Pressure ratio distributions for a gas turbine engine
A gas turbine engine 10 is provided in 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 (P 102 ) to the mean total pressure of the flow at the fan exit that subsequently flows through the bypass duct (P 104 ), is no greater than a certain value. The gas turbine engine 10 may provide 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 comprises a hub and a plurality of fan blades, each fan blade having a radial blade span extending from a root to a tip, and the plurality of fan blades being formed integrally with the hub such that the plurality of fan blades are fixed relative to the hub;
a gearbox that receives an input from the core shat 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 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:
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 root pressure ratio of the gas turbine engine, 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 in the range of from 1.15 to 1.25 at the cruise conditions;
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 fan root to tip pressure ratio is less than 0.88.
3. 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 .
4. 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) to the mass flow rate the core flow (A) at the cruise conditions, and the bypass ratio is greater than 10.
5. A gas turbine engine according to claim 1 , wherein the specific thrust at the cruise conditions is less than 100 Nkg −1 s.
6. 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.
7. A gas turbine engine according to claim 1 , wherein: a combustor 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 ); and
the ratio of the fan root pressure ratio to the overall pressure ratio is less than 0.04 at the cruise conditions.
8. A gas turbine engine according to claim 7 , wherein the ratio of the fan root pressure ratio to the overall pressure ratio is less than 0.03 at the cruise conditions.
9. A gas turbine engine according to claim 1 , wherein: the radial blade span of each fan blade extends from the root at a 0% span position to the tip at a 100% span position; and
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 .
10. A gas turbine engine according to claim 1 , wherein:
the radial blade span of each fan blade extends from the root at a 0% span position to the tip at a 100% span position; and 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 1.2 times the average camber of the radially innermost 10% of the blade span (α 2 −α 2 ) ave .
11. A gas turbine engine according to claim 1 , wherein the fan diameter is greater than 250 cm.
12. A gas turbine engine according to claim 1 , 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.
13. A gas turbine engine according to claim 1 , wherein the forward speed of the gas turbine engine at the cruise conditions is Mn 0.8.
14. A gas turbine engine according to claim 1 , 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 cruise conditions correspond to atmospheric conditions at an altitude of 11000 m.
16. A gas turbine engine according to claim 1 , wherein the cruise conditions correspond to:
a forward Mach number of 0.8;
a pressure of 23000 Pa; and
a temperature of −55 deg C.
17. 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 comprises a plurality of fan blades and a hub, each fan blade having a radial blade span extending from a root to a tip, 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 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;
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 root pressure ratio of the gas turbine engine, 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 in the range of from 1.15 to 1.25 at the cruise conditions;
a fan tip loading is defined as dH/Ut tip 2 , where dH is the enthalpy rise across the fan and Ut 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 .