Gas turbine engine for long range aircraft
A gas turbine engine comprises a fan for delivering air into a bypass duct as bypass flow, into a core housing as core flow, with the core housing containing an upstream compressor rotor and a downstream compressor rotor. An overall pressure ratio is defined across the upstream and downstream compressor rotors. A bypass ratio is defined as a volume of air delivered as bypass flow compared to a volume of air delivered into the core housing. The overall pressure ratio is greater than or equal to about 45.0, and the bypass ratio is greater than or equal to about 11.0.
1. A gas turbine engine comprising:
a fan for delivering air into a bypass duct as bypass flow and into a core housing as core flow, with said core housing containing an upstream compressor rotor and a downstream compressor rotor, and an overall pressure ratio defined across said upstream and downstream compressor rotors, and a bypass ratio defined as a volume of air delivered as bypass flow compared to a volume of air delivered into the core housing;
said overall pressure ratio is greater than or equal to 45.0, and said bypass ratio is greater than or equal to 11;
said gas turbine engine being designed for use on a long range aircraft, with said long range aircraft being defined by an aircraft with at least two passenger aisles; and
a fan drive turbine driving said fan through a gear reduction; and
wherein a ratio of a tip speed at said downstream compressor rotor compared to a tip speed at said upstream compressor rotor is less than or equal to 1.18 and greater than or equal to 1.0.
2. A gas turbine engine comprising:
a fan for delivering air into a bypass duct as bypass flow and into a core housing as core flow, with said core housing containing an upstream compressor rotor and a downstream compressor rotor, and an overall pressure ratio defined across said upstream and downstream compressor rotors, and a bypass ratio defined as a volume of air delivered as bypass flow compared to a volume of air delivered into the core housing;
said overall pressure ratio is greater than or equal to 45.0, and said bypass ratio is greater than or equal to 11;
said gas turbine engine being designed for use on a long range aircraft, with said long range aircraft being defined by an aircraft with a flight length equal to or greater than 6.0 hours; and
a fan drive turbine driving said fan through a gear reduction;
wherein said upstream compressor rotor has at least three stages; and
wherein said fan drive turbine is configured to drive said upstream compressor rotor and said fan rotor through said gear reduction, with said fan drive turbine having at least three stages; and
wherein a ratio of a tip speed at said downstream compressor rotor compared to a tip speed at said upstream compressor rotor is less than or equal to 1.18 and greater than or equal to 1.0.
3. The gas turbine engine as set forth in claim 2 , wherein the bypass ratio is greater than or equal to 14.0.
4. A gas turbine engine comprising:
a fan for delivering air into a bypass duct as bypass flow and into a core housing as core flow, with said core housing containing an upstream compressor rotor and a downstream compressor rotor, and an overall pressure ratio defined across said upstream and downstream compressor rotors, and a bypass ratio defined as a volume of air delivered as bypass flow compared to a volume of air delivered into the core housing;
said overall pressure ratio is greater than or equal to 45.0, and said bypass ratio is greater than or equal to 11;
said gas turbine engine being designed for use on a long range aircraft, with said long range aircraft being defined at least by one of an aircraft with at least two passenger aisles, or an aircraft with a flight length equal to, or greater than 6.0 hours; and
a fan drive turbine driving said fan through a gear reduction;
wherein a ratio of a tip speed at said downstream compressor rotor compared to a tip speed at said upstream compressor rotor is less than or equal to 1.18 and greater than or equal to 1.0.