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A ceramic matrix composite blade for use in a gas turbine engine having an airfoil with leading and trailing edges and pressure and suction side surfaces, a blade shank secured to the lower end of each airfoil, one or more interior fluid cavities within the airfoil having inlet flow passages at the

A ceramic matrix composite blade for use in a gas turbine engine having an airfoil with leading and trailing edges and pressure and suction side surfaces, a blade shank secured to the lower end of each airfoil, one or more interior fluid cavities within the airfoil having inlet flow passages at the lower end which are in fluid communication with the blade shank, one or more passageways in the blade shank corresponding to each one of the interior fluid cavities and a fluid pump (or compressor) that provides pressurized fluid (nominally cool, dry air) to each one of the interior fluid cavities in each airfoil. The fluid (e.g., air) is sufficient in pressure and volume to maintain a minimum fluid flow to each of the interior fluid cavities in the event of a breach due to foreign object damage.

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1. A turbine blade for use in a gas turbine engine, comprising: an airfoil comprised of a ceramic matrix composite material, said airfoil including a leading edge and trailing edge and having pressure and suction side surfaces;a blade shank secured to a lower end of said airfoil;one or more interior

1. A turbine blade for use in a gas turbine engine, comprising: an airfoil comprised of a ceramic matrix composite material, said airfoil including a leading edge and trailing edge and having pressure and suction side surfaces;a blade shank secured to a lower end of said airfoil;one or more interior cavities disposed within said airfoil, each of said one or more interior cavities being a dead end and having an inlet passage in fluid communication with said blade shank;one or more passageways formed in a lower end of said blade shank corresponding to each one of said interior cavities; anda gas pump for continuously providing a source of pressurized gas to each one of said interior gas cavities in said airfoil. 2. The turbine blade according to claim 1, wherein said one or more interior gas cavities extend from the lower end of said airfoil toward a tip of the airfoil. 3. The turbine blade according to claim 1, wherein said one or more interior gas cavities extend at least a portion a length of said airfoil to define a localized cavity corresponding to a predicted area of airfoil damage. 4. The turbine blade according to claim 1, wherein each of said one or more interior gas cavities is pressurized by air. 5. The turbine blade according to claim 4, wherein each of said one or more interior gas cavities is pressurized by air with a water content that is less than a water content of external air. 6. The turbine blade according to claim 1, wherein said blade shank includes a dovetail connection for assembly onto a turbine rotor wheel with other blades to form a circumferential array. 7. The turbine blade according to claim 1, wherein each of said one or more interior cavities is maintained at a prescribed air pressure and temperature. 8. The turbine blade according to claim 1, wherein said source of pressurized gas is sufficient in pressure and volume to maintain a minimum gas flow to each of said one or more interior cavities in the event of a breach of one or more of said cavities due to foreign object damage. 9. The turbine blade according to claim 1, wherein said airfoil comprises a preform of a ceramic fiber. 10. The turbine blade according to claim 9, wherein said preform comprises silicon carbide. 11. The turbine blade according to claim 1, wherein said gas passageways formed in the lower end of said blade shank extend from a bottom edge of said dovetail up through said blade shank and into said interior cavities. 12. The method of manufacturing a turbine blade for use in a gas turbine engine, comprising: a. forming an airfoil comprised of a ceramic composite material;b. forming one or more interior cavities within said airfoil, wherein each of said one or more interior cavities is a dead end and includes an inlet passage in fluid communication with said blade shank;c. providing pressurized gas to each of said one or more interior cavities in said airfoil by fluidly coupling each of the inlet passages to a source of the pressurized gas, andd. continuously applying the pressurized gas to each of said interior cavities from the source, wherein the pressurized gas is stagnant within each of the one or more interior cavities. 13. The method according to claim 12, wherein said step of forming an airfoil further includes preparing a preform using a silicon carbide ceramic fiber and weaving said fiber into the desired airfoil shape. 14. The method according to claim 13, wherein said step of preparing a preform further includes the step of infiltrating said preform with a matrix material. 15. The method according to claim 14, wherein said preform is coated for bonding with said matrix material using chemical vapor infiltration, slurry infiltration-sintering, slurry casting or melt infiltration. 16. The method according to claim 12, wherein said step of providing a pressurized gas is sufficient in to maintain a minimum gas flow to each of said interior gas cavities in the event of a breach of one or more of said cavities due to foreign object damage. 17. An apparatus for detecting a failure of a turbine blade in a gas turbine engine, the turbine blade comprising a ceramic matrix composite material airfoil including an outer surface, an inner cooling chamber and passages, the apparatus comprising: continuously applying a pressurized cooling gas to the inner cooling chamber and the passages which dead end in said airfoil, wherein the inner cooling chamber and passages are in fluid communication with a source of the pressurized cooling gas;maintaining a substantially constant static pressure in the inner cooling chamber and the passages by the continuous application of the pressurized cooling gas and due to the inner cooling chamber and the passages dead ending in the airfoil;one or more pressure transducers providing signals responsive to a change in the static pressure of the pressurized cooling gas in the cooling chamber and passages of said airfoil;a non-transitory storage device storing a computer code configured to correlate changes in said signals to indicate a change in the static pressure in the cooling chamber and passages of the airfoil;a processing unit operative with the computer code to configured to determine whether said changes in signals indicates a breach condition in the inner cooling chamber or the passages of said airfoil; andan output device providing an indication of at least one of said changes in the static pressure and the breach condition. 18. The apparatus according to claim 17, wherein said pressure transducers provide data indicating an extent of said change in pressure inside said airfoil. 19. The apparatus according to claim 17, wherein said output device comprises a warning light, an acoustic warning signal or warning message in a data recorder.