초록 ▼

A cooling system in an aircraft gas turbine engine includes a heat exchanger positioned within an annular nacelle space surrounding a bypass duct of the engine. The heat exchanger has a radial coolant passage extending between the bypass duct and ambient air surrounding the nacelle, and a flow passa

A cooling system in an aircraft gas turbine engine includes a heat exchanger positioned within an annular nacelle space surrounding a bypass duct of the engine. The heat exchanger has a radial coolant passage extending between the bypass duct and ambient air surrounding the nacelle, and a flow passage extending substantially normal to the radial passage for direction of a fluid to be cooled therethrough. A configuration of this nature may assist in defining a no-flow length of the heat exchanger in a third direction normal to the other two mentioned directions, which may allow for improved performance within a given radial envelope.

대표청구항 ▼

1. 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

1. 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. 2. The cooling system as defined in claim 1 wherein the heat exchanger no-flow length exceeds a radial height of the heat exchanger. 3. A method for cooling compressor bleed air 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 method comprising: directing bypass air through a first passage extending substantially radially through a nacelle space surrounding the bypass air duct, to thereby provide a radial coolant flow; anddirecting compressor bleed air through a second passage extending substantially tangentially in said nacelle space, the second passage arranged in heat transfer communication with the first passage and thereby configured for heat transfer between the first and second passages to cool the compressor bleed air, thereby resulting in no flow passing through said axial direction to define a no-flow length of the first and second passages substantially in said axial direction. 4. The method as defined in claim 3 further comprising a step of providing an opening in an outer wall of the bypass air duct to form an entry of the first passage for intake of said bypass air. 5. The method as defined in claim 3 further comprising a step of providing an opening in an outer nacelle wall to form an exit of the first passage for discharging said bypass air into ambient air. 6. The method as defined in claim 3 further comprising a step of connecting the second passage with a first tube line for input of a hot compressor bleed air flow to the second passage. 7. The method as defined in claim 3 further comprising a step of connecting the second passage of the heat exchanger with a second tube line for output of a cooled compressor bleed air flow from the second passage.