Gas turbine engine transfer efficiency

A gas turbine engine for an aircraft comprising:

• an engine core comprising a first turbine, a first compressor, and a first core shaft connecting the first turbine to the first compressor;
• and a second turbine, a second compressor, and a second core shaft connecting the second turbine to the second compressor, the second turbine being a higher pressure turbine than the first turbine and the second compressor being a higher pressure compressor than the first compressor, wherein a first turbine entrance temperature is defined as an average temperature of airflow at the entrance to the first turbine at cruise conditions and a first turbine exit temperature is defined as an average temperature of airflow at the exit from the first turbine at cruise conditions, and a low pressure turbine temperature change is defined as:
• the
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• first
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• turbine
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• exit
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• temperature
• the
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• first
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• turbine
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• entrance
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• temperature
• ,
• 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 splitter, form a streamsurface forming a radially inner boundary of a streamtube that contains all of the bypass flow;
• 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 the 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
• 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 turbine to fan tip temperature change ratio of:
• the
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• low
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• pressure
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• turbine
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• temperature
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• change
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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 1.46 to 2.0.