Rolls-Royce Patente

This disclosure relates to bidirectional current limiting circuits for use in aircraft electric power distribution systems. Example embodiments include a bidirectional current limiting circuit (500a-c) comprising first and second JFETs (501, 502) connected between first and second terminals (503, 504), wherein: the first terminal (503) is connected to a drain of the first JFET (501) and to a gate of the second JFET (502) via a first biasing element (505, 701); he second terminal (504) is connected to a drain of the second JFET (502) and to a gate of the first JFET (501) via a second biasing element (506, 702); and a source of the first JFET (501) is connected to a source of the second JFET (502).

The disclosure relates to bidirectional current limiting circuits for use in aircraft electric power distribution system and methods for operating such bidirectional current limiters. A bidirectional current limiting circuit (400) comprises first and second parallel current paths (401, 402) connected between first and second terminals (403, 404), the first current path (401) comprising a first current limiting device (405) connected in series with a measurement inductor (406), the second current path (402) comprising a damping resistor (407) in series with a second current limiting device (408), wherein the first current limiting device is a first JFET (405), the circuit (400) further comprising a controller (409) connected to a gate of the first JFET (405), the controller (409) connected to receive a voltage measurement across the measurement inductor (406) and provide a first gate voltage to the first JFET (405) to restrict current through the first path (401) if the voltage measurement exceeds a voltage threshold.

An aircraft gas turbine engine (103) comprises a fuel system comprising a first fuel line (108) configured to provide a first liquid hydrocarbon fuel to a combustor (203) of the gas turbine engine (103), and a second fuel line (109) configured to provide a second fuel, different to the first fuel, to the combustor (203) of the gas turbine engine (103). The gas turbine engine also comprises a hydraulic system configured to be driven by the first fuel during operation of the gas turbine combustor (203) when the gas turbine combustor utilises either or both of the first and second fuel.

This disclosure relates to cooling modules for power electronics converters. Example embodiments include a cooling module (100) for a power electronics converter, comprising: first face (101) having a cooling liquid inlet (102); an opposing second face (103) having a cooling liquid outlet (104); a plurality of side walls (105a-d) extending between the first and second faces (101, 103), each side wall (105a-d) being configured for cooling a semiconductor switching device (106<1-6>) of the power electronics converter; and a helical cooling channel (107) extending through the cooling module between the cooling liquid inlet (102) and the cooling liquid outlet (104).

A method (200) for determining grain structure in a component (102) includes disposing the component (102) on a support platform (110). The method (200) includes generating an imaging beam (106). The method (200) includes rotating and/or revolving an imaging beam source (104) and an imaging beam receiver (108) relative to the support platform (110). The method (200) includes acquiring, during the rotation and/or revolution a plurality of projections (116) each taken at the imaging beam receiver (108). The method (200) includes performing reconstruction, by a processor (114), of first projections (116-1) and second projections (116-2) from the plurality of projections (116) to obtain a first three-dimensional image (120) and at least one second three-dimensional image (122). The first projections (116-1) are acquired in at least one first angular range (AR1) and the second projections (116-2) are acquired in at least one second angular range (AR2).

An arrangement (1) for an electric machine (2, 2') comprises: a first winding system (WS1) and a second winding system (WS2), wherein at least one electrical conductor (120) of the first winding system (WS1) and at least one electrical conductor (121) of the second winding system (WS2) are at least partially wound around a common winding axis (A), wherein turns of the electrical conductors (120, 121) alternate in a direction of the winding axis (A); an inverter system (13) with a first inverter (134A) for the first winding system (WS1) and a second inverter (134B) for the second winding system (WS2); sensors (133) for sensing at least one electrical parameter of the first winding system (WS1) and at least one electrical parameter of the second winding system (WS2); and a control unit (131) configured to perform a test of the winding systems (WS1, WS2) by: controlling the first inverter (134A) to set the first winding system (WS1) to a first electrical potential (HV+) and/ot to control the second inverter (134B) to set the second winding system (WS2) to a second electrical potential (HV-) to create a potential difference between the winding systems (WS1, WS2); receiving corresponding sensor signals from the sensors (133); and analyzing the sensor signals.

A cleaning tool (100) for cleaning a component (50) located within a gas turbine engine (10) is disclosed. The cleaning tool (100) includes a tubular body (102) including: a first open end (104) and a second closed end (106); an engagement portion (108) that is proximal to the second closed end (106) and configured to engage with a support (110) within the gas turbine engine (10); a nozzle aperture (112) disposed between the first open end (104) and the engagement portion (108); and a flow channel (114) extending from the first open end (104) and in fluid communication with the nozzle aperture (112). The flow channel (114) is configured to transport a cleaning fluid (115) from the first open end (104) to the nozzle aperture (112). The cleaning fluid (115) is delivered to the component (50) through the nozzle aperture (112). The tubular body (102) is configured to be inserted through a first port (116) of the gas turbine engine (10) from the second closed end (106), such that the engagement portion (108) engages with the support (110).

There is disclosed an exhaust nozzle 30 for a gas turbine engine, comprising an exhaust structure 32 and a flap 34 configured to bound a passageway 36 through the exhaust structure. The flap 34 is coupled to the exhaust structure in a linkage arrangement configured to move the flap by compound translation and rotation with a single degree of freedom between a first orientation corresponding to a maximum throat area of the passageway and a second orientation corresponding to a minimum throat area of the passageway.

A climatic test conditioning apparatus (100) for conditioning a gas turbine engine (10) for climatic testing comprises a base (110) and an inflatable wall (120) connected to the base (110). The inflatable wall (120) has at least one port for at least partial inflation and deflation of the inflatable wall (120). The at least partially inflated inflatable wall (120) and base (110) define a volume (140) that envelopes the gas turbine engine (10). Also disclosed is a system and method for climatic test conditioning a gas turbine engine (10).

A fuel system for a gas turbine engine (103) is disclosed. The fuel system is configured to combust hydrogen fuel and comprises a fuel conduit (228), a first fuel pump (226) configured to operate on liquid hydrogen within the fuel conduit (228) and a second fuel pump (230) downstream in hydrogen fuel flow of the first fuel pump (226). The first fuel pump comprises one of a side channel and a regenerative pump (226) and the second fuel pump comprises a centrifugal pump (230). Methods of operation, a gas turbine engine comprising the fuel system, and an aircraft comprising the gas turbine engine are also disclosed.

There is provided a method of training an encoder model to map inputs representing nominally healthy assets to a cluster of data points. The method comprises providing an input representing a behaviour of an asset at a first point in time to the encoder model undergoing training to map the input to a data point in the feature space, wherein the encoder model is intended to cause data points relating to measured values of input parameters acquired from the same asset at similar time to cluster in feature space, providing the data point to a decoder model, to determine a reconstructed input, determining a difference between the input, and the reconstructed input, and training the encoder model based on the determined difference to map inputs representing nominally healthy assets acquired at similar times to a cluster of data points.

There is provided a method of monitoring the health of an asset. The method comprises providing an input representing the behaviour of the asset to an encoder model to map the input to a data point in a feature space, wherein the encoder model is configured to cause data points relating to measured values of input parameters acquired from the asset at a similar time to cluster in feature space, determining a distance of the data point relative to a cluster of data points in feature space, where the cluster of data points are determined by the encoder based on inputs representing the asset over a first time period, and in response to the determined distance exceeding a predetermined distance threshold, generating an abnormal asset behaviour prediction.

There is provided a method of monitoring the health of an asset. The method comprises obtaining a measured value of an input parameter representing an operational state of the asset, determining, based on the value, a residual that represents the behaviour of the asset, providing the determined residual to an encoder model to map the determined residual to a data point in a feature space, determining a distance of the data point relative to a cluster of data points in the feature space, wherein the cluster of data points are determined by the encoder based on residuals representing nominally healthy assets, and in response to the determined distance exceeding a predetermined distance threshold, generating an abnormal asset behaviour prediction.

A fuel system for a hydrogen fuelled gas turbine engine (201). The fuel system comprises a main hydrogen fuel line (206) configured to provide hydrogen fuel to a combustor (203) of the gas turbine engine (201) and a closed loop heat exchange fluid line (207) comprising a heat exchange fluid. The fuel system further comprises a first heat exchanger (208) configured to exchange heat between gas turbine engine compressor bleed air and heat exchange fluid in the heat exchange fluid line (207), a second heat exchanger (209) downstream of the first heat exchanger (208) in heat exchange fluid flow configured to exchange heat between the heated heat exchange fluid and hydrogen in the main hydrogen fuel line, and a third heat exchanger (215) downstream in main hydrogen fuel flow of the second heat exchanger (209) in main hydrogen fuel flow configured to exchange heat between a further engine fluid and heated hydrogen fuel.

A fuel system for a hydrogen fuelled gas turbine engine (201). The fuel system comprises a main hydrogen fuel line (206) configured to provide hydrogen fuel to a combustor (203) of the gas turbine engine (201) and a closed loop heat exchange fluid line (207) comprising a heat exchange fluid. The fuel system further comprises a first heat exchanger (208) configured to exchange heat between gas turbine engine compressor bleed air and heat exchange fluid in the heat exchange fluid line (207), a second heat exchanger (209) downstream of the first heat exchanger (208) in heat exchange fluid flow configured to exchange heat between the heated heat exchange fluid and hydrogen in the main hydrogen fuel line, and a third heat exchanger (215) downstream in main hydrogen fuel flow of the second heat exchanger (209) in main hydrogen fuel flow configured to exchange heat between a further engine fluid and heated hydrogen fuel.

A method (200) for scanning a plurality of components (102). The method (200) includes providing an imaging beam source (108), an imaging beam receiver (114), and a support platform (104). The support platform (104) is configured to rotate and/or revolve relative to the imaging beam source (108) and the imaging beam receiver (114) about one or more axes. The method (200) further includes disposing the plurality of components (102) on the support platform (104). The plurality of components (102) is positioned in a gap (118) between a first geometrical figure (120) and at least one second geometrical figure (122) in a configuration that reduces the variation in material thickness penetrated at different relative angles of rotation and/or revolution. The first geometrical figure (120) and the at least one second geometrical figure (122) are concentric.

A nosecone (30) for a gas turbine engine (10) includes a nosecone body (31). The nosecone body is substantially symmetrical around a central axis (X-X). The nosecone body includes a side wall which defines a diameter of the nosecone body that increases from a forward side of the nosecone body to an aft side of the nosecone body. The nosecone includes an attachment assembly (33). The attachment assembly is configured to attach the nosecone body to the gas turbine engine. At least one of the nosecone body or the attachment assembly comprises a tear-resistant fiber.

A nosecone for a gas turbine engine includes a nosecone body substantially symmetrical around a central axis. The nosecone body includes a side wall which defines a diameter of the nosecone body that increases from a forward side to an aft side of the nosecone body. The nosecone further includes an attachment assembly which attaches the nosecone body to the gas turbine engine. The nosecone includes a plurality of airstream protrusions attached to or integral with the nosecone body. A first airstream protrusion of the plurality of airstream protrusions is positioned upstream of a first fan blade to prepare an airstream flowing over the nosecone body for interaction with the first fan blade. The attachment assembly includes a plurality of attachment locations, and a first attachment location of the plurality of attachment locations is collocated with the first airstream protrusion of the plurality of airstream protrusions.

A gas turbine engine (10) including a rear attachment (400) comprising: a plurality of angularly distributed rotational pylon connections (221, 222, 223) for coupling to a wing pylon (200) and defining a pylon connection angular extent; a plurality of angularly distributed engine connections (121, 122, 123) for coupling to a rear mount structure (117) of the gas turbine engine 10 and defining an engine connection angular extent; and a link structure extending between the pylon connections (221, 222, 223) and the engine connections (121, 122, 123); wherein the link structure, pylon connections (221, 222, 223) and engine connections (121, 122, 123) are configured to transfer a roll torque from the rear mount structure (117) to the wing pylon (200); and wherein a ratio of the engine connection angular extent to the pylon connection angular extent is at least 1.75.

There is provided a propulsion machine 10 comprising a fluid duct defined by a wall 42, 44 and a moveable member 34, 36. The propulsion machine 10 also comprises a mounting structure 92 coupled to the moveable member 34, 36, and an extendable structure 94 having a sealing surface 96 and a sensing arrangement 83. The extendable structure 94 is moveable relative to the mounting structure 92 to provide a seal between the sealing surface 96 and an opposing surface 43 of the wall 42. The sensing arrangement 83 is configured to generate one or more signals indicative of a position of the extendable structure 94 relative to the mounting structure 92.

There is provided a fluid flow machine (10) comprising: a casing structure (24), a turbomachine blade (311) disposed within the casing structure (24), and a sealing arrangement (330) coupled to the casing structure (24). The sealing arrangement (330) comprises (330) an abrasion portion (334) including a lattice structure (400). The abrasion portion (334) is configured to provide a seal with a tip (321) of the turbomachine blade (311). The abrasion portion (334) includes the lattice structure (400) having a lattice density that increases along a radial direction (42) of the machine (10).There is also provided a method (600) of manufacturing such a fluid flow machine (10).

There is provided a reheat assembly (300A, 300B, 300C) for a gas turbine engine (10). The reheat assembly (300A, 300B, 300C) comprises a jetpipe casing (310), a support duct (340), a plurality of fuel discharge ports (362), and a fuel distribution passageway (342). The jetpipe casing (310) comprises a reheat core section (320) and a reheat bypass section (330). The support duct (340) radially separates the reheat core section (320) and the reheat bypass section (330). Each fuel discharge port (362) is configured to discharge fuel received from the fuel distribution passageway (342). The fuel distribution passageway (342) is defined by an interior channel embedded within the support duct (340).

There is provided a flameholder (370) for a reheat assembly (300) of a gas turbine engine (10). The flameholder comprises a fuel atomizer (379A, 379B), an air passageway (375A, 375B) and a fuel passageway (374, 374A, 374B). The air passageway (375A, 375B) is configured to convey a flow of air (521A, 521B) to the fuel atomizer (379A, 379B). The fuel passageway (374, 374A, 374B) is configured to convey a flow of fuel (531A) to the fuel atomizer (379A, 379B). The fuel atomizer (379A, 379B) includes a mixing chamber (430) and a swirler (420).

A turbine exit diffuser for a gas turbine engine, the turbine exit diffuser comprising: an exhaust duct arranged to receive flow from a turbine; and a plurality of fingers extending axially downstream and radially away from the exhaust duct, each finger of the plurality being circumferentially spaced apart from an adjacent finger of the plurality, wherein the fingers are arranged to contact a reheater during attachment of the reheater to the turbine exit diffuser.

An aerofoil component for a high-pressure or intermediate-pressure turbine (16, 17) of a multi-spool gas turbine engine (10). The component having an aerofoil member (31, 41) which, in use, spans a working gas annulus of the gas turbine engine. The aerofoil member has pressure side and suction side aerofoil surfaces which each extend from a leading edge to a trailing edge of the aerofoil member such that transverse cross sections through the aerofoil member provide respective aerofoil sections. The spacing between the leading and trailing edges on the midspan aerofoil section (MS) defines a midspan axial chord length (C<x>) of the aerofoil member. The aerofoil member are arranged, in use, in a circumferential row around the annulus with plural, equally spaced, identical aerofoil members whereby the circumferential spacing of the aerofoil members at the trailing edges of their midspan aerofoil sections (MS) defines a midspan pitch (S) of the aerofoil member. At midspan: the value of S/C<x> is in the range from 1.4 to 1.6; and the uncovered turning angle is 16° or more.

A fuel injector for a combustor (15, 300) of a gas turbine engine (10) is described. The fuel injector comprises a fuel injection device (360) having a fuel injection port (362) configured to inject fuel into the combustor, wherein the fuel injection device is configured to inhibit cavitation of the fuel. Also disclosed is a combustor (15) for a gas turbine engine (10), a reheat assembly (300) for a gas turbine engine (10), a gas turbine engine (10), and an aircraft (200).

There is provided a fluid flow machine (10) comprising a casing structure (24), a turbomachine blade (311) disposed within the casing structure (24), and a sealing arrangement (330) coupled to the casing structure (24). The sealing arrangement (330) comprises (330) an abrasion portion (334) including a lattice structure (400). The abrasion portion (334) is configured to provide a seal with a tip (321) of the turbomachine blade (311). The abrasion portion (334) includes a filler (500) within the lattice structure (400), the filler comprising one or more selected from: a polyester, a polyimide, a cyanate ester, a siloxane, a polysiloxane or a polyepoxide.There is also provided a vehicle (200) comprising such a fluid flow machine (10).There is also provided a method (600) of manufacturing such a fluid flow machine (10).

The disclosure relates to apparatus for fuel and air injection into a combustion chamber of a gas turbine engine and to a gas turbine engine comprising such an apparatus. Example embodiments include an apparatus (200) for fuel and air injection into a combustion chamber (201) of a gas turbine engine (10), the apparatus (200) comprising: back panel (202) having front and back faces (203, 204), an aperture (205) configured to receive a fuel injector (206) and a plurality of air inlets (207) arranged around the aperture (205); a heatshield panel (208) having front and back faces (209, 210), the heatshield panel (208) having a central aperture (212) and an annular side wall (211) mounted against the front face of the back panel to form a cavity (213) between the back face of the heatshield panel and the front face of the back panel (202); and an annular swirler assembly (214) disposed in the cavity (213) between the back panel (202) and the heatshield panel (208), the annular swirler assembly (214) having a central passage (215) for flow of an air and fuel mixture into the combustion chamber (201) and a flange (216) engaged with first and second grooves (217a, 217b) in the back face (210) of the heatshield panel (208) and the front face (203) of the back panel (202) respectively.

The foregoing describes a power converter apparatus (200) with at least a first power converter (101-1) and a second power converter (101-2) connected in parallel, so as to each receive an AC voltage generated by an electrical machine (104) and output a DC voltage on a DC bus (105P, 105N). Each of first and second power converters (101-1, 101-2) is coupled to respective first and second winding arrangements (102-1, 102-2) for receiving the AC voltages, and the first winding arrangement is grounded by a first impedance (201-1) and the second winding arrangement is grounded by a second impedance (201-2). The first impedance (201-1) has an effective impedance value which is different to that of the second impedance (202-1) such, in use, a zero-sequence current between the first and second power converters at least partly cancels a common-mode voltage component.

There is provided a fluid flow machine (10) comprising a casing structure (24) extending around an axial direction (41) of the fluid flow machine (10) and a turbomachine blade (311-315) disposed within the casing structure (24) The casing structure (24) includes an axially extending slot (331-335) having an axial extent (32) at least partially overlapping with an axial extent (31) of a tip (321-325) of the turbomachine blade (311-315), wherein the slot (331-335) has an angular extent (29) of no more than 36 degrees. The casing structure (24) includes a circumferentially extending groove (341, 342) having an angular extent (28) of at least 72 degrees. The groove (341, 342) is axially offset from the slot (331-335).

An apparatus (200) for in-situ application of an engineering coating (52) to components (50) has a head section (300) for applying the engineering coating (52) and a multi-part body section (400) for controlling the application process. The body section (400) has a first body member (402), a second body member (404), and a third body member (406) interconnected with each other. The first body member (402) includes a predetermined pathway (410) that guides movement and is connected to the head section (300). The second body member (404) is engaged with the first body member (402) and includes at least one hollow channel (412). The third body member (406) houses an actuator (420), an engagement pin (414) for connecting with the pathway, and at least one coupling member (418) for interacting with the at least one hollow channel (412). The actuator (420) ensures precise movement of the head section (300). The apparatus (200) further includes one or more gas channels (500) to provide necessary gases to the head section (300).

A method (100) of performing spectral analysis on a sample (220) identified within a gas turbine engine (10). The method (100) involves: obtaining (102) an optical spectrum (240) of the sample (220); normalising (104) the optical spectrum (240) to obtain a normalised optical spectrum (301); selecting (106) a characteristic parameter that characterises the normalised optical spectrum (301); determining (108) a characteristic parameter value associated with the normalised optical spectrum (301) based on the characteristic parameter; comparing (110) the characteristic parameter value with one or more predetermined reference values corresponding to the characteristic parameter; and determining (112) a presence of at least one compound in the sample (220) based on the comparison between the characteristic parameter value and the one or more predetermined reference values. The method can determine whether there is a need to disassemble the gas turbine engine for maintenance.

A method (200) for scanning a component (102) includes providing a filter (104) and the component, the filter at least partially surrounding the component. The filter comprises a multi-phase material (140) that has an attenuation coefficient (C1) that is lower than an attenuation coefficient (C2) of a material of the component (102) to be scanned. The method further includes disposing the component and the filter on a support platform (120), and providing an imaging beam source (108) and an imaging beam receiver (114). The support platform (120) is configured to rotate and/or revolve relative to the imaging beam source (108) and the imaging beam receiver (114) about one or more axes. The method includes generating, via the imaging beam source (108), an imaging beam (110) that passes through the component and the filter. The method includes attenuating, via the, a scatter beam (111) that is produced upon irradiation of the component (102) with the imaging beam (110).

An apparatus (200) for in-situ application of an engineering coating (52) to one or more components (50). The apparatus has a head section (300) adapted to selectively apply the engineering coating (52) to the components (50). The head section (300) has an applicator end (302) and a coupling end (304) opposite to the applicator end (302). The apparatus has a body section (400) coupled to the head section (300) at the coupling end (304) and adapted to actuate the head section (300). The body section (400) has a first stage actuator (402) adapted to actuate the head section (300) to swivel about a first axis (X-X') and a second stage actuator (404) adapted to actuate the head section (300) and the first stage actuator (402) together to rotate about a second axis (Y-Y'). The second axis (Y-Y') is orthogonal to the first axis (X-X'). The apparatus has one or more gas channels (500) provided along the body section (400) to the head section (300).

The foregoing describes a reheat assembly (300) for a gas turbine engine (100). The reheat assembly (300) comprises: a support duct section (321') comprising a core side (322') for facing a core reheat region (232), and a bypass side for facing a bypass reheat region (231); a support duct (21') comprising a plurality of circumferentially spaced inlets (380), each inlet (380) being configured to communicate the bypass reheat region (231) with the core reheat region (232); a fuel supply system (360) comprising a plurality of bypass fuel injection ports (363), each bypass fuel injection port (363) being associated with an inlet (380) of the plurality of inlets (380); wherein each bypass fuel injection port (363) is configured to discharge fuel into the bypass reheat region (231) for transit through the associated inlet into the core reheat region (232).

The foregoing describes a reheat assembly for a gas turbine engine. The reheat assembly comprises a support duct section, a vaporiser and a heat shield. The support duct section comprises a core side and a bypass side. During operation of the engine, the core side faces a core flow and the bypass side that faces a bypass flow. The vaporiser is attached to the bypass side of the support duct and comprises a plenum configured to cool core flow introduced into the plenum through an inlet to the vaporiser. The heat shield is configured to face the core side of the support duct section to shield the plenum from heat emitted by the core flow. The heat shield is spaced from core side of the support duct section to form a channel between the support duct section and the heat shield that extends upstream of the inlet.

A method of inspecting a component of a gas turbine engine for an aircraft, in which a housing circumferentially surrounds the component so as to define an annular space between the housing and the component. The method comprises guiding a guide line into the annular space through an opening to the annular space and around the component in a circumferential direction, and withdrawing the guide line through the opening, such withdrawal moving a probe that is connected to the guide line in a circumferential direction around the component for performing an inspection of the compressor.

An apparatus (100) for measuring a clearance (C) between a first component (102) and one or more areas of interest (104) of a second component (106). The apparatus (100) includes an imaging beam source (108) configured to generate an imaging beam (110) that passes between the first component (102) and the second component (106). The apparatus (100) includes a first tube (112) configured to guide the imaging beam (110) from the imaging beam source (108) to the one or more areas of interest (104) of the second component (106). The apparatus (100) includes an imaging beam receiver (114) configured to receive the imaging beam (110). The imaging beam receiver (114) is configured to generate an image (116) in response to receiving the imaging beam (110). The image (116) depicts the clearance (C) between the first component (102) and the one or more areas of interest (104) of the second component (106).

A device (100) for wear testing includes a frame (106), a rotary actuator (108), and a linear actuator (110) spaced apart from the rotary actuator (108) along a longitudinal axis (102) of the frame (106). The device (100) has first and second specimen holding units (112, 140). The first specimen holding unit (112) includes a seat (124) connected to the linear actuator (110), a plate (114) engaging with the seat (124), and a first thrust bearing (136) connected to the plate (114) and at least one first diaphragm plate (134) connected to the frame (106). The second specimen holding unit (140) includes a shaft (142) connected to the rotary actuator (108) and a second thrust bearing (150) connected the frame (106). The rotary actuator (108) is configured to rotate a second specimen (60) about the longitudinal axis (102). The linear actuator (110) is configured to move a first specimen (50) along the longitudinal axis (102), such that the first specimen (50) engages with the second specimen (60).

A power event manager controller for a gas turbine engine power system, the gas turbine engine power system comprising at least one gas turbine engine, a thermal management system and at least one of a generator, an energy storage system, the power event manager controller receiving an input power demand, and receiving input relating to the mission plan, and inputs from each of the systems within the gas turbine engine power system, based upon the inputs from the from the power demand and the mission plan the power event controller defines a series of constraints, and wherein the power event manager controller utilises an optimiser function to obtain control reference trajectories which minimise an objective cost function of modelled states, whilst being subject to the series of constraints.