GAS TURBINE ENGINE COMPRESSION SYSTEM WITH CORE COMPRESSOR PRESSURE RATIO

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United States of America Patent

APP PUB NO 20240317411A1
SERIAL NO

18675809

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Abstract

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A gas turbine engine has a compression system radius ratio defined as the ratio of the radius of the tip of a fan blade to the radius of the tip of the most downstream compressor blade in the range of from 5 to 9. This results in an optimum balance between installation benefits, operability, maintenance requirements and engine efficiency when the gas turbine engine is installed on an aircraft.

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Patent Owner(s)

Patent OwnerAddress
ROLLS-ROYCE PLCKINGS PLACE 90 YORK WAY LONDON N19FX

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Inventor(s)

Inventor Name Address # of filed Patents Total Citations
ARMSTRONG, Gareth M Nottingham, GB 16 29
HOWARTH, Nicholas Derby, GB 32 222

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