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.

A gas turbine engine is provided having: a turbomachine; a fan section having a fan rotatable by the turbomachine; a nacelle enclosing the fan; and an engine controller positioned within the nacelle. The nacelle defines an inner surface radius (r) along the radial direction inward of the engine controller, wherein the engine controller defines a radial height (r) along the radial direction, a total volume (V), and a normalized radius (r). The normalized radius (r) is a ratio of the inner surface radius (r) to the total volume (V) to cube root, and wherein these parameters are related by the following equation:

[00001]$0.1{\left(r^{}\right)}^{-1}<\frac{r}{r}<{K\left(r^{}\right)}^{-4/3},$

wherein the normalized radius (r) is between 1.25 and 8 and K is equal to 40%, or the normalized radius (r) is between 2.75 and 4.5 and K is equal to 65%.

A method of blending at least two fuels includes providing at least two fuels to a cyclonic mixer via a fuel supply system, the fuel supply system including a fuel supply circuit for each fuel of the at least two fuels, mixing, via at least one vortex formed in the cyclonic mixer, the at least two fuels to form a fuel mixture, determining, via one or more sensors, a measured interchangeability index of the fuel mixture, comparing the measured interchangeability index to a predetermined interchangeability index, adjusting, via the fuel supply system, one or more parameters of at least one of the at least two fuels based on the comparison between the measured interchangeability index and the predetermined interchangeability index, and providing the fuel mixture to a combustion system.

A compressor including one or more impellers configured to draw in a refrigerant in an axial direction and to compress the refrigerant in a centrifugal direction; a rotary shaft, to which the one or more impellers and a motor that rotates the one or more impellers are coupled; a thrust bearing that supports the rotary shaft; a bearing state sensor configured to detect a surface roughness of the thrust bearing; and a controller configured to control the motor based on the surface roughness of the thrust bearing.

Systems and methods to control gas turbine firing temperatures during air injection. A method of achieving a desired firing temperature of a gas turbine engine during air injection comprises injecting compressed air into the gas turbine engine using an external source. The external source includes a compressor and a recuperator. The method comprises using a controller of the gas turbine engine to: (a) determine an air injection exhaust bias gain using an inlet temperature of the gas turbine engine; (b) calculate, based on the determined air injection exhaust bias gain and a flow rate of the injected compressed air, an air injection exhaust curve bias; and (c) change a fuel flow of the gas turbine engine by adding the air injection exhaust curve bias to an existing exhaust curve of the gas turbine engine to thereby achieve the desired firing temperature during air injection.

A propulsion system includes at least two propulsors. The at least two propulsors each comprising a fan having a plurality of fan blades. A controller includes memory and one or more processors. The memory stores instructions that when executed by the one or more processors cause the system to perform the following: determine a pairwise phase difference between one propulsor of the at least two propulsors and another propulsor of the at least two propulsors; generate a reference phase angle; determine a target phase shift for each propulsor of the at least two propulsors; and adjust a speed of each propulsor of the at least two propulsors based on the target phase shift until the pairwise phase difference is equal to the reference phase angle.

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

A pitch control system includes: a coarse pitch pressure circuit configured to receive a pressurised fluid; a pitchlock mechanism configured to pitchlock a propeller blade of a propeller; and a controller configured to detect undesired feathering of a propeller blade. The controller is configured to, in response to detecting undesired feathering of the propeller blade, reduce the pressure of the pressurised fluid received by the coarse pitch pressure circuit so as to stop the propeller blade feathering and simultaneously cause the pitchlock mechanism to pitchlock the propeller blade.

A system for neural network compensated aero-thermodynamic gas turbine engine parameter/inlet condition synthesis. The system includes an aero-thermodynamic engine model configured to produce a real-time model-based estimate of engine parameters, a machine learning model configured to generate model correction errors indicating the difference between the real-time model-based estimate of engine parameters and sensed values of the engine parameters, and a comparator configured to produce residuals indicating a difference between the real-time model-based estimate of engine parameters and the sensed values of the engine parameters. The system also includes an inlet condition estimator configured to iteratively adjust an estimate of inlet conditions based on the residuals and adaptive control laws configured to produce engine control parameters for control of gas turbine engine actuators based on the inlet conditions.

Systems and methods to control gas turbine firing temperatures during air injection. A method of achieving a desired firing temperature of a gas turbine engine during air injection comprises injecting compressed air into the gas turbine engine using an external source. The external source includes a compressor and a recuperator. The method comprises using a controller of the gas turbine engine to: (a) determine an air injection exhaust bias gain using an inlet temperature of the gas turbine engine; (b) calculate, based on the determined air injection exhaust bias gain and a flow rate of the injected compressed air, an air injection exhaust curve bias; and (c) change a fuel flow of the gas turbine engine by adding the air injection exhaust curve bias to an existing exhaust curve of the gas turbine engine to thereby achieve the desired firing temperature during air injection.