Turbine engine
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.
1. A gas turbine engine for an aircraft, comprising:
an engine core having a core length and comprising 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; and
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,
wherein a ratio of the fan tip radius to the turbine diameter at the lowest pressure rotor stage is in a range from 1.2 to 2.0; and wherein the ratio of the fan tip radius to the core length is in a range from 0.3 to 1.
2. The gas turbine engine of claim 1 , wherein the ratio of the fan tip radius to the turbine diameter is in a range from 1.2 to 1.7.
3. The gas turbine engine of claim 1 , wherein the ratio of the fan tip radius to the turbine diameter is in a range from 1.2 to 1.5.
4. The gas turbine engine of claim 1 , wherein the fan tip radius is measured between a centreline of the gas turbine engine and an outermost tip of each fan blade at its leading edge, the fan face area being equal to 7 L multiplied by the square of the fan tip radius.
5. The gas turbine engine of claim 1 wherein the fan tip radius is 110 cm, 115 cm, 120 cm, 125 cm, 130 cm, 135 cm, 140 cm, 145 cm, 150 cm, 155 cm, 160 cm, 165 cm, 170 cm, 175 cm, 180 cm, 185 cm, 190 cm or 195 cm.
6. The gas turbine engine of claim 1 wherein the fan tip radius is in a range from 110 cm to 150 cm, and wherein the engine area ratio is in a range from 1.7 to 2.7.
7. The gas turbine engine of claim 1 wherein the fan tip radius is in a range from 110 cm to 150 cm.
8. The gas turbine engine of claim 1 , wherein the turbine diameter of the turbine at the lowest pressure rotor stage is measured:
(i) if the rotor is shrouded, to an underside of the shroud; or
(ii) if the rotor is unshrouded, to blade tips of the rotor.
9. The gas turbine engine of claim 1 , wherein the turbine diameter at the lowest pressure rotor stage is in a range from 70 cm to 170 cm.
10. The gas turbine engine of claim 1 , wherein the core length is defined as an axial distance between a forward region of the compressor and a rearward region of the turbine.
11. The gas turbine engine of claim 1 , wherein the core length is measured along a centreline of the gas turbine engine from a mean radius point of a first stage of a compressor blade leading edge to a mean radius point of the lowest pressure turbine rotor stage blade trailing edge of the turbine.
12. The gas turbine engine of claim 1 , wherein the engine core length is in a range from 150 cm to 320 cm.
13. The gas turbine engine according to claim 1 , wherein:
the turbine is a first turbine, the compressor is a first compressor, and the core shaft is a first core shaft;
the engine core further comprises a second turbine, a second compressor, and a second core shaft connecting the second turbine to the second compressor; and
the second turbine, second compressor, and second core shaft are arranged to rotate at a higher rotational speed than the first core shaft.