A power electronics cooling assembly with a housing for power electronics. The housing is mountable to a bulkhead for separating a fire zone from a non-fire zone of a gas turbine engine. In the non-fire zone, one or more ducts extend from the bulkhead to the housing. A fluid supply and return pipes extend through the one or more ducts from the bulkhead to the housing. The fluid supply and return pipes are for carrying an ignitable fluid coolant. A heat exchange structure arranged within the housing is configured to receive the ignitable fluid coolant from the fluid supply pipe, facilitate transfer of heat from the power electronics within the housing to the ignitable fluid coolant, and output the ignitable fluid coolant to the fluid return pipe. A system, gas turbine engine assembly and an aircraft comprising the power electronics cooling assembly are also provided.
A ducting arrangement for a gas turbine engine and a method of assembling such a ducting arrangement, wherein in the gas turbine engine includes an outer annular duct structure with an outer mount and an inner annular duct structure with an inner mount. The inner annular duct structure is configured to be received within the outer annular duct structure such that the inner mount and the outer mount are angularly aligned. In one aspect, the inner mount and the outer mount are coupled together by a tension tie to attach the outer annular duct structure to the inner annular duct structure, the tension tie being a cable or a wire. In another aspect, the inner mount and the outer mount are configured to be coupled together by a tension tie formed by a cable or a wire to attach the outer annular duct structure to the inner annular duct structure.
The disclosure relates to a ducted fan aircraft propulsion system and to an aircraft incorporating such a propulsion system. Example embodiments include a ducted fan aircraft propulsion system (300), comprising: a duct (301); a central body portion (302) having first and second ends (303, 304) and extending through the duct (301); a payload portion (305) extending from the first end (303) of the central body portion (302); and a rotor (306) extending across an internal volume (307) of the duct (301) from the central body portion (302), wherein the duct (301) comprises an outwardly flared inlet end (308) such that an inlet air flow passage (309) between the central body portion (302) and an inner surface (310) of the duct (301) has a sectional area that decreases from the inlet end (308) of the duct (301) to the rotor (306).
A method of manufacturing a component including a metal alloy comprises measuring crystallographic texture of a volume of a component, determining a risk factor of the component for cold dwell fatigue failure, and adjusting metallurgical processing of the component based on the risk factor. Such risk analysis and mitigation may aid in improving the usage and operation of components including materials that are susceptible to cold dwell fatigue failure.
A power reduction system for an energy storage system of an aircraft includes a controller configured to control power reduction of power supplied from the energy storage system to an aircraft engine supply bus; and an override switch configurable in an override state and a non-override state. The override switch is configured to: in the non-override state, permit the controller to control the power reduction according to a default configuration comprising one or more parameters that trigger the power reduction; and in the override state, control the power reduction to be performed according to a relaxed configuration that at least one of relaxes and omits the one or more parameters in the default configuration.
B64D 31/00 - Power plant control systemsArrangement of power plant control systems in aircraft
B60L 58/10 - Methods or circuit arrangements for monitoring or controlling batteries or fuel cells, specially adapted for electric vehicles for monitoring or controlling batteries
B64D 27/24 - Aircraft characterised by the type or position of power plants using steam or spring force
A gas turbine engine generates noise during use, and one particularly important flight condition for noise generation is take-off. A gas turbine engine is provided that has high efficiency together with low noise, in particular from the fan and the turbine that drives the fan. Values are defined for a combined contribution of the turbine noise and the noise from the fan emanating from the front of the engine to the Effective Perceived Noise Level (EPNL) at take-off to be in the range of from 2 EPNdB and 15 EPNdB lower than the EPNL of the whole engine at take-off.
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 method includes forming a moulded component that is axisymmetric about a component axis; segmenting the moulded component into a plurality of segments, each pair of adjacent segments including a pair of surfaces that is formed during segmentation of the moulded component; providing, via computerised numerical control machining, complementary finger joint profiles on the pair of surfaces of the each pair of adjacent segments; providing a plurality of slots on at least one of the pair of surfaces of the each pair of adjacent segments; positioning a plurality of blades partially within the plurality of slots; mating the complementary finger joint profiles provided on the pair of surfaces of the each pair of adjacent segments; and joining the each pair of adjacent segments to each other, such that the plurality of blades is retained within the plurality of slots.
A continuum arm robot segment that has a proximal end section and a distal end section. The continuum arm robot segment has at least one vertebra between the proximal end section and the distal end section, the vertebra being linked to the proximal end section and to the distal end section by linkages, wherein the proximal end section, the distal end section and the at least one vertebra are configured to have hollow cores to provide a passage within the continuum arm robot segment. The continuum arm robot segment also has at least two tendons that are connected to the distal end section and pass through the at least one vertebra and the proximal end section. The at least two tendons are routed helically so that they undergo substantially at least a 360° rotation between the proximal end section and the distal end section.
A turbine assembly includes a turbine case, a turbine shroud assembly including a carrier segment, and a locating plate. The locating plate is coupled with the turbine case axially forward of the carrier segment to block axially forward movement of the carrier segment and prevent separation of the carrier segment from the turbine case. The locating plate includes a main wall, a raised portion extending upwardly a first radial distance, and two circumferentially spaced apart extensions extending upwardly a second radial distance. The second radial distance is greater than the first radial distance such that, in a first arrangement, only the extensions contact the turbine case and such that, in a second arrangement, the raised portion is pulled toward the turbine case via a fastener so as to contact the turbine case in addition to the extensions.
A method of operating a gas turbine engine includes an engine core with a turbine, compressor, fuel combustor, and core shaft; 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 supplying oil to the gearbox; and a heat exchange system including: air-oil and fuel-oil heat exchangers, and wherein the oil loop system branches such that part of the oil flows along each branch and the air-oil and fuel-oil heat exchangers are parallel on different branches; and a modulation valve allows the oil sent via each branch to be varied, the method including controlling the heat exchange system wherein, under cruise conditions, a heat transfer ratio of:
is in the range from 0 to 0.67.
A method of operating a gas turbine engine comprising an engine core comprising a turbine, a compressor, a combustor arranged to combust a fuel, and a core shaft connecting the turbine to the compressor; a fan located 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; and a heat exchange system comprising an air-oil heat exchanger through which the oil flows, 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, is disclosed, the method comprising controlling the valve such that, under cruise conditions, an oil flow ratio of:
is in the range from 0 to 0.59.
An article includes a silicon-containing ceramic substrate and a coating system overlying the silicon-containing ceramic substrate. The coating system includes an intermediate coating overlying the silicon-containing ceramic substrate and a barrier coating overlying the intermediate coating. The intermediate coating includes silicon and hafnium disilicide. A coefficient of thermal expansion of the intermediate coat is less than about 7 parts per million (ppm) per degree Kelvin (K).
There is provided an air pressurisation system for an aircraft, the air pressurisation system comprising: a transmission for driving a rotor of a blower compressor for the air pressurisation system, wherein the transmission comprises: a first input configured to receive drive from a spool of a gas turbine engine; a second input; and an output configured to drive to the rotor of the blower compressor, wherein a speed of the output is determined by a function of speeds of the first and second inputs; and a brake coupled to the second input, wherein the brake is operable to engage and brake the transmission at the second input, in order to fix a transmission ratio of the transmission between the first input and the output.
A fuel system for a hydrogen fueled gas turbine engine comprises a main hydrogen fuel storage unit configured to store liquid hydrogen. The main fuel storage unit comprises an ullage space configured to store ullage fluid comprising gaseous or supercritical hydrogen. The system further comprises a liquid hydrogen drain line configured to provide liquid hydrogen from the main hydrogen fuel storage unit to a fuel conduit, the fuel conduit being configured to supply hydrogen fuel to a combustor of the gas turbine engine. A priming line is provided, which is configured to provide ullage fluid to the fuel conduit, and a priming valve is configured to selectively control flow through the priming line.
Gearboxes for aircraft gas turbine engines, in particular arrangements for journal bearings such gearboxes, and related methods of operating such gearboxes and gas turbine engines. A gearbox for an aircraft gas turbine engine includes: 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.
Rolls-Royce North American Technologies Inc. (USA)
Inventor
Lee, Andrew
Desai, Mihir
Kalyanasamy, Govindaraj
Collett, Mark
Abstract
A method of monitoring a fuel system in a gas turbine engine. The method may comprise pumping fuel to a combustor from a fuel tank with a pump. The method may comprise controlling a flow of the fuel to the combustor with a metering valve disposed downstream of the pump and closing a spill valve disposed downstream of the pump, wherein the spill valve is closed in fixed increments and closing the spill valve increases a pressure in the fuel system. The method may comprise opening a pressure valve in response to the pressure in the fuel system being equal to or greater than a predetermined value, and capturing a degree of closing of the spill valve when the pressure valve opens.
Rolls-Royce North American Technologies Inc. (USA)
Rolls-Royce Corporation (USA)
Inventor
Kalyanasamy, Govindaraj
Desai, Mihir
Lee, Andrew
Collett, Mark
Abstract
A method of mitigating uncommanded or uncontrollable high thrust in a gas turbine engine is provided. The method may comprise pumping fuel to a combustor from a fuel tank, controlling a flow rate of the fuel to the combustor with a metering valve, spilling a portion of the fuel pumped by the pump with a primary spill valve, controlling a pressure of the fuel flowing to the combustor via a pressure valve, detecting a pressure differential across the pressure valve with a pressure transducer, determining the flow rate of the fuel based on the detected pressure differential and the positional feedback of the pressure valve opening, comparing the determined flow rate with a demand flow rate, and opening a secondary spill valve when the determined flow rate exceeds the demand flow rate.
An inflatable probe for testing a component. The inflatable probe has a balloon formed of a dielectric material, the balloon having a neck and at least one electrode pair comprising an inner electrode and an outer electrode, the inner electrode being positioned on an internal surface of the balloon and the outer electrode being positioned on an external surface of the balloon. The inflatable probe also has a sealing plug that forms an air tight seal with neck of the balloon to retain a fluid within the balloon, the sealing plug at least having a seal electrode to connect to the inner electrode within the balloon, the sealing plug supporting a first wire to connect to a first seal electrode. The inflatable probe also has at least one tool that is connected to the balloon.
A tile for a gas turbine engine combustor. The tile has a base that has a hot-side surface, a cold-side surface, a first circumferential extremity, a second circumferential extremity and a local radial axis. The tile also has a plurality of cooling channels that have inlets on the cold-side surface and outlets on the hot-side surface, and one or more rail structure attached to the cold-side surface of the base.
A method for gathering inflight data during operation of an electric or hybrid-electric aircraft. The aircraft includes: a swappable battery pack, connected to an electrical system of the aircraft, the swappable battery pack including a data storage module accessible via a data interface of the swappable battery pack; and a control unit, connected to the swappable battery pack via the data interface. The method includes, while the aircraft is in flight, gathering, by the control unit, the inflight data from the aircraft and saving the gathered inflight data to the data storage module of the swappable battery pack.
Rolls-Royce North American Technologies Inc. (USA)
Rolls-Royce Corporation (USA)
Inventor
Oechsle, Victor L.
Pesyna, Kenneth M.
Lerg, Bryan H.
Monzella, Michael C.
Rauch, Zachary A.
Moser, Michael
Abstract
A gas turbine engine includes a bypass duct, a rotating detonation augmentor, and a flow valve. The bypass duct is configured to conduct air through a flow path arranged around an engine core of the gas turbine engine. The rotating detonation augmentor is located in the bypass duct and configured to be selectively operated to detonate fuel and a portion of the air to increase thrust for propelling the gas turbine engine. The flow valve is configured to vary selectively the portion of the air flowing into the rotating detonation augmentor to control a magnitude of the thrust increase provided by the rotating detonation augmentor during operation of the rotating detonation augmentor.
F02K 3/02 - 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
A method of using computer implemented neural network for simulation of aerodynamic performance of technical object having geometry, method includes: training neural network using plurality of sets of pre-computed computational fluid dynamics encodings, CFD, outputs, wherein training is generated using inputs including: geometry of at least one training technical object; spatial locations of neural network input nodes as node attributes; relationship between geometry of at least one training technical object and neural network input node locations; associated boundary conditions; operating conditions; and computed outputs including flow fields and aerodynamic performance parameters; training using loss function evaluating error between neural network and pre-computed CFD outputs to produce trained neural network; using trained neural network with new inputs to generate as output predicted aerodynamic performance of technical object, wherein relationship between geometry of at least one training technical object and neural network input node locations includes vector with at least two parameters.
G06F 30/28 - Design optimisation, verification or simulation using fluid dynamics, e.g. using Navier-Stokes equations or computational fluid dynamics [CFD]
The present application discloses 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 comprising: determining a mass of a fuel being loaded, or which has been loaded, onto the aircraft; determining a corresponding volume of the fuel; determining one or more fuel characteristics of the fuel based on the determined mass and volume. Also disclosed is a fuel characteristic determination system, a method of operating an aircraft, and an aircraft.
A gas turbine engine for an aircraft includes a staged combustion system having pilot fuel injectors and main fuel injectors. The gas turbine engine further includes a fuel delivery regulator arranged to control delivery of fuel to the pilot and main fuel injectors, and a fuel characteristic determination module configured to determine one or more fuel characteristics of the fuel being supplied to the staged combustion system. A controller is configured to determine a staging point defining the point at which the staged combustion system is switched between pilot-only operation and pilot-and-main operation, the staging point being determined based on the determined one or more fuel characteristics, the controller being configured to control the staged combustion system according to the determined staging point.
A turbine shroud assembly includes a first shroud segment, a second shroud segment, and a plurality of seals. The first shroud segment includes a first carrier segment and a first blade track segment having a first shroud wall. The second shroud segment includes a second carrier segment and a second blade track. The plurality of seals extend circumferentially into the first shroud segment and the second shroud segment to block gases from escaping the gas path radially between the first shroud segment and the second shroud segment.
The invention relates to a fluid supply assembly having a line assembly from a fluid reservoir to a target location for supplying a fluid mass flow in a variable supply quantity, in particular for supplying fuel from a fuel reservoir to a combustion chamber, having a primary mass flow supplied or able to be supplied via a primary line and a secondary mass flow supplied or able to be supplied via a secondary line. An advantageous adaptation of the fluid mass flow, in particular fuel mass flow, to variable operating conditions is achieved in that disposed in the primary line and/or the secondary line is a passively operating restrictor which is invariable in its geometry and is without moving parts, by which the mass flow ratio of primary mass flow and secondary mass flow is altered as a function of the fluid mass flow supplied via the line assembly.
A method for manufacturing a composite component includes providing a tool switchable between a deposition configuration and a flange-forming configuration. The tool includes a first portion including a first deposition surface, a second portion including a second deposition surface and a pre-form support surface, and a flange-forming block including a curved surface. The method further includes depositing a composite material at least partially on each of the first deposition surface and the second deposition surface in the deposition configuration to provide a composite pre-form. The method further includes moving the second portion relative to the first portion to transition the tool to the flange-forming configuration, which causes the composite pre-form to at least partially engage each of the pre-form support surface and the curved surface. The curved surface defines a radius of a flange of the composite component.
B29C 70/46 - Shaping or impregnating by compression for producing articles of definite length, i.e. discrete articles using matched moulds, e.g. for deforming sheet moulding compounds [SMC] or prepregs
The present disclosure relates to an airframe assembly comprising an airframe support structure defining an engine void and configured to support a gas turbine engine along an engine axis of the engine void; a thermal enclosure configured to surround the gas turbine within the engine void; a heat exchange channel formed between an outer flow-bounding wall of the engine void and an inner wall defined by the thermal enclosure, configured to convey a cooling flow through the engine void outside of the thermal enclosure for the gas turbine engine; and a heat exchanger mounted to the airframe support structure for heat exchange between the cooling flow and a working fluid provided to the heat exchanger.
The present disclosure relates to a combustor assembly for a gas turbine engine. The combustor assembly comprises a combustor liner, a combustor head, a cowl and a fastener. The combustor liner and the cowl define a cavity. The combustor liner has an integral lug that extends into the cavity from a wall of the combustor liner. The fastener extends into the combustor liner lug to fasten the combustor head to the combustor liner.
Rolls-Royce North American Technologies Inc. (USA)
Inventor
Costello, John
Dalley, Robert C.
Schetzel, Douglas
Abstract
A method of monitoring a health of a temperature system may comprise selecting a temperature sensor of a plurality of temperature sensors arranged in an array, where each temperature sensor corresponds to an address within the array. The method may comprise identifying the address corresponding to the single temperature sensor, sending the address to a multiplexer, and selecting the single temperature sensor using the identified address. The method may comprise testing the selected single temperature sensor calculating an average temperature detected by the plurality of temperatures sensors.
G01J 5/90 - Testing, inspecting or checking operation of radiation pyrometers
F01D 21/00 - Shutting-down of machines or engines, e.g. in emergencyRegulating, controlling, or safety means not otherwise provided for
F01D 21/12 - Shutting-down of machines or engines, e.g. in emergencyRegulating, controlling, or safety means not otherwise provided for responsive to temperature
G01J 5/00 - Radiation pyrometry, e.g. infrared or optical thermometry
Rolls-Royce North American Technologies Inc. (USA)
Inventor
Costello, John
Dalley, Robert C.
Schetzel, Douglas
Abstract
A system for detecting a failure in a thermocouple array may comprise the thermocouple array. The thermocouple array may comprise a plurality of thermocouples. The system may comprise an impedance determination circuit. The impedance determination circuit may include a capacitor that has a capacitance equal to an expected capacitance of one of the plurality of thermocouples. The one of the plurality of thermocouples may be connected to test nodes of the impedance determination circuit. The system may comprise a comparator circuit connected to the impedance determination circuit, where the comparator circuit includes an amplifier and a comparator. The system may comprise an excitation circuit connected to the impedance determination circuit, where the excitation circuit includes a waveform generator and an amplifier.
G01K 15/00 - Testing or calibrating of thermometers
G01K 7/02 - Measuring temperature based on the use of electric or magnetic elements directly sensitive to heat using thermoelectric elements, e.g. thermocouples
A combined cycle power generation and storage system is shown. A liquid air energy storage system uses excess power to liquefy air and stores it in a liquid state. A combustion engine produces power by combustion of a carbon-based fuel along with an exhaust stream containing carbon dioxide. A heat recovery system exchanges heat from the exhaust stream to air from the liquid air storage system, and thereby produce a cooled exhaust stream and heated air. An air expansion machine recovers power by expansion of heated air from the heat recovery system. A separation system separates carbon dioxide from ambient air prior to liquefaction during operation of the liquid air energy storage system, and separates carbon dioxide from the cooled exhaust stream during operation of the combustion engine prior to emission of the cooled exhaust stream to atmosphere.
F02B 63/04 - Adaptations of engines for driving pumps, hand-held tools or electric generatorsPortable combinations of engines with engine-driven devices for electric generators
F01N 3/02 - Exhaust or silencing apparatus having means for purifying, rendering innocuous, or otherwise treating exhaust for cooling, or for removing solid constituents of, exhaust
F01N 3/08 - Exhaust or silencing apparatus having means for purifying, rendering innocuous, or otherwise treating exhaust for rendering innocuous
A butterfly valve for a conduit defining a passage for a flow of a fluid therethrough in a flow direction. The butterfly valve includes a shaft rotatably mounted to the conduit and defining a longitudinal axis along its length. The butterfly valve includes a valve body coupled to the shaft, such that the valve body is rotatable along with the shaft about the longitudinal axis between a closed position and a fully open position. The valve body includes a first major surface, a second major surface opposite to the first major surface, a perimeter surface, a central plane, a first lobe, a second lobe, and a third lobe.
F16K 1/22 - Lift valves, i.e. cut-off apparatus with closure members having at least a component of their opening and closing motion perpendicular to the closing faces with pivoted closure members with pivoted discs or flaps with axis of rotation crossing the valve member, e.g. butterfly valves
A turbine shroud assembly includes a first shroud segment, a second shroud segment, and a seal assembly. The first shroud segment includes 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 about the central axis. The seal assembly is configured to block gases from escaping the gas path radially between the first shroud segment and the second shroud segment.
Rolls-Royce North American Technologies Inc. (USA)
Rolls-Royce Corporation (USA)
Inventor
Molnar, Jr., Daniel E.
Heeter, Robert W.
Rivers, Jonathan M.
Abstract
A gas turbine engine includes a fan and a fan case assembly. The fan includes a fan rotor configured to rotate about an axis of the gas turbine engine and a plurality of fan blades coupled to the fan rotor for rotation therewith. The fan case assembly extends circumferentially around the plurality of fan blades radially outward of the plurality of the fan blades.
A multilane power distribution system includes a plurality of DC power sources and a plurality of load devices. Each DC power source powers at least two of the load devices, and each load device is powered by at least two DC power sources. The system further includes a DC connection network including power buses for connecting the DC power sources and the load devices. The power buses having a high side voltage rail for the positive voltage and a low side voltage rail for the negative voltage. At least two power buses are connectable by switchable elements, namely, a first switchable element for the high side voltage rails and a second switchable element for the low side voltage rails. One of the switchable elements is a pulse-width modulated switch that electrically connects the respective voltage rails in accordance with the pulse-width modulation.
H02M 3/158 - Conversion of DC power input into DC power output without intermediate conversion into AC 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 with automatic control of output voltage or current, e.g. switching regulators including plural semiconductor devices as final control devices for a single load
H02M 3/157 - Conversion of DC power input into DC power output without intermediate conversion into AC 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 with automatic control of output voltage or current, e.g. switching regulators with digital control
37.
MOTOR AND MOTOR CONTROLLER THERMAL MANAGEMENT FEATURES
Rolls-Royce North American Technologies, Inc. (USA)
Inventor
Schenk, Peter
Hill, Mathew
Harral, Jacob Ward
Kabbes, Michael Joseph
Abstract
A system includes a gas turbine engine configured to provide propulsion to an aircraft and a starter system configured to start the gas turbine engine. The starter system comprises a motor controller and a closed-loop cooling system configured to cool the motor controller during an emergency in-flight restart operation of the gas turbine engine. The closed-loop cooling system includes a cooling fluid reservoir containing cooling fluid. The cooling fluid is configured to receive thermal energy from the motor controller during the emergency in-flight restart operation of the gas turbine engine.
An example system includes a first gas-turbine engine (GTE) of a plurality of GTEs that are configured to propel an aircraft, the first GTE comprising: a first electric starter of a plurality of electric starters, the first electric starter configured to rotate a spool of the first GTE, wherein the first electric starter is rotationally connected to the spool of the first GTE without a clutch; a second GTE of the plurality of GTEs, the second GTE comprising: a second electric starter of the plurality of electric starters, the second electric starter configured to rotate a spool of the second GTE, wherein the second electric starter is rotationally connected to the spool of the second GTE without a clutch; one or more controllers configured to control the plurality of GTEs; and a common electric starter controller configured to control the plurality of electric starters.
An example system includes a first gas-turbine engine (GTE) of a plurality of GTEs that are configured to propel an aircraft, the first GTE comprising: a first air-turbine starter (ATS) of a plurality of ATSs, the first ATS configured to rotate a spool of the first GTE; and a first electric starter of a plurality of electric starters, the first electric starter configured to rotate the spool of the first GTE; a second GTE of the plurality of GTEs, the second gas-turbine engine comprising: a second ATS of the plurality of ATSs, the second ATS configured to rotate a spool of the GTE; and a second electric starter of the plurality of electric starters, the second electric starter configured to rotate the spool of the GTE; and one or more controllers configured to control the plurality of GTEs.
Rolls-Royce North American Technologies, Inc. (USA)
Inventor
Lawrence, Jeffrey
Schenk, Peter
Harral, Jacob Ward
Huber, Brian Joseph
Abstract
An example system includes a first gas-turbine engine configured to propel an aircraft, the first gas-turbine engine comprising: a first air-turbine starter, the first air-turbine starter configured to rotate a spool of the first gas-turbine engine; and a first electric starter, the first electric starter configured to rotate the spool of the first gas-turbine engine; and one or more controllers collectively configured to: cause, during a time period, the first-air turbine starter and the first electric starter to start the first gas-turbine engine while the aircraft is on the ground; measure, during the time period, values of one or more parameters of the first gas-turbine engine; and determine, based on the values of the one or more parameters, whether the first electric starter is available for use in performing mid-air restart of the first gas-turbine engine.
Gas turbine engine for aircraft includes: an engine core including a turbine, compressor, and core shaft connecting the turbine to the compressor; a fan located upstream of the core; a gearbox; and a gearbox support arranged to at least partially support the gearbox. The fan has a mass in a range of 150 kg to 1200 kg. A moment of inertia of the fan is greater than or equal to 7.40×107 kgmm2. A radial bending stiffness to moment of inertia ratio of:
Gas turbine engine for aircraft includes: an engine core including a turbine, compressor, and core shaft connecting the turbine to the compressor; a fan located upstream of the core; a gearbox; and a gearbox support arranged to at least partially support the gearbox. The fan has a mass in a range of 150 kg to 1200 kg. A moment of inertia of the fan is greater than or equal to 7.40×107 kgmm2. A radial bending stiffness to moment of inertia ratio of:
the
radial
bending
stiffness
of
at
least
one
of
the
fan
shaft
at
the
output
of
the
gearbox
and
the
gearbox
support
the
moment
of
inertia
of
the
fan
Gas turbine engine for aircraft includes: an engine core including a turbine, compressor, and core shaft connecting the turbine to the compressor; a fan located upstream of the core; a gearbox; and a gearbox support arranged to at least partially support the gearbox. The fan has a mass in a range of 150 kg to 1200 kg. A moment of inertia of the fan is greater than or equal to 7.40×107 kgmm2. A radial bending stiffness to moment of inertia ratio of:
the
radial
bending
stiffness
of
at
least
one
of
the
fan
shaft
at
the
output
of
the
gearbox
and
the
gearbox
support
the
moment
of
inertia
of
the
fan
may be greater than or equal to 2.5×10−2 Nkg−1 m−1 mm−2. A tilt stiffness to moment of inertia ratio of:
Gas turbine engine for aircraft includes: an engine core including a turbine, compressor, and core shaft connecting the turbine to the compressor; a fan located upstream of the core; a gearbox; and a gearbox support arranged to at least partially support the gearbox. The fan has a mass in a range of 150 kg to 1200 kg. A moment of inertia of the fan is greater than or equal to 7.40×107 kgmm2. A radial bending stiffness to moment of inertia ratio of:
the
radial
bending
stiffness
of
at
least
one
of
the
fan
shaft
at
the
output
of
the
gearbox
and
the
gearbox
support
the
moment
of
inertia
of
the
fan
may be greater than or equal to 2.5×10−2 Nkg−1 m−1 mm−2. A tilt stiffness to moment of inertia ratio of:
the
tilt
stiffness
of
at
least
one
of
the
fan
shaft
at
the
output
of
the
gearbox
and
the
gearbox
support
the
moment
of
inertia
of
the
fan
Gas turbine engine for aircraft includes: an engine core including a turbine, compressor, and core shaft connecting the turbine to the compressor; a fan located upstream of the core; a gearbox; and a gearbox support arranged to at least partially support the gearbox. The fan has a mass in a range of 150 kg to 1200 kg. A moment of inertia of the fan is greater than or equal to 7.40×107 kgmm2. A radial bending stiffness to moment of inertia ratio of:
the
radial
bending
stiffness
of
at
least
one
of
the
fan
shaft
at
the
output
of
the
gearbox
and
the
gearbox
support
the
moment
of
inertia
of
the
fan
may be greater than or equal to 2.5×10−2 Nkg−1 m−1 mm−2. A tilt stiffness to moment of inertia ratio of:
the
tilt
stiffness
of
at
least
one
of
the
fan
shaft
at
the
output
of
the
gearbox
and
the
gearbox
support
the
moment
of
inertia
of
the
fan
may be greater than or equal to 4.0×10−4 Nmrad−1 kg−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 power electronics converter connectable, on a DC-side, to a DC electrical network and either, on an AC-side, to an electrical machine coupled to a drive shaft of an engine or propulsor, or, on a second DC-side, to a battery pack. The power electronics converter includes: a power conversion unit including a plurality of semiconductor switching elements and a DC-link; and a gate driver unit, configured to control the semiconductor switching elements so that the power conversion unit: inverts DC power received from the DC electrical network to AC power and provides the AC power to the electrical machine, rectifies AC power received from the electrical machine to DC power and provides the DC power to the DC electrical network, or performs DC-DC conversion between the DC-electrical network and the battery pack; wherein the gate driver unit includes an equipment health monitoring component.
G01R 31/00 - Arrangements for testing electric propertiesArrangements for locating electric faultsArrangements for electrical testing characterised by what is being tested not provided for elsewhere
A gas turbine engine configured with an engine core. A fan located upstream of the engine core, the fan comprising a plurality of fan blades; 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. The gearbox being an epicyclic gearbox comprising a sun gear, a plurality of planet gears, a ring gear, and a planet carrier on which the planet gears are mounted, the planet carrier having an effective linear torsional stiffness and the gearbox having a gear mesh stiffness between the planet gears and the ring gear. Additionally, the product of the effective linear torsional stiffness of the planet carrier and the gear mesh stiffness between the planet gears and the ring gear is greater than or equal to 5.0×1018 N2m−2.
A multilane power distribution system includes a plurality of DC power sources and a plurality of load devices. Each DC power source powers at least two load devices, and each load device is powered by at least two of the DC power sources. The system further includes a DC connection network including power buses for connecting the DC power sources and the load devices. The power buses have a high side voltage rail for the positive voltage and a low side voltage rail for the negative voltage. At least two power buses are connectable by switchable elements, namely, a first switchable element for the positive voltage rails and a second switchable element for the negative voltage rails. The switchable elements, when switched on, electrically connect the terminals of two of the DC power sources connected by the respective power buses. A bypass path is provided that bypasses a switchable element.
A hands-free smoking device is disclosed, which is a marijuana, cigarette, and/or cigar holder for use while gaming and other tasks device. The device is worn over the head like a headset and allows the user to smoke hands free. The hands-free smoking device comprises a headset component with a flexible arm component. The flexible arm component retains a smoking implement holder. The smoking implement holder is available in multiple shapes and sizes to hold any size cigarette, blunt, cigar, etc. The flexible arm component is bendable and can be bent in any position near a user's mouth or in any direction needed. Further, the headset component can comprise a cushioning component positioned near a user's head or neck when worn. Thus, the device enables users to maintain a comfortable and hands-free position while smoking.
An example system includes a plurality of transient voltage suppressors (TVSs) that are connected in series across an electrical bus of an aircraft, the electrical bus having a high side and a low side; a plurality of switches, each switch of the plurality of switches configured to selectively shunt a corresponding TVS of the plurality of TVSs to the low side of the electrical bus; and a controller configured to: determine a desired voltage suppression level; and control operation of the plurality of switches such that the plurality of TVSs provides the desired voltage suppression level.
Rolls-Royce North American Technologies, Inc. (USA)
Rolls-Royce Corporation (USA)
Inventor
Badger, Bradon
Steffen, Philip J.
Schenk, Peter
Abstract
A starting apparatus for a gas turbine engine of a plurality of gas turbine engines of an aircraft. The apparatus includes a fuel supply system, a combustor, and a controller. The controller is configured to cause fuel to be introduced to the combustor of the gas turbine engine at a first threshold rotational speed of the gas-turbine engine during a normal starting operation in which the aircraft is on the ground. The controller is configured to cause fuel to be introduced to the combustor at a second threshold rotational speed of the gas-turbine engine during an emergency in-flight restarting operation in which the aircraft is in-flight. The second threshold rotational speed is lower than the first threshold rotational speed. Introducing fuel at the second threshold rotational speed results in a higher temperature in a turbine of the gas turbine engine than introducing fuel at the first threshold rotational speed.
Rolls-Royce North American Technologies, Inc. (USA)
Inventor
Lawrence, Jeffrey
Schenk, Peter
Schetzel, Ii, Douglas Keith
Harral, Jacob Ward
Huber, Brian Joseph
Abstract
A starting apparatus for a first gas turbine engine of a plurality of gas turbine engines of an aircraft. The apparatus includes an air turbine starter, an electric machine, and a controller. The controller is configured to receive an emergency restart command for the first gas turbine engine while the aircraft is in-flight, determine whether the first gas turbine engine is in operation, determine whether at least a second gas turbine engine of the plurality of gas turbine engines is in operation, and, responsive to receiving the emergency restart command and determining that at least the second gas turbine engine of the plurality of gas turbine engines is in operation and that the first gas turbine engine is not in operation, perform an emergency restart of the first gas turbine engine.
There is disclosed an energy storage system for an electric aircraft, the energy storage system comprising: at least one battery pack configured to be disposed onboard the aircraft, and a thermal management system. The thermal management system comprises a first circulation loop configured to be disposed onboard the aircraft and configured to contain a first working fluid, the first circulation loop including: a variable speed pump configured to pump the first working fluid around the first circulation loop, a battery heat exchanger configured to provide a thermal interface between the at least one battery pack and the first working fluid, and a controller configured to control operation of the variable speed pump to intermittently pump the first working fluid to distribute heat around the thermal management system.
B60L 58/26 - Methods or circuit arrangements for monitoring or controlling batteries or fuel cells, specially adapted for electric vehicles for monitoring or controlling batteries for controlling the temperature of batteries by cooling
An exhaust nozzle for a gas turbine engine comprises an exhaust duct and a first flap. The exhaust duct is configured to receive an exhaust flow of gas from a combustor of the gas turbine engine. The first flap is rotatably coupled to the exhaust duct for rotation about a first axis of rotation. Further, the first flap at least in part defines an exhaust gas passageway configured to convey the exhaust flow of gas to an exterior of the gas turbine engine. Additionally, the first flap comprises a first pin. The exhaust nozzle comprises a first moveable cam having a first moveable slot configured to slidably receive the first pin. The exhaust nozzle is configured such that movement of the first moveable cam causes the first flap to be moved about the first axis of rotation between a first inner position and a first outer position.
A gas turbine engine for an aircraft and a method of operating a gas turbine engine on an aircraft. Embodiments disclosed include a gas turbine engine for an aircraft including: an engine core has a turbine, a compressor, and a core shaft; a fan located upstream of the engine core, the fan has a plurality of fan blades; a nacelle surrounding the engine core and defining a bypass duct and bypass exhaust nozzle; and a gearbox that receives an input from the core shaft and outputs drive to the fan wherein the gas turbine engine is configured such that a jet velocity ratio of a first jet velocity exiting from the bypass exhaust nozzle to a second jet velocity exiting from an exhaust nozzle of the engine core at idle conditions is greater by a factor of 2 or more than the jet velocity ratio at maximum take-off conditions.
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 9/18 - Control of working fluid flow by bleeding, by-passing or acting on variable working fluid interconnections between turbines or compressors or their stages
A gas turbine engine (10) for an aircraft comprises an engine core (11) comprising a turbine (19), a compressor (14), a core shaft (26), and a core exhaust nozzle (20), the core exhaust nozzle (20) having a core exhaust nozzle pressure ratio calculated using total pressure at the core nozzle exit (56); a fan (23) comprising a plurality of fan blades; and a nacelle (21) surrounding the fan (23) and the engine core (11) and defining a bypass duct (22), the bypass duct (22) comprising a bypass exhaust nozzle (18), the bypass exhaust nozzle (18) having a bypass exhaust nozzle pressure ratio calculated using total pressure at the bypass nozzle exit;
A gas turbine engine (10) for an aircraft comprises an engine core (11) comprising a turbine (19), a compressor (14), a core shaft (26), and a core exhaust nozzle (20), the core exhaust nozzle (20) having a core exhaust nozzle pressure ratio calculated using total pressure at the core nozzle exit (56); a fan (23) comprising a plurality of fan blades; and a nacelle (21) surrounding the fan (23) and the engine core (11) and defining a bypass duct (22), the bypass duct (22) comprising a bypass exhaust nozzle (18), the bypass exhaust nozzle (18) having a bypass exhaust nozzle pressure ratio calculated using total pressure at the bypass nozzle exit;
wherein a bypass to core ratio of:
A gas turbine engine (10) for an aircraft comprises an engine core (11) comprising a turbine (19), a compressor (14), a core shaft (26), and a core exhaust nozzle (20), the core exhaust nozzle (20) having a core exhaust nozzle pressure ratio calculated using total pressure at the core nozzle exit (56); a fan (23) comprising a plurality of fan blades; and a nacelle (21) surrounding the fan (23) and the engine core (11) and defining a bypass duct (22), the bypass duct (22) comprising a bypass exhaust nozzle (18), the bypass exhaust nozzle (18) having a bypass exhaust nozzle pressure ratio calculated using total pressure at the bypass nozzle exit;
wherein a bypass to core ratio of:
bypass
exhaust
nozzle
pressure
ratio
core
exhaust
nozzle
pressure
ratio
A gas turbine engine (10) for an aircraft comprises an engine core (11) comprising a turbine (19), a compressor (14), a core shaft (26), and a core exhaust nozzle (20), the core exhaust nozzle (20) having a core exhaust nozzle pressure ratio calculated using total pressure at the core nozzle exit (56); a fan (23) comprising a plurality of fan blades; and a nacelle (21) surrounding the fan (23) and the engine core (11) and defining a bypass duct (22), the bypass duct (22) comprising a bypass exhaust nozzle (18), the bypass exhaust nozzle (18) having a bypass exhaust nozzle pressure ratio calculated using total pressure at the bypass nozzle exit;
wherein a bypass to core ratio of:
bypass
exhaust
nozzle
pressure
ratio
core
exhaust
nozzle
pressure
ratio
is configured to be in the range from 1.1 to 2.0 under aircraft cruise conditions.
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
F01D 17/16 - Final actuators arranged in stator parts varying effective cross-sectional area of nozzles or guide conduits by means of nozzle vanes
F02K 1/00 - Plants characterised by the form or arrangement of the jet pipe or nozzleJet pipes or nozzles peculiar thereto
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
53.
Adjustable fan track liner with dual slotted array active fan tip treatment for distortion tolerance
Rolls-Royce North American Technologies Inc. (USA)
Rolls-Royce Corporation (USA)
Inventor
Molnar, Jr., Daniel E.
Heeter, Robert W.
Rivers, Jonathan M.
Abstract
A gas turbine engine includes a fan and a fan case assembly. The fan includes a fan rotor configured to rotate about an axis of the gas turbine engine and a plurality of fan blades coupled to the fan rotor for rotation therewith. The fan case assembly extends circumferentially around the plurality of fan blades radially outward of the plurality of the fan blades.
Rolls-Royce North American Technologies Inc. (USA)
Rolls-Royce Corporation (USA)
Inventor
Heeter, Robert W.
Molnar, Jr., Daniel E.
Rivers, Jonathan M.
Abstract
A gas turbine engine includes a fan and a fan case assembly. The fan includes a fan rotor configured to rotate about an axis of the gas turbine engine and a plurality of fan blades coupled to the fan rotor for rotation therewith. The fan case assembly extends circumferentially around the plurality of fan blades radially outward of the plurality of the fan blades.
F04D 29/68 - Combating cavitation, whirls, noise, vibration, or the likeBalancing by influencing boundary layers
F01D 11/08 - Preventing or minimising internal leakage of working fluid, e.g. between stages for sealing space between rotor blade tips and stator
F01D 11/22 - Actively adjusting tip-clearance by mechanically actuating the stator or rotor components, e.g. moving shroud sections relative to the rotor
F04D 29/52 - CasingsConnections for working fluid for axial pumps
55.
METHOD AND FURNACE FOR TREATING AN IRON-COBALT COMPONENT
A method of treating an iron-cobalt component is disclosed. The method comprises: heat treating the component using a temperature of at least 700° C.; applying a static magnetic field of at least 3000 A/m to the component during the heat treatment. Also disclosed is a furnace for treating the iron-cobalt component, and a method of forming a stator for a transverse flux electric machine.
C21D 1/04 - General methods or devices for heat treatment, e.g. annealing, hardening, quenching or tempering with simultaneous application of supersonic waves, magnetic or electric fields
C21D 9/00 - Heat treatment, e.g. annealing, hardening, quenching or tempering, adapted for particular articlesFurnaces therefor
C21D 11/00 - Process control or regulation for heat treatments
H02K 15/02 - Processes or apparatus specially adapted for manufacturing, assembling, maintaining or repairing of dynamo-electric machines of stator or rotor bodies
56.
SOLID STATE POWER CONTROLLER, POWER MANAGEMENT SYSTEM AND POWER CONVERTER
A solid state power controller includes at least one semiconductor switch with a control terminal, a controller for generating a pulsed signal, and a gate driver circuit for receiving the pulsed signal and generating a pulsed driver signal. The gate driver circuit includes a gate driver for receiving the pulsed signal and providing a pulsed output signal, and driver signal generating device operable in a first state and a second state. The driver signal generating device receives the gate driver pulsed output signal and generates in the first state a first pulsed driver signal that operates a semiconductor switch in an active region and generates in the second state a second pulsed driver signal that operates the semiconductor switch in a saturated region. The controller is configured to set the driver signal generating device in the first or second state depending on a voltage level signal received by the controller.
A power protection system includes a first DC power source, a first load, a first power bus connecting the first power source and the first load, and a first solid state circuit breaker circuit integrated in the first power bus. The first solid state circuit breaker circuit includes a first semiconductor switch, a first capacitor arranged between the high side voltage rail and the low side voltage rail closer to the power source or to the load than the semiconductor switch, and a first inductor located such that current generated by the first capacitor when unloading in case of a short circuit at the power source or at the load passes the first inductor. The controller is configured to measure a voltage change over the first inductor and trigger an actuation signal for the first semiconductor switch if a voltage change surpasses a predetermined threshold voltage.
H02H 3/26 - Emergency protective circuit arrangements for automatic disconnection directly responsive to an undesired change from normal electric working condition, with or without subsequent reconnection responsive to difference between voltages or between currentsEmergency protective circuit arrangements for automatic disconnection directly responsive to an undesired change from normal electric working condition, with or without subsequent reconnection responsive to phase angle between voltages or between currents
H02H 1/00 - Details of emergency protective circuit arrangements
AC-DC converter circuits which may include: first and second rectifier circuits each having a plurality of input connections for connection to respective first and second sets of windings of a generator, each input connection connected between a pair of series-connected rectifier diodes connected between first and second output terminals, an output capacitor connected between the first and second output terminals; a first output diode connected between the second output terminals of the first and second rectifier circuits; and a first output switch connected between the second output terminal of the first rectifier circuit and the first output terminal of the second rectifier circuit.
H02M 7/06 - Conversion of AC power input into DC power output without possibility of reversal by static converters using discharge tubes without control electrode or semiconductor devices without control electrode
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
59.
METHOD OF MANUFACTURING A COMPONENT OF A GAS TURBINE ENGINE
There is provided a method of manufacturing a component of a gas turbine engine, comprising: providing a precursor component having at least one internal cooling passage configured to receive a flow of cooling air therethrough; estimating a predicted temperature profile for the component based on one or more operating parameters of the gas turbine engine, the predicted temperature profile indicating a predicted operating temperature of at least one gas-washed surface of the component; determining a thermal barrier coating (TBC) configuration for the component based on the predicted temperature profile, comprising setting a TBC thickness to be below a threshold thickness in a region of the at least one gas-washed surface of the component based on the predicted operating temperature of the at least one gas-washed surface of the component exceeding a threshold temperature; and applying a TBC to the precursor component according to the TBC configuration.
A fluid control valve is disclosed, comprising: a valve body defining an outer wall for a curved flow path through the fluid control valve; and a valve element defining an opposing inner wall and rotatable relative to the valve body about a pivot point through between a closed position a range of open positions. A separation between the first inner wall and the outer wall along a restrictor portion of the flow path varies between open positions. The first inner wall and the outer wall are cooperatively defined so that for at least some open positions the separation between the first inner wall and the outer wall is constant along the respective restrictor portion of the flow path.
F16K 1/24 - Lift valves, i.e. cut-off apparatus with closure members having at least a component of their opening and closing motion perpendicular to the closing faces with valve members that, on opening of the valve, are initially lifted from the seat and next are turned around an axis parallel to the seat
A measurement apparatus for measuring a flow rate of a powder includes a casing, a nozzle configured to dispense the powder, a fixture plate, a weighing scale, and a powder collector. The fixture plate includes a plurality of pinhole members. Each pinhole member includes a tip, a cylindrical hole extending from the tip, and a discharge passage. The cylindrical hole of each pinhole member has a diameter. The diameters of the cylindrical holes of the plurality of pinhole members are different from each other. The nozzle is configured to dispense the powder selectively into the cylindrical hole, and the powder collector is configured to receive at least a portion of the powder from the discharge passage.
The invention relates to a method for producing a coil device (1), in particular a multi-layer coil device for an electric machine (10), wherein a) at least one electrically non-conductive fastening element (2) combined with an adhesive in the form of a wrappable tape is placed in and/or on a coil body (3) having Litz wires (5), b) at least one electrical insulating means (4) is placed on the outside of the coil body (3), in particular also on the fastening element (2), and c) the entirety of fastening element (2), coil body (3) and electrical insulating means (4) is integrally joined together as a result of pressure action (P) and/or thermal action (T), by adhesive bonding or melting of the adhesive, so as to form the coil device (1). The invention also relates to a coil device (1) produced by means of said method and to an electric machine (10) having corresponding coil devices (1).
H02K 3/34 - Windings characterised by the shape, form or construction of the insulation between conductors or between conductor and core, e.g. slot insulation
H02K 15/04 - Processes or apparatus specially adapted for manufacturing, assembling, maintaining or repairing of dynamo-electric machines of windings prior to their mounting into the machines
H02K 15/10 - Applying solid insulation to windings, stators or rotors, e.g. applying insulating tapes
H02K 15/12 - Impregnating, moulding insulation, heating or drying of windings, stators, rotors or machines
A fuel system for a gas turbine engine. The fuel system comprises a fuel offtake configured to divert a portion of hydrogen fuel from a main fuel conduit, a burner configured to burn the portion of hydrogen fuel diverted from the main fuel conduit and at least first and second heat exchangers. The first heat exchanger is configured to transfer heat from exhaust gasses produced by the burner to hydrogen fuel in the main fuel conduit and the second heat exchanger is provided upstream in hydrogen flow of the first heat exchanger and is configured to transfer heat from a further heat exchange fluid to hydrogen fuel in the main fuel conduit. In an embodiment, the further heat exchange fluid is compressor bleed air bled from a core compressor of the gas turbine engine.
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 system for inserting a flexible robotic arm into a workspace, the system having: a feed mechanism having a housing containing a drive mechanism, the drive mechanism gripping the flexible robotic arm using at least one drive wheel which is coupled to a drive motor, the feed mechanism is provided with at least one force sensor; and a haptic controller coupled to the feed mechanism, the haptic controller having input controller for instructing the drive motor within the feed mechanism to activate in response to the signal from the input controller, and wherein the input controller is provided with haptic feedback to provide an operator an indication of the state of operation of the feed mechanism based on signals output from signals provided by the force sensors.
A system for inspecting a complex component, the system comprising: a flexible inspection device having a driver for controlling the motion of the flexible inspection device by manipulating the relative positions of a plurality of joints within the device, the flexible device having a sensor on its distal end; a feed mechanism for controlling the insertion or retraction of the flexible device into or from the complex component; a component driver for driving the movement of an aspect of the complex component from at least a first position to a second position; and a computer running a computer program, the computer interfacing with the flexible inspection device, the feed mechanism, and the component driver, so as to link the operation of the flexible inspection device, the feed mechanism and the component driver through a single program.
A feed mechanism for inserting and/or retracting a flexible robotic arm from an area, the feed mechanism including a housing, a passageway extending about a central longitudinal axis through the length thereof, a drive portion and rotational portion, the drive portion not being fixedly connected within the housing and having at least a pair of drive wheels coupled to a drive motor, the drive wheels are connected to a mounting frame with the mounting frame being connected to pivotable linkages with at least one of the linkages being connected to a spring, such that the expansion or contraction of the spring allows the drive wheels to move relative to the longitudinal axis, and the rotational portion being coupled to a motor, the rotational motor having a gear system that connects to the drive portion and causes a rotation of the drive portion about the central longitudinal axis within the housing.
A method of operating an aircraft gas turbine engine including: a core including a turbine, a compressor, a combustor to combust fuel, and a core shaft connecting the turbine to the compressor; a fan upstream of the core; a fan shaft; a bearing supporting the fan shaft; an oil loop system supplying oil to the bearing; and a heat exchange system including: an air-oil heat exchanger through which oil in the oil loop system flows; and a fuel-oil heat exchanger through which oil in the oil loop system and fuel flow to transfer heat between the oil and the fuel; and a bypass pipe to allow a proportion of the oil to flow past one of the air-oil and the fuel-oil heat exchanger; and a bypass valve to allow the proportion of the oil sent through the bypass pipe to be varied.
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
An annulus filler for mounting to a rotor disc of a gas turbine engine includes a coupling portion and an outer lid. The outer lid includes a leading edge, a trailing edge, a pair of longitudinal edges, and an outer radial surface. Each longitudinal edge includes a first edge portion and a second edge portion. At least one longitudinal edge includes a recessed portion that extends at least axially between the first edge portion and the second edge portion, such that the recessed portion connects the first edge portion to the second edge portion. The recessed portion further extends circumferentially inwards from each of the first edge portion and the second edge portion towards the other longitudinal edge.
A gas turbine engine component that has a web provided with an array of cooling holes distributed with respect to a first direction, wherein the cooling holes have a cross-sectional shape that varies along the first direction. The component may be configured so that the first direction corresponds to a loading distribution of the component which increases along the first direction from a low loading position to a high loading position.
A propulsion device comprises: a casing structure to surround a fan of the propulsion device; a core structure to support a core of the propulsion device; a thrust reverser unit comprising two thrust reverser halves, each thrust reverser half being pivotable about a respective hinge line between an open position for access to the core structure, and a closed position. Upper and lower support members extend from the casing structure at diametrically opposing sides of a centreline axis. For each thrust reverser half there is at least one locating arrangement comprising an upper locating arrangement with cooperating portions configured to engage each other as the thrust reverser half moves towards the closed position in a closing operation; and/or a lower locating arrangement with cooperation portions configured to engage each other as the thrust reverser half moves towards the closed position in the closing operation.
A gas turbine engine includes a staged combustion system having pilot fuel injectors and main fuel injectors. A fuel delivery regulator controls delivery of fuel to the pilot and main fuel injectors, receives fuel from a first fuel source containing a first fuel having a first fuel characteristic and a second fuel source containing a second fuel having a second fuel characteristic. In a transition range of operation between the pilot-only and the pilot-and-main ranges of operation, fuel is delivered to both the pilot and main fuel injectors at a transition staging ratio different from the pilot-and-main staging ratio. The fuel delivery regulator delivers fuel to one or both the pilot and main fuel injectors during the transition range of operation having a different fuel characteristic from fuel delivered to one or both the pilot and main fuel injectors during at least part of the pilot-and-main range of operation.
Rolls-Royce High Temperature Composites Inc. (USA)
Inventor
Thomas, David J.
Downie, Christopher
Sippel, Aaron D.
Freeman, Ted J.
Snyder, Clark
Abstract
A turbine shroud assembly adapted for use with a gas turbine engine includes a shroud segment. The shroud segment includes a heat shield, an attachment flange, and a multi-layer coating. The heat shield extends circumferentially partway around the axis to define a portion of gas path for the gas turbine engine. The attachment feature extends radially outward from the heat shield. The multi-layer coating is applied to different surfaces of the heat shield and the attachment feature of the shroud segment.
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
F01D 25/00 - Component parts, details, or accessories, not provided for in, or of interest apart from, other groups
F01D 25/24 - CasingsCasing parts, e.g. diaphragms, casing fastenings
F01D 5/28 - Selecting particular materialsMeasures against erosion or corrosion
74.
Electric starter verification during gas-turbine engine barring
Rolls-Royce North American Technologies, Inc. (USA)
Inventor
Lawrence, Jeffrey
Huber, Brian Joseph
Munevar, Erik A.
Schetzel, Ii, Douglas Keith
Abstract
An example system includes a first gas-turbine engine configured to propel an aircraft, the first gas-turbine engine comprising a first electric starter, the first electric starter configured to rotate a spool of the first gas-turbine engine; and one or more controllers collectively configured to: cause, following operation of the first gas-turbine engine, the first electric starter to perform barring of the first gas-turbine engine; measure, during the barring of the first gas-turbine engine, values of one or more parameters of the first gas-turbine engine; and determine, based on the values of the one or more parameters, whether the first electric starter is available for use in performing mid-air restart of the first gas-turbine engine.
An electrical system (1A-1C) for an electric machine (2) comprises: at least three connections (A1-A3) for electrical connection to in each case one phase of an at least three-phase AC voltage; a first star circuit (10A-10C) with three strands (100A-100C) via each of which one of the at least three connections (A1-A3) is electrically connected to a common star point (101) of the first star circuit (10A-10C) and which each have at least (11A-11C) with three strands (110A-110C) via each of which one of the at least three connections (A1-A3) is electrically connected to a common star point (111) of the second star circuit (11A-11C) and which each have at least two tooth windings (Z) which are connected in series; and a measuring device (12), which is electrically connected between the first star circuit (10A-10C) and the second star circuit (11A-11C), for measuring an electrical variable.
H02P 25/16 - Arrangements or methods for the control of AC motors characterised by the kind of AC motor or by structural details characterised by the circuit arrangement or by the kind of wiring
B64D 27/32 - Aircraft characterised by electric power plants within, or attached to, fuselages
76.
METHOD OF MANUFACTURING A COMPOSITE ARTICLE PRECURSOR
A method of manufacturing a composite article precursor, e.g. from which to make an aerofoil. The method of manufacturing involves: laying down a first grouping of at least one layer of composite material upon a substantially upwardly-facing first surface of a preform of the composite article precursor; disposing a first deformable support over the first grouping of at least one layer of composite material; securing the first deformable support to the preform via a first securing fixing; rotating the preform about an axis such the first surface is no longer substantially upwardly-facing and a second surface of the preform is substantially upwardly-facing; laying down a second grouping of at least one layer of composite material upon the second surface of the preform; and vacuum debulking an assembly comprising the first grouping composite material, the second grouping of composite material, the first deformable support and the preform.
B29C 70/54 - Component parts, details or accessoriesAuxiliary operations
B29C 70/44 - Shaping or impregnating by compression for producing articles of definite length, i.e. discrete articles using isostatic pressure, e.g. pressure difference-moulding, vacuum bag-moulding, autoclave-moulding or expanding rubber-moulding
B29L 31/08 - Blades for rotors, stators, fans, turbines or the like, e.g. screw propellers
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.
F02C 7/30 - Preventing corrosion in gas-swept spaces
B08B 3/00 - Cleaning by methods involving the use or presence of liquid or steam
B08B 7/02 - Cleaning by methods not provided for in a single other subclass or a single group in this subclass by distortion, beating, or vibration of the surface to be cleaned
B08B 13/00 - Accessories or details of general applicability for machines or apparatus for cleaning
78.
SUPER-COOLED ICE IMPACT PROTECTION FOR A GAS TURBINE ENGINE
A gas turbine engine comprises a fan mounted to rotate about a main longitudinal axis; an engine core, comprising in axial flow series a compressor, a combustor, and a turbine coupled to the compressor through a shaft; a reduction gearbox that receives an input from the shaft and outputs drive to the fan so as to drive the fan at a lower rotational speed than the shaft; wherein the compressor comprises a first stage at an inlet and a second stage, downstream of the first stage, comprising respectively a first rotor with a row of first blades and a second rotor with a row of second blades, the first and second blades comprising respective leading edges, trailing edges and tips, and wherein the ratio of a maximum leading edge radius of the first blades to a maximum leading edge radius of the second blades is greater than 2.8.
F02C 7/05 - Air intakes for gas-turbine plants or jet-propulsion plants having provisions for obviating the penetration of damaging objects or particles
F02C 7/36 - Power transmission between the different shafts of the gas-turbine plant, or between the gas-turbine plant and the power user
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
80.
HIGH POWER EPICYCLIC GEARBOX AND OPERATION THEREOF
A gas turbine engine for an aircraft including: an engine core including a turbine, a compressor, and a core shaft connecting the turbine to the compressor; a fan located upstream of the engine core, the fan including a plurality of fan blades; a gearbox that can receive an input from the core shaft, and can 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; and a fan shaft mounting structure arranged to mount the fan shaft within the engine, the fan shaft mounting structure including at least two supporting bearings connected to the fan shaft.
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
There is provided a method of operating a gas turbine engine including a combustor. The combustor includes a combustion chamber and a plurality of fuel spray nozzles configured to inject fuel into the chamber. The nozzles include a first and second subset of fuel spray nozzles. The combustor is operable in a condition wherein the first subset is supplied with more fuel than the second. A ratio of the number of nozzles in the first subset to the second is 1:2 to 1:5. The method includes operating the engine so a reduction of 20-80% in an average of particles/kg of nvPM in the exhaust when the engine is operating at 85% available thrust for given operating conditions and when the engine is operating at 30% is obtained when fuel provided to the nozzles is a sustainable aviation fuel instead of a fossil-based hydrocarbon fuel. Also provided is a gas turbine engine for an aircraft.
F23R 3/34 - Feeding into different combustion zones
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
F02C 9/40 - Control of fuel supply specially adapted to the use of a special fuel or a plurality of fuels
F23R 3/28 - Continuous combustion chambers using liquid or gaseous fuel characterised by the fuel supply
An electrical system includes: a battery; a fuel-cell pack; a load electrically coupled to the fuel-cell pack; a switching arrangement electrically coupled to the battery, the fuel-cell pack and the load; a DC-DC converter; and a control system. The switching arrangement configures the electrical system in at least one of a battery-charge mode and a combined-drive mode. In the battery-charge mode the battery is coupled in series to the fuel-cell pack and the load via the DC-DC converter for simultaneous charging of the battery and driving of the load by the fuel-cell pack. In the combined-drive mode, the battery is coupled in series to the fuel-cell pack and the load via the DC-DC converter for driving of the load by both the battery and the fuel-cell pack. The control system is configured to: monitor a parameter of an electrical power provided to the load; and control the DC-DC converter.
An electrical system includes: a battery; a fuel-cell pack; a load; a switching arrangement; and a control system. The switching arrangement selectively configures the electrical system in a battery-isolation mode and at least one of a battery-charge mode and a combined-drive mode. In the battery-isolation mode, the battery is decoupled from the fuel-cell pack and the load, and the fuel-cell pack is coupled to the load for driving of the load by the fuel-cell pack. In the battery-charge mode, the battery is coupled in series to the fuel-cell pack and the load for simultaneous charging of the battery and driving of the load by the fuel-cell pack. In the combined-drive mode, the battery is coupled in series to the fuel-cell pack and the load for driving of the load by both the battery and the fuel-cell pack.
The disclosure relates to fault protection in a DC-DC electric power converter. The converter comprises first, second, third and fourth switches connected either side of an inductor. Fifth and sixth switches are connected between respective input and output terminals and a common line, the fifth and sixth switches providing protection in a fault event.
H02M 3/158 - Conversion of DC power input into DC power output without intermediate conversion into AC 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 with automatic control of output voltage or current, e.g. switching regulators including plural semiconductor devices as final control devices for a single load
This invention concerns a fuel delivery system for an aircraft engine, comprising a fuel delivery regulator configured to receive fuel from a plurality of fuel sources for supply to the engine. An engine operating condition sensor reading is received by a control unit configured to control operation of the regulator. The control unit is configured to actuate the regulator based on a received signal from the engine operating condition sensor in order to vary the volume of fuels from the plurality of fuel sources supplied to the engine during at least a portion of a cruise phase relative to a further portion of an aircraft flight. The different fuels may comprise kerosene and an alternative fuel, such as a biofuel.
An apparatus for determining the health of a battery including a plurality N of series-connected cells. The apparatus includes an AC signal source operable to cause an AC current to flow through one or more cells of the plurality of series-connected cells, a current measurement unit operable to measure the AC current, a voltage measurement unit operable to measure an AC voltage induced across the one or more cells by the AC current, and a health determination unit operable to determine a measurement of the health of the battery based on the measured AC current and the measured AC voltage.
G01R 31/392 - Determining battery ageing or deterioration, e.g. state of health
G01R 19/02 - Measuring effective values, i.e. root-mean-square values
G01R 31/36 - Arrangements for testing, measuring or monitoring the electrical condition of accumulators or electric batteries, e.g. capacity or state of charge [SoC]
G01R 31/385 - Arrangements for measuring battery or accumulator variables
G01R 31/389 - Measuring internal impedance, internal conductance or related variables
A ceramic matrix composite (CMC) component is provided that includes: a CMC body in which an environmental protection layer is completely embedded within a CMC material of the CMC body, the environmental protection layer comprising a ceramic that has a higher impact and/or environmental resistance than the CMC material. Methods for manufacturing the CMC component are also provided.
B32B 3/08 - Layered products comprising a layer with external or internal discontinuities or unevennesses, or a layer of non-planar shapeLayered products comprising a layer having particular features of form characterised by features of form at particular places, e.g. in edge regions characterised by added members at particular parts
B32B 5/06 - Layered products characterised by the non-homogeneity or physical structure of a layer characterised by structural features of a layer comprising fibres or filaments characterised by a fibrous layer needled to another layer, e.g. of fibres, of paper
A fixture for removably securing a plurality of components includes a hollow body including a plurality of channels and a plurality of sets of holes. The fixture further includes a plurality of hollow wedges and a plurality of sets of balls. Upon an axial movement of an elongate member in a first direction, the elongate member axially moves a first hollow wedge, thereby causing one or more elastic members to be sequentially compressed and move the subsequent hollow wedges. Upon the axial movement, a frustoconical surface portion of each hollow wedge moves the corresponding set of balls radially outwards within the corresponding set of holes, such that each ball extends partially into the corresponding channel, thereby moving the corresponding component into locking engagement with the corresponding channel.
F16B 21/16 - Means without screw-thread for preventing relative axial movement of a pin, spigot, shaft, or the like and a member surrounding itStud-and-socket releasable fastenings without screw-thread by separate parts with grooves or notches in the pin or shaft
G01M 15/14 - Testing gas-turbine engines or jet-propulsion engines
A power routing circuit having first and second input nodes VIN1, VIN2 for connection to first and second power supplies and output node VOUT1 for load connection; and includes a first main switching circuit connected between the first input node VIN1 and output node VOUT1 and conductively connects the first input node VIN1 to the output node VOUT1 if a voltage at the first input node VIN1 is higher than a voltage at the output node VOUT1, and a first auxiliary switching circuit connected between the second input node VIN2 and output node VOUT1 and conductively connects the second input node VIN2 to output node VOUT1 if a voltage at the second input node VIN2 is higher than the voltage at the output node VOUT1 and the voltage at the first input node VIN1 is at least a threshold amount lower than the voltage at the second input node VIN2.
Rolls-Royce North American Technologies Inc. (USA)
Rolls-Royce Corporation (USA)
Inventor
Heeter, Robert W.
Molnar, Jr., Daniel E.
Rivers, Jonathan M.
Abstract
A gas turbine engine includes a fan and a fan case assembly. The fan includes a fan rotor configured to rotate about an axis of the gas turbine engine and a plurality of fan blades coupled to the fan rotor for rotation therewith. The fan case assembly extends circumferentially around the plurality of fan blades radially outward of the plurality of the fan blades.
F04D 29/52 - CasingsConnections for working fluid for axial pumps
F01D 11/22 - Actively adjusting tip-clearance by mechanically actuating the stator or rotor components, e.g. moving shroud sections relative to the rotor
The disclosure relates to a heat exchanger for a cooling system, the heat exchanger comprising: a first part having a first surface for connecting to a component to be cooled and a second opposing surface, the first part having a fluid flow channel extending from a fluid inlet through the first part between the first and second surfaces; a second part extending from the second surface to an external third surface, the second part having an open-cell porous structure in fluid communication with the fluid flow channel such that fluid flowing through the fluid flow channel passes into the second part and exits the heat exchanger at the external third surface.
The disclosure relates to power electronics converters having circuit breaker protection for use in case of faults. Example embodiments include a power electronics converter for converting an input AC supply having a plurality of phases to an output DC supply, the converter comprising: a plurality of AC input terminals connectable to the plurality of phases of the AC supply; first and second DC output terminals; a capacitor connected between the first and second DC output terminals; a first MOSFET connected between each of the plurality of AC input terminals and the first DC output terminal; a second MOSFET connected between each of the plurality of AC input terminals and the second DC output terminal; and a third MOSFET reverse connected in series with the first MOSFET between the first DC output terminal and each of the plurality of AC input terminals.
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
H02M 1/32 - Means for protecting converters other than by automatic disconnection
A combustor assembly for a gas turbine engine includes a combustor shell extending along a shell axis. The combustor shell includes a shell wall extending around the shell axis and a meter panel connected to an upstream end of the shell wall. The shell wall and the meter panel together define a combustion space therebetween. The combustor assembly further includes: a tile disposed within the combustor shell and including a tile flange portion extending radially with respect to the shell axis; a heatshield connected to the meter panel and extending at least radially towards the tile; a fastener connecting the tile flange portion to the meter panel; and a compartment thermally shielded from the combustion space. Either the heatshield and the tile together form the compartment or the tile alone forms the compartment. The fastener is spaced apart from the combustion space and at least partially disposed within the compartment.
A resin composition that is useful in the preparation of protective coatings. The resin composition comprises a polyfunctional cyanate ester, a phenol-end-modified PDMS oligomer, and an imidazolium dicyanamide catalyst. The method of preparing a resin blend by mixing a polyfunctional cyanate ester, a phenol-end-modified PDMS oligomer and an imidazolium dicyanamide catalyst to form a resin composition and curing the resin composition to form the resin blend.
C08L 65/00 - Compositions of macromolecular compounds obtained by reactions forming a carbon-to-carbon link in the main chainCompositions of derivatives of such polymers
C09D 165/00 - Coating compositions based on macromolecular compounds obtained by reactions forming a carbon-to-carbon link in the main chainCoating compositions based on derivatives of such polymers
A bearing holding device with an annular inner region and a flange region, extending at least approximately outwards in the radial direction, with recesses, which flange region is designed in one piece with the inner region and is narrower in the axial direction than the annular inner region, is described. The flange region is, in the circumferential region, surrounded at least in certain regions by a connection region, which is formed in one piece with the flange region and via which the inner region and the flange region can be firmly connected to a housing. Furthermore, a rolling bearing apparatus with the bearing holding device is proposed.
F16C 35/04 - Rigid support of bearing unitsHousings, e.g. caps, covers in the case of ball or roller bearings
F16C 19/06 - Bearings with rolling contact, for exclusively rotary movement with bearing balls essentially of the same size in one or more circular rows for radial load mainly with a single row of balls
F16C 35/077 - Fixing them on the shaft or housing with interposition of an element between housing and outer race ring
A method of operating a gas turbine engine and a gas turbine engine includes a fuel delivery system arranged to provide fuel, a combustor arranged to combust at least a proportion of the fuel, a primary fuel-oil heat exchanger arranged to have up to 100% of the fuel provided by the fuel delivery system flow therethrough, and a secondary fuel-oil heat exchanger arranged to have a proportion of the fuel from the primary fuel-oil heat exchanger flow therethrough. Fuel is arranged to flow from the primary fuel-oil heat exchanger to the secondary fuel-oil heat exchanger whereas oil is arranged to flow from the secondary fuel-oil heat exchanger to the primary fuel-oil heat exchanger. A fuel viscosity is adjusted to a maximum of 0.58 mm2/s on entry to the combustor at cruise conditions.
A method of manufacturing a component includes forming a mould assembly including an initial mould unit, providing a seed crystal including a primary growth direction, determining an optimal angular orientation of the unit, rotating the unit to dispose the unit's optimal angular orientation, encasing the unit in a refractory material, and forming a refractory mould unit having a component mould including a mould wall defining a mould cavity, and a seed holder. In the optimal angular orientation, the seed crystal's primary growth direction is angled away from the wall, thereby forming a converging disposition with the wall in a of the wall's first region facing the central sprue and a diverging disposition with the wall in the wall's second region facing a mould heater. The method includes receiving the seed crystal within the seed holder and filling the mould cavity with molten castable material to form the component.
A bearing holding device with an annular inner region and a flange region, extending at least approximately outwards in the radial direction, with recesses, which flange region is designed in one piece with the inner region and is narrower in the axial direction than the annular inner region, is described. The flange region is, in the circumferential region, surrounded at least in certain regions by a connection region, which is formed in one piece with the flange region and via which the inner region and the flange region can be firmly connected to a housing. Furthermore, a rolling bearing apparatus with the bearing holding device is proposed.
A circuit board includes a plurality of circuit board layers arranged one on top of the other, wherein through-holes are integrated in the circuit board, wherein the through-holes are configured to receive metal screws that screw the circuit board to a heat sink. Insulating sleeves made of an insulating material are integrated into the circuit board, wherein the through-holes are formed in the circuit board in the region of the insulating sleeves.
The invention relates to a method for controlling a multi-phase electrical machine (2) comprising a stator (20) and a rotor (21), comprising the steps: identifying (S1) a short-circuit (K) in a turn of a winding (201) of the electrical machine (2) and at least one parameter describing the short-circuit (K) in the turn; setting (S2) an angle (η) of a current vector to the positive q-axis in the dq coordinate system of the rotor (21), having a negative d portion and a q portion which does not equal zero, on the basis of the at least one parameter describing the short-circuit (K) in the turn; and controlling (S3) the electrical machine (2) according to the set angle (η).
