Turbine engine core and bypass flows

A gas turbine engine for an aircraft comprising:

• an engine core comprising a turbine, a compressor, a core shaft connecting the turbine to the compressor, and a core exhaust nozzle having a core exhaust nozzle exit, the core exhaust nozzle having a core exhaust nozzle pressure ratio calculated using total pressure at the core exhaust nozzle exit;
• the turbine comprises a lowest pressure turbine stage having a row of rotor blades, each of the rotor blades extending radially and having a leading edge and a trailing edge;
• a fan located upstream of the engine core, the fan comprising a plurality of fan blades, wherein a fan tip radius of the fan is measured between a centreline of the gas turbine engine and an outermost tip of one of the plurality of fan blades at its leading edge, the fan tip radius being in the range 95 cm to 150 cm;
• a gearbox connected between the core shaft and the fan, the gearbox receiving an input from the core shaft and providing an output to drive the fan at a lower rotational speed than the core shaft, wherein the gearbox has a gear ratio in the range of from 3.2 to 3.8;
• and
• a nacelle surrounding the fan and the engine core and defining a bypass duct located radially outside of the engine core, the bypass duct comprising a bypass exhaust nozzle having a bypass exhaust nozzle exit, the bypass exhaust nozzle having an outer radius measured as a radial distance between the centreline of the gas turbine engine and an inner surface of the nacelle at an axial position of a rearmost tip of the nacelle, wherein an outer bypass to fan ratio of:
• the
• ⁢
• outer
• ⁢
• radius
• ⁢
• of
• ⁢
• the
• ⁢
• bypass
• ⁢
• exhaust
• ⁢
• nozzle
• the
• ⁢
• fan
• ⁢
• tip
• ⁢
• radius
• is in the range from 0.60 to 1.05, the fan tip radius being in the range from 155 cm to 200 cm, and
• the bypass exhaust nozzle having a bypass exhaust nozzle pressure ratio calculated using total pressure at the bypass exhaust nozzle exit;
• wherein a bypass to core ratio of:
• bypass
• ⁢
• exhaust
• ⁢
• nozzle
• ⁢
• pressure
• ⁢
• ratio
• core
• ⁢
• exhaust
• ⁢
• nozzle
• ⁢
• pressure
• ⁢
• ratio
• is configured to be in the range from 1.0 to 1.4 under aircraft cruise conditions, and
• wherein a fan-turbine radius difference is defined as a radial distance between:
• a point on a circle swept by a radially outer tip of the trailing edge of one of the rotor blades of the lowest pressure turbine stage;
• and a point on a circle swept by the outermost tip of the leading edge of the one of the plurality of fan blades and a fan speed to fan-turbine radius ratio defined as:
• the
• ⁢
• maximum
• ⁢
• take
• -
• off
• ⁢
• rotational
• ⁢
• speed
• ⁢
• of
• ⁢
• the
• ⁢
• fan
• ⁢
• (
• in
• ⁢
• rpm
• )
• fan
• -
• turbine
• ⁢
• radius
• ⁢
• difference
• ⁢
• (
• in
• ⁢
• mm
• )
• is in the range between 3.4 rpm/mm to 5 rpm/mm.

A gas turbine engine for an aircraft comprising:

• an engine core comprising a turbine, a compressor, a core shaft connecting the turbine to the compressor, and a core exhaust nozzle having a core exhaust nozzle exit, the core exhaust nozzle having a core exhaust nozzle pressure ratio calculated using total pressure at the core exhaust nozzle exit;
• the turbine comprises a lowest pressure turbine stage having a row of rotor blades, each of the rotor blades extending radially and having a leading edge and a trailing edge;
• a fan located upstream of the engine core, the fan comprising a plurality of fan blades, wherein a fan tip radius of the fan is measured between a centreline of the gas turbine engine and an outermost tip of one of the plurality of fan blades at its leading edge, the fan tip radius being in the range 95 cm to 150 cm;
• a gearbox connected between the core shaft and the fan, the gearbox receiving an input from the core shaft and providing an output to drive the fan at a lower rotational speed than the core shaft, wherein the gearbox has a gear ratio in the range of from 3.2 to 3.8;
• and
• a nacelle surrounding the fan and the engine core and defining a bypass duct located radially outside of the engine core, the bypass duct comprising a bypass exhaust nozzle having a bypass exhaust nozzle exit, the bypass exhaust nozzle having an outer radius measured as a radial distance between the centreline of the gas turbine engine and an inner surface of the nacelle at an axial position of a rearmost tip of the nacelle, and
• the bypass exhaust nozzle having a bypass exhaust nozzle pressure ratio calculated using total pressure at the bypass exhaust nozzle exit;
• wherein a bypass to core ratio of:
• bypass
• ⁢
• exhaust
• ⁢
• nozzle
• ⁢
• pressure
• ⁢
• ratio
• core
• ⁢
• exhaust
• ⁢
• nozzle
• ⁢
• pressure
• ⁢
• ratio
• is configured to be in the range from 1.1 to 2.0 under aircraft cruise conditions, and
• wherein a fan-turbine radius difference is defined as a radial distance between:
• a point on a circle swept by a radially outer tip of the trailing edge of one of the rotor blades of the lowest pressure turbine stage;
• and a point on a circle swept by the outermost tip of the leading edge of the one of the plurality of fan blades and a fan speed to fan-turbine radius ratio defined as:
• the
• ⁢
• maximum
• ⁢
• take
• -
• off
• ⁢
• rotational
• ⁢
• speed
• ⁢
• of
• ⁢
• the
• ⁢
• fan
• ⁢
• (
• in
• ⁢
• rpm
• )
• fan
• -
• turbine
• ⁢
• radius
• ⁢
• difference
• ⁢
• (
• in
• ⁢
• mm
• )
• is in the range between 0.8 rpm/mm to 5 rpm/mm;
• and
• wherein the temperature of the flow at the entry to the exit of the combustor at cruise conditions is in a range of 1600K to 1650K.