Gas turbine engine compression system
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 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 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 is in the range of from 5 to 9,
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 in a range of 1.35 to 1.43 at cruise conditions, and
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 in a range of 1.18 to 1.3 at cruise conditions.
2. A gas turbine engine according to claim 1 , wherein the compression system radius ratio is in the range of from 5.2 to 8.5.
3. The gas turbine engine according to claim 1 , wherein the radius of the tip of the fan blade is in the range of from 120 cm to 210 cm.
4. The gas turbine engine according to claim 1 , wherein:
the radius of the tip of the fan blades is in the range of from 160 cm to 190 cm and the compression system radius ratio is in the range of from 5.3 to 7.7; or
the radius of the tip of the fan blades is in the range of from 120 cm to 140 cm and the compression system radius ratio is in the range of from 5.1 to 6.6.
5. The gas turbine engine according to claim 1 , wherein the gearbox has a reduction ratio in the range of from 3.2 to 3.8, such that the rotational speed of the gearbox input core shaft is in the range of from 3.2 to 3.8 times the rotational speed of the fan.
6. The gas turbine engine according to claim 1 , wherein:
a bypass duct is defined radially outside the core, with the leading edge of a splitter defining the point at which flow splits into core flow and bypass flow;
an engine core radius ratio is defined as the ratio of the radius of the tip of the most downstream turbine blade in the engine to the radius of the leading edge of the splitter; and
the compression system radius ratio divided by the engine core radius ratio is in the range of from 5.5 to 10.
7. The gas turbine engine according to claim 1 , wherein:
a bypass duct is defined radially outside the core, with the leading edge of a splitter defining the point at which flow splits into core flow and bypass flow in use;
a core compressor aspect ratio is defined as the ratio of the axial distance between the leading edge of the splitter and the leading edge of the tip of the most downstream compressor blade to the radius of the leading edge of the splitter; and
the compression system radius ratio divided by the core compressor aspect ratio is in the range of from 1.7 to 4.2.
8. The gas turbine engine according to claim 7 , wherein:
the radius of the tip of the fan blades is in the range of from 160 cm to 190 am and the compression system radius ratio divided by the core compressor aspect ratio is in the range of from 1.9 to 2.9; or
the radius of the tip of the fan blades is in the range of from 120 cm to 140 cm and the compression system radius ratio divided by the core compressor aspect ratio is in the range of from 1.7 to 2.4.
9. The gas turbine engine according to claim 1 , wherein the ratio of the radius of a fan blade at its hub to the radius of the fan blade at its tip is in the range of from 0.23 to 0.33.
10. The gas turbine engine according to claim 1 , wherein the fan blades comprise a main body attached to a leading edge sheath, the main body and the leading edge sheath being formed using different materials, wherein, optionally the leading edge sheath material comprises titanium and/or the main body material comprises carbon fibre or an aluminium alloy.
11. The gas turbine engine according to claim 1 , further comprising an intake that extends upstream of the fan blades, wherein:
an intake length L is defined as the axial distance between the leading edge of the intake and the leading edge of the tip of the fan blades;
the fan diameter D is the diameter of the fan at the leading edge of the tips of the fan blades; and
the ratio L/D is in the range of from 0.2 to 0.45.
12. The gas turbine engine according to claim 1 , wherein:
a compression system speed ratio is defined as the ratio of the rotational speed of the most downstream compressor blade to the rotational speed of the fan at cruise conditions; and
the product of the compression system radius ratio and the compression system speed ratio is in the range of from 25 to 80.
13. The gas turbine engine according to claim 12 , wherein the compression system speed ratio is in the range of from 6 to 10.
14. The gas turbine engine according to claim 12 , wherein the cruise conditions correspond to:
a forward Mach number of 0.85; and
international standard atmospheric conditions at 35000 ft (10668 m).
15. The gas turbine engine according to claim 1 , wherein:
a bypass ratio, defined as the ratio of the mass flow rate through a bypass duct defined outside the engine core and the mass flow rate through the engine core at cruise conditions, is in the range of from 10 to 20.
16. The gas turbine engine according to claim 1 , wherein:
a fan tip pressure ratio is defined as the ratio of the mean total pressure of the flow at the fan exit that subsequently flows through the bypass duct defined radially outside the engine core to the mean total pressure of the flow at the fan inlet; and
the ratio between the fan root pressure ratio to the fan tip pressure ratio at cruise conditions is less than 0.95.
17. The gas turbine engine according to claim 1 , wherein:
the core comprises a first turbine connected to a first compressor by a first core shaft;
a second turbine connected to a second compressor by a second core shaft;
the second turbine, second compressor, and second core shaft are arranged to rotate at a higher rotational speed than the first core shaft; and
the most downstream compressor blade is a part of the second compressor.
18. An aircraft comprising at least one gas turbine engine according to claim 1 .