Buffer system that communicates buffer supply air to one or more portions of a gas turbine engine

1. A gas turbine engine, comprising:

a compressor section;

a combustor in fluid communication with said compressor section;

a turbine section in fluid communication with said combustor;

at least one shaft that interconnects at least a portion of said compressor section and said turbine section;

a bearing structure that supports said at least one shaft, wherein said bearing structure includes a bearing compartment;

a buffer system that selectively communicates a buffer supply air to pressurize said bearing compartment, wherein the buffer supply air comprises a low pressure bleed supply air and a first high pressure bleed supply air, wherein the low pressure bleed supply air is communicated to said bearing compartment in response to take-off conditions, climb conditions, and cruise conditions of the gas turbine engine, and wherein the first high pressure bleed supply air is communicated to said bearing compartment in response to ground conditions, idle conditions, and descent idle conditions of the gas turbine engine;

wherein said buffer system includes a first circuit that provides said low pressure bleed supply air and said first high pressure bleed supply air, and a second circuit that provides a second high pressure bleed supply air;

a first valve of said first circuit selectively controlling a flow of said buffer supply air to said bearing compartment; and

a second valve of said second circuit selectively controlling a flow of said second high pressure bleed supply air to a second bearing compartment.

2. The gas turbine engine as recited in claim 1 , comprising a fan section driven by a geared architecture.

3. The gas turbine engine as recited in claim 2 , wherein the fan section includes a low Fan Pressure Ratio of less than 1.45 during operation of the gas turbine engine.

4. The gas turbine engine as recited in claim 3 , wherein said low Fan Pressure Ratio is a pressure ratio measured across a fan blade alone, and the fan blade is a component of the fan section.

5. The gas turbine engine as recited in claim 4 , wherein said geared architecture includes an epicyclic gear train.

6. The gas turbine engine as recited in claim 2 , wherein said bearing compartment is located aft of said geared architecture.

7. The gas turbine engine as recited in claim 1 , wherein the gas turbine engine is a high bypass geared aircraft engine having a bypass flow path, a core flow path, and a bypass ratio of the bypass flow path to the core flow path is between six (6) and ten (10) during operation of the gas turbine engine.

8. The gas turbine engine as recited in claim 1 , wherein the compressor section includes a low pressure compressor and a high pressure compressor, and wherein the low pressure bleed supply air and the first high pressure bleed supply air are each sourced from the high pressure compressor.

9. The gas turbine engine as recited in claim 8 , wherein the low pressure bleed supply air is sourced from an upstream stage of said high pressure compressor and the first high pressure bleed supply air is sourced from one of a middle stage and a downstream stage of said high pressure compressor.