Turbine engine core compression temperatures

The method according to claim 20, wherein the core compressor to fan root temperature rise ratio is in the range from 2.8 to 3.2.

The method according to claim 20, wherein the fan root temperature rise is in the range from 1.03 to 1.09.

The method according to claim 20, wherein the core compressor temperature rise is in the range from 2.9 to 4.0.

The method according to claim 20, wherein the engine core has a core radius defined between the centreline of the engine and a forwardmost tip of the engine core, wherein the fan rotor entry temperature is defined as the average temperature of airflow across a radially inner portion of the leading edge of each fan blade at cruise conditions, and wherein the radially inner portion of each fan blade is or comprises the portion of each fan blade at a distance from the centreline of the engine less than the core radius.

The method according to claim 20, wherein the engine comprises more than one compressor, and wherein the compressor exit temperature is measured at the exit from the highest pressure compressor.

The method according to claim 20, wherein the engine core comprises an annular splitter at which the flow is divided between a core flow that flows through the engine core, and a bypass flow that flows along a bypass duct, wherein stagnation streamlines around the circumference of the engine, stagnating on a leading edge of the annular splitter, form a streamsurface forming a radially outer boundary of a streamtube that contains all of the core flow, and wherein each fan blade has a radially inner portion lying within the streamtube that contains the core flow, and wherein the core entry temperature is defined as an average temperature of airflow across the trailing edge of the radially inner portion of each fan blade at cruise conditions.

The method according to claim 20, wherein a specific thrust of the engine at cruise conditions, defined as net engine thrust divided by mass flow rate through the engine, is in the range from 50 to 100 Nkg−1s.

The method according to claim 20, wherein a fan tip loading at cruise conditions is defined as dH/Utip2, where dH is the enthalpy rise across the fan and Utip is the (translational) velocity of the fan tip is in the range from 0.25 to 0.4.

The method according to claim 20, 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.

The method according to claim 20, wherein a curvature of a root portion of each fan blade is between 40% and 60% less than a curvature across a tip portion of the blade.

The method according to claim 20, wherein the cruise conditions correspond to:

• (i) atmospheric conditions defined by the International Standard Atmosphere at an altitude of 11582 m and a forward Mach Number of 0.8;
• or
• (ii) atmospheric conditions defined by the International Standard Atmosphere at an altitude of 10668 m and a forward Mach Number of 0.85.

The method according to claim 20, wherein the cruise conditions correspond to atmospheric conditions at an altitude that is in the range of from 10500 m to 11600 m.

The method according to claim 20, wherein the Overall Pressure Ratio (OPR) at cruise is greater than 40 and lower than 80.

The method of according to claim 20, wherein:

• (i) the fan tip radius is in the range from 110 cm to 150 cm;
• or
• (ii) the fan tip radius is in the range from 155 cm to 200 cm.

The method according to claim 20, further comprising 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 wherein the gearbox has a gear ratio in the range of from 3.2 to 5.