Patent ID: 8266888

Claim:
A cooling system in an aircraft gas turbine engine having a bypass air duct around a core engine for directing a bypass air flow therethrough, the engine having a main axis of rotation defining axial, radial and tangential directions, the radial and tangential directions being normal to the axial direction the cooling system comprising a heat exchanger disposed between an annular outer wall of the bypass air duct and an annular outer nacelle wall surrounding the annular outer wall or the bypass duct, the heat exchanger having a first passage extending in a substantially radial direction through the outer wall of the bypass air duct to communicate with the bypass air duct and through the outer nacelle wall to communicate with ambient air surrounding the nacelle, thereby configured to direct a bypass air flow radially though the heat exchanger for discharge into the ambient environment, the heat exchanger having a second passage extending in a substantially tangential direction and communicating with a source of compressor bleed air and configured to direct said bleed air through the heat exchanger adjacent the first passage for cooling by the bypass air flow passing though the first passage, and the heat exchanger having no flow passing through said axial direction to define a no-flow length of the heat exchanger substantially in said axial direction.