the first electrical machine 522 being reversed. This may be done if CO CM a desired reduction in the rotational power extracted from the spool 440 by the blower system 400 cannot be attained by the first electrical machine 522 continuing to operate as a generator. Although the examples described above each relate to controlling the first 5 electrical machine 522 to have the effect of reducing the rotational power extracted from the spool 440, in other examples the same effect may be achieved in other ways. For example, as described previously, a valve in fluid communication with the rotor 410 (e.g., a recirculation valve or overboard bleed valve) may be controlled to reduce a compressive load on the rotor 410 (e.g., a torque applied to the rotor 410 due to the 10 work required to compress air provided thereto). Such control may be employed instead of controlling the first electrical machine 522, or to oppose any control of the first electrical machine 522 which in isolation would have the effect of increasing the power extracted from the spool 440. Accordingly, the control of the first electrical machine 522 in order to provide a desired reduction in the rotational power extracted 15 from the spool 440 may be considered to be a function of the operating point of the rotor 410 and a bleed state of any valve in fluid communication with the rotor 410. Each of these options described above can selectively be employed by the controller 490 to reduce an amount of rotational power drawn from the accessory gearbox 514 for driving the rotor 410 to rotate, which in turn reduces the rotational 20 power extracted from the spool 440 by the accessory gearbox 514 (and therefore reduced the rotational power extracted from the spool 440 by the blower system 400 as a whole). Although it has been described that the variable transmission 430 comprises an accessory gearbox 514 and a second electrical machine 524, this may not be the 25 case. For instance, the variable transmission 430 may comprise only the first electrical machine 522 and the differential gearbox 512, with the differential gearbox being otherwise mechanically coupled to the rotor 410 and the spool 440. Further, although the electrical machines 522, 524 have been shown and described as three-phase electrical machines in the specific example configuration of Figure 5, various types 30 and configuration of electrical machine are suitable for use in the variable transmission 430 as will be apparent to those skilled in the art. The DC bus 540 could also be replaced by an AC bus, and / or the power converters 532, 534 and / or the energy storage system 550 may be omitted. Figure 6 shows a schematic plan view of an aircraft 600 comprising a gas 35 turbine engine 10. The gas turbine engine 10 comprises a blower system 400 in accordance with the example blower system 400 described above with reference to Figures 4 and 5. The gas turbine engine 10 also comprises a controller configured to carry out a method of operating the gas turbine engine 10 in accordance with the example method described below with reference to Figure 8. Otherwise, the gas 5 turbine engine 10 may generally be in accordance with the gas turbine engine 10 described above with reference to Figures 1-3. Figure 7 shows, schematically, a non-transitory machine readable medium 700 comprising instructions 701 which, when executed by a processor 790 in data communication with the non-transitory machine readable medium, cause the 10 processor 790 to carry out a method of operating a gas turbine engine in accordance with the example gas turbine engine of Figure 4. The method of operating the gas turbine engine 10 carried out by the processor 790 is in accordance with the example method described below with reference to Figure 8. The processor 790 may be the controller 490 described above with reference to Figure 4. 15 Figure 8 is a flowchart which shows an example method 800 of controlling a gas turbine engine, such as the gas turbine engine 10 described above with reference to Figure 4. The method 800 may generally be a computer-implemented method or include computer-implemented control steps. Performance of the method 800 is generally associated with a reduced risk of a compressor 14, 15 of the gas turbine 20 engine 10 exhibiting (e.g., suffering from) engine compressor surge. As will be appreciated by those skilled in the art, the compressor 14, 15 exhibiting engine compressor surge may occur following the development of aerodynamic instabilities in the compressor 14, 15. The compressor 14, 15 is any axial compressor within the engine core 11 of the gas turbine engine 10, and may in particular be the low pressure 25 compressor 14 or the high pressure compressor 15 described above with reference to Figures 1 and 2. The compressor 14, 15 of the gas turbine engine 10 may be at an elevated risk of exhibiting engine compressor surge during transient operations of the gas turbine engine 10. The term “transient operation” of the gas turbine engine 10 as 30 employed herein refers to operation of the gas turbine engine 10 in which a rotational speed of the spool changes with respect to time, for example an intentional or controlled operation (e.g., “transient control operation”). The use of the term “transient operation” as used in the present disclosure does not refer to operations which go beyond the design limitations of the gas turbine engine 10 and therefore is not 35 intended to refer to abnormal operations during, for instance, a shaft break. On the other hand, the use of the term “transient operation” may in particular refer to a rapid acceleration of the gas turbine engine 10 or a demand to rapidly increase a thrust generating by the gas turbine engine 10. This may occur, for example, when the gas turbine engine 10 is accelerating from idle, such as during a take-off roll of an aircraft 5 to which the gas turbine engine is provided. The method 800 may generally be performed during a transient operation of the gas turbine engine 10, such that the elevated risk of exhibiting engine compressor surge during the transient operation is mitigated. The method 800 includes, at block 805, operating the blower system 400 to 10 extract rotational power from the spool 440 to drive the rotor 410 to rotate and compress air received from the gas turbine engine 10. In other words, block 805 comprises operating the blower system 400 in the compressor configuration and extracting rotational power from the spool 440 for this purpose. The method 800 includes continuing to extract the rotational power from the spool 440 for driving the 15 rotor 410 to rotate as the method proceeds through the remaining blocks which will now be described, subject to the step at block 830 (and, if present block 830’) as described below. Following block 805, the method 800, at block 810 includes receiving a signal associated with an operating parameter of the gas turbine engine 10. In an example, 20 the controller 490 may receive a signal from a thrust lever system of an aircraft to which the gas turbine engine 10 is provided or from an electronic engine controller (EEC) of the gas turbine engine 10. The signal from the thrust lever system or the EEC may relate to a thrust demand from the gas turbine engine 10 (e.g., a demand for the gas turbine engine 10 to produce thrust). In another example, the controller 25 490 may receive a signal from a rotational speed sensor configured to monitor a rotational speed of a rotary component of the gas turbine engine 10. The rotary component of the gas turbine engine 10 may be, for example, the low pressure compressor 14, the high pressure compressor 15, the high pressure turbine 17, the low pressure turbine 19 or the fan 23. Accordingly, the signal may relate to a rotational 30 speed of the low pressure compressor 14, the high pressure compressor 15, the high pressure turbine 17, the low pressure turbine 19 or the fan 23. Ina further example, the controller 490 may receive a signal from a fuel flow sensor configured to monitor a fuel flow rate to the combustor 16. Consequently, the signal may relate to a fuel flow to the combustor 16 of the gas turbine engine 10. In an additional example, the 35 controller 490 may receive a signal from a pressure sensing arrangement adapted to CO monitor a pressure difference (a pressure ratio) across a rotary component of the gas turbine engine 14, 15, 17, 19, 23. In a yet further example, the controller 490 may receive a signal from an acceleration sensor configured to monitor an acceleration of the gas turbine engine 10. The acceleration of the gas turbine engine 10 may be 5 defined as a rate-of-change in ground-speed of the gas turbine engine 10 or a rate-of-change in air-speed of the gas turbine engine 10. Each of the signals referred to in the examples above are associated with a respective operating parameter of the gas turbine engine 10 and may be used to determine whether a compressor 14, 15 of the gas turbine engine 10 is at risk of 10 exhibiting engine compressor surge. To this end, the method 800 also comprises, at block 820, evaluating a surge criterion relating to engine compressor surge, wherein the evaluation is based on the signal received at block 810. Following block 820, the method 800 may either proceed to block 830 or return to block 810. Whether the method 800 proceeds to block 830 or returns to block 810 is based on an output of 15 the surge criterion evaluation at block 810, as explained in further detail below. By way of example, the surge criterion may be defined with reference to an assessment parameter. If so, evaluating the surge criterion at block 820 includes determining whether the assessment parameter exceeds a predetermined range or threshold value based on the signal received at block 810. The assessment parameter 20 and the predetermined threshold value are each related to the signal associated with the operating parameter of the gas turbine engine 10 as received at block 810. The assessment parameter may be a magnitude of the operating parameter to which the signal received at block 810 relates or a rate-of-change of the operating parameter to which the signal received at block 810 relates. 25 The predetermined range or threshold value may be selected to correspond to a value of the assessment parameter above which a compressor 14, 15 of the gas turbine engine 10 is at risk of exhibiting engine compressor surge, such that the assessment parameter exceeding the predetermined range or threshold value is indicative of a risk of engine compressor surge in the gas turbine engine 10. During a 30 transient operation of the gas turbine engine 10, a working line of the compressor 14, 15 rises and approaches a surge line of the compressor. Accordingly, a surge margin of the compressor 14,15 is typically reduced in a transient operation of the gas turbine engine 10. The surge margin of the compressor 14, 15 may be reduced by an especially large degree when the gas turbine engine 10 is accelerating rapidly (such 35 as when accelerating from an idle state). In many of the examples described herein, the predetermined threshold value corresponds to a value of the assessment parameter which is indicative of the gas turbine engine 10 performing a transient operation and is therefore at risk of exhibiting engine compressor surge in the compressor 14, 15 thereof. The predetermined threshold value may be stored in a 5 memory provided to the controller 490. The relationship between the assessment parameter, the predetermined threshold value and the signal is explained further by way of various examples below. If the signal (received at block 810) relates to a thrust demand for the gas turbine engine 10, the controller 490 may determine (at block 820) whether a rate-of-10 change in the thrust demand for the gas turbine engine exceeds a predetermined thrust demand rate-of-change threshold value. In a similar way, if the signal (received at block 810) relates to a rotational speed of a rotary component 14, 15, 17, 19, 23 of the gas turbine engine 10, the controller 490 may determine (at block 820) whether a rate-of-change in the rotational speed of the rotary component 14, 15, 17, 19, 23 of 15 the gas turbine engine 10 exceeds a predetermined rotational speed rate-of-change threshold value. Otherwise, if the signal (received at block 810) relates to the fuel flow to the combustor 16, the controller 490 may determine (at block 820) whether a rate-of-change in the fuel flow to the combustor 16 exceeds a fuel flow rate-of-change predetermined threshold value. Further, if the operating parameter is the pressure 20 ratio across the compressor 14, 15, the controller 490 may determine (at block 820) whether the pressure ratio across the compressor 14, 15 exceeds a predetermined pressure ratio threshold value. The predetermined pressure ratio threshold value may be defined with respect to the rotational speed of the compressor 14, 15. Additionally, if the operating parameter is the acceleration of the gas turbine engine 10, the 25 controller 490 may determine (at block 820) whether the acceleration of the gas turbine engine 10 exceeds a predetermined acceleration threshold value. If it is determined, at block 820, that the assessment parameter does not exceed the predetermined threshold value (that is, it is determined that there is low risk or no risk of engine compressor surge), the method 800 returns to block 810. 30 Conversely, if it is determined, at block 820, that the assessment parameter exceeds the predetermined threshold value (that is, it is determined that there is risk or high risk of engine compressor surge), the method proceeds to block 830. The operating parameter may be one or more operating parameters of the gas turbine engine and there may be one or more associated signals. The surge criterion 35 may be evaluated based on one or more received signals associated with one or more operating parameters. For instance, block 810 may comprise receiving a plurality of signals, with each signal being associated with a respective operating parameter of the gas turbine engine 10. If so, the surge criterion may be evaluated at block 820 based on a plurality of received signals associated with a respective operating 5 parameters. As an example, it may be that block 810 comprises receiving a first signal and a second signal, with the first signal being associated with a first operating parameter of the gas turbine engine 10 and the second signal being associated with a second operating parameter of the gas turbine engine 10. The surge criterion evaluated at 10 block 820 may be defined with reference to a first assessment parameter and a second assessment parameter. Evaluating the surge criterion relating to engine compressor surge may include determining whether the first assessment parameter exceeds a first predetermined threshold value and also determining whether the second assessment parameter exceeds a second predetermined threshold value. 15 The method 800 may only proceed to block 830 if it is determined that the first assessment parameter exceeds the first predetermined threshold value and the second assessment parameter exceeds the second predetermined threshold value. Each received signal, operating parameter, assessment parameter and predetermined threshold value may generally be in accordance with the examples 20 described above. However, the first operating parameter and the second operating parameter may differ. For instance, the first operating parameter may be a rotational speed of the compressor 14, 15, a turbine 17, 19 or a fan 23 of the gas turbine engine 10 and the second operating parameter may be a pressure ratio across the compressor 14,15 or the turbine 17,19 of the gas turbine engine 10. In a similar way, 25 the first assessment parameter and the second assessment parameter may differ. For example, the first assessment parameter may be a rate-of-change in a rotational speed of the compressor 14, 15, a turbine 17, 19 or a fan 23 of the gas turbine engine 10 and the second assessment parameter may be a pressure ratio across the compressor 14, 15 or the turbine 17, 19 of the gas turbine engine 10. 30 Evaluating the surge criterion, at block 820, based on a plurality of received signals may reduce a probability of the method 800 (unnecessarily) proceeding to block 830 when the gas turbine engine 10 is not at substantial risk of exhibiting engine compressor surge, which in turn is associated with improved operability of the gas turbine engine 10. In other words, the use of at least two parameters and associated 35 signals may permit a more accurate evaluation of a surge risk. At block 830, the method includes causing a reduction of the rotational power extracted from the spool 440 by the blower system 400. The method 800 may include, at block 830, controlling the variable transmission 430 to cause the reduction of the rotational power extracted from the spool 440 by the blower system 400. If the variable 5 transmission 430 comprises an electrical machine and a differential gearbox such as the first electrical machine 522 and the differential gearbox 512 described above with reference to Figure 5, the method 800 may include, at block 830, controlling the first electrical machine 522 to cause the reduction of the rotational power extracted from the spool 440 by the blower system 400. In particular, the method 800 may include, 10 at block 830, controlling the first electrical machine 512 to reverse the torque applied to the differential gearbox 522 by the first electrical machine 522 and thereby cause the reduction of the rotational power extracted from the spool 440 by the blower system 400. This may generally include switching the first electrical machine 522 from operating as a generator to operating as motor, as described above with reference to 15 Figure 5. Additionally or alternatively, the method 800 may include, at block 830, controlling a valve in fluid communication with the rotor 410 to cause the reduction of the rotational power extracted from the spool 440 by the blower system 400, as described above with reference to Figure 4. In general, extracting power from the spool 440 of the gas turbine engine 10 20 has the effect of reducing the surge margin of the compressor 14,15 of the gas turbine engine 10. Therefore, by reducing the rotational power extracted from the spool 440 by the blower system 400 when the assessment parameter exceeds the predetermined threshold value, the method 800 increases the surge margin of the compressor 14, 15 and thereby reduces a risk of the compressor 14, 15 of the gas 25 turbine engine 10 exhibiting (e.g., suffering from) engine compressor surge. As a result, the method 800 is associated with a reduced risk of malfunction of the gas turbine engine 10 in use (e.g., due to flameout), as well as reduced wear on and / or damage to rotary components of the gas turbine engine 10 due to internal aerodynamic instabilities and vibration. 30 Further, because of the reduction in power extraction from the spool 440 of the gas turbine engine 10 during transient operations, the method 800 is associated with an improved transient thrust response of the gas turbine engine 10. Additionally, the increase in surge margin of the compressor 14, 15 enables the gas turbine engine 10 to safely respond to a larger acceleration demand or rate-of-change in thrust 35 demand, which in turn enables a lower idle thrust setting to be used for the gas turbine 10 CO 15 20 25 30 35 engine in use. This is associated with a reduced fuel consumption of the gas turbine engine 10 and / or an improved glideslope capture capability (e.g., a higher glideslope angle potential) of an aircraft to which the gas turbine engine 10 is provided. These effects are linked with increased operational flexibility of the aircraft and / or an improved noise signature of the aircraft due to the aircraft having a higher average altitude on approach. The method 800 may also include blocks 810’, 820’ and 830’, which relate to the subsequent monitoring of the engine and control to permit or cause an increase of the rotational energy extracted by the blower system, following the reduction of the rotational energy extracted from the spool by the blower system. If so, the method 800 proceeds from block 830 to block 810’, such that the contents of block 810’ is performed subsequent to (e.g., chronologically after) the reduction in the rotational power extracted from the spool 440 by the blower system 400 in block 830. The method steps of each of blocks 810’ and 820’ are generally similar to (e.g., substantially identical to) those described above with respect to blocks 810 and 820, respectively. For example, both blocks 810 and block 810’ include receiving a signal associated with the operation parameter of the gas turbine engine 10, and both blocks 820 and 820’ include evaluating a surge criterion relating to engine compressor surge. The signal(s) described with reference to blocks 810 and 810’ may be a signal relating to the operating parameter of the gas turbine engine 10 which is continuously monitored or sampled. In some examples, the signal(s) received in blocks 810 and 810’ may be the same signal. In other examples, the signal(s) received at blocks 810 and 810’ may be separate signals. In addition, the response to the output of the surge criterion evaluation at block 820’ may differ to the response to the output of the surge criterion evaluation at block 820 described above. For instance, if it is determined, at block 820’, that the assessment parameter does not exceed the predetermined threshold value (that is, it is determined that there is low risk or no risk of engine compressor surge), the method 800 proceeds to block 830’ rather than returning to block 810’. Conversely, if it is determined, at block 820’, that the assessment parameter exceeds the predetermined threshold value (that is, it is determined that there is risk or high risk of engine compressor surge), the method returns to block 810’ rather than proceeding to block 830’. In contrast to block 830, the method 800 at block 830’ comprises includes causing an increase of the rotational power extracted from the spool 440 by the blower CM system 400. In particular, at block 830’ the method 800 may comprise reversing any operations previously carried out at block 830 such that after performance of block 830’, the power extracted from the spool 440 by the blower system 400 is the same or closer to the amount as it was before block 830 was performed. The method 800 5 may then return to block 805 and then to block 810 as described above. Accordingly, the method 800 may include, at block 830’, controlling the variable transmission 430 to cause the increase of the rotational power extracted from the spool 440 by the blower system 400. If the variable transmission 430 comprises an electrical machine and a differential gearbox such as the first electrical machine 10 522 and the differential gearbox 512 described above with reference to Figure 5, the method 800 may include, at block 830’, controlling the first electrical machine 522 to cause the increase of the rotational power extracted from the spool 440 by the blower system 400. In particular, the method 800 may include, at block 830’, controlling the first electrical machine 512 to reverse the torque applied to the differential gearbox 15 522 by the first electrical machine 522 and thereby cause the increase of the rotational power extracted from the spool 440 by the blower system 400. This may generally include switching the first electrical machine 522 from operating as a motor to operating as generator, as described above with reference to Figure 5. Additionally or alternatively, the method 800 may include, at block 830’, controlling a valve in fluid 20 communication with the rotor 410 to cause the increase of the rotational power extracted from the spool 440 by the blower system 400, as described above with reference to Figure 4. It will be understood that the invention is not limited to the embodiments above described and various modifications and improvements can be made without 25 departing from the concepts described herein. Except where mutually exclusive, any of the features may be employed separately or in combination with any other features and the disclosure extends to and includes all combinations and sub-combinations of one or more features described herein.

Claims

1. A method (800) of controlling a gas turbine engine (10) comprising a blower system (400) including a rotor (410) drivingly coupled to a spool (440) of the gas turbine engine 5 via a variable transmission (430), the variable transmission including an electrical machine (522) operable as a motor and a differential gearbox (512) mechanically coupled to the electrical machine, the spool (440) of the gas turbine engine (10) and the rotor (410) of the blower system (400), the method comprising:operating (805) the blower system to extract rotational power from the spool to 10 drive the rotor to rotate and compress air received from the gas turbine engine;receiving (810) a signal associated with an operating parameter of the gas turbine engine;evaluating (820) a surge criterion relating to engine compressor surge based on the received signal; and15 controlling (830, 832) the electrical machine to cause a reduction (830) of the rotational power extracted from the spool by the blower system in response to an output of the surge criterion evaluation.

2. The method (800) of claim 1, wherein the method is performed during a transient 20 operation of the gas turbine engine (10) in which a rotational speed of the spool changes over time.

3. The method (800) of claim 1 or claim 2, comprising controlling (832) the electrical machine (522) to reverse a torque applied to the differential gearbox (512) by the 25 electrical machine and thereby cause the reduction (830) of the rotational power.

4. The method (800) of any preceding claim, comprising controlling (834) a valve in fluid communication with the rotor (410) to cause the reduction (830) of the rotational power.

305. The method (800) of any preceding claim, wherein the operating parameter is selected from a group consisting of:a thrust demand for the gas turbine engine (10);a rotational speed of a compressor (14, 15), a turbine (17, 19) or a fan (23) of the gas turbine engine;a fuel flow to a combustor (16) of the gas turbine engine;a pressure ratio across a compressor (14, 15) or a turbine (17, 19) of the gas 5 turbine engine; andan acceleration of the gas turbine engine.10CO1520256. The method (800) of any preceding claim, wherein the surge criterion is defined with reference to an assessment parameter selected from a group consisting of:a rate-of-change in a thrust demand for the gas turbine engine (10);a rate-of-change in a rotational speed of a compressor (14, 15), a turbine (17, 19) or a fan (23) of the gas turbine engine;a rate-of-change in a fuel flow to a combustor (16) of the gas turbine engine;a pressure ratio across a compressor (14, 15) or a turbine (17, 19) of the gas turbine engine; andan acceleration of the gas turbine engine.

7. The method (800) of any preceding claim, wherein subsequent to the reduction (830) of the rotational power, the method comprises:receiving (810’) a signal associated with the operating parameter of the gas turbine engine (10);evaluating (820’) a surge criterion relating to engine compressor surge based on the received signal; andcausing an increase (830’) of the rotational power extracted from the spool by the blower system in response to an output of the surge criterion evaluation.

8. The method (800) of claim 7, wherein the method comprises controlling (832’) the variable transmission to cause the increase (830’) of the rotational power.30 9. The method (800) of claim 8, wherein the method comprises controlling (832’) theelectrical machine to cause the increase (830’) of the rotational power.

10. The method (800) of claim 9, wherein the method comprises controlling (832’) the electrical machine (522) to reverse a torque applied to the differential gearbox (512) by 35 the electrical machine and thereby cause the increase (830’) of the rotational power.

11. The method (800) of any preceding claim, comprising controlling (834’) a valve in fluid communication with the rotor (410) to cause the increase (830’) of the rotational power.

512. A controller (490) configured to perform the method (800) of any of claims 1 to13.

13. A non-transitory machine readable medium (700) comprising instructions (701) which, when executed by the controller (490) of claim 12, cause the controller to carry 10 out the method (800) of any of claims 1 to 11.

14. A gas turbine engine (10) comprising: the controller (490) of claim 12;CO a blower system (400) including a rotor (410) drivingly coupled to a spool (440) of£\J 15 the gas turbine engine via a variable transmission (430), the variable transmissionincluding an electrical machine (522) operable as a motor and a differential gearbox (512) mechanically coupled to the electrical machine, the spool (440) of the gas turbine engine (10) and the rotor (410) of the blower system (400).CM20 15. The gas turbine engine (10) of claim 14, comprising:an engine core (11) comprising a turbine (19), a compressor (14), and a core shaft (26) connecting the turbine to the compressor;a fan (23) located upstream of the engine core, the fan comprising a plurality of fan blades; and25 a gearbox (30) that receives an input from the core shaft (26) and outputs drive to the fan so as to drive the fan at a lower rotational speed than the core shaft.

16. The gas turbine engine (10) of claim 15, wherein:the turbine is a first turbine (19), the compressor is a first compressor (14), and 30 the core shaft is a first core shaft (26);the engine core (11) further comprises a second turbine (17), a second compressor (15), and a second core shaft (27) connecting the second turbine to the second compressor; andthe second turbine, second compressor, and second core shaft are arranged to 35 rotate at a higher rotational speed than the first core shaft.

17. An aircraft (600) comprising the gas turbine engine (10) of any of claims 14 to 16.COCM

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