Compression in a gas turbine engine

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, wherein a compressor exit temperature is defined as an average temperature of airflow at the exit from the compressor, the engine core further comprising 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 slitter, form a streamsurface forming a radially inner boundary of a streamtube that contains all of the bypass flow;
• and
• a fan located upstream of the engine core, the fan comprising a plurality of fan blades extending from a hub, each fan blade having a leading edge and a trailing edge, each fan blade having a radially outer portion lying within the streamtube that contains the bypass flow, and wherein a fan rotor entry temperature is defined as an average temperature of airflow across the leading edge of each fan blade at cruise conditions and a fan tip rotor exit temperature is defined as an average temperature of airflow across a radially outer portion of each fan blade at the trailing edge at cruise conditions and a fan tip temperature rise is defined as:
• the
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• fan
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• tip
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• rotor
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• exit
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• temperature
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• the
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• fan
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• rotor
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• entry
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• temperature
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• ,
• and a core temperature rise is defined as:
• the
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• compressor
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• exit
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• temperature
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• the
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• fan
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• rotor
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• entry
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• temperature
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• ,
• and wherein a core to fan tip temperature rise ratio of:
• the
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• core
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• temperature
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• rise
• the
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• fan
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• tip
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• temperature
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• rise
• is in the range from 2.845 to 3.8.