Gas turbine engine compression system with core compressor pressure ratio
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.
1. A gas turbine engine for an aircraft, the gas turbine engine comprising:
an engine core comprising a turbine, a compressor, and a combustor;
a fan comprising a plurality of fan blades; and
a gearbox that receives an input from a core shaft that is connected to at least a part of the turbine, the gearbox outputting drive to the fan so as to drive the fan at a lower rotational speed than the core shaft, wherein
a compression system speed ratio is defined as a ratio of a rotational speed of a most downstream compressor blade to a rotational speed of the fan at cruise conditions;
the compression system speed ratio is greater than 6;
a diameter of the fan is greater than 280 cm;
a gear ratio of the gearbox is greater than 4.2; and
the maximum net thrust of the engine at standard atmospheric conditions at sea level plus 15 degrees C. with the engine static is no greater than 160 kN.
2. The gas turbine engine according to claim 1 , wherein an overall pressure ratio defined as the ratio of the stagnation pressure before entry into the combustor to the stagnation pressure upstream of the fan is in a range of 35 to 50 at cruise conditions.
3. The gas turbine engine according to claim 1 , wherein an overall pressure ratio defined as the ratio of the stagnation pressure before entry into the combustor to the stagnation pressure upstream of the fan is in a range of 35 to 45 at cruise conditions.
4. The gas turbine engine according to claim 1 , wherein a fan root pressure ratio, defined as the ratio of the mean total pressure of the flow at the fan exit that subsequently flows through the engine core to the mean total pressure of the flow at the fan inlet, is no greater than 1.2 at cruise conditions.
5. The gas turbine engine according to claim 3 , wherein a fan root pressure ratio, defined as the ratio of the mean total pressure of the flow at the fan exit that subsequently flows through the engine core to the mean total pressure of the flow at the fan inlet, is no greater than 1.2 at cruise conditions.
6. The gas turbine engine according to claim 1 , wherein the fan comprises no more than 14 fan blades.
7. The gas turbine engine according to claim 1 , wherein the diameter of the fan is in the range 370 cm to 420 cm.
8. The gas turbine engine according to claim 1 , wherein a fan pressure ratio, defined as the ratio of the mean total pressure of the flow at the fan exit to the mean total pressure of the flow at the fan inlet, is no greater than 1.35 at cruise conditions.
9. A gas turbine engine for an aircraft, the gas turbine engine comprising:
an engine core comprising a turbine, a compressor, and a combustor;
a fan comprising a plurality of fan blades; and
a gearbox that receives an input from a core shaft that is connected to at least a part of the turbine, the gearbox outputting drive to the fan so as to drive the fan at a lower rotational speed than the core shaft, wherein
a compression system speed ratio is defined as a ratio of a rotational speed of a most downstream compressor blade to a rotational speed of the fan at cruise conditions;
the compression system speed ratio is greater than 6;
a diameter of the fan is greater than 280 cm;
an overall pressure ratio defined as the ratio of the stagnation pressure before entry into the combustor to the stagnation pressure upstream of the fan is in a range of 35 to 45 at cruise conditions; and
a compression system radius ratio, defined as a ratio of a radius of a tip of a fan blade of the plurality of fan blades to a radius of a tip of the most downstream compressor blade, is at least 6.
10. The gas turbine engine according to claim 9 , wherein the overall pressure ratio is in a range of 35 to 40 at cruise conditions.
11. The gas turbine engine according to claim 9 , wherein a fan root pressure ratio, defined as the ratio of the mean total pressure of the flow at the fan exit that subsequently flows through the engine core to the mean total pressure of the flow at the fan inlet, is no greater than 1.2 at cruise conditions.
12. The gas turbine engine according to claim 9 , wherein a specific thrust of the engine, defined as the net thrust of the engine divided by the total mass flow through the engine, is less than 80 Nkg −1 s at cruise conditions.
13. The gas turbine engine according to claim 9 , wherein a fan pressure ratio, defined as the ratio of the mean total pressure of the flow at the fan exit to the mean total pressure of the flow at the fan inlet, is no greater than 1.35 at cruise conditions.
14. The gas turbine engine according to claim 9 , wherein the diameter of the fan is in the range 350 cm to 420 cm.
15. A gas turbine engine for an aircraft, the gas turbine engine comprising:
an engine core comprising a turbine, a compressor, and a combustor;
a fan comprising a plurality of fan blades; and
a gearbox that receives an input from a core shaft that is connected to at least a part of the turbine, the gearbox outputting drive to the fan so as to drive the fan at a lower rotational speed than the core shaft, wherein
a compression system speed ratio is defined as a ratio of a rotational speed of a most downstream compressor blade to a rotational speed of the fan at cruise conditions;
the compression system speed ratio is greater than 6;
a diameter of the fan is greater than 280 cm;
a fan root pressure ratio, defined as the ratio of the mean total pressure of the flow at the fan exit that subsequently flows through the engine core to the mean total pressure of the flow at the fan inlet, is no greater than 1.2 at cruise conditions; and
a compression system radius ratio, defined as a ratio of a radius of a tip of a fan blade of the plurality of fan blades to a radius of a tip of the most downstream compressor blade, is at least 6.
16. The gas turbine engine according to claim 15 , wherein:
a gear ratio of the gearbox is greater than 4.2;
the diameter of the fan is in the range 350 cm to 420 cm; and
an overall pressure ratio defined as the ratio of the stagnation pressure before entry into the combustor to the stagnation pressure upstream of the fan is in a range of 35 to 45 at cruise conditions.
17. The gas turbine engine according to claim 15 , wherein a specific thrust of the engine, defined as the net thrust of the engine divided by the total mass flow through the engine, is less than 80 Nkg −1 s at cruise conditions.
18. The gas turbine engine according to claim 15 , wherein the fan blades are surrounded by a nacelle.