A gas turbine engine for an aircraft includes an engine core with a turbine, a combustor, a compressor, and a core shaft connecting the turbine to the compressor. The engine includes a fan upstream of the engine core and driven by the core shaft. The engine includes a nacelle surrounding the fan and engine core and defining a bypass duct radially outside of the engine core. The bypass ratio, defined as the ratio of the mass flow rate through the bypass duct to the mass flow rate through the core at cruise conditions, is at least 4. The engine includes a generator heat management system and an actuator to actuate a valve within the generator heat management system. The engine comprises a fuel supply system arranged to supply fuel to fueldraulically drive the actuator. The actuator enables non-binary position adjustment between an open valve position and a closed valve position.
F02C 7/14 - Cooling of plants of fluids in the plant
F01D 17/26 - Devices dealing with sensing elements or final actuators or transmitting means between them, e.g. power-assisted the operation or power assistance being predominantly non-mechanical fluid, e.g. hydraulic
F02C 6/06 - Gas-turbine plants providing heated or pressurised working fluid for other apparatus, e.g. without mechanical power output providing compressed gas
F02C 7/224 - Heating fuel before feeding to the burner
F02C 7/232 - Fuel valvesDraining valves or systems
F02C 9/28 - Regulating systems responsive to plant or ambient parameters, e.g. temperature, pressure, rotor speed
B64D 13/06 - Arrangements or adaptations of air-treatment apparatus for aircraft crew or passengers, or freight space the air being conditioned
A method of operating a gas turbine engine with a heat exchange system including an air-oil heat exchanger through which oil flows; a fuel-oil heat exchanger through which the oil and fuel flow to transfer heat between the oil and fuel; and a temperature sensor to indicate fuel temperature downstream of the fuel-oil heat exchanger; and a valve to allow at least one of oil and air flow rate through at least one of the fuel-oil heat exchanger and the air-oil heat exchanger to be varied. The method includes determining if the fuel temperature has increased above a set threshold at cruise conditions; and in response to determining that the fuel temperature has increased above the set threshold at cruise conditions, controlling the at least one valve to change the flow rate through heat exchanger.
Rolls-Royce North American Technologies Inc. (USA)
Inventor
Marengo, Giovanni A.
Remmert, Ariane
Kimmerle, Maria
Dos Santos Ramos Lopes E Paiva, Ricardo Miguel
Peters, Jeffrey T.
Abstract
A climatic test conditioning apparatus for conditioning a machine for climatic testing comprises a base and an inflatable wall connected to the base. The inflatable wall has at least one port for at least partial inflation and deflation of the inflatable wall. The at least partially inflated inflatable wall and base define a volume that envelopes the machine. Also disclosed is a system and method for climatic test conditioning a machine.
A fuel system for a gas turbine engine is disclosed. The fuel system is configured to combust hydrogen fuel and comprises a fuel conduit, a first fuel pump configured to operate on liquid hydrogen within the fuel conduit and a second fuel pump downstream in hydrogen fuel flow of the first fuel pump. The first fuel pump comprises one of a side channel and a regenerative pump and the second fuel pump comprises a centrifugal pump. Methods of operation, a gas turbine engine comprising the fuel system, and an aircraft comprising the gas turbine engine are also disclosed.
F02C 3/22 - Gas-turbine plants characterised by the use of combustion products as the working fluid using a special fuel, oxidant, or dilution fluid to generate the combustion products the fuel or oxidant being gaseous at standard temperature and pressure
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 gas turbine engine for aircraft includes a combustor with a combustion chamber and fuel spray nozzles to inject fuel into the combustion chamber. The fuel spray nozzles include a first subset and a second subset of nozzles. Each of the first subset is supplied with fuel at a greater rate than each of the second subset. A ratio of the first subset to the second subset is 1:2 to 1:5. A MTO nvPM emissions index ratio is
A gas turbine engine for aircraft includes a combustor with a combustion chamber and fuel spray nozzles to inject fuel into the combustion chamber. The fuel spray nozzles include a first subset and a second subset of nozzles. Each of the first subset is supplied with fuel at a greater rate than each of the second subset. A ratio of the first subset to the second subset is 1:2 to 1:5. A MTO nvPM emissions index ratio is
EI
maxTO
,
SAF
EI
maxTO
,
FF
.
A gas turbine engine for aircraft includes a combustor with a combustion chamber and fuel spray nozzles to inject fuel into the combustion chamber. The fuel spray nozzles include a first subset and a second subset of nozzles. Each of the first subset is supplied with fuel at a greater rate than each of the second subset. A ratio of the first subset to the second subset is 1:2 to 1:5. A MTO nvPM emissions index ratio is
EI
maxTO
,
SAF
EI
maxTO
,
FF
.
EImaxTO,SAF is nvPM emissions index in mg/kg of the engine when operating at around 100% available thrust if fuel provided to the fuel spray nozzles includes sustainable aviation fuel. EImaxTO,FF is nvPM emissions index in mg/kg of the engine when operating at around 100% available thrust if fuel provided to the fuel spray nozzles is fossil-based hydrocarbon fuel. The MTO nvPM emissions index ratio is less than 1.
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.
A method for scanning a plurality of components includes providing an imaging beam source, an imaging beam receiver, and a support platform. The support platform is configured to rotate and/or revolve relative to the imaging beam source and the imaging beam receiver about one or more axes. The method further includes disposing the plurality of components on the support platform. The plurality of components is positioned in a gap between a first geometrical figure and at least one second geometrical figure in a configuration that reduces the variation in material thickness penetrated at different relative angles of rotation and/or revolution. The first geometrical figure and the at least one second geometrical figure are concentric.
G01N 23/046 - Investigating or analysing materials by the use of wave or particle radiation, e.g. X-rays or neutrons, not covered by groups , or by transmitting the radiation through the material and forming images of the material using tomography, e.g. computed tomography [CT]
G01N 23/083 - Investigating or analysing materials by the use of wave or particle radiation, e.g. X-rays or neutrons, not covered by groups , or by transmitting the radiation through the material and measuring the absorption the radiation being X-rays
A fuel system for a hydrogen fuelled gas turbine engine. The fuel system having a main hydrogen fuel line configured to provide hydrogen fuel to a combustor of the gas turbine engine and a closed loop heat exchange fluid line having a heat exchange fluid. The fuel system further having a first heat exchanger configured to exchange heat between gas turbine engine compressor bleed air and heat exchange fluid in the heat exchange fluid line, a second heat exchanger downstream of the first heat exchanger 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 downstream in main hydrogen fuel flow of the second heat exchanger 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. The fuel system having a main hydrogen fuel line configured to provide hydrogen fuel to a combustor of the gas turbine engine and a closed loop heat exchange fluid line having a heat exchange fluid. The fuel system further having a first heat exchanger configured to exchange heat between gas turbine engine compressor bleed air and heat exchange fluid in the heat exchange fluid line, a second heat exchanger downstream of the first heat exchanger 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 downstream in main hydrogen fuel flow of the second heat exchanger in main hydrogen fuel flow configured to exchange heat between a further engine fluid and heated hydrogen fuel.
An engine for an aircraft includes an engine core having a turbine, a compressor, and a core shaft connecting the turbine to the compressor; a fan located upstream of the engine core, the fan having a plurality of fan blades; and a gearbox. The gearbox is an epicyclic gearbox and comprises a sun gear, a plurality of planet gears, a ring gear, and a planet carrier on which the planet gears are mounted. The radial bending stiffness of the planet carrier is equal to or greater than 1.20×109 N/m, and/or the tilt stiffness of the planet carrier is greater than or equal to 6.00×108 Nm/rad. A method of operation of such an engine is also disclosed.
A method of operating a gas turbine engine for an aircraft is disclosed. The aircraft includes a turbine, a compressor, a combustor, and a core shaft connecting the turbine to the compressor; a fan upstream of the engine core; a gearbox that receives an input from the core shaft and outputs drive to the fan; an oil loop system arranged to supply oil to the gearbox; a heat exchange system having an air-oil heat exchanger; a fuel-oil heat exchanger through which the oil and the fuel flow; and a valve arranged to allow a proportion of the oil sent via at least one of the heat exchangers to be varied. The method includes controlling the valve such that, under idle conditions, an oil flow ratio of:
A method of operating a gas turbine engine for an aircraft is disclosed. The aircraft includes a turbine, a compressor, a combustor, and a core shaft connecting the turbine to the compressor; a fan upstream of the engine core; a gearbox that receives an input from the core shaft and outputs drive to the fan; an oil loop system arranged to supply oil to the gearbox; a heat exchange system having an air-oil heat exchanger; a fuel-oil heat exchanger through which the oil and the fuel flow; and a valve arranged to allow a proportion of the oil sent via at least one of the heat exchangers to be varied. The method includes controlling the valve such that, under idle conditions, an oil flow ratio of:
rate
of
oil
flow
into
air
-
oil
heat
exchanger
r
ate
of
oil
flow
into
fuel
-
oil
heat
exchanger
A method of operating a gas turbine engine for an aircraft is disclosed. The aircraft includes a turbine, a compressor, a combustor, and a core shaft connecting the turbine to the compressor; a fan upstream of the engine core; a gearbox that receives an input from the core shaft and outputs drive to the fan; an oil loop system arranged to supply oil to the gearbox; a heat exchange system having an air-oil heat exchanger; a fuel-oil heat exchanger through which the oil and the fuel flow; and a valve arranged to allow a proportion of the oil sent via at least one of the heat exchangers to be varied. The method includes controlling the valve such that, under idle conditions, an oil flow ratio of:
rate
of
oil
flow
into
air
-
oil
heat
exchanger
r
ate
of
oil
flow
into
fuel
-
oil
heat
exchanger
is in the range from 0.62 to 5.29.
Gearboxes for aircraft gas turbine engines, in particular to arrangements for journal bearings such gearboxes, and to related methods of operating such gearboxes and gas turbine engines. Example embodiments include a gearbox for an aircraft gas turbine engine, the gearbox including: a sun gear; a plurality of planet gears surrounding and engaged with the sun gear; and a ring gear surrounding and engaged with the plurality of planet gears, each of the plurality of planet gears being rotatably mounted around a bearing.
Gearboxes for aircraft gas turbine engines, in particular to arrangements for journal bearings such gearboxes, and to related methods of operating such gearboxes and gas turbine engines. Example embodiments include a gearbox for an aircraft gas turbine engine, the gearbox including: a sun gear; a plurality of planet gears surrounding and engaged with the sun gear; and a ring gear surrounding and engaged with the plurality of planet gears, each of the plurality of planet gears being rotatably mounted around a bearing.
Gearboxes for aircraft gas turbine engines, in particular to arrangements for journal bearings such gearboxes, and to related methods of operating such gearboxes and gas turbine engines. Example embodiments include a gearbox for an aircraft gas turbine engine, the gearbox including: a sun gear; a plurality of planet gears surrounding and engaged with the sun gear; and a ring gear surrounding and engaged with the plurality of planet gears, each of the plurality of planet gears being rotatably mounted around a bearing.
F01D 25/16 - Arrangement of bearingsSupporting or mounting bearings in casings
F02C 3/113 - Gas-turbine plants characterised by the use of combustion products as the working fluid having a turbine driving a compressor with two or more rotors connected by power transmission with variable power transmission between rotors
F02C 7/36 - Power transmission between the different shafts of the gas-turbine plant, or between the gas-turbine plant and the power user
F02K 3/06 - Plants including a gas turbine driving a compressor or a ducted fan in which part of the working fluid by-passes the turbine and combustion chamber the plant including ducted fans, i.e. fans with high volume, low-pressure outputs, for augmenting jet thrust, e.g. of double-flow type with front fan
F16H 57/04 - Features relating to lubrication or cooling
F16H 57/08 - General details of gearing of gearings with members having orbital motion
A method of operating a gas turbine engine. The gas turbine engine includes an accessory gearbox arranged to power one or more engine accessories, and a combustor in which fuel is combusted and a fuel-oil heat exchanger arranged to transfer heat between oil and fuel that is provided to the combustor. An average temperature of lubricating oil is at least 180° C. on exit from the accessory gearbox at cruise conditions. The method includes transferring heat from the oil to the fuel before the fuel enters the combustor so as to raise the fuel temperature to an average of at least 135° C. on entry to the combustor at cruise conditions. Also provided is a gas turbine engine.
An aerofoil component for a high-pressure or intermediate-pressure turbine of a multi-spool gas turbine engine having an aerofoil member which 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 to a trailing edge 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 defines a midspan axial chord length. The aerofoil member are arranged 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 defines a midspan pitch of the aerofoil member. At midspan: the value of S/Cx is in the range from 1.4 to 1.6; and the uncovered turning angle is 16° or more.
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.
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 promote 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).
A geared gas turbine engine comprises a heat management system configured to provide lubrication and cooling to a power gearbox and turbomachinery bearings, and comprising a pipe assembly adapted to provide a lubricant flow to the power gearbox and turbomachinery bearings to remove the heat generated by the power gearbox and turbomachinery bearings, an air-lubricant heat exchanger to dissipate a first amount of heat, and a fuel-lubricant heat exchanger to dissipate a second amount of heat wherein the heat management system is configured to provide the first amount of heat and the second amount of heat such that at cruise conditions a proportion of heat generated by the gearbox and the turbomachinery and dissipated to air is in the range of from 0.35 to 0.80.
The foregoing describes a power converter apparatus with at least a first power converter and a second power converter connected in parallel, so as to each receive an AC voltage generated by an electrical machine and output a DC voltage on a DC bus. Each of first and second power converters is coupled to respective first and second winding arrangements for receiving the AC voltages, and the first winding arrangement is grounded by a first impedance and the second winding arrangement is grounded by a second impedance. The first impedance has an effective impedance value which is different to that of the second impedance such, in use, a zero-sequence current between the first and second power converters at least partly cancels a common-mode voltage component.
H02M 7/23 - Conversion of AC power input into DC power output without possibility of reversal by static converters using discharge tubes with control electrode or semiconductor devices with control electrode using devices of a triode or transistor type requiring continuous application of a control signal using semiconductor devices only arranged for operation in parallel
B64D 33/00 - Arrangement in aircraft of power plant parts or auxiliaries not otherwise provided for
H02M 1/12 - Arrangements for reducing harmonics from AC input or output
H02M 7/217 - Conversion of AC power input into DC power output without possibility of reversal by static converters using discharge tubes with control electrode or semiconductor devices with control electrode using devices of a triode or transistor type requiring continuous application of a control signal using semiconductor devices only
A gas turbine engine including a rear attachment comprising: a plurality of angularly distributed rotational pylon connections for coupling to a wing pylon and defining a pylon connection angular extent; a plurality of angularly distributed engine connections for coupling to a rear mount structure of the gas turbine engine 10 and defining an engine connection angular extent; and a link structure extending between the pylon connections and the engine connections; wherein the link structure, pylon connections and engine connections are configured to transfer a roll torque from the rear mount structure to the wing pylon; 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 comprising a fluid duct defined by a wall and a moveable member. The propulsion machine also comprises a mounting structure coupled to the moveable member, and an extendable structure having a sealing surface and a sensing arrangement. The extendable structure is moveable relative to the mounting structure to provide a seal between the sealing surface and an opposing surface of the wall. The sensing arrangement is configured to generate one or more signals indicative of a position of the extendable structure relative to the mounting structure.
F02K 1/12 - Varying effective area of jet pipe or nozzle by means of pivoted flaps
G01D 5/20 - Mechanical means for transferring the output of a sensing memberMeans for converting the output of a sensing member to another variable where the form or nature of the sensing member does not constrain the means for convertingTransducers not specially adapted for a specific variable using electric or magnetic means influencing the magnitude of a current or voltage by varying inductance, e.g. by a movable armature
There is provided a flameholder for a reheat assembly of a gas turbine engine. The flameholder comprises a fuel atomizer, an air passageway and a fuel passageway. The air passageway is configured to convey a flow of air to the fuel atomizer. The fuel passageway is configured to convey a flow of fuel to the fuel atomizer.
There is provided a reheat assembly for a gas turbine engine. The reheat assembly comprises a jetpipe casing, a support duct, a plurality of fuel discharge ports, and a fuel distribution passageway. The jetpipe casing comprises a reheat core section and a reheat bypass section. The support duct radially separates the reheat core section and the reheat bypass section. Each fuel discharge port is configured to discharge fuel received from the fuel distribution passageway. The fuel distribution passageway is defined by an interior channel embedded within the support duct.
A fuel injector for a combustor of a gas turbine engine is described. The fuel injector includes a fuel injection device having a fuel injection port configured to inject fuel into the combustor, wherein the fuel injection port includes an elongate aperture. A combustor for a gas turbine engine, a reheat assembly for a gas turbine engine, a gas turbine engine, and an aircraft.
A hyper-redundant manipulator spray system for depositing an engineering coating material on a component located within a gas turbine engine. The system has a hyper-redundant manipulator arm, an end effector, an actuator pack and a tank for holding the engineering coating material to be deposited on the component. The end effector has an atomizing nozzle unit that atomizes a suspension of the engineering coating material and directs it into a flame of a combusted flammable fluid from a burner head that sprays the atomized engineering coating material on to the component.
B25J 19/00 - Accessories fitted to manipulators, e.g. for monitoring, for viewingSafety devices combined with or specially adapted for use in connection with manipulators
The foregoing describes a reheat assembly for a gas turbine engine. The reheat assembly comprises: a support duct section comprising a core side for facing a core reheat region, and a bypass side for facing a bypass reheat region; the support duct comprising a plurality of circumferentially spaced inlets, each inlet being configured to communicate the bypass reheat region with the core reheat region; a fuel supply system comprising a plurality of bypass fuel injection ports, each bypass fuel injection port being associated with an inlet of the plurality of inlets; wherein each bypass fuel injection port is configured to discharge fuel into the bypass reheat region for transit through the associated inlet into the core reheat region.
A flameholder for a reheat assembly of a gas turbine engine. The flameholder includes a main structure and a baffle structure. The baffle structure is suspended within the main structure at one or more mounting locations. The baffle structure is coextensive with the main structure along a coextensive extent. The baffle structure has a detached portion spaced apart from the one or more mounting locations.
F23R 3/18 - Flame stabilising means, e.g. flame holders for after-burners of jet-propulsion plants
F02K 3/10 - Plants including a gas turbine driving a compressor or a ducted fan with supplementary heating of the working fluidControl thereof by after-burners
A method of performing spectral analysis on a sample identified within a gas turbine engine. The method involves: obtaining an optical spectrum of the sample; normalising the optical spectrum to obtain a normalised optical spectrum; selecting a characteristic parameter that characterises the normalised optical spectrum; determining a characteristic parameter value associated with the normalised optical spectrum based on the characteristic parameter; comparing the characteristic parameter value with one or more pre-determined reference values corresponding to the characteristic parameter; and determining a presence of at least one compound in the sample based on the comparison between the characteristic parameter value and the one or more pre-determined reference values. The method can determine whether there is a need to disassemble the gas turbine engine for maintenance.
G01N 21/31 - Investigating relative effect of material at wavelengths characteristic of specific elements or molecules, e.g. atomic absorption spectrometry
G01M 15/14 - Testing gas-turbine engines or jet-propulsion engines
A method for scanning a component having one or more regions of interest includes disposing at least one element adjacent to and at least partially engaging the one or more regions of interest of the component. The component has an aspect ratio that is higher than an aspect ratio of the at least one element. The method further includes providing an imaging beam source and an imaging beam receiver; generating, via the imaging beam source, an imaging beam that passes through the component and the at least one element; receiving the imaging beam at the imaging beam receiver; and generating, via the imaging beam receiver, an image.
G01N 23/04 - Investigating or analysing materials by the use of wave or particle radiation, e.g. X-rays or neutrons, not covered by groups , or by transmitting the radiation through the material and forming images of the material
G01N 23/083 - Investigating or analysing materials by the use of wave or particle radiation, e.g. X-rays or neutrons, not covered by groups , or by transmitting the radiation through the material and measuring the absorption the radiation being X-rays
A titanium alloy comprising Al 3.00 to 5.50 wt. %; V 3.50 to 5.00 wt. %; Mo 1.50 to 3.00 wt. %; Cr 1.00 to 2.50 wt. %; Fe 0.00 to 0.50 wt. %; Sn 0.00 to 1.25 wt. %; Zr 0.00 to 2.00 wt. %; C 0.01 wt. % to 0.50 wt. %; Si 0.00 wt. % to 0.50 wt. %; N up to 500 ppm; O up to 2500 ppm and H up to 150 ppm; the balance being Ti and incidental elements and unavoidable impurities. Such a titanium alloy is useful for manufacturing gas turbine engine components including compressor disks, bladed disks, or casings.
B22D 21/00 - Casting non-ferrous metals or metallic compounds so far as their metallurgical properties are of importance for the casting procedureSelection of compositions therefor
B22F 3/105 - Sintering only by using electric current, laser radiation or plasma
B22F 10/66 - Treatment of workpieces or articles after build-up by mechanical means
B33Y 40/20 - Post-treatment, e.g. curing, coating or polishing
B33Y 80/00 - Products made by additive manufacturing
C22C 1/04 - Making non-ferrous alloys by powder metallurgy
The disclosure relates to an 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. The apparatus comprising; a back panel, an aperture configured to receive a fuel injector and a plurality of air inlets arranged around the aperture; a heatshield panel having front and back faces, the heatshield panel having a central aperture and an annular side wall mounted against the back panel to form a cavity between the heatshield panel and the back panel; and an annular swirler assembly disposed in the cavity between the back panel and the heatshield panel, the annular swirler assembly having a central passage for flow of an air and fuel mixture into the combustion chamber and a flange engaged with grooves in the heatshield panel and the back panel, respectively.
F23R 3/14 - Air inlet arrangements for primary air inducing a vortex by using swirl vanes
F02C 3/22 - Gas-turbine plants characterised by the use of combustion products as the working fluid using a special fuel, oxidant, or dilution fluid to generate the combustion products the fuel or oxidant being gaseous at standard temperature and pressure
A titanium alloy having 1.50 to 7.00 wt. % aluminium, 3.00 to 5.00 wt. % vanadium, 1.00 to 3.00 wt. % molybdenum; 0.50 to 2.50 wt. % zirconium, 0.05 to 0.40 wt. % oxygen, 0.05 to 2.00 wt. % tin, 0.00 to 1.00 wt. % iron, 0.00 to 0.3 wt % silicon, 0.01 to 0.15 wt. % carbon, and 0.001 to 0.05 wt. % nitrogen; the balance being titanium and incidental elements and unavoidable impurities.
A seal arrangement providing a seal between three or more relatively moveable panels having gaps therebetween. A seal element comprises a nexus and plurality of legs each of the legs radiating from the nexus along a respective gap. Each leg comprises a cover spaced from the panels and straddling a gap and a first flange connected to a first of the panels separated by the respective gap and a second flange connected to a second of the panels separated by the respective gap.
An apparatus for measuring a clearance between a first component and one or more areas of interest of a second component. The apparatus includes an imaging beam source configured to generate an imaging beam that passes between the first component and the second component. The apparatus includes a first tube configured to guide the imaging beam from the imaging beam source to the one or more areas of interest of the second component. The apparatus includes an imaging beam receiver configured to receive the imaging beam. The imaging beam receiver is configured to generate an image in response to receiving the imaging beam. The image depicts the clearance between the first component and the one or more areas of interest of the second component.
G01B 15/00 - Measuring arrangements characterised by the use of electromagnetic waves or particle radiation, e.g. by the use of microwaves, X-rays, gamma rays or electrons
G01M 15/14 - Testing gas-turbine engines or jet-propulsion engines
G01N 23/041 - Phase-contrast imaging, e.g. using grating interferometers
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 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). There is also provided a method (600) of manufacturing such a fluid flow machine (10).
F01D 11/12 - Preventing or minimising internal leakage of working fluid, e.g. between stages for sealing space between rotor blade tips and stator using a rubstrip, e.g. erodible, deformable or resiliently biased part
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 for in-situ application of an engineering coating to one or more components. The apparatus has a head section adapted to selectively apply the engineering coating to the components. The head section has an applicator end and a coupling end opposite to the applicator end. The apparatus has a body section coupled to the head section at the coupling end and adapted to actuate the head section. The body section has a first stage actuator adapted to actuate the head section to swivel about a first axis and a second stage actuator adapted to actuate the head section and the first stage actuator together to rotate about a second axis. The second axis is orthogonal to the first axis. The apparatus has one or more gas channels provided along the body section to the head section.
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).
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 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). There is also provided a method (600) of manufacturing such a fluid flow machine (10).
F01D 11/12 - Preventing or minimising internal leakage of working fluid, e.g. between stages for sealing space between rotor blade tips and stator using a rubstrip, e.g. erodible, deformable or resiliently biased part
B33Y 80/00 - Products made by additive manufacturing
43.
APPARATUS, SYSTEM, AND METHOD FOR IN-SITU APPLICATION OF AN ENGINEERING COATING
An apparatus for in-situ application of an engineering coating to components has a head section for applying the engineering coating and a multi-part body section for controlling the application process. The body section has a first body member, a second body member, and a third body member interconnected with each other. The first body member has a pre-determined pathway that guides movement and is connected to the head section. The second body member is engaged with the first body member and has at least one hollow channel. The third body member houses an actuator, an engagement pin for connecting with the pathway, and at least one coupling member for interacting with the at least one hollow channel. The actuator ensures precise movement of the head section. The apparatus has one or more gas channels to provide necessary gases to the head section.
B05B 13/02 - Means for supporting workArrangement or mounting of spray headsAdaptation or arrangement of means for feeding work
B05B 13/04 - Means for supporting workArrangement or mounting of spray headsAdaptation or arrangement of means for feeding work the spray heads being moved during operation
44.
METHOD AND SCANNING APPARATUS FOR SCANNING A COMPONENT
A method for scanning a component includes providing a filter and the component, the filter at least partially surrounding the component. The filter comprises a multi-phase material that has an attenuation coefficient that is lower than an attenuation coefficient of a material of the component to be scanned. The method further includes disposing the component and the filter on a support platform, and providing an imaging beam source and an imaging beam receiver. The support platform is configured to rotate and/or revolve relative to the imaging beam source and the imaging beam receiver about one or more axes. The method includes generating, via the imaging beam source, an imaging beam that passes through the component and the filter. The method includes attenuating, via the, a scatter beam that is produced upon irradiation of the component with the imaging beam.
G01N 23/04 - Investigating or analysing materials by the use of wave or particle radiation, e.g. X-rays or neutrons, not covered by groups , or by transmitting the radiation through the material and forming images of the material
G01N 23/083 - Investigating or analysing materials by the use of wave or particle radiation, e.g. X-rays or neutrons, not covered by groups , or by transmitting the radiation through the material and measuring the absorption the radiation being X-rays
Gas turbine engine includes: an engine core including a turbine, compressor, and core shaft connecting the turbine to the compressor; a fan located upstream of the core; and a gearbox that is configured to receive an input from the core shaft, and output drive to a fan shaft via an output of the gearbox so as to drive the fan at a lower rotational speed than the core shaft. A fan shaft radial bending stiffness to moment of inertia ratio defined as:
Gas turbine engine includes: an engine core including a turbine, compressor, and core shaft connecting the turbine to the compressor; a fan located upstream of the core; and a gearbox that is configured to receive an input from the core shaft, and output drive to a fan shaft via an output of the gearbox so as to drive the fan at a lower rotational speed than the core shaft. A fan shaft radial bending stiffness to moment of inertia ratio defined as:
a
radial
bending
stiffness
of
the
fan
shaft
at
the
output
of
the
gearbox
the
moment
of
inertia
of
the
fan
Gas turbine engine includes: an engine core including a turbine, compressor, and core shaft connecting the turbine to the compressor; a fan located upstream of the core; and a gearbox that is configured to receive an input from the core shaft, and output drive to a fan shaft via an output of the gearbox so as to drive the fan at a lower rotational speed than the core shaft. A fan shaft radial bending stiffness to moment of inertia ratio defined as:
a
radial
bending
stiffness
of
the
fan
shaft
at
the
output
of
the
gearbox
the
moment
of
inertia
of
the
fan
is in a range from 2.5×10−2 Nkg−1m−1mm−2 to 6.0 Nkg−1m−1 mm−2.
F02C 7/36 - Power transmission between the different shafts of the gas-turbine plant, or between the gas-turbine plant and the power user
F02C 3/107 - Gas-turbine plants characterised by the use of combustion products as the working fluid having a turbine driving a compressor with two or more rotors connected by power transmission
F02K 3/06 - Plants including a gas turbine driving a compressor or a ducted fan in which part of the working fluid by-passes the turbine and combustion chamber the plant including ducted fans, i.e. fans with high volume, low-pressure outputs, for augmenting jet thrust, e.g. of double-flow type with front fan
A device for wear testing includes a frame, a rotary actuator, and a linear actuator spaced apart from the rotary actuator along a longitudinal axis of the frame. The device further includes first and second specimen holding units. The first specimen holding unit includes a seat connected to the linear actuator, a plate engaging with the seat, and a first thrust bearing connected to the plate and at least one first diaphragm plate connected to the frame. The second specimen holding unit includes a shaft connected to the rotary actuator and a second thrust bearing connected the frame. The rotary actuator is configured to rotate a second specimen about the longitudinal axis. The linear actuator is configured to move a first specimen along the longitudinal axis, such that the first specimen engages with the second specimen.
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.
A method for facilitating maintenance management of a propulsion system, such as an engine, for a vehicle is disclosed. The method comprises obtaining, for each of a plurality of propulsion systems, records of maintenance events experienced by the propulsion system, and records of traversals of transport routes by the propulsion system during a period of propulsion system operation. The method further comprises using a Machine Learning model to classify the recorded maintenance events into a plurality of maintenance categories. The method then comprises identifying, from the classified recorded maintenance events and the records of traversals of transport routes, a correlation between a given maintenance category and the transport routes traversed by propulsion systems during operational periods preceding maintenance events classified into the maintenance category. The correlation may be used in maintenance and/or route planning for propulsion systems.
A gas turbine engine has a compression system blade ratio defined as the ratio of the height of a fan blade to the height of the most downstream compressor blade in the range of from 45 to 95. This results in an optimum balance between installation benefits, operability, maintenance requirements and engine efficiency when the gas turbine engine is installed on an aircraft.
F02C 3/113 - Gas-turbine plants characterised by the use of combustion products as the working fluid having a turbine driving a compressor with two or more rotors connected by power transmission with variable power transmission between rotors
A fuel system (200) for a hydrogen-fuelled gas turbine engine (201) comprises a main fuel conduit (217) configured to conduct hydrogen fuel from a hydrogen storage unit (204) to a core combustor (206) of the gas turbine engine (201), a pre-heater (218) comprising an auxiliary combustor (222) configured to combust hydrogen fuel diverted from the main fuel conduit (217) to heat hydrogen fuel in the main fuel conduit (217). A first exhaust passage (228) is configured to conduct combustion products from the auxiliary combustor (222) to the core combustor (206) of the gas turbine engine (201).
F02C 7/224 - Heating fuel before feeding to the burner
F02C 3/22 - Gas-turbine plants characterised by the use of combustion products as the working fluid using a special fuel, oxidant, or dilution fluid to generate the combustion products the fuel or oxidant being gaseous at standard temperature and pressure
F02C 9/32 - Control of fuel supply characterised by throttling of fuel
A spray nozzle for thermal spraying a coating material on a substrate includes a central port configured to deliver the coating material along a spray axis in the form of an atomized suspension. The spray nozzle further includes a plurality of gas ports arranged symmetrically and peripherally around the central port with respect to the spray axis. Each gas port is configured to deliver a pressurized combustible gas for entraining the coating material after the coating material has exited the central port. The pressurized combustible gas is configured to heat the coating material in a heating zone for generating a heated gas stream. The heated gas stream is directed toward a target surface of the substrate.
B05B 7/20 - Spraying apparatus for discharge of liquids or other fluent materials from two or more sources, e.g. of liquid and air, of powder and gas incorporating means for heating the material to be sprayed by flame or combustion
B05B 13/00 - Machines or plants for applying liquids or other fluent materials to surfaces of objects or other work by spraying, not covered by groups
A fluid flow machine including a first rotor and a stator alongside the first rotor, each of the first rotor and the stator including circumferentially distributed turbomachine blades, wherein the stator includes a stator shroud structure including a stator shroud surface and the first rotor includes a first rotor shroud structure including a first rotor shroud surface, the stator shroud surface and first rotor shroud surface together at least partially defining a radially inner flow surface of the fluid flow machine, wherein the first rotor shroud structure defines a first cavity, the first cavity being disposed radially inward of the first rotor shroud surface, and wherein the stator shroud structure includes a first protrusion, the first protrusion being disposed radially inward of the radially inner flow surface, wherein the first protrusion extends in a direction with an axial component and into the first cavity of the first rotor shroud structure.
There is provided a busbar (300, 301, 302) comprising a plurality of tabs (312, 314, 316), a fuse link (323, 325) and a cartridge (333, 335). The plurality of tabs (312, 314, 316) are offset from one another along a separation direction (502), with each tab (312, 314, 316) being configured to electrically couple with at least one energy storage device (400). The fuse link (323, 325) extends between two adjacent tabs (312, 314, 316) of the plurality of tabs (312, 314, 316), with the fuse link (323, 325) being configured to form a gap (355) when a fault current flows along the fuse link (323, 325). The cartridge (333, 335) surrounds the fuse link (323, 325). The cartridge (333, 335) is configured to inhibit electrical arcing originating from the gap (355).
H01M 50/588 - Means for preventing undesired use or discharge for preventing incorrect connections inside or outside the batteries outside the batteries, e.g. incorrect connections of terminals or busbars
H01M 50/503 - Interconnectors for connecting terminals of adjacent batteriesInterconnectors for connecting cells outside a battery casing characterised by the shape of the interconnectors
H01M 50/507 - Interconnectors for connecting terminals of adjacent batteriesInterconnectors for connecting cells outside a battery casing comprising an arrangement of two or more busbars within a container structure, e.g. busbar modules
H01M 50/583 - Devices or arrangements for the interruption of current in response to current, e.g. fuses
A system for monitoring and optimizing a repair operation being performed on a component includes at least one light sensor disposed proximal to the component to collect heat signature data emitted from the component. The at least one light sensor is disposed in an oblique orientation relative to the component. The system further includes a controller communicably coupled with the at least one light sensor. The controller is configured to receive the heat signature data from the at least one light sensor, compare the heat signature data with a nominal range for the heat signature, and generate an output signal if the heat signature data is outside of the nominal range for the heat signature.
B23K 31/12 - Processes relevant to this subclass, specially adapted for particular articles or purposes, but not covered by any single one of main groups relating to investigating the properties, e.g. the weldability, of materials
B23K 31/02 - Processes relevant to this subclass, specially adapted for particular articles or purposes, but not covered by any single one of main groups relating to soldering or welding
55.
HIGH POWER EPICYCLIC GEARBOX AND OPERATION THEREOF
A gas turbine engine for an aircraft includes an engine core with a turbine, a compressor, and a core shaft connecting the turbine and compressor; a fan upstream of the engine core including a plurality of fan blades; and a gearbox that receives an input from a gearbox input shaft portion of the core shaft and outputs drive to a fan shaft so as to drive the fan at a lower rotational speed than the core shaft, the gearbox being an epicyclic gearbox including a sun gear, a plurality of planet gears, a ring gear, and a planet carrier arranged to have the plurality of planet gears mounted thereon, and wherein the sun gear receives input from the core shaft. At cruise conditions the torque on the core shaft is greater than 10,000 Nm and a ratio of core shaft stiffness to core shaft torque is within a specified range.
F16H 1/28 - Toothed gearings for conveying rotary motion with gears having orbital motion
F16H 1/32 - Toothed gearings for conveying rotary motion with gears having orbital motion in which the central axis of the gearing lies inside the periphery of an orbital gear
A blower compressor assembly comprises: a flow modifier comprising an array of nozzle guide vanes and an array of diffuser vanes, wherein the flow modifier is movable relative to a rotor between a turbine configuration and a compressor configuration; and an actuator assembly for moving the flow modifier relative to the rotor between the turbine configuration and the compressor configuration, wherein the actuator assembly comprises: a first actuator comprising a first chamber, a first valve configured to selectively supply pressurised air to the first chamber or vent pressurised air from the first chamber, and a piston arranged to one side of the first chamber and coupled to the flow modifier such that the flow modifier is moved into the turbine configuration or the compressor configuration based on the pressure of air within the first chamber.
A method of hot forming a pair of hollow components. The method involves heating a die stack comprising a first die, an intermediate die and a second die, in which the first die and the intermediate die form a first die set, and the intermediate die and the second die form a second die set. The die stack is opened and loaded with a first component preform in the first space and a second component preform in the second space. Once loaded, the die stack is closed and a pressurised fluid is then provided via a first conduit to an internal cavity of the first component preform, while a pressurised fluid is also provided via a second conduit to an internal cavity of the second component preform. The die stack is opened and the first component and the second component removed from the respective first space and the second space.
The disclosure relates to monitoring of power electronics modules for multiphase electrical machines, in particular to determine the presence or absence of a fault. Example embodiments include a method of monitoring a system (100) comprising first and second multiphase electrical machines (1011, 1012) driven by respective first and second inverters (1021, 1022) from respective first and second DC electrical supplies (1031, 1032), the method comprising measuring DC link currents and voltages (Idc1, Idc2, Vdc1, Vdc2), stator currents and voltages (Ia1, Ib1, Ic1, Ia2, Ib2, Ic2, Va1, Vb1, Vc1, Va2, Vb2, Vb3), calculating on-state voltage, on-state resistance and turn-off delay for the inverters (1021, 1022) and an equivalent series resistance of the inverters (1021, 1022), calculating ratios of the calculated values and determining the presence or absence of a fault in the inverters (1021, 1022) based on a combination of two or more of the ratios.
A gas turbine engine has a compression system radius ratio defined as the ratio of the radius of the tip of a fan blade to the radius of the tip of the most downstream compressor blade in the range of from 5 to 9. This results in an optimum balance between installation benefits, operability, maintenance requirements and engine efficiency when the gas turbine engine is installed on an aircraft.
B64D 27/10 - Aircraft characterised by the type or position of power plants of gas-turbine type
B64D 33/02 - Arrangement in aircraft of power plant parts or auxiliaries not otherwise provided for of combustion air intakes
F01D 5/28 - Selecting particular materialsMeasures against erosion or corrosion
F02C 3/04 - Gas-turbine plants characterised by the use of combustion products as the working fluid having a turbine driving a compressor
F02C 3/06 - Gas-turbine plants characterised by the use of combustion products as the working fluid having a turbine driving a compressor the compressor comprising only axial stages
F02C 3/107 - Gas-turbine plants characterised by the use of combustion products as the working fluid having a turbine driving a compressor with two or more rotors connected by power transmission
F02C 7/04 - Air intakes for gas-turbine plants or jet-propulsion plants
F02K 3/06 - Plants including a gas turbine driving a compressor or a ducted fan in which part of the working fluid by-passes the turbine and combustion chamber the plant including ducted fans, i.e. fans with high volume, low-pressure outputs, for augmenting jet thrust, e.g. of double-flow type with front fan
F02K 3/068 - Plants including a gas turbine driving a compressor or a ducted fan in which part of the working fluid by-passes the turbine and combustion chamber the plant including ducted fans, i.e. fans with high volume, low-pressure outputs, for augmenting jet thrust, e.g. of double-flow type being characterised by a short axial length relative to diameter
A fuel system for a hydrogen-fueled gas turbine engine includes a main fuel conduit configured to conduct hydrogen fuel from a hydrogen storage unit to a core combustor of the gas turbine engine, a pre-heater including an auxiliary combustor configured to combust hydrogen fuel diverted from the main fuel conduit to heat hydrogen fuel in the main fuel conduit. A first exhaust passage is configured to conduct combustion products from the auxiliary combustor to the core combustor of the gas turbine engine.
An electrical machine having a monitoring apparatus the monitoring apparatus comprising at least one first fibre optic cable having a periodic structure within, at least, a section of the first fibre optic cable, the first fibre optic cable being coupled to an outcoupling optical set up comprising at least one first lens, and an interrogator connected to a second fibre optic cable and at least one second lens, the light transmitted through the first fibre optic cable is manipulated by the first lens and transmitted to the second lens which is coupled to the second fibre optical cable which transfers a light signal to the interrogator, and wherein a portion of the first fibre optical cable is connected to a component within the electrical machine.
A connector for connecting a sensor to a data recorder, the connector comprising: circuitry configured to receive a signal from the sensor and to transmit a digital signal to the data recorder; and a memory storing one or more of: sensor configuration data or sensor calibration data.
G01D 11/30 - Supports specially adapted for an instrumentSupports specially adapted for a set of instruments
G01D 18/00 - Testing or calibrating apparatus or arrangements provided for in groups
H01R 13/514 - BasesCases formed as a modular block or assembly, i.e. composed of co-operating parts provided with contact members or holding contact members between them
A data recorder comprising: at least one processor; at least one memory comprising computer readable instructions; the at least one processor being configured to read the computer readable instructions to cause performance of: receiving one or more of: sensor configuration data; or sensor calibration data from a memory of a connector connected between the data recorder and a sensor; reading the connector to receive a digital signal; processing the received digital signal; and controlling storage of data in the processed digital signal in the at least one memory of the data recorder.
G07C 5/08 - Registering or indicating performance data other than driving, working, idle, or waiting time, with or without registering driving, working, idle, or waiting time
G01M 15/14 - Testing gas-turbine engines or jet-propulsion engines
G01R 31/66 - Testing of connections, e.g. of plugs or non-disconnectable joints
There is provided a gas turbine engine, the gas turbine engine comprising a compressor assembly, a combustor assembly having a first circumferential array of burner assemblies, a turbine assembly, and an exhaust assembly. The turbine assembly comprising a second circumferential array of high-pressure nozzle guide vanes, a high-pressure turbine assembly, a third circumferential array of outlet guide vanes, and a fourth circumferential array of temperature measurement rakes. Each temperature measurement rake comprises a plurality of temperature measurement sensors. A clocking position is defined by a relative circumferential orientation of the first circumferential array with the fourth circumferential array. A quantity of burner assemblies are positioned in the first circumferential array, and a quantity of temperature measurement rakes are positioned in the fourth circumferential array, such that there is a minimum quantity of five clocking positions between the fourth circumferential array, and the first circumferential array.
F23R 3/42 - Continuous combustion chambers using liquid or gaseous fuel characterised by the arrangement or form of the flame tubes or combustion chambers
F02C 7/00 - Features, component parts, details or accessories, not provided for in, or of interest apart from, groups Air intakes for jet-propulsion plants
A method of operating a gas turbine engine including an engine core including a turbine, compressor, combustor to combust a fuel, and core shaft connecting the turbine and compressor; a fan upstream of the engine core; a fan shaft; a gearbox that receives an input from the core shaft and outputs drive to the fan via the fan shaft; a primary oil loop system to supply oil to the gearbox; and a heat exchange system. The method includes controlling the heat exchange system to adjust fuel viscosity to be lower than or equal to 0.58 mm2/s on entry to the combustor at cruise conditions.
A gas turbine engine, gas turbine engine including compressor assembly, combustor assembly having a first circumferential array of burner assemblies, turbine assembly, and exhaust assembly. The turbine assembly including a second circumferential array of high-pressure nozzle guide vanes, high-pressure turbine assembly, third circumferential array of low-pressure nozzle guide vanes, low-pressure turbine assembly, and fourth circumferential array of outlet guide vanes. A clocking position is defined by relative circumferential orientation of first circumferential array with any one of second circumferential array, third circumferential array, and fourth circumferential array. A quantity of burner assemblies are positioned in first circumferential array, and quantity of guide vanes are positioned in at least one of the second circumferential array, third circumferential array, and fourth circumferential array, such that there is a minimum quantity of five clocking positions between at least one of second circumferential array, third circumferential array, fourth circumferential array, and first circumferential array.
A highly efficient gas turbine engine is a system wherein the fan of the gas turbine engine is driven from a turbine via a gearbox, such that the fan has a lower rotational speed than the driving turbine, thereby providing efficiency gains. The efficient fan system is mated to a core that has low cooling flow requirements and/or high temperature capability, and which may have particularly low mass for a given power.
F02K 3/06 - Plants including a gas turbine driving a compressor or a ducted fan in which part of the working fluid by-passes the turbine and combustion chamber the plant including ducted fans, i.e. fans with high volume, low-pressure outputs, for augmenting jet thrust, e.g. of double-flow type with front fan
68.
THERMAL BARRIER COATING FOR GAS TURBINE ENGINE COMPONENTS
A thermal barrier coating for a substrate that has a major surface. The thermal barrier coating comprises: a bond coat layer disposed on the major surface of the substrate; a thermally grown oxide (TGO) layer disposed on the bond coat layer; and a top coat layer disposed on the TGO layer. The top coat layer comprises a ceramic top coat material that comprises: a matrix phase comprising at least one compound of formula ABO4, AB3O7 or AB3O9, where A is a rare-earth element, B is selected from the group consisting of niobium and tantalum, and O is oxygen; and an inclusion phase that is dispersed in the matrix phase and comprising one or more of alumina and a compound of formula A3AlO12, AAlO3 or A4Al2O9, where A is a rare-earth element, Al is aluminium, and O is oxygen.
C23C 28/04 - Coating for obtaining at least two superposed coatings either by methods not provided for in a single one of main groups , or by combinations of methods provided for in subclasses and only coatings of inorganic non-metallic material
69.
SYSTEM AND METHOD FOR REPAIRING A DAMAGED BLADE OF A GAS TURBINE ENGINE
A system for repairing a damaged blade of a gas turbine engine includes a sacrificial build platform configured to be positioned and fused with the damaged blade proximal to at least one stub portion of the damaged blade. The at least one stub portion protrudes from a root of the damaged blade and defines a deposition surface. The sacrificial build platform includes a base defining a pair of cavities. The at least one stub portion includes a pair of projections, each cavity from the pair of cavities is configured to receive a corresponding projection from the pair of projections. The system further includes an additive manufacturing system that is configured to deposit material on the base of the sacrificial build platform and the deposition surface of the at least one stub portion to form a base structure of a repaired blade.
A gas turbine engine includes a first and a second subset of fuel spray nozzles. A combustor is operable with the first subset of fuel spray nozzles supplied with fuel at a greater fuel flow rate than each of the second subset of fuel spray nozzles. A ratio of the first to the second subset of fuel spray nozzles is 1:2 to 1:5. A MTO nvPM emissions index ratio is:
A gas turbine engine includes a first and a second subset of fuel spray nozzles. A combustor is operable with the first subset of fuel spray nozzles supplied with fuel at a greater fuel flow rate than each of the second subset of fuel spray nozzles. A ratio of the first to the second subset of fuel spray nozzles is 1:2 to 1:5. A MTO nvPM emissions index ratio is:
EI
maxTO
,
SAF
EI
maxTO
,
FF
.
A gas turbine engine includes a first and a second subset of fuel spray nozzles. A combustor is operable with the first subset of fuel spray nozzles supplied with fuel at a greater fuel flow rate than each of the second subset of fuel spray nozzles. A ratio of the first to the second subset of fuel spray nozzles is 1:2 to 1:5. A MTO nvPM emissions index ratio is:
EI
maxTO
,
SAF
EI
maxTO
,
FF
.
EImaxTO,SAF is the system loss corrected nvPM emissions index in mg/kg operating at around 100% available thrust if fuel provided to the fuel spray nozzles includes sustainable aviation fuel. EImaxTO,FF is the system loss corrected nvPM emissions index in mg/kg operating at around 100% available thrust if fuel provided to the plurality of fuel spray nozzles is fossil-based hydrocarbon fuel. The MTO nvPM emissions index ratio of the gas turbine engine is less than 1.
A fuel system for a gas turbine engine comprises a fuel offtake configured and arranged to divert a portion of hydrogen fuel from a main fuel conduit, a burner configured and arranged to burn the portion of hydrogen fuel diverted from the main fuel conduit, a heat exchanger configured and arranged to transfer heat from exhaust gasses produced by the burner to hydrogen fuel in the main fuel conduit, and an outlet baffle positioned between the burner and the heat exchanger. The outlet baffle is configured to introduce turbulence to combustion gases entering the heat exchanger.
F02C 7/224 - Heating fuel before feeding to the burner
F02C 3/22 - Gas-turbine plants characterised by the use of combustion products as the working fluid using a special fuel, oxidant, or dilution fluid to generate the combustion products the fuel or oxidant being gaseous at standard temperature and pressure
A gas turbine engine comprises an engine core having first and second engine core exhaust paths arranged to pass first and second portions respectively of the mass flow of the engine’s core exhaust mass. A heat-exchange system comprises a recuperator system disposed within the first engine core exhaust path and arranged to transfer heat from said first portion to a buffer fluid, and a heat exchanger arranged to transfer heat from the buffer fluid to fuel within a fuel path arranged to convey fuel to the engine’s combustor. The engine provides for heat to be recovered from the engine’s core exhaust flow to the engine’s fuel supply, thus improving thermal efficiency, but without significantly impeding the engine core exhaust flow or presenting the significant fire or explosion risk associated with a recuperator arranged to heat fuel directly.
F02C 7/224 - Heating fuel before feeding to the burner
F01D 25/24 - CasingsCasing parts, e.g. diaphragms, casing fastenings
F02C 3/22 - Gas-turbine plants characterised by the use of combustion products as the working fluid using a special fuel, oxidant, or dilution fluid to generate the combustion products the fuel or oxidant being gaseous at standard temperature and pressure
A gas turbine engine for an aircraft includes a combustor with a combustion chamber and fuel spray nozzles to inject fuel into the chamber. The nozzles include a first and second subset. The combustor is operable so each of the first subset is supplied with fuel at a greater rate than each of the second subset. A ratio of nozzles in the first subset to the second subset is 1:3 to 1:6. A fuel-flow nvPM emissions index ratio is
A gas turbine engine for an aircraft includes a combustor with a combustion chamber and fuel spray nozzles to inject fuel into the chamber. The nozzles include a first and second subset. The combustor is operable so each of the first subset is supplied with fuel at a greater rate than each of the second subset. A ratio of nozzles in the first subset to the second subset is 1:3 to 1:6. A fuel-flow nvPM emissions index ratio is
EI
idle
×
W
f
,
idle
EI
maxTO
×
W
f
,
maxTO
.
A gas turbine engine for an aircraft includes a combustor with a combustion chamber and fuel spray nozzles to inject fuel into the chamber. The nozzles include a first and second subset. The combustor is operable so each of the first subset is supplied with fuel at a greater rate than each of the second subset. A ratio of nozzles in the first subset to the second subset is 1:3 to 1:6. A fuel-flow nvPM emissions index ratio is
EI
idle
×
W
f
,
idle
EI
maxTO
×
W
f
,
maxTO
.
EIidle is the system loss corrected nvPM emissions index in mg/kg of the engine at around 7% available thrust. EImaxTO is the index at around 100% available thrust. Wf,idle is the rate of fuel to the nozzles in kg/s at around 7% available thrust. Wf,maxTO is the rate of fuel to the nozzles in kg/s at around 100% available thrust. The index ratio is between 0.357 and 8.
F23R 3/28 - Continuous combustion chambers using liquid or gaseous fuel characterised by the fuel supply
F02C 3/20 - Gas-turbine plants characterised by the use of combustion products as the working fluid using a special fuel, oxidant, or dilution fluid to generate the combustion products
F02C 7/224 - Heating fuel before feeding to the burner
F02C 7/228 - Dividing fuel between various burners
A gas turbine engine for an aircraft includes a combustor with a combustion chamber and fuel spray nozzles to inject fuel into the chamber. The nozzles include a first and second subset. The combustor is operable so each of the first subset is supplied with fuel at a greater rate than each of the second subset. A ratio of nozzles in the first subset to nozzles in the second subset is 1:3 to 1:6. A thrust nvPM emissions index ratio is EImaxTO/FmaxTO/EIidle/Fidle. EIidle is system loss corrected nvPM emissions index in mg/kg of the engine operating at around 7% available thrust. EImaxTO is system loss corrected nvPM emissions index in mg/kg of the engine at around 100% available thrust. FmaxTO is thrust of the engine at around 100% available thrust. Fidle is thrust at around 7% available thrust. The thrust nvPM emissions index ratio is between 0.0009 and 0.02.
A power system for an aircraft includes at least one gas turbine engine arranged to burn a fuel so as to provide power to the aircraft; at least one first fuel tank arranged to be used to power ground-based operation of the aircraft; at least one secondary fuel tank arranged to contain a fuel to be used to power the aircraft in flight; and a fuel manager arranged to control fuel supply so as to take fuel from only the at least one first fuel tank to power at least the majority of ground-based operations.
There is provided a reheat assembly 300, 300A for a gas turbine engine 10. The reheat assembly 300, 300A comprises a support duct 340 and a flameholder 370. The flameholder 370 comprises a flange portion 32 defining an inlet aperture 374 to an interior of the flameholder 370 and a boss 36 extending away from the flange portion 32 into the interior of the flameholder 370. The flameholder 370 is mounted to the support duct 340 by a fastener 38 extending through the support duct 340 into a hole 37 defined by the boss 36. The flameholder 370 is configured to receive a flow of air via the inlet aperture 374 to cool the boss 36.
F02K 3/11 - Heating the by-pass flow by means of burners or combustion chambers
F02K 3/10 - Plants including a gas turbine driving a compressor or a ducted fan with supplementary heating of the working fluidControl thereof by after-burners
There is provided a flameholder 370 for a reheat assembly 300, 300A of a gas turbine engine 10. The flameholder comprises an internal flow passageway 376 extending from an inlet aperture 374 and defining a flow direction F for flow through the flameholder 370. The internal flow passageway 376 is defined by an internal surface 34 of the flameholder 370. A grid of recesses 31 is formed in the internal surface 34 of the flameholder 370.
A gas turbine engine for an aircraft 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; a bypass duct delimited by a bypass duct inner wall and a bypass duct outer wall and located radially outwardly from the engine core and downstream of the fan; and an outlet guide vane assembly, located within the bypass duct and, comprising a plurality of outlet guide vanes distributed circumferentially within the bypass duct, each outlet guide vane extending radially along a span between the bypass duct inner wall and the bypass duct outer wall, wherein a space-chord ratio of at least one outlet guide vane, at 50% of the span length from the bypass duct inner wall, is less than 0.72.
A fuel system for a hydrogen fuelled gas turbine engine includes a main hydrogen fuel storage unit, a hydrogen fuel pump configured to be supplied with hydrogen from the hydrogen storage unit, a hydrogen fuel preheater configured to be supplied with high pressure hydrogen from the hydrogen fuel pump, and configured to supply heated gaseous hydrogen to a combustor of the gas turbine engine, and a hydrogen priming tank configured to store compressed gaseous hydrogen and to deliver gaseous hydrogen to at least the hydrogen fuel pump and preheater. The fuel system comprises a high-pressure gaseous hydrogen fuel offtake downstream in hydrogen fuel flow of the fuel pump in fluid communication with the hydrogen priming tank, and configured to fill the hydrogen priming tank with high pressure gaseous hydrogen.
A method of starting a liquid hydrogen fuelled gas turbine engine of an aircraft propulsion system, wherein the aircraft propulsion system includes a hydrogen storage tank configured to store liquid hydrogen, a liquid hydrogen pump configured to pump hydrogen in at least a liquid state, a core combustor configured to receive hydrogen fuel from the hydrogen fuel pump, and a hydrogen fuel vent provided downstream of the liquid hydrogen pump, and configured to selectively vent hydrogen fuel. The method includes, in a liquid priming step, flowing hydrogen from the hydrogen storage tank through the liquid hydrogen pump and venting hydrogen through the hydrogen fuel vent until the hydrogen pump is primed with liquid hydrogen, then, in a liquid pumping step, operating the liquid hydrogen pump to pump liquid hydrogen to the core combustor at a required flow rate and pressure for engine ignition in an engine ignition step.
An apparatus for inspecting a gas turbine engine component having a plurality of cooling apertures includes a main body including a plurality of body apertures configured to at least partially align with the plurality of cooling apertures. Each body aperture from the plurality of body apertures extends through the main body. The apparatus further includes a seal connected or connectable to the main body and configured to engage with the gas turbine engine component. The seal includes a plurality of seal apertures corresponding to the plurality of body apertures. Each seal aperture from the plurality of seal apertures extends through the seal. The plurality of seal apertures is at least partially aligned with the plurality of body apertures of the main body, such that the plurality of seal apertures is disposed or disposable in fluid communication with the plurality of body apertures.
A method of determining one or more fuel characteristics of an aviation fuel suitable for powering a gas turbine engine of an aircraft. The method includes: exposing the surface of a piezoelectric crystal to the fuel; measuring a vibration parameter of the piezoelectric crystal; and determining one or more fuel characteristics of the fuel based on the vibration parameter. Also disclosed is a fuel characteristic determination system, a method of operating an aircraft, and an aircraft.
G01N 21/33 - Investigating relative effect of material at wavelengths characteristic of specific elements or molecules, e.g. atomic absorption spectrometry using ultraviolet light
B64D 27/10 - Aircraft characterised by the type or position of power plants of gas-turbine type
G01N 21/3577 - Investigating relative effect of material at wavelengths characteristic of specific elements or molecules, e.g. atomic absorption spectrometry using infrared light for analysing liquids, e.g. polluted water
B64D 37/00 - Arrangements in connection with fuel supply for power plant
G01N 21/35 - Investigating relative effect of material at wavelengths characteristic of specific elements or molecules, e.g. atomic absorption spectrometry using infrared light
A nuclear power system wherein: a nuclear heat source (10), turbine (15), heat dissipator (20), compressor (25), and electric heater (35) form a fluid flow circuit for channelling a working fluid; the turbine is connected to the compressor and a generator (40) by a connecting shaft (30) such that when the turbine rotates electrical energy is generated; a sensor system (55) is connected to a controller (60), and is configured to acquire data from at least the electric heater, the nuclear heat source, the generator, a battery (50), and a load (45); and the controller is configured to control the distribution of the electrical energy between a load, the battery, and the electric heater, based on the data provided by the sensor system, such that the electrical energy output to the grid can be varied, whilst maintaining a constant temperature, pressure, and massflow of the working fluid entering the turbine.
A nosecone for a ducted fan gas turbine engine is shown. The nosecone includes a body with an outer surface that tapers in axial extent from an apex to a base of the body, the apex defining a position of 0 percent axial extent and the base defining a position of 100 percent axial extent, and in which the outer surface includes a plurality of undulations extending between a first location on the outer surface and a second location on the outer surface, wherein the first location is positioned at from 0 to 50 percent of axial extent and the second location is positioned at from 85 to 100 percent of axial extent.
F02C 7/00 - Features, component parts, details or accessories, not provided for in, or of interest apart from, groups Air intakes for jet-propulsion plants
An aircraft propulsion system comprises a core gas turbine engine (201) comprising a core compressor (202, 204) configured to provide core air to a core combustor (206) and a core turbine (208, 209) in fluid flow series. An auxiliary compressor (220) is provided, which is separate to the core compressor (202, 204), and configured to provide air to an auxiliary air system (218). A heat exchanger (218) is provided, which is configured to transfer heat from air from the auxiliary compressor (220) to fuel in the main fuel conduit (217) prior to provision to the core combustor (206).
A crawler robot comprising a proximal end section, a distal end section and a central core section, through each section passes a central channel; the distal end section comprises a rigid core surrounding a deformable foot, the proximal end section comprises a rigid core surrounding a deformable foot, and the central core section comprises a deformable annular body surrounding an extension spring, surrounding the annular body is at least one balloon, and wherein fluid carrying conduits are connected to the balloon in the central core section, and to the deformable foot in the distal end section and the deformable foot in the proximal end section.
B62D 57/02 - Vehicles characterised by having other propulsion or other ground-engaging means than wheels or endless track, alone or in addition to wheels or endless track with ground-engaging propulsion means, e.g. walking members
B08B 9/043 - Cleaning the internal surfacesRemoval of blockages using cleaning devices introduced into and moved along the pipes moved by externally powered mechanical linkage, e.g. pushed or drawn through the pipes
F16L 55/26 - Pigs or moles, i.e. devices movable in a pipe or conduit with or without self-contained propulsion means
A crawler robot has a proximal end section, a distal end section and a core section. Each of the proximal end section, the distal end section and core section has a hollow core. The core section has a flexible body and is surrounded by at least one balloon that is supplied by a fluid carrying conduit. The proximal end section and the distal end section are connected to fluid carrying conduits and have at least three thrust vents each for venting the supplied fluid to generate a thrust force.
A method of operating a gas turbine engine is disclosed, the gas turbine engine including a combustor arranged to combust a fuel and a fuel management system arranged to provide the fuel to the combustor. The fuel management system includes two fuel-oil heat exchangers through which oil and the fuel flow, the heat exchangers arranged to transfer heat to the fuel and comprising a primary fuel-oil heat exchanger and a secondary fuel-oil heat exchanger; and a fuel pump arranged to deliver the fuel to the combustor, wherein the fuel pump is located between the two heat exchangers. The method includes controlling the fuel management system so as to raise the fuel temperature to at least 135° C. on entry to the combustor at cruise conditions.
A method of operating an electrical power system for an aircraft. The electrical power system includes a semiconductor-based active power converter and a contactor coupled to and controllable by the semiconductor-based active power converter. The method includes, in response to a determination that there is a fault within the electrical power system, operating in a fault mode which includes: maintaining the semiconductor-based active power converter in a blocked configuration (optionally followed by maintaining the semiconductor-based active power converter in a crow-bar configuration); and subsequently causing the contactor to be opened.
H02H 7/22 - Emergency protective circuit arrangements specially adapted for specific types of electric machines or apparatus or for sectionalised protection of cable or line systems, and effecting automatic switching in the event of an undesired change from normal working conditions for distribution gear, e.g. bus-bar systemsEmergency protective circuit arrangements specially adapted for specific types of electric machines or apparatus or for sectionalised protection of cable or line systems, and effecting automatic switching in the event of an undesired change from normal working conditions for switching devices
There is provided a propulsion machine 10 comprising a fluid duct defined by a wall 42, 44 and a moveable member 34, 36 with a sealing module 90, 90′ therebetween. The moveable member 34, 36 is moveable relative to the wall 42, 44. The sealing module 90, 90′ comprises a mounting structure 92 coupled to the moveable member 34, 36 and an extendable structure 94 having a sealing surface 96. A chamber 98 is defined between the mounting structure 92 and the extendable structure 94 throughout a travel of the extendable structure 94 relative to the mounting structure 92. The sealing module 90, 90′ is configured to receive a pressurized actuation fluid into the chamber 98 to load the sealing surface 96 against an opposing surface 43 of the wall 42, 44 to provide a seal with the opposing surface 43.
Apparatus for active remote detection of leaking hydrogen comprises (i) pressure tubing for enclosing a length of pipework from which leaking hydrogen is to be detected; (ii) a detection vessel containing a hydrogen sensor; and (iii) connecting tubing connecting the interior of the pressure tubing to the interior of the detection vessel.
G01M 3/28 - Investigating fluid tightness of structures by using fluid or vacuum by measuring rate of loss or gain of fluid, e.g. by pressure-responsive devices, by flow detectors for pipes, cables, or tubesInvestigating fluid tightness of structures by using fluid or vacuum by measuring rate of loss or gain of fluid, e.g. by pressure-responsive devices, by flow detectors for pipe joints or sealsInvestigating fluid tightness of structures by using fluid or vacuum by measuring rate of loss or gain of fluid, e.g. by pressure-responsive devices, by flow detectors for valves
G01K 7/02 - Measuring temperature based on the use of electric or magnetic elements directly sensitive to heat using thermoelectric elements, e.g. thermocouples
G01K 7/22 - Measuring temperature based on the use of electric or magnetic elements directly sensitive to heat using resistive elements the element being a non-linear resistance, e.g. thermistor
A robotic shaping and forming system comprising a plurality of opposing robotic arms, a work platform for supporting a non-planar workpiece and a computer system, the plurality of opposing robotic arms having multiple degrees of freedom and an end effector for holding a tool, and wherein at least one robotic arm being mounted on a radially extending rail, the computer system being connected to the plurality of robotic arms, the computer system controlling the movement of the robotic arms, so that at least a pair of robotic arms work together to shape and form a non-planar workpiece that is mounted upon a work platform.
A method of forming a plurality of components on a build-plate, the method comprising the steps of providing a forming system comprising a plurality of high-energy beam generators mounted on a common support and configured to direct energy onto the build-plate; the build-plate comprising at least one segregation zones segregating the build plate into a plurality of build zones, each build zone associated with one of the high-energy beam generators; forming within each build zone at least one datum from material melted by energy from its associated high-energy beam generator; and forming within each build zone at least one component from material melted by energy from its associated high-energy beam generator; cutting or machining the build plate at the segregation zone or segregation zones to separate the build zones.
A turbine shroud assembly for use with a gas turbine engine includes a first shroud segment, a second shroud segment, and a damping strip seal assembly. The first shroud segment has a first carrier segment arranged circumferentially at least partway around a central axis and a first blade track segment supported by the first carrier segment. The second shroud segment is arranged circumferentially adjacent the first shroud segment. The damping strip seal assembly includes a body segment and a damping segment that extends along a curvilinear path.
A gas turbine includes an engine core with a turbine, a compressor, a combustor to combust a fuel, and a core shaft connecting the turbine to the compressor; a fan located upstream of the engine core; a fan shaft; a main gearbox that receives an input from the core shaft and outputs drive to the fan via the fan shaft; a primary oil loop system arranged to supply oil to lubricate the main gearbox; and a heat exchange system arranged to transfer heat between the oil and the fuel, the oil having an average temperature of at least 180° C. on entry to the heat exchange system at cruise conditions. A method of operating the turbine includes transferring heat from the oil to the fuel so as to lower the fuel viscosity to a value of less than or equal to 0.58 mm2/s on entry to the combustor at cruise conditions.
A highly efficient gas turbine engine is provided. The fan of the gas turbine engine is driven from a turbine via a gearbox, such that the fan has a lower rotational speed than the driving turbine, thereby providing efficiency gains. The efficient fan system is mated to a core that has low cooling flow requirements and/or high temperature capability, and which may have particularly low mass for a given power.
F02C 3/073 - Gas-turbine plants characterised by the use of combustion products as the working fluid having a turbine driving a compressor the compressor comprising only axial stages the compressor and turbine stages being concentric
A gas turbine engine for an aircraft includes a rich burn, quick quench, lean burn combustor having 14-22 fuel spray nozzles or 2-6 fuel spray nozzles per unit engine core size. A thrust nvPM emissions index ratio is
A gas turbine engine for an aircraft includes a rich burn, quick quench, lean burn combustor having 14-22 fuel spray nozzles or 2-6 fuel spray nozzles per unit engine core size. A thrust nvPM emissions index ratio is
EI
maxTO
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maxTO
EI
idle
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A gas turbine engine for an aircraft includes a rich burn, quick quench, lean burn combustor having 14-22 fuel spray nozzles or 2-6 fuel spray nozzles per unit engine core size. A thrust nvPM emissions index ratio is
EI
maxTO
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F
maxTO
EI
idle
/
F
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EIidle is the nvPM emissions index in mg/kg of the gas turbine engine operating at around 7% available thrust for given operating conditions. EImaxTO is the nvPM emissions index in mg/kg of the gas turbine engine operating at around 100% available thrust for the given operating conditions. FmaxTO is the thrust of the gas turbine engine at around 100% available thrust in kN. Fidle is the thrust of the gas turbine engine at around 7% available thrust in kN. The thrust nvPM emissions index ratio is greater than 0.09. The gas turbine engine is configured to provide fuel comprising a sustainable aviation fuel to the fuel spray nozzles.
A hybrid transmission component includes a composite tube defining a central axis along its length and extending between a first end and a second end opposite to the first end. The composite tube includes a tube inner surface extending circumferentially about the central axis and a tube outer surface that is radially spaced apart from the tube inner surface with respect to the central axis. The composite tube further includes a first tube axial end surface extending between the tube inner surface and the tube outer surface at the first end and a second tube axial end surface extending between the tube inner surface and the tube outer surface at the second end. The tube inner surface comprises a wedge portion disposed at the first end.
F16D 1/08 - Couplings for rigidly connecting two coaxial shafts or other movable machine elements for attachment of a member on a shaft or on a shaft-end with clamping hubCouplings for rigidly connecting two coaxial shafts or other movable machine elements for attachment of a member on a shaft or on a shaft-end with hub and longitudinal key
A hydrogen fuel delivery system (300) comprises a fuel line (312) having an inlet (315) and an outlet (316), a liquid fuel pump (307) configured to provide a flow of liquid hydrogen fuel from a hydrogen fuel storage tank (308) to the fuel line inlet (315), a heat exchanger (306) having first and second fluid paths (313, 314), the fuel line (312) passing through the first fluid path (313), a pre-heater line (317) having an inlet (318) connected to the fuel line (312) between the fuel line inlet (315) and the heat exchanger (306), the pre-heater line (317) comprising a first control valve (301) and a burner (305) between the pre-heater line inlet (318) and the heat exchanger (306), the pre-heater line (317) passing through the second fluid path (314) of the heat exchanger (306) towards a pre-heater line outlet (319), a second control valve (302) in the fuel line (312) between the heat exchanger (306) and the fuel line outlet (316), a first temperature sensor (321) configured to measure a first fuel temperature (T1) in the fuel line (312) between the heat exchanger (306) and the second control valve (302), and a control system (400) configured provide a first control signal (CV1) to control operation of the first control valve (301) dependent on an input target temperature (T1Target) compared to the first fuel temperature (T1) and on a measure of fuel flow (mbfuel) through the pre-heater line (317).
A current limiting device and an electrical power system including a current limiting device are provided. The current limiting device includes: a primary current path extending between a first node N1 and a second node N2 and having a primary current limiter connected therein, the primary current limiter including at least one JFET 101, 1011-N; and a secondary current path extending between the first node N1 and the second node N2 in parallel with the primary current path, the secondary current path including a Transient Voltage Suppressor (TVS). The primary current limiter is configured so that a voltage drop across the primary current limiter increases as a current flowing through the primary current path increases. When the current flowing through the primary current path passes a threshold, the voltage drop across the primary current limiter passes a breakdown voltage of the TVS.
