Gas turbine engine
A gas turbine engine 10 is provided in which a fan having fan blades in which the camber distribution along the span allows the gas turbine engine to operate with improved efficiency when compared with conventional engines, whilst retaining an acceptable flutter margin.
1. A gas turbine engine for an aircraft comprising:
an engine core comprising a turbine, a compressor, and a core shaft connecting the turbine to the compressor;
a fan located upstream of the engine core, the fan comprising a hub and a plurality of fan blades, each fan blade having a radial blade span extending from a root at a 0% span position to a tip at a 100% span position, and the plurality of fan blades being fixed pitch fan blades that are unable to rotate relative to the hub;
a gearbox that receives an input from the core shaft and outputs drive to the fan so as to drive the fan at a lower rotational speed than the core shaft; and
an annular splitter at which flow is divided between a core flow (A) that flows through the engine core, and a bypass flow (B) that flows along a bypass duct,
wherein:
the average camber (α 2 −α 1 ) ave of each fan blade taken over the radially innermost 10% of the blade span is no greater than 75% of the average camber of the 10% blade span portion that has the maximum average camber (θ 2 −θ 1 ) max ;
a fan root to tip pressure ratio of the gas turbine engine, defined as the ratio of the mean total pressure of flow at a fan exit that subsequently flows through the engine core (P 102 ) to the mean total pressure of flow at the fan exit that subsequently flows through the bypass duct (P 104 ), is in the range of from 0.83 to 0.9 at cruise conditions; and
a fan tip loading is defined as dH/U tip 2 , where dH is the enthalpy rise across the fan and U tip is the velocity of the fan at the tip of each fan blade, and the fan tip loading at the cruise conditions is greater than 0.3 JKg −1 /(ms −1 ) 2 .
2. A gas turbine engine according to claim 1 , wherein the average camber (α 2 −α 1 ) ave of each fan blade taken over the radially innermost 10% of the blade span is no greater than 70% of the average camber of the 10% blade span portion that has the maximum average camber (θ 2 −θ 1 ) max .
3. A gas turbine engine according to claim 1 , wherein the average camber (α 2 −α 1 ) ave of each fan blade taken over the radially innermost 10% of the blade span is no greater than 65% of the average camber of the 10% blade span portion that has the maximum average camber (θ 2 −θ 1 ) max .
4. A gas turbine, engine according to claim 1 , wherein the fan tip loading at the cruise conditions is in the range of from 0.3 to 0.4 JKg −1 /(ms −1 ) 2 .
5. A gas turbine engine according to claim 1 , wherein a bypass ratio is defined as the ratio of the mass flow rate of the bypass flow (B) that flows along the bypass duct to the mass flow rate of the core flow (A) at the cruise conditions, and the bypass ratio is greater than 10.
6. A gas turbine engine according to claim 1 , wherein the specific thrust at the cruise conditions is less than 100 NKg −1 s.
7. A gas turbine engine according to claim 1 , wherein:
the ratio of the radius of each fan blade at the root (r root ) to the radius of each fan blade at the tip (r tip ) is less than 0.33.
8. A 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.
9. A gas turbine engine according to claim 1 , wherein the fan diameter is greater than 250 cm.
10. A gas turbine engine according to claim 1 , wherein a fan root pressure ratio, defined as the ratio of the mean total pressure of flow at the fan exit that subsequently flows through the engine core (P 102 ) to the mean total pressure of flow at a fan inlet (P 100 ), is no greater than 1.25 at the cruise conditions.
11. A gas turbine engine according to claim 10 , wherein: a combustor ( 17 ) is provided downstream of the fan and the compressor and upstream of the turbine; an overall pressure ratio is defined as the ratio of the mean total pressure of flow at an inlet to the combustor (P 106 ) to the mean total pressure of flow at the fan inlet (P 100 ) at the cruise conditions; and the ratio of the fan root pressure ratio to the overall pressure ratio is less than 0.04.
12. A gas turbine engine according to claim 11 , wherein the forward speed of the gas turbine engine at the cruise conditions is in the range of from Mn 0.75 to Mn 0.85, and the cruise conditions correspond to atmospheric conditions at an altitude that is in the range of from 10500 m to 11600 m.
13. A gas turbine engine according to claim 10 , wherein the forward speed of the gas turbine engine at the cruise conditions is in the range of from Mn 0.75 to Mn 0.85.
14. A gas turbine, engine according to claim 10 , wherein the cruise conditions correspond to atmospheric conditions at an altitude, that is in the range of from 10500 m to 11600 m.
15. A gas turbine engine according to claim 1 , wherein the fan is a blisk such that the plurality of fan blades are formed integrally with the hub.
16. A gas turbine engine for an aircraft comprising:
an engine core comprising a turbine, a compressor, and a core shaft connecting the turbine to the compressor;
a fan located upstream of the engine core, the fan comprising a hub and a plurality of fan blades, each fan blade having a radial blade span extending from a root at a 0% span position to a tip at a 100% span position, the plurality of fan blades being fixed pitch fan blades that are unable to rotate relative to the hub;
a gearbox that receives an input from the core shaft and outputs drive to the fan so as to drive the fan at a lower rotational speed than the core shaft; and
an annular splitter at which flow is divided between a core flow (A) that flows through the engine core, and a bypass flow (B) that flows along a bypass duct, wherein:
the average camber (β 2 −β 1 ) ave of each fan blade taken over a portion of the blade between the 30% span position and the 40% span position is at least 1.2 times the average camber of the radially innermost 10% of the blade span (α 2 −α 1 ) ave ; and
a fan root to tip pressure ratio, defined as the ratio of the mean total pressure of flow at a fan exit that subsequently flows through the engine core (P 102 ) to the mean total pressure of flow at the fan exit that subsequently flows through the bypass duct (P 104 ), is in the range of from 0.83 to 0.9 at cruise conditions; and
a fan tip loading is defined as dH/U tip 2 , where dH is the enthalpy rise across the fan and U tip is the velocity of the fan at the tip of each fan blade, and the fan tip loading at the cruise conditions is greater than 0.3 JKg −1 /(ms −1 ) 2 .
17. A gas turbine engine accord ng to claim 16 , wherein the average camber (β 2 −β 1 ) ave of each fan blade taken over a portion of the fan blade between the 30% span position and the 40% span position is at least 1.3 times the average camber of the radially innermost 10% of the blade span (α 2 −α 1 ) ave .
18. A gas turbine engine according to claim 16 , wherein the average camber (β 2 −β 1 ) ave of each tan blade taken over a portion of the fan blade between the 30% span position and the 40% span position is at least 1.4 times the average camber of the radially innermost 10% of the blade span (α 2 −α 1 ) ave .