BACKGROUND

A gas turbine engine typically includes a fan section, a compressor section, a combustor section and a turbine section. Air entering the compressor section is compressed and delivered into the combustion section where it is mixed with fuel and ignited to generate a high-speed exhaust gas flow. The high-speed exhaust gas flow expands through the turbine section to drive the compressor and the fan section.

In one type of engine, the compressor section typically includes low and high pressure compressors, and the turbine section includes low and high pressure turbines. The high pressure turbine drives the high pressure compressor through an outer shaft to form a high spool, and the low pressure turbine drives the low pressure compressor through an inner shaft to form a low spool. The fan section may also be driven by the low inner shaft. A speed reduction device such as an epicyclical gear assembly may be utilized to drive the fan section such that the fan section may rotate at a speed different than the turbine section so as to increase the overall propulsive efficiency of the engine. In such engine architectures, a shaft driven by one of the turbine sections provides an input to the epicyclical gear assembly that drives the fan section at a reduced speed such that both the turbine section and the fan section can rotate at closer to optimal speeds.

Other types of engines have been used, such as direct drive configurations in which no gear train is used, or such as three spool configurations in which three discrete turbine are provided. Although geared architectures have improved propulsive efficiency, turbine engine manufacturers continue to seek further improvements to engine performance including improvements to thermal, transfer and propulsive efficiencies.

SUMMARY

In one exemplary embodiment, a gas turbine engine has a fan which includes a plurality of fan blades that are rotatable about an axis, and a compressor section, where the combustor section includes a first compressor and a second compressor aft of the first compressor. At least one first variable guide vane controls operation of the first compressor and at least one second variable guide vane controls operation of the second compressor. A combustor is in fluid communication with the compressor section and a turbine section is in fluid communication with the combustor. A geared architecture is driven by the turbine section for rotating the fan about the axis.

In a further embodiment of the above, the pressure ratio of the first compressor is in the range of 4:1 to 7:1.

In a further embodiment of any of the above, the pressure ratio of the second compressor is in the range of 8:1 to 15:1.

In a further embodiment of any of the above, the compressor section includes an overall pressure ratio in the range of 40:1 to 70:1.

In a further embodiment of any of the above, the pressure ratio of the second compressor is in the range of 8:1 to 15:1.

In a further embodiment of any of the above, the turbine section includes a first turbine and a second turbine arranged aft of the first turbine. The first compressor and the second turbine are mounted to a first spool. The second compressor and the first turbine are mounted on a second spool.

In a further embodiment of any of the above, the geared architecture is coupled between the first spool and the fan section.

In a further embodiment of any of the above, the geared architecture is arranged forward of the combustor.

In a further embodiment of any of the above, the gas turbine engine is a high bypass geared aircraft engine that has a bypass ratio of greater than about 6:1.

In a further embodiment of any of the above, the gas turbine engine includes a fan pressure ratio of less than about 1.45:1.

In a further embodiment of any of the above, the second turbine has a pressure ratio that is greater than about 5:1.

In a further embodiment of any of the above, the gas turbine engine includes multiple stages of first variable guides in the first compressor.

In a further embodiment of any of the above, the gas turbine engine includes multiple stages of second variable guides in the second compressor.

In a further embodiment of any of the above, the compressor section includes an overall pressure ratio in the range of 40:1 to 70:1.

In a further embodiment of any of the above, the compressor section includes an overall pressure ratio of greater than 70:1.

