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A cooling system of a gas turbine engine, includes a heat exchanger having a common wall shared by a first air passage for directing a portion of a compressor air flow to be used as cooling air, and a second air passage for directing a portion of a bypass air flow, the portion of the compressor air

A cooling system of a gas turbine engine, includes a heat exchanger having a common wall shared by a first air passage for directing a portion of a compressor air flow to be used as cooling air, and a second air passage for directing a portion of a bypass air flow, the portion of the compressor air flow being thereby cooled by the portion of the bypass air flow through the common wall.

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1. A gas turbine engine comprising: a gas path having a core flow portion and a bypass flow portion, the core flow portion including at least a compressor, combustor and turbine, the compressor providing compressor air and the bypass flow portion conducting bypass air;a heat exchanger incorporated w

1. A gas turbine engine comprising: a gas path having a core flow portion and a bypass flow portion, the core flow portion including at least a compressor, combustor and turbine, the compressor providing compressor air and the bypass flow portion conducting bypass air;a heat exchanger incorporated with at least one portion of turbine case surrounding the turbine, the heat exchanger having first and second air passages, the at least one portion of the turbine case forming a common wall separating and shared by the first and second air passages; andthe first air passage in fluid communication with the compressor air for directing a portion of the compressor air through the heat exchanger to cool a turbine system, the second air passage in fluid communication with the bypass air for directing a portion of the bypass air through the heat exchanger to cool the portion of compressor air passing through the heat exchanger, the common wall in the heat exchanger configured to transfer heat from the portion of the compressor air flowing in the first air passage to the portion of the bypass air flowing in the second air passage, to cool the portion of the compressor air, the first air passage being defined between the common wall formed by the at least one portion of the turbine case and an annular inner wall positioned radially, inwardly spaced apart from and supported by the turbine case, the annular inner wall surrounding a turbine shroud casing and including a plurality of holes extending radially through the annular inner wall for directing the cooled portion for the compressor air to impinge on the turbine shroud casing;an apparatus configured for selectively adding an uncooled portion of the compressor air into the cooled portion of the compressor air after being discharged through the holes in the inner wall, in response to a temperature change of the turbine shroud casing of the turbine system, wherein the apparatus comprises an annular inlet introducing the uncooled portion of the compressor air into an annulus defined between the annular inner wall and the turbine shroud casing for receiving the cooled portion of the compressor air discharged from the holes in the inner wall, the inlet being defined between an axial section of the annular inner wall and an axial section of the turbine shroud casing, the inlet thereby changing an opening size when the axial section of the turbine shroud casing changes a size in a radial dimension in response to temperature changes of the turbine shroud casing. 2. The gas turbine engine as defined in claim 1 wherein the first air passage communicates with an active tip clearance control (ATCC) system downstream of the heat exchanger. 3. The gas turbine engine as defined in claim 1 further comprising a plurality of fins radially inwardly extending from an inner side of the common wall into the first air passage. 4. The gas turbine engine as defined in claim 1 wherein the second air passage is defined between the common wall formed by the at least one portion of the turbine case and an annular outer wall surrounding and radially spaced apart from the turbine case. 5. The gas turbine engine as defined in claim 1 further comprising a plurality of fins radially outwardly extending from an outer side of the common wall into the second passage. 6. The gas turbine engine as defined in claim 1 further comprising a plurality of fins radially outwardly extending from an outer side of the common wall into the second passage. 7. The gas turbine engine as defined in claim 1 further comprising means for regulating the portion of the bypass air passing through the heat exchanger, thereby controlling heat transfer between the first and second air passages and thereby controlling a temperature of the cooled portion of the compressor air. 8. A gas turbine engine having a fan assembly, a compressor assembly, a combustion gas generator assembly, a turbine assembly and a bypass duct, the engine further comprising a heat exchanger for an active tip clearance control (ATCC) system using a portion of a compressor air flow as cooling air of the ATCC, the heat exchanger including at least one portion of a turbine case which surrounds the turbine assembly, the at least one portion of the turbine case forming a common wall shared by and separating in the heat exchanger, a first air passage for directing a portion of the compressor air flow and a second air passage for directing a portion of the bypass air flow introduced from the bypass duct to thereby cool the portion of the compressor air flow; wherein the first air passage is defined between the common wall formed by the at least one portion of the turbine case and an annular inner wall positioned radially, inwardly spaced apart from and supported by the turbine case; wherein the annular inner wall surrounds a turbine shroud casing and comprises a plurality of holes extending radially through the annular inner wall for directing the cooled portion of the compressor air to impinge on the turbine shroud casing; and wherein an apparatus is configured for selectively adding an uncooled portion of the compressor air into the cooled portion of the compressor air after being discharged through the holes in the inner wall, in response to a temperature change of the turbine shroud casing of the turbine system, the apparatus is including an annular inlet introducing the uncooled portion of the compressor air into an annulus defined between the annular inner wall and the turbine shroud casing for receiving the cooled portion of the compressor air discharged from the holes in the inner wall, the inlet being defined between an axial section of the annular inner wall and an axial section of the turbine shroud casing, the inlet thereby changing an opening size when the axial section of the turbine shroud casing changes a size in a radial dimension in response to temperature changes of the turbine shroud casing. 9. The gas turbine engine as defined in claim 8 wherein the shared common wall formed by the at least one portion of the turbine case comprises a first group of fins affixed to an inner side of the common wall and extending into the first air passage and a second group of fins affixed to an outer side of the common wall and extending into the second air passage. 10. The gas turbine engine as defined in claim 8 wherein the ATCC system further comprises a regulator for controlling the portion of the bypass air flow passing through the heat exchanger, thereby controlling a temperature of the cooled portion of the compressor air flow in the first air passage. 11. The gas turbine engine as defined in claim 8 wherein the apparatus further comprises an inlet for selectively introducing a varying portion of an uncooled compressor air flow into an annulus in response to temperature changes of a turbine shroud casing, the annulus surrounding the turbine shroud casing and receiving a cooled portion of the compressor air flow discharged from the first air passage. 12. A method for active tip clearance control (ATCC) of a turbine, the method comprising the steps of: a) directing a portion of a bypass air flow to cool a portion of a compressor air flow;b) directing the cooled portion of the compressor air flow to cool a surface of an active tip clearance control (ATCC) apparatus; andc) adjusting a temperature of the surface of ATCC apparatus by selectively introducing an uncooled portion of the compressor air flow into the cooled portion of the compressor air flow to increase a temperature of air provided to the ATCC apparatus, wherein the uncooled portion of the compressor air flow is introduced through an annular inlet into an annulus defined between an annular inner wall and a turbine shroud casing for receiving the cooled portion of the compressor air discharged from holes in the inner wall, the inlet being defined between an axial section of the annular inner wall and an axial section of the turbine shroud casing, the inlet thereby changing an opening size when the axial section of the turbine shroud casing changes a size in a radial dimension in response to temperature changes of the turbine shroud casing. 13. The method as defined in claim 12 wherein the portion of the bypass air flow used to cool the portion of the compressor air flow, is regulated to control a temperature of the cooled portion of the compressor air flow. 14. The method as defined in claim 12 wherein step (a) is conducted through a heat exchanger having first and second air passages sharing a common wall defined by at least one portion of a turbine case, the first air passage directing the portion of the compressor air flow and the second air passage directing the portion of the bypass air flow. 15. The method as defined in claim 12 wherein the selective introduction of the uncooled portion of the compressor air flow is automatically conducted in response to temperature changes of the surface of the ATCC apparatus.