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;
a bypass duct defined radially outside the core; 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 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 no greater than 6.5;
a core compressor pressure ratio, defined as a pressure ratio across the compressor at cruise conditions excluding a pressure rise due to the fan, is in a range of 35 to 38 at cruise conditions; and
a fan root to tip 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 exit that subsequently flows through the bypass duct, is in a range 0.8 to 0.95 at cruise conditions.
2 . The gas turbine engine according to claim 1 , wherein the fan root to tip pressure ratio is in a range of 0.83 to 0.88.
3 . The gas turbine engine according to claim 1 , wherein
an engine core radius ratio is defined as the ratio of the radius of a tip of a 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 no greater than 7.
4 . The gas turbine engine according to claim 1 , wherein: an engine core radius ratio, defined as the ratio of the radius of a tip of a most downstream turbine blade in the engine to the radius of the leading edge of the splitter, is in a range of 0.8 to 0.98.
5 . The gas turbine engine according to claim 1 , wherein
an engine core radius ratio, defined as the ratio of the radius of a tip of a most downstream turbine blade in the engine to the radius of the leading edge of the splitter, is less than 1; and
the compression system radius ratio divided by the engine core radius ratio is no greater than 7.
6 . The gas turbine engine according to claim 1 , wherein the compression system radius ratio is 5.3.
7 . The gas turbine engine according to claim 1 , wherein:
the gearbox is a planetary gearbox in which a planet carrier is coupled to an output shaft and a ring gear is fixed;
a gear ratio of the gearbox is in the range 3.3 to 3.9; and
a diameter of the fan is no greater than 230 cm.
8 . The gas turbine engine according to claim 1 , wherein the compression system radius ratio is in a range of 5.6 to 6.5.
9 . The gas turbine engine according to claim 1 , wherein:
the bypass duct has a fixed area nozzle; and
each fan blade comprises a carbon-fibre body with a protective leading edge.
10 . The gas turbine engine according to claim 1 , wherein the radius of the tip of the most downstream compressor blade is in a range 20 cm to 25 cm.
11 . 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;
a bypass duct defined radially outside the core, with a leading edge of a splitter defining the point at which flow splits into a core flow and a bypass flow; 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 is 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;
an engine core radius ratio is defined as the ratio of the radius of a tip of a most downstream turbine blade in the engine to the radius of the leading edge of the splitter;
the compression system radius ratio divided by the engine core radius ratio is no greater than 7;
a core compressor pressure ratio, defined as a pressure ratio across the compressor at cruise conditions excluding a pressure rise due to the fan, is in a range of 34 to 52 at cruise conditions;
a fan root to tip 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 exit that subsequently flows through the bypass duct, is in a range 0.8 to 0.95 at cruise conditions;
a fan tip pressure ratio is in a range of 1.30 to 1.45 at cruise conditions; and
a bypass ratio, defined as the ratio of the mass flow rate of the flow through the bypass duct to the mass flow rate of the flow through the core, is in a range of 10 to 18 at cruise conditions.
12 . The gas turbine engine according to claim 11 , wherein the fan root to tip pressure ratio is in a range of 0.82 to 0.89.
13 . The gas turbine engine according to claim 11 , wherein the compression system radius ratio divided by the engine core radius ratio is 5.6.
14 . The gas turbine engine according to claim 11 , wherein the radius of the leading edge of the splitter is 50 cm to 70 cm.
15 . The gas turbine engine according to claim 11 , wherein the radius of the tip of the most downstream turbine blade is 40 cm to 60 cm.
16 . The gas turbine engine according to claim 11 , wherein:
the radius of the leading edge of the splitter is 50 cm to 70 cm;
the radius of the tip of the most downstream turbine blade is 40 cm to 60 cm; and
a height of each fan blade is 75 cm to 100 cm.
17 . 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;
a bypass duct defined radially outside the core, with a leading edge of a splitter defining the point at which flow splits into a core flow and a bypass flow; 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:
an engine core radius ratio, defined as the ratio of the radius of a tip of a most downstream turbine blade in the engine to the radius of the leading edge of the splitter, is in a range of 0.8 to 0.98;
a core compressor pressure ratio, defined as a pressure ratio across the compressor at cruise conditions excluding a pressure rise due to the fan, is in a range of 34 to 52 at cruise conditions;
a fan root to tip 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 exit that subsequently flows through the bypass duct, is in a range 0.8 to 0.95 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 no greater than 1.24.
18 . The gas turbine engine according to claim 17 , wherein:
the rotational speed of the fan at cruise conditions is in a range 2200 rpm to 2700 rpm; and
a compression system speed ratio, defined as a ratio of a rotational speed of a most downstream compressor blade to a rotational speed of the fan, is in a range 6 to 8 at cruise conditions.
19 . The gas turbine engine according to claim 17 , wherein the engine core radius ratio is in a range 0.82 to 0.93.
20 . The gas turbine engine according to claim 17 , wherein the core compressor has exactly 12, 13 or 14 stages.