Abstract: 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.

Abstract: 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).

Abstract: 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).

Abstract: 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).

Abstract: 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.

Abstract: An aircraft including first and second like gas turbine engines having like combustion apparatus each of which includes a respective plurality of fuel injectors arranged in an annular array and having a first and second sets of fuel-emitting apertures arranged to emit fuel normally to the plane of the array and in a direction having a component in the plane of the array directed towards an adjacent fuel injector. The aircraft further includes a fuel system which for any given engine increases the proportion of the total fuel flow rate to that engine which is provided to the second set of apertures during lighting or re-lighting of that engine's combustion apparatus, or upon detection of aircraft manoeuvring likely to increase the risk of flame-out. The aircraft provides faster and more reliable lighting and re-lighting, and additional resistance to flame-out, especially for use of hydrogen fuel.

Abstract: A gas turbine engine for an aircraft is disclosed. The gas turbine engine comprises: a combustor, comprising a combustion chamber and a plurality of fuel spray nozzles configured to inject fuel into the combustion chamber, wherein the plurality of fuel spray nozzles comprises a first subset of fuel spray nozzles and a second subset of fuel spray nozzles, wherein the combustor is operable in a condition in which each of the fuel spray nozzles of the first subset of fuel spray nozzles is supplied with fuel at a greater fuel flow rate than each of the fuel spray nozzles of the second subset of fuel spray nozzles, wherein a ratio of the number of fuel spray nozzles in the first subset of fuel spray nozzles to the number of fuel spray nozzles in the second subset of fuel spray nozzles is in the range of 1:3 to 1:6.

Abstract: A DC:DC converter includes: a DC:AC converter circuit having DC and AC sides; an AC:DC converter circuit having DC and sides; an AC link connecting the AC side of the DC:AC circuit and the AC side of the AC:DC converter circuit, the AC link including a transformer having a first winding connected to the AC side of the DC:AC converter circuit and a second winding connected to the AC side of the AC:DC converter; a capacitor; and a switch arrangement having a first state and a second state, wherein: in the first state, the capacitor is connected in series between the AC side of the DC:AC converter circuit and the first winding of the transformer; and in the second state, there is a current path between the AC side of the DC:AC converter circuit and the first winding of the transformer that does not include the capacitor.

Abstract: A decoking cleaning system for cleaning at least one component in-situ within a combustion engine system, the decoking system including at least one ultrasonic transducer with a means of connecting the at least one ultrasonic transducer to a component to be cleaned within the combustion engine system, the at least one ultrasonic transducer being connected to a power and voltage supply.

Abstract: A computer-based control system for a gas turbine engine includes: a controller including control logic, the controller obtains a set of inputs formed by measurements of one or more process variables and values of one or more engine operation set points, and the control logic determines one or more process command values gas turbine engine operation in response to the set of inputs; and an event detection unit obtaining further measurements of one or more process variables and determine whether an abnormal event has occurred based on the further measurements of one or more process variables. The controller also updates the tuning variables using an event accommodation data array, a given tuning variable being changed in response to detection of a given abnormal event when the respective element in the event accommodation data array is assigned an intervention value.

Abstract: A thermal management system for cooling at least one electronic component, the thermal management system including: a heat exchanger connectable to at least one electronic component, the heat exchanger configured to dissipate heat generated by the at least one electronic component; a plurality of cooling channels each connected to a respective pump and the heat exchanger, wherein: each cooling channel is configured to carry fluid between the heat exchanger and the respective pump, and each pump is configured to pump fluid along the respective cooling channel such that fluid carried by each cooling channel flows through the heat exchanger.

Abstract: A system for cooling one or more components associated with a gas turbine engine includes a main duct, a first duct that receives and directs a portion of airflow from the main duct towards a core zone cooling arrangement of the gas turbine engine, and a second duct that receives and directs a portion of the airflow from the main duct towards a turbine case cooling arrangement or an oil cooling unit of the gas turbine engine. The system includes a valve unit including a first valve member disposed in the first duct and a second valve member disposed in the second duct. The first and second valve members control a fluid flow through the first and second ducts, respectively. The system includes at least one controller configured to control the valve unit to modulate the portion of the airflow through each of the first and second ducts.

Abstract: 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.

Abstract: 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).

Abstract: 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.

Abstract: 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.

Abstract: 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).

Abstract: 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.

Abstract: 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.

Abstract: 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 pin of a planet gear carrier with a journal bearing having an internal sliding surface on the planet gear and an external sliding surface on the pin.