RAMPED STIFFENER AND APPARATUS AND METHOD FOR FORMING THE SAME

A die tool for forming a C-section component having radiused shoulders, the includes a cylindrical inner die and an outer die having a cylindrical central portion connected to opposing end flanges by respective radiused concave portions. A portion of the inner die is arranged to be disposed between the end flanges of the outer die and spaced apart therefrom to define a cavity corresponding to the desired cross-section of the C-section component to be formed. The radiused convex and concave portions have a radius of curvature that varies about the circumference of the respective inner and outer dies. The inner and outer dies are rotatable such that the radius of curvature of the radiused convex and concave portions where the inner and outer dies are adjacent to one another varies as the inner and outer dies are rotated. 18-. (canceled)9. A die tool for forming a top-hat section stiffener , the die tool comprising:a outer die having at least two axially aligned first cylindrical body sections, each first body section being connected to an opposing end face by a radiused shoulder, and at least one second cylindrical body section connected between the opposing end faces of adjacent first body sections, the diameter of the first cylindrical body sections being greater than the diameter of the second cylindrical body section; andan inner die having at least one third cylindrical body section having opposing end faces, each end face being connected to a respective axially aligned fourth cylindrical body section by a radiused concave portion, the diameter of each third cylindrical body section being greater than the diameter of the fourth cylindrical body sections, wherein:the inner and outer dies are disposed relative to one another such that each third body section of the inner die is disposed between adjacent first body sections of the outer die, the inner and outer dies being spaced apart from one another, characterised in that:the radiused shoulders of the outer die ...

ASSEMBLY OF A SERVO PUMP AND A HYDRAULIC MOTOR

An assembly of a servo pump and a hydraulic motor. The assembly has a housing which contains the pump and the motor. The motor has a rotating body which rotates under the motive power of a pressurised motor liquid flow. The motor has a high pressure region which receives the pressurised motor liquid flow, and a low pressure region through which the motor liquid flow leaves the motor. The pump also has a rotating body. The pump has a low pressure region which receives servo liquid flow to be pumped by its rotating body, and a high pressure region through which pressurised servo liquid flow leaves the pump. Each rotating body is mounted on a respective journal. 1. An assembly of a servo pump and a hydraulic motor , wherein:the assembly has a housing which contains the servo pump and the hydraulic motor;{'sub': min', 'mout, 'the hydraulic motor has a rotating body which rotates under the motive power of a pressurised motor liquid flow, whereby the hydraulic motor has a high pressure region (P) which receives the pressurised motor liquid flow, and a low pressure region (P) through which the motor liquid flow leaves the hydraulic motor;'}{'sub': pin', 'pout, 'the servo pump has a rotating body, whereby the servo pump has a low pressure region (P) which receives servo liquid flow to be pumped by its rotating body, and a high pressure region (P) through which pressurised servo liquid flow leaves the servo pump;'}each rotating body is mounted on a respective journal, each journal being supported on one side of its rotating body by a respective first bearing block and being supported on the opposite side of its rotating body by a respective second bearing block, each bearing block having a running face adjacent to its rotating body and a non-running face distal from its rotating body;the assembly has a mechanical linkage having a coupling at one end to the journal of the motor rotating body and a coupling at an opposite end to the journal of the pump rotating body such that ...

Gear pump bearing

The invention provides a gear pump bearing block and method of manufacturing gear pump bearing block. Bearing block includes bush formed of antifriction alloy, having a cylindrical portion providing bore adapted to receive bearing shaft of a gear of pump, and further having a flange portion extending radially outwardly at end of cylindrical portion to provide thrust face adapted to slidingly engage with side surface of gear. Bearing block also has backing layer covering radially outer surface of cylindrical portion and rear face of flange portion, backing layer being formed of relatively less dense alloy compared to antifriction alloy. Furthermore, there is an annular formation of one or more stiffening members, formation surrounding bore, the, or each, stiffening member being embedded in flange portion, and the, or each, stiffening member being formed of material having higher elastic modulus than antifriction alloy.

GEAR PUMP BEARING

Invention provides gear pump bearing block and method of manufacturing gear pump bearing block. Bearing block includes a bush formed of antifriction alloy, bush having cylindrical portion providing bore adapted to receive bearing shaft of gear of pump, and further having flange portion extending radially outwardly at end of cylindrical portion to provide thrust face adapted to slidingly engage with side surface of gear. Bearing block also has backing layer covering radially outer surface of cylindrical portion and rear face of flange portion, backing layer being formed of less dense alloy compared to antifriction alloy. Furthermore, there is an annular gallery embedded in flange portion such that gallery is spaced from surface of bore and from thrust face, gallery surrounding bore, with inlet to and outlet from annular gallery such that, in use, fluid flows from inlet, through annular gallery to provide cooling of thrust face, and then to outlet. 1. A gear pump bearing block having:a bush formed of antifriction alloy, the bush having a cylindrical portion providing a bore adapted to receive a bearing shaft of a gear of the pump, and further having a flange portion extending radially outwardly at an end of the cylindrical portion to provide a thrust face adapted to slidingly engage with a side surface of the gear;a backing layer covering a radially outer surface of the cylindrical portion and a rear face of the flange portion, the backing layer being formed of relatively less dense alloy compared to the antifriction alloy;an annular gallery embedded in the flange portion such that the gallery is spaced from the surface of the bore and from the thrust face, the gallery surrounding the bore; andan inlet to and an outlet from the annular gallery such that, in use, fluid flows from the inlet, through the annular gallery to provide cooling of the thrust face, and then to the outlet.2. A gear pump bearing block according to claim 1 , wherein the gear pump bearing block ...

GEAR PUMP BEARING

The present invention provides a gear pump having one or more gears with bearing shafts supported by respective pressure-loaded gear pump bearing blocks. Each pressure-loaded bearing block has: a bore adapted to receive the bearing shaft of a gear of the pump; a thrust face at one end adapted to slidingly engage with a side surface of the gear, and an opposing rear face at the other end, the pump being adapted such that, in use, pressurised fluid is supplied to the rear face to provide a hydraulic load urging the thrust face of the bearing block towards the side surface of the gear; and a seal carried by the rear face, the seal, in use, partitioning higher fluid pressure and lower fluid pressure portions of the rear face. The gear pump further has a pressure regulating valve arranged such that, in use, the fluid pressure difference between the higher and lower fluid pressure portions of the rear face of the bearing block is adjustable by the valve to vary said hydraulic load. 1. A gear pump having at least one gear stage including a pair of meshing gears with bearing shafts of the meshing gears supported by respective pressure-loaded gear pump bearing blocks , each pressure-loaded bearing block having:a bore adapted to receive the bearing shaft of a gear of the pump;a thrust face at one end adapted to slidingly engage with a side surface of the gear, and an opposing rear face at the other end, the pump being adapted such that, in use, pressurised fluid is supplied to the rear face to provide a hydraulic load urging the thrust face of the bearing block towards the side surface of the gear; anda seal carried by the rear face, the seal, in use, partitioning higher fluid pressure and lower fluid pressure portions of the rear face;wherein the gear pump is configured to port fluid at an inlet pressure of the gear stage to the lower fluid pressure portion of the rear face of each bearing block; andwherein the gear pump further has a pressure regulating valve arranged such ...

PUMP

A pump has a brushless electrical motor including a stator assembly and a rotor ring arranged externally of the stator assembly. The pump further has a centrifugal impeller operatively connected to the rotor ring such that the electrical motor rotates the impeller. The impeller is configured to raise the pressure of a fluid which flows through the impeller between an inlet and an outlet thereof. The pump is configured such that a portion of the raised pressure fluid flow is diverted to form a recirculation flow which circulates through and cools the electrical motor before being returned to the impeller for re-pressurisation thereby. 1. A fuel pumping unit for an aero-engine , the unit having a low pressure pump and a high pressure pump , the low pressure pump supplying fuel at a raised pressure to the high pressure pump for onward supply to a fuel metering unit of the engine , wherein the low pressure pump has:a brushless electrical motor including a stator assembly and a rotor ring arranged externally of the stator assembly; anda centrifugal impeller operatively connected to the rotor ring such that the electrical motor rotates the impeller, the impeller being configured to raise the pressure of a fluid which flows through the impeller between an inlet and an outlet thereof;wherein the low pressure pump is configured such that a portion of the raised pressure fluid flow is diverted to form a recirculation flow which circulates through and cools the electrical motor before being returned to the impeller for re-pressurisation thereby.2. A fuel pumping unit according to claim 1 , wherein the centrifugal impeller is coaxial with the motor.3. A fuel pumping unit according to claim 1 , wherein the electrical motor is fully immersed in the pumped fluid.4. A fuel pumping unit according to claim 1 , wherein the centrifugal impeller has an inducer section which extends from an inlet for the fluid flow towards an outlet for the fluid flow.5. A fuel pumping unit according to ...

FUEL PUMPING UNIT

A fuel pumping unit has a low pressure centrifugal pump and a high pressure centrifugal pump. In use, the low pressure pump supplies fuel at a boosted pressure to the high pressure pump for onward supply to a fuel metering unit. The pumping unit further has a drive input which drives the low and high pressure pumps. A gear arrangement is operatively located between the drive input and the low and high pressure pumps such that the low and high pressure pumps are driven at different speeds by the drive input. 1. A fuel pumping unit having a low pressure centrifugal pump and a high pressure centrifugal pump , in use , the low pressure pump supplying fuel at a boosted pressure to the high pressure pump for onward supply to a fuel metering unit;the pumping unit further having a drive input which drives the low and high pressure pumps, a gear arrangement being operatively located between the drive input and the low and high pressure pumps such that the low and high pressure pumps are driven at different speeds by the drive input; andthe pumping unit further having a selector valve which is configured to selectably provide a high pressure mode and a low pressure mode, wherein:in the high pressure mode the low pressure pump supplies fuel at the boosted pressure to the high pressure pump for onward supply to the fuel metering unit; andin the low pressure mode the fuel supply from the low pressure pump to the high pressure pump is shut off such that the high pressure pump runs dry and the low pressure pump supplies fuel to the fuel metering unit.2. A fuel pumping unit according to claim 1 , wherein the high pressure pump is driven at a higher speed than the low pressure pump.3. A fuel pumping unit according to claim 1 , wherein the gear arrangement includes a bevel gear drive.4. A fuel pumping unit according to further having an axial flow inducer at the inlet to the low pressure pump.5. A fuel pumping unit according to claim 1 , wherein the axes of the low and high pressure ...

Gear pump

A gear pump has first and second meshing gears for pumping a fluid. The gear pump further has a first electrical motor having a first rotor, a first stator and first coil windings. The gear pump further has a second electrical motor having a second rotor, a second stator and second coil windings. The first rotor is operatively connected to the first gear and the second rotor is operatively connected to the second gear such that each gear is rotated by its respective rotor. The first and second coil windings are energised by respective and separate first and second electrical power circuits such that the first and second meshing gears can be driven independently of each other.

AERO-ENGINE LOW PRESSURE PUMP

An aero-engine low pressure pump is provided for supplying fuel at a raised pressure to a high pressure pump. The low pressure pump has a pumping mechanism which raises the pressure of fuel flowing though the mechanism. The low pressure pump further has electrical motor which drives the pumping mechanism. The low pressure pump further has a variable frequency motor drive which supplies electrical power to the electrical motor. The variable frequency motor drive measures the electrical power supplied to the electrical motor. The low pressure pump further has a control unit which compares the measured electrical power to a reference power, and, when the measured electrical power is less than the reference power by a predetermined amount, controls the motor drive to increase the power supplied to the electrical motor thereby increasing the pressure rise produced by the pumping mechanism. 1. An aero-engine low pressure pump for supplying fuel at a raised pressure to a high pressure pump , the low pressure pump having;a pumping mechanism which raises the pressure of fuel flowing though the mechanism:an electrical motor which drives the pumping mechanism; anda variable frequency motor drive which supplies electrical power to the electrical motor;wherein the variable frequency motor drive measures the electrical power supplied to the electrical motor;wherein the pump further has a control unit which compares the measured electrical power to a reference power, and, when the measured electrical power is less than the reference power by a predetermined amount, controls the motor drive to increase the power supplied to the electrical motor thereby increasing the pressure rise produced by the pumping mechanism; andwherein the measurement of the supplied electrical power is synchronised to the frequency of a variable pulse-width modulation used by the variable frequency motor drive to supply the electrical power.2. A low pressure pump according to claim 1 , wherein the reference ...

AXIAL PISTON PUMP

An axial piston pump has a fluid inlet, a fluid outlet, and a rotatable, one-piece piston housing carrying one or more pistons movable within one or more respective sleeves formed by the housing. A circumferential row of fixed field members is mounted to the housing at a mounting location formed by the housing, the mounting location being radially outward of the sleeve(s). The pump further has a stator surrounding the piston housing and including a circumferential row of armature windings such that the piston housing and the stator form an electro-magnetic motor operable to rotate the piston housing, and a swashplate engaged with the piston(s) such that rotation of the piston housing relative to the swashplate around an axis of rotation produces reciprocating movement of the piston(s) for the pressurisation of fluid received into the sleeve(s) from the fluid inlet and then discharged from the sleeve(s) to the fluid outlet. 1. An axial piston pump having:a fluid inlet, and a fluid outlet;a rotatable, one-piece piston housing carrying one or more pistons movable within one or more respective sleeves formed by the housing, a circumferential row of fixed field members being mounted to the housing at a mounting location formed by the housing, the mounting location being radially outward of the sleeve(s);a stator surrounding the piston housing and including a circumferential row of armature windings such that the piston housing and the stator form an electro-magnetic motor operable to rotate the piston housing; anda swashplate engaged with the piston(s) such that rotation of the piston housing relative to the swash plate around an axis of rotation produces reciprocating movement of the piston(s) for the pressurisation of fluid received into the sleeve(s) from the fluid inlet and then discharged from the sleeve(s) to the fluid outlet;wherein the fixed-field members and the armature windings are configured such that, when an electrical current is passed through the armature ...

COMPRESSOR GEOMETRY CONTROL

A system for controlling the geometry of a variable geometry compressor. The system having: a mechanical linkage operable to vary the compressor geometry; a first fluid-powered actuator arranged to operate the linkage and configured to be powered by a pressurised supply of an incompressible fluid; and a second fluid-powered actuator arranged to operate the linkage and configured to be powered by a pressurised supply of a compressible fluid. Wherein the first and second actuators are further arranged to operate in combination on the mechanical linkage such that a first actuator force exerted by the first actuator on the mechanical linkage sums with a second actuator force exerted by the second actuator on the mechanical linkage. 1. A system for controlling the geometry of a variable geometry compressor , the system having:a mechanical linkage operable to vary the compressor geometry;a first fluid-powered actuator arranged to operate the linkage and configured to be powered by a pressurised supply of an incompressible fluid; anda second fluid-powered actuator arranged to operate the linkage and configured to be powered by a pressurised supply of a compressible fluid;wherein the first and second actuators are further arranged to operate in combination on the mechanical linkage such that a first actuator force exerted by the first actuator on the mechanical linkage sums with a second actuator force exerted by the second actuator on the mechanical linkage.2. The system of claim 1 , wherein the mechanical linkage only operates to vary the compressor geometry when an actuator force equal to or greater than a threshold actuator force is exerted on the linkage claim 1 , and wherein the second actuator is arranged such that the second actuator force is less than the threshold actuator force.3. The system of claim 1 , further including a controller which is arranged to control the second actuator such that the second actuator force exerted by the second actuator on the ...

COMBUSTION STAGING SYSTEM

A combustion staging system has: a splitter; pilot and mains fuel manifolds; mains flow control valves; and fuel servo line. Each mains flow control valve has a chamber containing a piston, the chamber to a piston mains side fed by the mains fuel manifold, and the chamber to a piston servo side fed by the servo line. The piston has an open pilot-and-mains position allowing flow from the chamber mains side to the respective injector mains discharge orifice. The piston prevents flow from the chamber mains side. The piston is movable under a pressure change in the servo line relative to the mains fuel manifold. The system has a servo pump and a hydraulic motor driving it. The servo pump changes fuel pressure. Motive power for the hydraulic motor is fuel diverted from a fuel pump high pressure output, the motor returning the diverted fuel to a low pressure input. 1. A combustion staging system for fuel injectors of a multi-stage combustor of a gas turbine engine , the system having:a splitting unit which receives a metered total fuel flow and controllably splits the metered total fuel flow into out-going pilot and mains fuel flows to perform pilot-only and pilot-and-mains staging control of the combustor;pilot and mains fuel manifolds which respectively receive the pilot and mains fuel flows;a plurality of mains flow control valves which distribute the mains fuel flow from the mains fuel manifold to mains discharge orifices of respective injectors of the combustor; anda fuel servo line which extends through each of the mains flow control valves in series;wherein each mains flow control valve has a chamber containing a movable piston, the chamber to a mains side of the piston being fed by the mains fuel manifold, the chamber to a servo side of the piston being fed by the servo line, the piston having an open pilot-and-mains position which allows flow out of the mains side of the chamber to the mains discharge orifice of the respective injector, the piston being biased ...

Ramped stiffener and apparatus and method for forming the same

A die tool for forming a C-section component having radiused shoulders, the includes a cylindrical inner die and an outer die having a cylindrical central portion connected to opposing end flanges by respective radiused concave portions. A portion of the inner die is arranged to be disposed between the end flanges of the outer die and spaced apart therefrom to define a cavity corresponding to the desired cross-section of the C-section component to be formed. The radiused convex and concave portions have a radius of curvature that varies about the circumference of the respective inner and outer dies. The inner and outer dies are rotatable such that the radius of curvature of the radiused convex and concave portions where the inner and outer dies are adjacent to one another varies as the inner and outer dies are rotated.

Ramped stiffener and apparatus and method for forming the same

A die tool for forming a C-section component having radiused shoulders, the includes a cylindrical inner die and an outer die having a cylindrical central portion connected to opposing end flanges by respective radiused concave portions. A portion of the inner die is arranged to be disposed between the end flanges of the outer die and spaced apart therefrom to define a cavity corresponding to the desired cross-section of the C-section component to be formed. The radiused convex and concave portions have a radius of curvature that varies about the circumference of the respective inner and outer dies. The inner and outer dies are rotatable such that the radius of curvature of the radiused convex and concave portions where the inner and outer dies are adjacent to one another varies as the inner and outer dies are rotated.

Reconfigurable service provision via a communication network

A services provision system provides information services over one or more communications networks and has a software infrastructure divided into domains. Each domain has an intelligent software agent and this community of agents sits in a computing environment represented in each domain by a DPE kernel. The community of agents co-operates to provide service and service management functionality to user. At least one of the agents is reconfigurable to change the functionality the system makes available. Reconfigurability is based on the use of a plurality of reusable software modules, the agent reconfiguring by selecting a new combination of modules. The software modules themselves incorporate rules, or policies, which determine process steps offered by the modules at run-time. These policies are external to the modules and may be loaded at run-time, allowing dynamic modification to functionality of the system. The system as a whole offers functionality associated with using services, providing them and managing them and the reconfigurability allows it to offer the different types of functionality in an efficient way. It also allows access control to functionality at different levels with particularly good security against fraudulent use.

Fuel supply system

A fuel supply system for a gas turbine engine including first and second positive displacement pumps operated simultaneously to supply fuel under pressure from a low pressure source, a combining spill valve controlling the output flows from the first and second pumps to combine the outputs of the first and second pumps for supply to a metering valve of the system, or to spill some or all of the output of one or both pumps back to the low pressure supply, a pressure raising and shut off valve arrangement downstream of the metering valve for isolating the fuel system from an associated engine until the fuel pressure upstream of the pressure raising and shut off valve arrangement exceeds a predetermined pressure, and, a control system dependent upon the position of the combining spill valve for reducing said predetermined pressure at which said pressure raising and shut off valve arrangement opens.

Gear pump bearing

Information services provision and management

Fuel pumping system

A fuel pumping system comprises a high pressure pump having a drive input, a drive gear driven for rotation, in use, by the drive input, an idler gear driven by the drive gear, and an output gear driven by the idler gear, wherein the output gear is arranged to drive a rotor of an electrical generator and the idler gear is arranged to drive a rotor of a low pressure pump.

Aero-engine low pressure pump, a pumping unit of an aero-engine, a pumping system of an aero-engine, and an aero-engine

Apparatus and method for redirecting camera flash light so as to avoid red-eye

Apparatus for redirecting light beams (66) emanating from a camera flash (26) to avoid 'red-eye' in a photograph. The apparatus redirects the light beams in a direction parallel (68) to the central axis of the lens (32) at a distance further therefrom. The apparatus includes a main body (40) which forms an enclosure with an inlet opening (55) and outlet opening (48) and is attached to the camera (22) such that the inlet opening is aligned with and covers the camera flash. The device can be releasably mounted to existing flash cameras. It can also be built into new compact cameras and act as a sliding lens cover and power switch. With the built-in design, the flash of the camera can be directed either vertically or horizontally and a built-in light shifter with single reflective surface is provided to redirect the light parallel and far removed from the central axis of the lens.

Provision and management of information services

Apparatus and method for storing diagram data

An apparatus for storing data for use in displaying diagrams comprises a computer and a file server. Each diagram includes boxes and flow lines connecting the boxes. The software includes a database and a main program which is responsible for storing and retrieving diagram data as well as performing predetermined operations on the data. In the database, data describing a box is stored as an instance of a first software object class, data describing a flow line is stored as an instance of a second software object class, general data relating to a diagram is stored as an instance of a third software object class, and data describing the location of each entity (box or flow line) on a diagram is stored as an instance of a fourth software object class. Each instance of the fourth software object class includes a pointer to the instance of the third software object class which contains general data for the diagram and also to a pointer to the instance of the first or second software object class which describe the entity.

Pump

Exhaust content

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 comprises: determining, during use of the gas turbine engine, one or more exhaust content parameters by performing a sensor measurement on an exhaust of the gas turbine engine; and determining one or more fuel characteristics of the fuel based on the one or more exhaust parameters. Also disclosed is a fuel characteristic determination system, a method of operating an aircraft, and an aircraft.

Combustion staging system

A combustion staging system has: a splitter; pilot and mains fuel manifolds; mains flow control valves; and fuel servo line. Each mains flow control valve has a chamber containing a piston, the chamber to a piston mains side fed by the mains fuel manifold, and the chamber to a piston servo side fed by the servo line. The piston has an open pilot-and-mains position allowing flow from the chamber mains side to the respective injector mains discharge orifice. The piston prevents flow from the chamber mains side. The piston is movable under a pressure change in the servo line relative to the mains fuel manifold. The system has a servo pump and a hydraulic motor driving it. The servo pump changes fuel pressure. Motive power for the hydraulic motor is fuel diverted from a fuel pump high pressure output, the motor returning the diverted fuel to a low pressure input.

Fuel supply system

A fuel supply system for a gas turbine engine including first and second positive displacement pumps (11, 12) operated simultaneously to supply fuel under pressure from a low pressure source (13), and a combining spill valve (16) controlling the output flows from the first and second pumps to combine the outputs of the first and second pumps for supply to a metering valve (33) of the system, or to spill some or all of the output of one or both pumps back to the low pressure supply, pressure raising and shut off valve means (34) downstream of the metering valve for isolating the fuel system from an associated engine until the fuel pressure upstream of the pressure raising and shut off valve means exceeds a predetermined pressure, and, means (43, 42; 62) dependent upon the position of the combining spill valve for reducing said predetermined pressure at which said pressure raising and shut off valve means (34) opens.

Fuel supply system

A fuel supply system for a gas turbine engine comprises a first positive displacement pump ( 16 ), a second positive displacement pump ( 27 ), said first and second positive displacement pumps ( 16, 27 ) being operable simultaneously for the supply of fuel from a low pressure source, and switching valve means ( 14 ) downstream of said first positive displacement pump ( 16 ) for changing a connection mode between said first and second positive displacement pumps ( 16, 27 ) between a series mode and a parallel mode.

Client server model

A client-side intermediary (30) is provided to balance the loading of Web service requests between a plurality of servers (32). The status of the Web service servers (32) is monitored by a monitoring server (35) which provides status updates to the intermediary (30) upon request. The intermediary then uses the information on the status of the servers (32) to decide where to send web service requests. Additionally, the intermediary is able to direct requests for Web service descriptions to the least busy server on the basis of status information. The intermediary (30) substitutes its own identifier for the service name and port in the Web service description before passing it to the client so that all requests are directed through it, thus allowing the continual provision of service for the client even in the event that one of the servers fails.

Information services provision and management

A services provision system, for use in providing information services over one or more communications networks, has a software infrastructure divided into domains (101, 103, 104, 106). Each domain has an intelligent software agent (102, 107, 109, 110) and this community of agents sits in a computing environment represented in each domain by a DPE kernel (105). The community of agents co-operates to provide service and service management functionality t o a user. At least one of the agents (102, 107, 109, 110) is reconfigurable to change the functionality the system makes available. Reconfigurability is based on the use of a plurality of reusable software modules, the agent reconfiguring by selecting a new combination of modules. The software module s themselves incorporate rules, or policies, which determine process steps offered by the modules at run-time. These policies are external to the modul es and may be loaded at run-time, allowing dynamic modification to functionalit y of the system. The system as a whole offers functionality associated with using services, providing them and managing them and the reconfigurability allows it to offer the different types of functionality in an efficient way. It also allows access control to functionality at different levels with particularly good security against fraudulent use.

Pump assembly

A pump assembly comprises first and second gear pumps (10, 13), each of which has a respective driver gear (11, 14) and a respective driven gear (12, 15), wherein the driven gear (12) of the first gear pump (10) is arranged to drive the driver gear (14) of the second gear pump (13).

Fuel supply system

Un sistema de suministro de combustible para un motor de turbina de gas, el sistemacomprendiendo una primera bomba de desplazamiento positivo (16), una segunda bomba dedesplazamiento positivo (27), dichas primera y segunda bombas de desplazamiento positivo (16, 27)siendo accionadas simultáneamente para el suministro de combustible a partir de una fuente de bajapresión, y un medio de válvula de conmutación (14) aguas abajo de dicha primera bomba dedesplazamiento positivo (16) para cambiar un modo de conexión entre dichas primera y segundabombas de desplazamiento positivo (16, 27) entre un modo de serie y un modo paralelo, ycaracterizado por una válvula de regulación de división de presión (29) para mantener la diferenciade presión a través de ambas bombas (16, 27) en un ratio requerido. A fuel supply system for a gas turbine engine, the system comprising a first positive displacement pump (16), a second positive displacement pump (27), said first and second positive displacement pumps (16, 27) being driven simultaneously for supplying fuel from a low pressure source, and a switching valve means (14) downstream of said first positive displacement pump (16) to change a connection mode between said first and second positive displacement pumps ( 16, 27) between a series mode and a parallel mode, and characterized by a pressure division regulating valve (29) to maintain the pressure difference across both pumps (16, 27) at a required ratio.

Exhaust content

A method of determining one or more fuel characteristics of an aviation fuel F comprises determining one or more exhaust content parameters by performing a sensor 151 measurement on an exhaust of the gas turbine engine 10, and determining one or more fuel characteristics based on the one or more exhaust parameters. The sensor may measure nvPM content of the exhaust, using Laser Induced Incandescence (LII), or Condensation Particle count (CPC) using Laser scattering. Alternatively, the sensor may comprise a non-dispersive infra-red absorption (NDIR) measurement device, an aerosol mass spectrometer (AMS) device, a Scanning Mobility Particle Sizer (SMPS) or Differential Mobility Particle Sizer (DMS). The determination may determine a proportion of Sustainable Aviation Fuel (SAF).

Transaction authentication

In a payment validation system, a mobile phone or other communications terminal associated with a particular user is used. A vending node communicates with a validation platform which either returns a telephone number to be displayed for the user to call or which calls the alleged users phone or terminal for confirmation.

METHOD FOR DETERMINING ONE OR MORE FUEL CHARACTERISTICS OF AN AVIATION FUEL

TENEUR EN GAZ D’ÉCHAPPEMENT La présente demande divulgue un procédé (1050) de détermination d’une ou plusieurs caractéristiques de carburant d’un carburant d’aviation (F) adapté pour alimenter un moteur à turbine à gaz (10) d’un aéronef (1). Le procédé comprend : la détermination (1051), pendant l’utilisation du moteur à turbine à gaz, d’un ou plusieurs paramètres de teneur en gaz d’échappement en réalisant (1052) une mesure de capteur sur un gaz d’échappement du moteur à turbine à gaz (10) ; et la détermination (1053) d’une ou plusieurs caractéristiques de carburant du carburant sur la base du ou des plusieurs paramètres de gaz d’échappement. L’invention concerne également un système de détermination de caractéristiques de carburant (150), un procédé (1065) de fonctionnement d’un aéronef et un aéronef (1). Figure 25 EXHAUST GAS CONTENT This application discloses a method (1050) of determining one or more fuel characteristics of an aviation fuel (F) suitable for powering a gas turbine engine (10) of a aircraft (1). The method includes: determining (1051), during use of the gas turbine engine, one or more exhaust gas content parameters by performing (1052) a sensor measurement on an exhaust gas of the gas turbine engine (10); and determining (1053) one or more fuel characteristics of the fuel based on the one or more exhaust gas parameters. The invention also relates to a system for determining fuel characteristics (150), a method (1065) of operating an aircraft and an aircraft (1). Picture 25

Control System

A control system is described that comprises a manifold, at least one valve device operable to control the delivery of fluid from the manifold, and a drive member located within the manifold and rotatable to control the operation of the at least one valve device.

Aircraft refuelling

A method of refuelling an aircraft comprising a gas turbine engine and a fuel tank arranged to provide fuel to the gas turbine engine comprises obtaining an amount of energy required for an intended flight profile; obtaining a calorific value of fuel available to the aircraft for refuelling; calculating the amount of the available fuel needed to provide the required energy; and refuelling the aircraft with the calculated amount of the available fuel. The calculating the amount of the available fuel needed to provide the required energy may comprise obtaining an energy content of fuel already in the fuel tank and subtracting that from the determined amount of energy required for the intended flight profile.

Rampenförmige versteifung sowie vorrichtung und verfahren zu ihrer herstellung

Determination of fuel characteristics

A method 1014 of determining one or more fuel characteristics of an aviation fuel F suitable for powering a gas turbine engine 10 of an aircraft 1 comprises exposing 1015 the surface of a piezoelectric crystal of a sensor 115 to the fuel F; measuring 1016 a vibration parameter of the piezoelectric crystal; and determining 1017 one or more fuel characteristics of the fuel F based on the vibration parameter. The system may determine an oxygen content or thermal stability of the fuel, and/or a coking level of the fuel. Other fuel characteristics may be determined or inferred such as the hydrocarbon distribution of the fuel, the percentage of sustainable aviation fuel, or the aromatic hydrocarbon content of the fuel.

Aircraft operation

A gas turbine engine includes: a combustor that combust the fuel and having an exit, a combustor exit temperature (T40) is the average temperature of flow and a combustor exit pressure (P40) is the total pressure there; a turbine including a rotor having a leading edge and a trailing edge, and wherein a turbine rotor entry temperature (T41) is an average temperature of flow at the leading edge and a turbine rotor entry pressure (P41) is the total pressure there; and a compressor having an exit, wherein a compressor exit temperature (T30) is the average temperature of flow at the exit from the compressor and a compressor exit pressure (P30) is the total pressure there (all at cruise conditions). A method of determining at least one fuel characteristic includes changing a fuel supplied to the engine; and determining a change in a relationship between T30 or P30, T40 and T41, or of P40 and P41, respectively.

Aircraft refuelling

A method 2020 of refuelling an aircraft comprising a gas turbine engine and a fuel tank arranged to provide fuel to the gas turbine engine comprises obtaining 2022 an amount of energy required for an intended flight profile; obtaining 2024 a calorific value of fuel available to the aircraft for refuelling; calculating 2026 the amount of the available fuel needed to provide the required energy; and refuelling 2029 the aircraft with the calculated amount of the available fuel. The calculating 2026 the amount of the available fuel needed to provide the required energy may comprise obtaining an energy content of fuel already in the fuel tank and subtracting that from the determined amount of energy required for the intended flight profile.

Aircraft fuel management

A propulsive gas turbine engine 10 of an aircraft 1 comprises a combustor 16 having an exit temperature T40 or exit pressure P40, a turbine having a turbine rotor entry temperature T41 and entry pressure P41, and a compressor having an exit temperature T30 and exit pressure P30. A method 2010 of determining at least one fuel characteristic of a fuel provided to the gas turbine engine comprises changing 2012 a fuel supplied to the gas turbine engine and determining 2016 a fuel characteristic based on one of T30, P30, T40, P40, T41 and P41. Preferably the determination is based on a change in a relationship between T30 and one of T40 and T41, or P30 and one of P40 and P41. The relationship may correspond to a difference which is indicative of a rise in temperature across the combustor. The determination may be used to change a variable inlet guide vane (VIGV).

Aircraft fuel management

Determination of fuel characteristics

A method of determining one or more fuel characteristics of an aviation fuel F for powering a gas turbine engine of an aircraft, comprises: measuring one or more trace substance parameters of the fuel, the parameters each associated with a respective trace substance in the fuel; and determining one or more characteristics of the fuel based on the trace substance parameters. Also disclosed is a fuel characteristic determination system 126 comprising a sensor 127 for monitoring fuel parameters, based on fluorescence measurement or spectroscopy (Infrared, FTIR, or UV-Vis spectroscopy). The trace substances may be sulphur or aromatic hydrocarbons. Sustainable aviation fuel (SAF) does not contain sulphur or aromatics thus determining parameters such as presence/concentration may indicate what percentage the fuel is of SAF to fossil-based fuel such as kerosene.

Procédé de détermination d’une ou plusieurs caractéristiques de carburant d’un carburant aviation

La présente demande décrit un procédé (1008) de détermination d’une ou plusieurs caractéristiques de carburant d’un carburant aviation adapté pour alimenter un moteur à turbine à gaz  d’un aéronef, le moteur à turbine à gaz) ayant une chambre de combustion à laquelle du carburant est apporté d’un système de carburant, le procédé comprenant : la détermination (1009) d’une masse du carburant étant apporté à la chambre de combustion ; la détermination (1010) d’un volume correspondant du carburant étant apporté à la chambre de combustion ; et la détermination (1011) dʼune ou plusieurs caractéristiques de carburant en fonction de la masse et du volume déterminés. Il est aussi décrit un système de détermination de caractéristiques de carburant, un procédé d’opération d’un aéronef, et un aéronef. Figure pour l’abrégé : 19

Aubes directrices d’entrée orientables

Un procédé 3010 de commande d’un système de propulsion d’un aéronef, le système de propulsion comprenant un moteur à turbine à gaz agencé pour être alimenté par un carburant et au moins une aube directrice d’entrée orientable - VIGV, comprend l’obtention 3012 d’au moins une caractéristique de carburant du carburant fourni au moteur à turbine à gaz ; et la réalisation 3014 d’un changement de programmation de l’au moins une VIGV sur la base de l’au moins une caractéristique de carburant obtenue. Figure 5

Gestion de carburant d’aéronef

Un procédé 2090 de détermination d’au moins une caractéristique de carburant d’un carburant fourni à un moteur à turbine à gaz d’un aéronef comprend la réalisation 2092 d’un changement de fonctionnement, le changement de fonctionnement étant effectué par un composant pouvant être commandé d’un système de propulsion dont fait partie le moteur à turbine à gaz, et étant agencé pour affecter le fonctionnement du moteur à turbine à gaz, la détection 2094 d’une réponse au changement de fonctionnement ; et la détermination 2096 de l’au moins une caractéristique de carburant sur la base de la réponse au changement de fonctionnement. Figure 7

Procede de determination d’une ou plusieurs caracteristiques de carburant d’un carburant d’aviation

L’invention concerne un procédé (1040) de détermination d’une ou plusieurs caractéristiques de carburant d’un carburant d’aviation (F) adapté pour alimenter un moteur à turbine à gaz (10) d’un aéronef (1). Le procédé comprend : la détermination (1041), pendant l’utilisation du moteur à turbine à gaz (10), d’un ou plusieurs paramètres de traînée de condensation liés à la formation de traînée de condensation par le moteur à turbine à gaz (10), dans lequel la détermination (1041) du ou des plusieurs paramètres de traînée de condensation comprend la réalisation (1042) d’une mesure de capteur sur une région derrière le moteur à turbine à gaz dans laquelle une traînée de condensation est ou peut être formée ; et la détermination (1043) d’une ou plusieurs caractéristiques de carburant du carburant (F) sur la base du ou des plusieurs paramètres de traînée de condensation. L’invention concerne également un système de détermination de caractéristiques de carburant (140), un procédé (1065) de fonctionnement d’un aéronef et un aéronef (1). Figure 23

Détermination de caractéristiques de carburant

La présente demande divulgue un procédé (1021) de détermination d’une ou plusieurs caractéristiques de carburant d’un carburant d’aviation (F) pour alimenter un moteur à turbine à gaz (10) d’un aéronef (1). Le procédé comprend : l’exposition (1022) de la surface d’un composant de capteur formé à partir d’un matériau d’étanchéité en nitrile au carburant (F) ; la mesure (1023) d’un paramètre de gonflement du matériau d’étanchéité ; et la détermination (1024) d’une ou plusieurs caractéristiques de carburant du carburant (F) sur la base du paramètre de gonflement. L’invention concerne également un système de détermination de caractéristiques de carburant (119), un procédé (1065) de fonctionnement d’un aéronef et un aéronef (1). Figure 9

Procede de fonctionnement d’un aeronef avec determination de parametres de performance

La présente demande divulgue un procédé (1058) de détermination d’une ou plusieurs caractéristiques de carburant d’un carburant d’aviation (F) utilisé pour alimenter un moteur à turbine à gaz (10) d’un aéronef (1). Le procédé comprend : la détermination (1059) d’un ou plusieurs paramètres de performance du moteur à turbine à gaz (10) pendant une première période de temps de fonctionnement du moteur à turbine à gaz (10) ; et la détermination (1060) d’une ou plusieurs caractéristiques de carburant du carburant (F) sur la base du ou des plusieurs paramètres de performance. L’invention concerne également un procédé de fonctionnement d’un aéronef, un système de détermination de caractéristiques de carburant (155) et un aéronef. Figure 27

Procédé de génération d’un programme de maintenance pour un aéronef

DÉTERMINATION DE CARACTÉRISTIQUES DE CARBURANT La présente demande divulgue un procédé (1080) de génération d’un programme de maintenance pour un aéronef (1) ayant un ou plusieurs moteurs à turbine à gaz (10) alimentés par un carburant d’aviation (F). Le procédé comprend : la détermination (1081) d’une ou plusieurs caractéristiques de carburant du carburant (F) ; et la génération (1082) d’un programme de maintenance selon la ou les plusieurs caractéristiques de carburant. L’invention concerne également un procédé (1085) de maintenance d’un aéronef, un système de génération de programme de maintenance (160) et un aéronef (1). Figure 32  

Détermination de caractéristiques de carburant

DÉTERMINATION DE CARACTÉRISTIQUES DE CARBURANT La présente demande fournit un procédé (1034) de détermination d’une ou plusieurs caractéristiques de carburant d’un carburant d’aviation (F) adapté pour alimenter un moteur à turbine à gaz (10) d’un aéronef (1). Le procédé comprend : le passage (1035) d’une lumière à spectre UV-visible à travers le carburant (F) ; la mesure (1036) d’un paramètre de transmittance indiquant la transmittance de la lumière à travers le carburant (F) ; la détermination (1037) d’une ou plusieurs caractéristiques de carburant du carburant (F) sur la base du paramètre de transmittance ; et la communication (1038) de la ou des plusieurs caractéristiques de carburant à un module de commande du moteur à turbine à gaz ou de l’aéronef. L’invention concerne également un système de détermination de caractéristiques de carburant (130), un procédé (1065) de fonctionnement d’un aéronef (1) et un aéronef (1). Figure 19

Contrail parameters

A method of determining one or more fuel characteristic of an aviation fuel suitable for powering a gas turbine engine of an aircraft is disclosed. The method includes: determining, during use of the gas turbine engine, one or more contrail parameters related to contrail formation by the gas turbine engine. The determining of the one or more contrail parameters includes performing a sensor measurement on a region behind the gas turbine engine in which a contrail is or can be formed. The method also includes determining one or more fuel characteristics of the fuel based on the one or more contrail parameters. A fuel characteristic determination system, a method of operating an aircraft and an aircraft are also disclosed.

Determination of fuel characteristics

Fuel characteristics

A method 1002 of determining one or more fuel characteristics of an aviation fuel suitable for powering a gas turbine engine (10, Fig. 16) of an aircraft (1, Fig. 16), the method comprising: determining 1003 a mass of a fuel being loaded, or which has been loaded, onto the aircraft; determining 1004 a corresponding volume of the fuel; determining 1005 one or more fuel characteristics of the fuel based on the determined mass and volume; Also disclosed is a fuel characteristic determination system (102, Fig. 16), a method (1065, Fig. 20) of operating an aircraft, and an aircraft.

Fuel delivery

The present application discloses a method 1008 of determining one or more fuel characteristics of an aviation fuel suitable for powering a gas turbine engine (10, Fig. 1) of an aircraft, the gas turbine engine (10, Fig. 1) having a combustor (16, Fig. 1) supplied with fuel from a fuel system, the method comprising: determining 1009 a mass of the fuel being supplied to the combustor; determining 1010 a corresponding volume of the fuel being supplied to the combustor; and determining 1011 one or more fuel characteristics based on the determined mass and volume.; Also disclosed is a fuel characteristic determination system (106, Fig. 18), a method (1065, Fig. 20) of operating an aircraft, and an aircraft.

Calorific value

A method of checking refuelling of an aircraft comprising a gas turbine engine and a fuel tank arranged to provide fuel to the gas turbine engine comprises: receiving an input of calorific value data for fuel provided to the aircraft on refuelling; independently determining at least one of: (i) the calorific value of fuel supplied to the gas turbine engine in use; and (ii) the calorific value of the fuel provided to the aircraft on refuelling; and providing an alert if the determined calorific value is inconsistent with the calorific value data input received

Fuel delivery

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 gas turbine engine having a combustor supplied with fuel from a fuel system, the method comprising: determining a mass of the fuel being supplied to the combustor; determining a corresponding volume of the fuel being supplied to the combustor; and determining one or more fuel characteristics based on the determined mass and volume. Also disclosed is a fuel characteristic determination system, a method of operating an aircraft, and an aircraft.

Zahnradpumpe mit lager

Determination of fuel characteristics

Calorific value

An aircraft has a fuel tank 50 which provides fuel to a gas turbine engine 10, an input of calorific value data for fuel available for refuelling the aircraft is received, preferably by a fuel tracking system 203. Independently, the calorific value of fuel supplied to the gas turbine engine in use, and/or the calorific value of the fuel provided to the aircraft on refuelling, is determined. An alert is generated if the determined calorific value is inconsistent with the received calorific value data. Determining the calorific value of the fuel provided to the aircraft may: use a fuel testing unit provided off-wing at a refuelling site; be performed during aircraft operations prior to take-off; burn fuel taken from the tank in an auxiliary power unit; be based upon monitored engine parameters during a time period; identify a tracer in the fuel and lookup a corresponding calorific value; be inferred from a physical or chemical fuel property. The input of calorific value data may be received via a user interface or communicated electronically to the aircraft. Checking the calorific value of the fuel may be particularly relevant for aircraft engines which use a sustainable aviation fuel (SAF) rather than kerosene.

Propulsion system

A propulsion system (fig.10,2) for an aircraft ((fig.10,1) has a number of fuel tanks 50,53 storing fuel powering a gas turbine engine (fig.10,10) with a fuel pump 108, an inlet (fig.14,108a) receives fuel from the tanks and an outlet (fig.14,108b) provides fuel to a combustor 16. The fuel contains sustainable aviation fuel (SAF) with a density 90-98% of kerosene (ρK), and a calorific value 101-105% of kerosene (CVK); the fuel pumped has between 30-100% SAF with any remainder being kerosene, with fuel mix density (ρF) and calorific value (CVF). The pump fuels the combustor at energy flow rate C, outputing fuel at a volumetric flow rate Q; the pump has a spill percentage - the percentage of fuel passing through the pump which is recirculated and not provided to the combustor. A fuel-change spill ratio Rs is the spill percentage at cruise using kerosene over the spill percentage at cruise using the SAF fuel blend, and is equal to: (Q – (C / CVK x ρK)) / (Q – (C / CVF x ρF)) The percentage of SAF may be as low as 5%. The value of Rs may be 1.003, when the percentage of SAF is at least 30%.

Engine parameters

A power system 2 for an aircraft 1 comprising a gas turbine engine 10, a fuel tank 53a, 53b and a fuel tracking system arranged to sense an engine parameter during a first time period of aircraft operation during which the engine uses fuel, and determine a calorific value of the fuel based on the engine parameter. The power system may be associated with an auxiliary power unit 44. A sensor (204, fig. 8) may be arranged to sense the engine parameter. Also provided is a method of determining a calorific value of fuel supplied to a gas turbine engine of an aircraft.

Determination of fuel characteristics

A method of generating a maintenance schedule for an aircraft including one or more gas turbine engines powered by an aviation fuel. The method includes: determining one or more fuel characteristics of the fuel; and generating a maintenance schedule according to the one or more fuel characteristics. Also disclosed is a method of maintaining an aircraft, a maintenance schedule generation system and an aircraft.

Transmittance parameters

The present application provides a method of determining one or more fuel characteristics of an aviation fuel suitable for powering a gas turbine engine of an aircraft. The method includes: passing UV-visual spectrum light through the fuel; measuring a transmittance parameter indicating the transmittance of light through the fuel; determining one or more fuel characteristics of the fuel based on the transmittance parameter; and communicating the one or more fuel characteristic to a control module of the gas turbine engine or the aircraft. Also disclosed is a fuel characteristic determination system, a method of operating an aircraft, and an aircraft.

Exhaust content

Gear pump

Procédé de contrôle d’avitaillement d’un aéronef

Un procédé (2030) de contrôle de lʼavitaillement dʼun aéronef comprenant un moteur à turbine à gaz et un réservoir de carburant agencé pour fournir du carburant au moteur à turbine à gaz comprend : la réception (2032) dʼune entrée de données de pouvoir calorifique de carburant fourni à lʼaéronef lors de l’avitaillement ; la détermination (2034) indépendamment dʼau moins lʼun parmi : (i) le pouvoir calorifique de carburant apporté au moteur à turbine à gaz en cours dʼutilisation ; et (ii) le pouvoir calorifique du carburant fourni à lʼaéronef lors de l’avitaillement ; et la fourniture (2036) dʼune alerte si le pouvoir calorifique déterminé est incohérent avec lʼentrée de données de pouvoir calorifique reçue. Figure pour l’abrégé : 7 

Verfahren zur kontrolle der betankung eines luftfahrzeugs

Verfahren (2030) zur Kontrolle der Betankung eines Luftfahrzeugs, das ein Gasturbinentriebwerk und einen Kraftstofftank aufweist, der so angeordnet ist, dass er dem Gasturbinentriebwerk Kraftstoff zuführt, mit den folgenden Schritten: Empfangen (2032) einer Eingabe von Heizwertdaten für Kraftstoff, der dem Luftfahrzeug bei der Betankung zugeführt wird; unabhängiges Bestimmen (2034) von mindestens einem der Folgenden: (i) des Heizwerts von Kraftstoff, der dem Gasturbinentriebwerk (10) im Betrieb zugeführt wird; und (ii) des Heizwerts des Kraftstoffs, der dem Luftfahrzeug bei der Betankung zugeführt wird; und Bereitstellen (2036) einer Warnmeldung, wenn der bestimmte Heizwert nicht mit den empfangenen eingegebenen Heizwertdaten übereinstimmt.

Aircraft refuelling

AIRCRAFT REFUELLING A method 2020 of refuelling an aircraft comprising a gas turbine engine and a fuel tank arranged to provide fuel to the gas turbine engine comprises obtaining 2022 an amount of energy required 5 for an intended flight profile; obtaining 2024 a calorific value of fuel available to the aircraft for refuelling; calculating 2026 the amount of the available fuel needed to provide the required energy; and refuelling 2029 the aircraft with the calculated amount of the available fuel. The calculating 2026 the amount of the available fuel needed to provide the required energy may comprise obtaining an energy content of fuel already in the fuel tank and subtracting that from the determined .0 amount of energy required for the intended flight profile. To be reproduced with Figure 5 when published. Fig 0 _2Z _ 2 "TT

Fuel viscosity

A method of operating a gas turbine engine including an engine core including a turbine, compressor, combustor to combust a fuel, and core shaft connecting the turbine and compressor; a fan upstream of the engine core; a fan shaft; a gearbox that receives an input from the core shaft and outputs drive to the fan via the fan shaft; a primary oil loop system to supply oil to the gearbox; and a heat exchange system. The method includes controlling the heat exchange system to adjust fuel viscosity to be lower than or equal to 0.58 mm 2 /s on entry to the combustor at cruise conditions.

Fuel viscosity

A method of operating a gas turbine engine including an engine core including a turbine, compressor, combustor to combust a fuel, and core shaft connecting the turbine and compressor; a fan upstream of the engine core; a fan shaft; a gearbox that receives an input from the core shaft and outputs drive to the fan via the fan shaft; a primary oil loop system to supply oil to the gearbox; and a heat exchange system. The method includes controlling the heat exchange system to adjust fuel viscosity to be lower than or equal to 0.58 mm2/s on entry to the combustor at cruise conditions.

Bestimmung von kraftstoffmerkmalen

Die vorliegende Anmeldung offenbart ein Verfahren (1014) zur Bestimmung eines oder mehrerer Kraftstoffmerkmale eines Flugkraftstoffs (F), der zum Antrieb eines Gasturbinentriebwerks (10) eines Luftfahrzeugs (1) geeignet ist. Das Verfahren umfasst: Aussetzen (1015) der Oberfläche eines piezoelektrischen Kristalls dem Kraftstoff (F); Messen (1016) eines Schwingungsparameters des piezoelektrischen Kristalls; und Bestimmen (1017) eines oder mehrerer Kraftstoffmerkmale des Kraftstoffs (F) basierend auf dem Schwingungsparameter. Ebenfalls offenbart sind ein System (114) zur Bestimmung von Kraftstoffmerkmalen, ein Verfahren (1065) zum Betreiben eines Luftfahrzeugs und ein Luftfahrzeug.

Determination of fuel characteristics

Aircraft operation

OF THE DISCLOSURE AIRCRAFT OPERATION A gas turbine engine of an aircraft forms part of a propulsion system, and comprises: a combustor arranged to combust the fuel and having an exit, and wherein a combustor exit temperature - T40 - is 5 defined as an average temperature of flow at the combustor exit and a combustor exit pressure - P40 is defined as the total pressure there; a turbine comprising a rotor having a leading edge and a trailing edge, and wherein a turbine rotor entry temperature - T41 - is defined as an average temperature of flow at the leading edge of the rotor of the turbine and a turbine rotor entry pressure - P41 - is defined as the total pressure there; and a compressor having an exit, wherein a compressor exit temperature .0 T30 - is defined as an average temperature of flow at the exit from the compressor at cruise conditions and a compressor exit pressure - P30 - is defined as the total pressure there (all at cruise conditions). A method of determining at least one fuel characteristic of a fuel provided to the gas turbine engine comprises changing a fuel supplied to the gas turbine engine; and determining the at least one fuel characteristic of the fuel based on a change in a relationship between T30 or P30 and one of T40 and .5 T41, or of P40 and P41, respectively. To be reproduced with Figure 9 when published. .0 1/6 Fig. 1 21 10 14 1 2 18 B- - 15 1 23-- 17 A-- -20 27 Fig.2 23 24 38 40 32 28 34 2

Aircraft operation

A method of determining at least one fuel characteristic of a fuel provided to a gas turbine engine of an aircraft includes making an operational change, the operational change being effected by a controllable component of a propulsion system of which the gas turbine engine forms a part, and being arranged to affect operation of the gas turbine engine, sensing a response to the operational change; and determining the at least one fuel characteristic based on the response to the operational change.

Performance parameters

A method and system 155 for determining one or more fuel characteristics of an aviation fuel for an aircraft gas turbine engine 10. The method comprises determining one or more performance parameters of the engine during a first time period of operation and determining one or more fuel characteristics based on the performance parameter(s). The characteristic(s) may be determined during a second later period of operation, and the first period may be a first flight phase and the second period may be a second flight phase, different from the first. The performance parameter(s) may include a fan rotation speed, a turbine entry temperature (TET), and/or a combustor fuel-to-air ratio. Each determined performance parameter may be compared with a reference performance parameter corresponding to operation of an engine with a fuel having a known fuel characteristic. The fuel characteristic(s) may include a hydrocarbon distribution, a percentage of sustainable aviation fuel (SAF), an aromatic hydrocarbon content and/or an indication that the fuel is a fossil fuel e.g. kerosene.

Leistungsparameter

Die vorliegende Anmeldung offenbart ein Verfahren (1058) zur Bestimmung eines oder mehrerer Kraftstoffmerkmale eines Flugkraftstoffs (F), der für den Antrieb eines Gasturbinentriebwerks (10) eines Luftfahrzeugs (1) verwendet wird. Das Verfahren umfasst: Bestimmen (1059) eines oder mehrerer Leistungsparameter des Gasturbinentriebwerks (10) während eines ersten Betriebszeitraums des Gasturbinentriebwerks (10); und Bestimmen (1060) eines oder mehrerer Kraftstoffmerkmale des Kraftstoffs (F) auf der Grundlage des einen oder der mehreren Leistungsparameter. Ein Verfahren zum Betreiben eines Luftfahrzeugs, ein System zur Bestimmung von Kraftstoffmerkmalen (155) und ein Luftfahrzeug (1) werden ebenfalls offenbart.

Fuel management system

A method of operating a gas turbine engine is disclosed, the gas turbine engine comprising an engine core comprising a combustor arranged to burn a fuel; a turbine, the turbine comprising a plurality of turbine blades; a compressor arranged to be used as a source of cooling air for the turbine blades; and an inducer arranged to accelerate and direct the cooling air onto the turbine blades and comprising a plurality of airflow passageways and a modulating valve arranged to control cooling air flow into a subset of the passageways; and a fuel management system arranged to provide the fuel to the combustor, the fuel having a sulphur content of less than 30 ppm. The fuel management system comprises two fuel-oil heat exchangers through which oil and the fuel flow. The method comprises using the modulating valve to adjust the cooling air flow based on turbine inlet temperature.

Procédé de détermination d’une ou plusieurs caractéristiques de carburant d’un carburant aviation d’un aéronef

La présente demande décrit un procédé (1002) de détermination d’une ou plusieurs caractéristiques de carburant d’un carburant aviation adapté pour alimenter un moteur à turbine à gaz d’un aéronef, le procédé comprenant : la détermination (1003) d’une masse d’un carburant étant chargé, ou qui a été chargé, sur l’aéronef ; la détermination (1004) d’un volume correspondant du carburant ; la détermination (1005) d’une ou plusieurs caractéristiques de carburant du carburant en fonction de la masse et du volume déterminés. Il est aussi décrit un système de détermination de caractéristiques de carburant, un procédé (1065) d’opération d’un aéronef, et un aéronef. Figure pour l’abrégé : Fig. 17

Determination of fuel characteristics

A fuel characteristic determination system and method for determining one or more fuel characteristics of an aviation fuel for powering a gas turbine engine of an aircraft. The system includes a sensor component that is formed from a nitrile seal material and includes a surface that can be exposed to the fuel; a sensor for measuring a swell parameter of the seal material; and a fuel characteristics determination module for determining one or more fuel characteristics of the fuel based on the swell parameter. The method includes: exposing one or more seals of a fuel system of an aircraft to fuel within the fuel system; exposing a sensor component, made from the same material as the one or more seals, to the fuel; and measuring a swell parameter of the seal material.

Aviation fuel

AVIATION FUEL 5 The present application discloses a method (1030) of determining one or more fuel characteristics of an aviation fuel (F) for powering a gas turbine engine (10) of an aircraft (1). The method comprises: measuring (1031) one or more trace substance parameters of the fuel (F), the one or more trace substance parameters each associated with a respective trace substance in the fuel (F); and determining (1032) one or more fuel characteristics of the fuel (F) based on the one or more .0 trace substance parameters. Also disclosed is a fuel characteristic determination system (126), a method (1065) of operating an aircraft (1), and an aircraft. [Figure 15] .5 1/17 Fig. 1 10 B 14 1 11 22 1 8 23 A 1 20 30 26 16' 27 Fig.2 23 2 38 40 32 28 34 2

Système de Propulsion pour un aéronef

Un système de propulsion (2) pour un aéronef (1) comprend un moteur à turbine à gaz (10) et un réservoir de carburant (50, 53) agencé pour stocker du carburant pour alimenter le moteur à turbine à gaz (10), dans lequel le carburant stocké comprend au moins une proportion d’un carburant aviation durable – SAF – ayant une densité comprise entre 90 % et 98 % de la densité, ρK, de kérosène et un pouvoir calorifique compris entre 101 % et 105 % le pouvoir calorifique PCK, de kérosène. Le moteur à turbine à gaz 10 comprend une chambre de combustion 16 ; et une pompe à carburant 108 agencée pour apporter un carburant à la chambre de combustion 16 à un débit d’énergie, C, la pompe 108 étant agencée pour délivrer en sortie du carburant à un débit volumétrique, Q, le pourcentage de carburant passant à travers la pompe qui n’est pas fourni à la chambre de combustion 16 étant appelé un pourcentage de déversement. Le carburant comprend X % de SAF, où X % est dans la plage allant de 5 % à 100 %, et a une densité, ρF, et un pouvoir calorifique PCF. Le système de propulsion 2 est agencé de telle sorte que : le rapport de déversement de changement de carburant, Rd : est égal à : Figure pour l’abrégé : 13 

Bestimmung von kraftstoffmerkmalen

Die vorliegende Anmeldung offenbart ein Verfahren (1080) zur Erstellung eines Wartungsplans für ein Luftfahrzeug (1) mit einem oder mehreren Gasturbinentriebwerken (10), die mit einem Flugkraftstoff (F) betrieben werden. Das Verfahren umfasst: Bestimmen (1081) eines oder mehrerer Kraftstoffmerkmale des Kraftstoffs (F); und Erzeugen (1082) eines Wartungsplans gemäß dem einen oder den mehreren Kraftstoffmerkmalen. Ebenfalls offenbart werden ein Verfahren (1085) zur Wartung eines Luftfahrzeugs, ein System (160) zur Erstellung eines Wartungsplans und ein Luftfahrzeug (1).

Bestimmung von kraftstoffmerkmalen

Ein Verfahren (1040) zur Bestimmung eines oder mehrerer Kraftstoffmerkmale eines Flugkraftstoffs (F), der zum Antrieb eines Gasturbinentriebwerks (10) eines Luftfahrzeugs (1) geeignet ist, wird offenbart. Das Verfahren umfasst: Bestimmen (1041), während der Verwendung des Gasturbinentriebwerks (10), eines oder mehrerer Kraftstoffmerkmale, die sich auf die Bildung von Kondensstreifen durch das Gasturbinentriebwerk (10) beziehen, wobei das Bestimmen (1041) des einen oder der mehreren Kondensstreifenparameter das Durchführen (1042) einer Sensormessung in einem Bereich hinter dem Gasturbinentriebwerk umfasst, in dem ein Kondensstreifen gebildet wird oder gebildet werden kann; und Bestimmen (1043) eines oder mehrerer Kraftstoffmerkmale des Kraftstoffs (F) auf der Grundlage des einen oder der mehreren Kondensstreifenparameter. Ein System (140) zur Bestimmung von Kraftstoffmerkmalen, ein Verfahren (1065) zum Betrieb eines Luftfahrzeugs und ein Luftfahrzeug (1) werden ebenfalls offenbart.

Bestimmung von kraftstoffmerkmalen

Die vorliegende Anwendung stellt ein Verfahren (1034) zur Bestimmung eines oder mehrerer Kraftstoffmerkmale eines Flugkraftstoffs (F) bereit, der für den Antrieb eines Gasturbinentriebwerks (10) eines Luftfahrzeugs (1) geeignet ist. Das Verfahren umfasst: Durchleiten (1035) von Licht aus dem sichtbaren/UV Spektrum durch den Kraftstoff (F); Messen (1036) eines Transmissionsparameters, der die Lichtdurchlässigkeit des Kraftstoffs (F) anzeigt; Bestimmen (1037) eines oder mehrerer Kraftstoffmerkmale des Kraftstoffs (F) auf der Grundlage des Transmissionsparameters; und Übermitteln (1038) des einen oder der mehreren Kraftstoffmerkmale an ein Steuermodul des Gasturbinentriebwerks oder des Luftfahrzeugs. Ebenfalls offenbart werden ein System (130) zur Bestimmung von Kraftstoffmerkmalen, ein Verfahren (1065) zum Betrieb eines Luftfahrzeugs (1) und ein Luftfahrzeug (1).

Procédure de rallumage de chambre de combustion

L'invention concerne un procédé (1100) de fonctionnement d'un moteur à turbine à gaz (10). Le moteur à turbine à gaz (10) comprend une chambre de combustion étagée (16) ayant un agencement de buses de pulvérisation de carburant (403, 404) dans lequel un écoulement de carburant est sollicité vers un sous-ensemble des buses (403, 404) adjacent à un ou plusieurs allumeurs (405) pendant une procédure de rallumage. Le procédé (1100) comprend le transfert de chaleur (1201) de l'huile vers le carburant avant que le carburant n'entre dans la chambre de combustion (16) de façon à abaisser la viscosité de carburant à 0,58 mm2/s ou moins lors de l'entrée dans la chambre de combustion (16) aux conditions de croisière. L'invention concerne également un moteur à turbine à gaz (10). Figure pour l’abrégé : Fig. 9

Magerverbrennungssystem

Es wird ein Verfahren (400) zum Betreiben eines Gasturbinentriebwerks (10) mit einem Verbrennungssystem mit Magerbrennstufe mit einer Brennkammer (16) bereitgestellt, in der Kraftstoff verbrannt wird. Das Gasturbinentriebwerk (10) umfasst einen Kraftstoff-Öl-Wärmetauscher (114), der angeordnet ist, um Wärme zwischen Öl und Kraftstoff zu übertragen, welcher der Brennkammer (16) bereitgestellt wird. Das Verfahren (400) umfasst ein Übertragen von Wärme (401) aus dem Öl auf den Kraftstoff, bevor der Kraftstoff in die Brennkammer (16) eintritt, um die Kraftstoffviskosität auf 0,58 mm2/s oder niedriger bei Eintritt in die Brennkammer (16) unter Reiseflugbedingungen zu verringern. Ebenfalls bereitgestellt wird ein Gasturbinentriebwerk (10).

Gasturbinenbetrieb

Es wird ein Verfahren (1100) zum Betreiben eines Gasturbinentriebwerks (10) bereitgestellt, wobei das Gasturbinentriebwerk (10) umfasst: eine Brennkammer für fettes Verbrennen, schnelles Abkühlen, mageres Verbrennen (RQL-Brennkammer) (16), die eine Anzahl von Kraftstoffsprühdüsen (403, 404) im Bereich von 14-22 oder eine Anzahl von Kraftstoffsprühdüsen pro Einheitstriebwerkskerngröße im Bereich von 2 bis 6 aufweist. Das Verfahren (1000) umfasst ein Betreiben (1101) des Gasturbinentriebwerks (10) derart, dass eine Reduzierung von 10-70 % eines Durchschnitts an Partikeln/kg nvPM im Abgas des Gasturbinentriebwerks (10), wenn das Triebwerk 10 mit 85 % verfügbarem Schub bei gegebenen Betriebsbedingungen betrieben wird, und Partikeln/kg nvPM im Abgas des Gasturbinentriebwerks (10), wenn das Triebwerk (10) mit 30 % verfügbarem Schub bei den gegebenen Betriebsbedingungen betrieben wird, erhalten wird, wenn ein Kraftstoff, der an die Brennkammer (16) bereitgestellt wird, ein nachhaltiger Flugzeugkraftstoff anstelle eines Kohlenwasserstoffkraftstoffs auf fossiler Basis ist.

A gas turbine engine

A method (12000) of operating a gas turbine engine (10) is disclosed, the gas turbine engine (10) comprising an engine core comprising a turbine, a compressor, a combustor (16) arranged to combust a fuel, and a core shaft connecting the turbine to the compressor; a fan located upstream of the engine core; a fan shaft (42); a gearbox (30) that receives an input from the core shaft and outputs drive to the fan via the fan shaft (42); a primary oil loop system (2000) arranged to supply oil to the gearbox (30); and a heat exchange system. The heat exchange system comprises an air-oil heat exchanger (2020) through which oil in the primary oil loop system (2000) flows; and a fuel-oil heat exchanger (1006) through which the oil in the primary oil loop system (2000) and the fuel flow such that heat is transferred between the oil and the fuel, and wherein the primary oil loop system (2000) branches such that a proportion of the oil can flow along each branch and the air-oil and fuel-oil heat exchangers are arranged in a parallel configuration on different branches of the primary oil loop system (2000); and a modulation valve (2016) arranged to allow the proportion of oil sent via each branch to be varied. The method comprises controlling (12200) the heat exchange system so as to raise the fuel temperature to at least 135°C on entry to the combustor (16) at cruise conditions.

Gas turbine engine heat exchange

A method (10000) of operating a gas turbine engine (10) is disclosed, the gas turbine engine (10) comprising a combustor (16) arranged to combust a fuel; and a fuel management system arranged to provide the fuel to the combustor (16), wherein the fuel management system comprises two fuel-oil heat exchangers through which oil and the fuel flow, the heat exchangers arranged to transfer heat between the oil and the fuel and comprising a primary fuel-oil heat exchanger (1004) and a secondary fuel-oil heat exchanger (1006); a fuel pump (1003) arranged to deliver the fuel to the combustor (16), wherein the fuel pump (1003) is located between the two fuel-oil heat exchangers; and a recirculation valve (6010) located downstream of the primary fuel-oil heat exchanger (1004), the recirculation valve (6010) arranged to allow a controlled amount of fuel which has passed through the primary fuel-oil heat exchanger (1004) to be returned to the inlet of the primary fuel-oil heat exchanger (1004). The method comprises selecting (10100) one or more fuels such that the calorific value of the fuel provided to the gas turbine engine (10) is at least 43.5 MJ/kg.

Gas turbine engine heat exchange

A method of operating a gas turbine engine including: a combustor arranged to combust a fuel; and a fuel management system arranged to provide the fuel to the combustor, wherein the fuel management system includes two fuel-oil heat exchangers through which oil and fuel flow, which are arranged to transfer heat between the oil and fuel and include primary and secondary fuel-oil heat exchangers; a fuel pump arranged to deliver the fuel to the combustor, wherein the fuel pump is located between the heat exchangers; and a recirculation valve located downstream of the primary heat exchanger, the recirculation valve arranged to allow a controlled amount of fuel which has passed through the primary heat exchanger to be returned to the inlet. The method includes selecting one or more fuels such that the calorific value of the fuel provided to the gas turbine engine is at least 43.5 MJ/kg.

Gas turbine engine and method of operating gas turbine engine

A method of operating a gas turbine engine is disclosed, the gas turbine engine comprising a combustor arranged to combust a fuel; and a fuel management system arranged to provide the fuel to the combustor. The fuel management system comprises two fuel-oil heat exchangers through which oil and the fuel flow, the heat exchangers arranged to transfer heat to the fuel and comprising a primary fuel-oil heat exchanger and a secondary fuel-oil heat exchanger; and a fuel pump arranged to deliver the fuel to the combustor, wherein the fuel pump is located between the two heat exchangers. The method comprises controlling the fuel management system so as to transfer between 200 and 600 KJ/m3 of heat to the fuel from the oil in the primary fuel-oil heat exchanger at cruise conditions.

Verfahren zum Betreiben eines Gasturbinentriebwerks

Es wird ein Verfahren (1100) zum Betreiben eines Gasturbinentriebwerks (10) bereitgestellt. Das Gasturbinentriebwerk (10) umfasst eine gestufte Brennkammer (16), die eine Anordnung von Kraftstoffsprühdüsen (403, 404) aufweist, in der ein Kraftstoffstrom während einer Neuzündungsprozedur an eine Teilmenge der Düsen (403, 404) angrenzend an einen oder mehrere Zünder (405) vorgespannt ist. Das Verfahren (1100) umfasst das Bereitstellen von Kraftstoff (1101) an die Brennkammer (16), der aromatische Inhaltsstoffe von 10 Vol.-% oder weniger aufweist. Ebenfalls offenbart ist ein Gasturbinentriebwerk (10).

Staged combustor

There is provided a method of operating a gas turbine engine. The gas turbine engine comprises a staged combustor comprising an arrangement of fuel spray nozzles in which fuel flow is biased to a subset of the nozzles adjacent one or more ignitors during a re-light procedure. The method comprises providing fuel to the combustor having a calorific value of at least 43.5 MJ/kg. Also disclosed is a gas turbine engine.

Staged combustor

There is provided a method (1000) of operating a gas turbine engine (10). The gas turbine engine (10) comprises a staged combustor (16) comprising an arrangement of fuel spray nozzles (403, 404) in which fuel flow is biased to a subset of the nozzles (403, 404) adjacent one or more ignitors (405) during a re-light procedure. The method (1000) comprises providing fuel (1001) to the combustor (16) having a calorific value of at least 43.5 MJ/kg. Also disclosed is a gas turbine engine (10).

Combustor re-light procedure

There is provided a method of operating a gas turbine engine. The gas turbine engine comprises a staged combustor having an arrangement of fuel spray nozzles in which fuel flow is biased to a subset of the nozzles adjacent one or more ignitors during a re-light procedure. The method comprises transferring heat from the oil to the fuel before the fuel enters the combustor so as to lower the fuel viscosity to 0.58 mm 2 /s or lower on entry to the combustor at cruise conditions. Also disclosed is a gas turbine engine.

Gas turbine engine for an aircraft and method of operation

A method (13000) of operating a gas turbine engine (10) is disclosed, the gas turbine engine (10) comprising an engine core (11) comprising a turbine, a compressor, a combustor (16) arranged to combust a fuel, and a core shaft connecting the turbine to the compressor; a fan located upstream of the engine core (11); a fan shaft (42); a gearbox (30) that receives an input from the core shaft and outputs drive to the fan via the fan shaft (42); a heat exchange system comprising one or more heat exchangers arranged to transfer heat to the fuel; and a fuel pump arranged to deliver the fuel to the combustor (16), wherein the fuel pump is located downstream of the one or more heat exchangers. The method (13000) comprises controlling (13200) the heat exchange system so as to raise the fuel temperature to at least 135°C on entry to the combustor (16) at cruise conditions.