A gas turbine engine has a compression system radius ratio defined as the ratio of the radius of the tip of a fan blade to the radius of the tip of the most downstream compressor blade in the range of from 5 to 9. This results in an optimum balance between installation benefits, operability, maintenance requirements and engine efficiency when the gas turbine engine is installed on an aircraft.

A gas turbine engine has a compression system radius ratio defined as the ratio of the radius of the tip of a fan blade to the radius of the tip of the most downstream compressor blade in the range of from 5 to 9. This results in an optimum balance between installation benefits, operability, maintenance requirements and engine efficiency when the gas turbine engine is installed on an aircraft.

An exemplary aircraft includes a turbine engine having a gas generator spool and a power spool, the power spool operational to drive a rotor, a first generator coupled to the gas generator spool, and a controller operable to increase a load on the gas generator spool when the gas generator spool is on a speed limit thereby increasing a speed limit margin in order to increase power available from the turbine engine.

An exemplary aircraft includes a turbine engine having a gas generator spool and a power spool, the power spool operational to drive a rotor, a first generator coupled to the gas generator spool, and a controller operable to increase a load on the gas generator spool when the gas generator spool is on a speed limit thereby increasing a speed limit margin in order to increase power available from the turbine engine.

A method of determining one or more fuel characteristics of an aviation fuel for powering a gas turbine engine of an aircraft includes: measuring one or more trace substance parameters of the fuel, the one or more trace substance parameters each associated with a respective trace substance in the fuel; and determining one or more fuel characteristics of the fuel based on the one or more trace substance parameters. Further, a fuel characteristic determination system, a method of operating an aircraft, and an aircraft.

A method of determining one or more fuel characteristics of an aviation fuel for powering a gas turbine engine of an aircraft includes: measuring one or more trace substance parameters of the fuel, the one or more trace substance parameters each associated with a respective trace substance in the fuel; and determining one or more fuel characteristics of the fuel based on the one or more trace substance parameters. Further, a fuel characteristic determination system, a method of operating an aircraft, and an aircraft.

A cabin blower for an aircraft, the system comprising: a cabin blower compressor; an electric machine; and a controller configured to control the cabin blower system so that: in a cabin blower mode of operation, the cabin blower compressor is driven by power extracted from one or more spools of a gas turbine engine of the aircraft and provides a flow of air to a cabin of the aircraft. The controller may be further configured to control the system so that: in a rotor bow mitigation mode of operation, the cabin blower compressor is driven by the electric machine using electrical power from an electrical power source and provides a flow of air through a core of the gas turbine engine to remove heat from the core. A method of operating a cabin blower system of an aircraft is also provided.

The invention relates to a device for the rapid reactivation of a helicopter turbine engine (6), characterised in that it comprises a pneumatic turbine (7) mechanically connected to said turbine engine (6) so as to be able to rotate it and ensure reactivation thereof; a pneumatic storage (9) connected to said pneumatic turbine (7) by means of a pneumatic circuit (10) for supplying pressurised gas to said pneumatic turbine (7); a controlled fast-opening pneumatic valve (11) arranged on the pneumatic circuit (10) between said storage (9) and said pneumatic turbine (7) and suitable for being on demand placed at least in an open position in which the gas can supply said pneumatic turbine (7), or in a closed position in which said pneumatic turbine (7) is no longer supplied with pressurised gas.

The invention relates to a device for the rapid reactivation of a helicopter turbine engine (6), characterised in that it comprises a pneumatic turbine (7) mechanically connected to said turbine engine (6) so as to be able to rotate it and ensure reactivation thereof; a pneumatic storage (9) connected to said pneumatic turbine (7) by means of a pneumatic circuit (10) for supplying pressurised gas to said pneumatic turbine (7); a controlled fast-opening pneumatic valve (11) arranged on the pneumatic circuit (10) between said storage (9) and said pneumatic turbine (7) and suitable for being on demand placed at least in an open position in which the gas can supply said pneumatic turbine (7), or in a closed position in which said pneumatic turbine (7) is no longer supplied with pressurised gas.

An oil tank system (100) for a gas turbine engine is provided. The oil tank system (100) includes an oil tank (102) having an upper tank portion (112) and a lower tank portion (114), a waisted section (118) being provided between the upper tank portion (112) and the lower tank portion (114). Oil is received by a de-aerator (104) of the system (100) which supplies de-aerated oil to the upper tank portion (112). The waisted section (118) includes an upper face (119) configured to catch oil drips from above the waisted section (118) and to guide oil to a lower face (121) of the waisted section (118).