Abstract: 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.

Abstract: 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.

Abstract: 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.

Abstract: 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.

Abstract: 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

Abstract: A method of controlling a propulsion system of an aircraft, the propulsion system comprising a gas turbine engine arranged to be powered by a fuel and at least one variable inlet guide vane—VIGV, comprises obtaining at least one fuel characteristic of the fuel being provided to the gas turbine engine; and making a change to scheduling of the at least one VIGV based on the at least one obtained fuel characteristic.

Abstract: 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.

Abstract: A method of determining a calorific value of fuel supplied to a gas turbine engine of an aircraft comprises sensing at least one engine parameter during a first time period of aircraft operation during which the gas turbine engine uses the fuel; and, based on the at least one sensed engine parameter, determining a calorific value of the fuel. The sensing may be repeated such that the at least one engine parameter is monitored over time. The gas turbine engine may be a propulsive gas turbine engine of the aircraft or a gas turbine engine of an auxiliary power unit of the aircraft.

Abstract: A propulsion system for an aircraft includes a gas turbine engine and a fuel tank, wherein the fuel includes at least a proportion of a sustainable aviation fuel—SAF—having a density between 90% and 98% of the density, ?K, of kerosene and a calorific value between 101% and 105% the calorific value CVK, of kerosene. The engine includes a combustor; and a fuel pump arranged to supply a fuel thereto at an energy flow rate, C, the pump being arranged to output fuel at a volumetric flow rate, Q, the percentage of fuel passing through the pump not provided to the combustor being referred to as a spill percentage. The fuel include X % SAF, where X % is in the range from 5% to 100%, and has a density, ?F, and a calorific value CVF.

Abstract: There is disclosed a gas cooler 20 for providing high-pressure sealing gas to a bearing chamber. The cooler comprises a turbine 22; a turbine inlet 24 arranged to receive gas to drive the turbine and a turbine outlet 26 arranged to deliver gas output from the turbine; a compressor 28 arranged to be driven by the turbine 22; a compressor inlet 30 arranged to receive gas to be compressed by the compressor and a compressor outlet 32 arranged to deliver gas output from the compressor 28; and a cooler outlet 36 in fluid communication with the turbine outlet 26 and the compressor outlet 32 so as to deliver high-pressure sealing gas comprising gas merged from the turbine outlet 26 and the compressor outlet 32.

Abstract: There is disclosed a gas cooler 20 for providing high-pressure sealing gas to a bearing chamber. The cooler comprises a turbine 22; a turbine inlet 24 arranged to receive gas to drive the turbine and a turbine outlet 26 arranged to deliver gas output from the turbine; a compressor 28 arranged to be driven by the turbine 22; a compressor inlet 30 arranged to receive gas to be compressed by the compressor and a compressor outlet 32 arranged to deliver gas output from the compressor 28; and a cooler outlet 36 in fluid communication with the turbine outlet 26 and the compressor outlet 32 so as to deliver high-pressure sealing gas comprising gas merged from the turbine outlet 26 and the compressor outlet 32.

Abstract: An arrangement for controlling flow of fluid to at least one component of a gas turbine engine, the arrangement includes a first conduit for providing fluid to the component; a flow valve, for controlling the flow of fluid in the first conduit, having first and second configurations; a second conduit for providing fluid to the flow valve to the control the configuration of the flow valve; and a magnetic valve, for controlling the flow of fluid in the second conduit, having first and second configurations and including at least one ferromagnet forming a portion of a magnetic circuit; the member comprising ferromagnetic material is thermally coupled to one of the group comprising the fluid in the first conduit, the fluid in the second conduit and the component; and the configuration of the flow valve is dependent on the configuration of the magnetic valve.

Abstract: In order to improve cooling of a blade 31, 41/disc 32, 42 combination, a pre-swirler feature 40, 50 is utilised in order to present coolant air through presentational pathways typically incorporating holes 37 or grooves/slots such that coolant air displaces hotter air leaked past a seal 35, 45. The swirler features 40, 50 are located within stationary housing components 34, 44 such that the coolant air is propelled at substantially the same rotational speed as the disc 32, 42 whilst presentation of the coolant air flow A, C, E is such that hot air displacement is achieved at least adjacent front faces 36, 46 of blade roots 33, 43.

Abstract: An arrangement for controlling flow of fluid to at least one component of a gas turbine engine, the arrangement comprising: a first conduit for providing fluid to the component; a flow valve, for controlling the flow of fluid in the first conduit, having first and second configurations; a second conduit for providing fluid to the flow valve to the control the configuration of the flow valve; and a magnetic valve, for controlling the flow of fluid in the second conduit, having first and second configurations and including at least one ferromagnet forming a portion of a magnetic circuit, whereby the configuration of the magnetic valve depends on the temperature of the ferromagnet; the member comprising ferromagnetic material (68) is thermally coupled to one of the group comprising the fluid in the first conduit (24), the fluid in the second conduit (54) and the component; and the configuration of the flow valve is dependent on the configuration of the magnetic valve.

Abstract: A mounting arrangement 31 incorporates a blade pocket cavity 35, 135, 235 between a blade platform 34, 134, 234 and a rotor disc post 53. Coolant air is presented within the cavity 35, 135, 235 in order to provide cooling of the platform 34, 134, 234. In order to inhibit ingress of hot air from adjacent wheel spaces 36, an air curtain is created across an open end 45, 145 whereby coolant air is contained within the cavity 35, 135, 235. Air flow from the air curtain itself may pass along a surface 48, 148 to provide more direct cooling for the platform 34, 134, 234. Apertures 49, 50 may be provided in order to bleed coolant air from the cavity 35 and also draw an air flow 51 across the surface 48 for greater cooling efficiency.

Abstract: A gas turbine engine (10) is provided with a stage of turbine blades (22) surrounded by hollow shroud segments (28). Cooling of the shroud segments (28) is achieved by providing a flow of engine leakage air from space volume (48) centrally of the engine (10) via an annular compartment (46) wherein it is diverted away from the blade cooling system and passed via piping (44) within each guide vane (20) to shroud segments (28). The flow impingement cools the shroud segments (28) and then passes into the gas annulus via slots (60).

Abstract: A gas turbine engine (10) is provided with a stage of turbine blades (22) surrounded by hollow shroud segments (28). Cooling of the shroud segments (28) is achieved by providing a flow of engine leakage air from space volume (48) centrally of the engine (10) via an annular compartment (46) wherein it is diverted away from the blade cooling system and passed via piping (44) within each guide vane (20) to shroud segments (28). The flow impingement cools the shroud segments (28) and then passes into the gas annulus via slots (60).