Gear system architecture for gas turbine engine

A gas turbine engine comprising:

• a fan including a plurality of fan blades rotatable about an axis;
• a fan pressure ratio across a fan blade alone of less than 1.45;
• a bypass duct;
• a compressor section;
• a bypass ratio greater than ten (10), the bypass ratio being defined as the portion of air delivered into the bypass duct divided by the amount of air delivered into the compressor section;
• a combustor in fluid communication with the compressor section;
• a fan drive turbine in communication with the combustor, the fan drive turbine comprising a plurality of fan drive turbine rotors and a pressure ratio greater than about five (5), wherein the fan drive turbine further includes an inlet having an inlet pressure, and an outlet that is prior to any exhaust nozzle and having an outlet pressure, and the pressure ratio of the fan drive turbine is a ratio of the inlet pressure to the outlet pressure;
• a gear system including a plurality of gears providing a speed reduction between the fan drive turbine and the fan and transferring power input from the fan drive turbine to the fan at an efficiency greater than 98%, wherein the gear system further comprises a gear reduction ratio of greater than 2.3;
• a gear support system including a spring rate that accommodates deflections of the gear system;
• a lubrication system providing lubricant to the gear system and removing thermal energy from the gear system, wherein the lubrication system has a maximum capacity for removing thermal energy equal to no more than about 2% of power input into the gear system by the fan drive turbine during operation of the engine;
• and
• wherein the plurality of fan blades is less than 26 and the plurality of fan drive turbine rotors is less than 6 fan drive turbine rotors, and a ratio between the number of fan blades and the number of fan drive turbine rotors is between 3.3 and 8.6.