A gas turbine engine including: a high pressure turbine, a low pressure turbine, a high pressure compressor coupled to the high pressure turbine by a high pressure shaft, a propulsor and a low pressure compressor coupled to the low pressure turbine via a low pressure shaft and a reduction gearbox; wherein the high pressure compressor defines an average stage pressure ratio at cruise conditions of between 1.25 and 1.35 and consists of 10 or 11 stages; and the high pressure compressor and low pressure compressor together define a core overall pressure ratio at cruise conditions of between 40:1 and 60:1.

A gas turbine engine including an engine core with high and low pressure compressors and high and low pressure turbines, the high pressure compressor and high pressure turbine being coupled by a high pressure shaft, and the low pressure compressor and low pressure turbine being coupled by a low pressure shaft, the engine further including a fan coupled to the low pressure shaft by a reduction gearbox. The high pressure compressor consists of 8 or 9 stages and has an average stage loading at cruise conditions of between 1.39 and 1.42, and the low and high pressure compressor together define an overall core pressure ratio at cruise conditions of between 40 and 60.

A gas turbine engine includes an engine core having high and low pressure compressors and high and low pressure turbines. The engine further includes a fan coupled to a low pressure shaft and the low pressure turbine by a reduction gearbox. The low pressure compressor includes no more than two compressor stages, and the low and high pressure compressor together define a cruise overall core pressure ratio of between 30 and 50.

A gas turbine engine for an aircraft includes an engine core including an engine core, a turbine, a compressor, and a core shaft connecting the turbine to the compressor, the turbine comprising a lowest pressure rotor stage, the turbine having a turbine diameter at the lowest pressure rotor stage. A fan located upstream of the engine core, the fan comprising a plurality of fan blades extending from a hub, the hub and fan blades together defining a fan face having a fan face area and a fan tip radius. A ratio of the fan tip radius to the turbine diameter at the lowest pressure rotor stage is in the range from 1.2 to 2.0.

A gas turbine engine for an aircraft includes an engine core including a turbine, compressor, and core shaft connecting turbine to compressor; a fan located upstream of the engine core and including a plurality of fan blades each having a leading and trailing edge. The turbine includes a lowest pressure turbine stage having a row of rotor blades, each rotor blades extending radially and having a leading and trailing edge. The engine has a fan tip axis that joins a radially outer tip of the leading edge of a fan blade and the radially outer tip of the trailing edge of a rotor blade of the lowest pressure stage. The fan tip axis lies in a longitudinal plane which contains a centreline of engine. A fan axis angle is defined as the angle between fan tip axis and centreline, and is in a range between 10 and 20 degrees.

A power generation system includes a shroud that defines a fluid flow path. A compressor is in the fluid flow path, and a combustor is in the fluid flow path downstream from the compressor. A turbine is in the fluid flow path downstream from the compressor and the combustor. An electric generator is in the fluid flow path upstream from the compressor, and the electric generator includes a rotor coaxially aligned with the turbine.

A small satellite propulsion system using a gaseous oxidizer and a gaseous fuel as primary propellants with a liquid as a film coolant for the inner surface of the rocket motor. The gaseous fuel is also used as a pressurant for the coolant and as a cold gas propellant for attitude control system (hereinafter ACS) thrusters. The oxidizer, fuel, and coolant tanks, as well as most valves and plumbing, are integrated into a single core unit along with the rocket motor, rocket motor plumbing, and safety valves. Attitude control thrusters may be remotely located with plumbing to the fuel tank. The core unit is four inches high and less than four inches deep and wide. The small satellite propulsion system uses no pyrotechnics and no hazardous toxic materials.

A gas turbine engine for an aircraft includes an engine core, fan, air intake, nacelle, and gearbox. The core includes a turbine, compressor, and core shaft connecting the turbine and compressor. The fan is upstream of the core and includes a plurality of fan blades, and has a diameter greater than 2.0 m. The air intake is upstream of the fan and has ratio of intake length to fan diameter of 0.20 to 0.60 and defines highlight, throat and diffuser areas. The nacelle at least partially surrounds the core and fan. The gearbox receives input from the core shaft and outputs drive to the fan to drive the fan at a lower rotational speed than the core shaft. The engine has a bypass ratio greater than 10. The nacelle has a length and the ratio of the length of the nacelle to the fan diameter is 0.4 to 2.5.

A turbine generator contains a turbine part and a generator part. The turbine contains a turbine wheel. A sealing arrangement is arranged between the turbine wheel and the generator part, the sealing effect of which varies during operation. The generator part further has a generator shaft, which is supported by an axial bearing configured as a magnetic bearing with two coils axially spaced apart from each other. A bearing ring is arranged between these coils with an axial distance from the coils. To ensure a safe operation, a setpoint value for the axial distance is varied to change the sealing effect of the sealing arrangement. Alternatively or additionally, it is provided that when a current threshold of a coil current is exceeded, a control signal is emitted to control the flow of the medium or the rotational speed.

A gas turbine engine for an aircraft includes an engine core including a turbine, compressor, and core shaft connecting turbine to compressor; a fan located upstream of the engine core and including a plurality of fan blades each having a leading and trailing edge. The turbine includes a lowest pressure turbine stage having a row of rotor blades, each rotor blades extending radially and having a leading and trailing edge. The engine has a fan tip axis that joins a radially outer tip of the leading edge of a fan blade and the radially outer tip of the trailing edge of a rotor blade of the lowest pressure stage. The fan tip axis lies in a longitudinal plane which contains a centreline of engine. A fan axis angle is defined as the angle between fan tip axis and centreline, and is in a range between 10 and 20 degrees.