Abstract: A fuel system for a hydrogen fueled aircraft propulsion system comprises a fuel tank configured to store hydrogen fuel, a main fuel conduit configured to provide fuel to a combustor of a gas turbine engine, a fuel heater comprising a catalytic combustor configured to catalytically combust a portion of the hydrogen fuel prior to delivery to the combustor and a heat exchanger configured to exchange heat between exhaust gases from the fuel heater and hydrogen fuel in the fuel conduit.

Abstract: An electric machine includes a stator having a phase arrangement, and a rotor. The phase arrangement includes legs connected in parallel at first and second primary junctions. Each leg includes a plurality of coils connected in series through a respective intermediate junction. Each leg is connected to at least one other leg by a branch at the respective intermediate junctions such that the phase arrangement is a bridge circuit. The phase arrangement conducts a motor current through the respective coils between the first and second primary junctions. The phase arrangement permits an alignment current to flow between the first and second primary junctions along an alignment current path which passes through at least one coil of two different legs via a respective branch. The motor currents cause a torque to be applied to the rotor, and the alignment current causes a translational force to be applied to the rotor.

Abstract: A gas turbine engine for an aircraft comprising an engine core comprising a turbine, a compressor, and a core shaft connecting the turbine to the compressor. A fan located upstream of the engine core, the fan comprising a plurality of fan blades and having a fan diameter and a gearbox arranged to receive an input from the core shaft and to output drive to the fan so as to drive the fan at a lower rotational speed than the core shaft, thereby defining a gear ratio. The gearbox being an epicyclic gearbox. A gearbox diameter is defined as the pitch circle diameter of the ring gear; and a geared-fan value defined by a product of the fan diameter and gear ratio, which are divided by the gearbox diameter, being greater than 12.

Abstract: A coating system includes a bond layer including CoNiCrAlY; a thermal barrier layer at least partially disposed on the bond layer and including yttria stabilised zirconia; a mesh layer connecting the thermal barrier layer to the bond layer and including CoNiCrAlY; and an abradable layer disposed on the thermal barrier layer and including magnesium spinel. The mesh layer includes a plurality of cells connected to each other and is at least partially embedded in the thermal barrier layer. Each cell defines a cell opening therethrough and the cell opening of each of the plurality of cells at least partially receives the thermal barrier layer therein. The abradable layer and the thermal barrier layer are at least partially removable from the mesh layer and the bond layer without damaging them.

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: 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.

Abstract: 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.

Abstract: 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.

Abstract: 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.

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: 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.

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 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.

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.