Patent ID: 8256229

Claim:
A gas turbine engine comprising: a high pressure compressor including a last stage having a last stage compressor blade and a last stage vane; a turbine; a first flow path, wherein bleed air flows to the turbine through the first flow path, and the first flow path includes a plurality of pipes; a second flow path, wherein air downstream of the high pressure compressor flows along the second flow path; and a heat exchanger, wherein the bleed air in the first flow path exchanges heat with the air in the second flow path to cool the air in the second flow path to cooled air, and the cooled air in the second flow path is returned to the high pressure compressor to cool a portion of the high pressure compressor, wherein the heat exchanger is located on each of the plurality of pipes, wherein a portion of the air enters the heat exchanger through a diffuser case and exits the heat exchanger through a strut, wherein a flow of the portion of the air through the heat exchanger is substantially parallel to the diffuser case.