Coolable gas turbine engine component

Abstract

A gas turbine engine component coupled with a temperature member configured to assist in maintaining a temperature of the gas turbine engine component below a predetermined temperature. The temperature member can take the form of a phase changeable material that can change phase by melting or vaporization, among potential others. A variety of gas turbine engine components can be coupled with the temperature member.

1 . An apparatus comprising:
 a gas turbine engine device including a gas turbine engine component having an envelope of operability and thermodynamically coupled with a thermal portion having size sufficient to prolong a useful life of the gas turbine engine component by keeping the gas turbine engine component within the envelope of operability during a relatively high heating rate mission event, the thermal portion operable to change phases from one physical state to another physical state during at least a portion of the relatively high heating rate mission event.   
     
     
         2 . The apparatus of  claim 1 , wherein the thermal portion is activated by temperature and is operable to be sacrificed during the relatively high heating rate mission event. 
     
     
         3 . The apparatus of  claim 1 , wherein the gas turbine engine device includes a conduit for the passage of a working fluid, the thermal portion coupled with the conduit and operable to change phase when a temperature of a working fluid traversing the conduit is at or above a phase transition temperature. 
     
     
         4 . The apparatus of  claim 3 , wherein the thermal portion is coupled to an outside wall of the conduit. 
     
     
         5 . The apparatus of  claim 1 , wherein the gas turbine engine device is an engine case, the thermal portion operative to maintain the temperature of the engine case during an engine thermal condition. 
     
     
         6 . The apparatus of  claim 5 , wherein the engine thermal condition occurs during an exigent condition. 
     
     
         7 . The apparatus of  claim 1 , wherein the gas turbine engine device is a turbine vane having an internal cavity, the thermal portion occupying at least part of the internal cavity. 
     
     
         8 . The apparatus of  claim 1 , wherein the gas turbine engine device is a portion of an actuation device, the thermal portion operative to prevent the actuation device from overheating during at least a portion of an engine thermal condition. 
     
     
         9 . An apparatus comprising:
 a gas turbine engine including a component subject to a heat stress; and   a temperature activated portion covering the gas turbine engine component to protect the component from high temperatures during the heat stress, the temperature activated portion operable to be sacrificed by changing phases from one physical state to another.   
     
     
         10 . The apparatus of  claim 9 , wherein the temperature activated portion is structured to change phase through melting, vaporization, or sublimation. 
     
     
         11 . The apparatus of  claim 9 , wherein the gas turbine engine is designed and manufactured as a single-mission, expendable gas turbine engine, the component subject to the heat stress during at least a portion of the single-mission of the gas turbine engine. 
     
     
         12 . The apparatus of  claim 9 , wherein the temperature activated portion has a size sufficient to extend the lifetime of the gas turbine engine component during an operational exigency. 
     
     
         13 . The apparatus of  claim 12 , wherein the operational exigency is characterized by a failure of a thermal transfer mechanism intended to maintain the gas turbine engine component within a temperature limit. 
     
     
         14 . The apparatus of  claim 12 , wherein the temperature activated portion is sacrificed by having a portion that is separated from the remaining portion of the temperature activated portion. 
     
     
         15 . An apparatus comprising:
 a gas turbine engine component operable to be located in a high temperature region of a gas turbine engine; and   means for sacrificially prolonging the life expectancy of the gas turbine engine component, the means coupled with the gas turbine engine component.   
     
     
         16 . A method comprising:
 orienting a gas turbine engine component in preparation for coupling the component to a sacrificial temperature sensitive structure;   identifying a portion of the gas turbine engine component to be coupled with the sacrificial temperature sensitive structure;   selecting a quantity of the phase transition material sufficient for the gas turbine engine component to endure at least a portion of a heat flux event; and   thermodynamically coupling the sacrificial temperature sensitive structure to the gas turbine engine component.   
     
     
         17 . The method of  claim 16 , wherein the selecting includes applying a thickness of the sacrificial temperature sensitive structure appropriate to protect the gas turbine engine component from a heating source for a duration. 
     
     
         18 . The method of  claim 16 , wherein the thermodynamically coupling further includes affixing an amount of the sacrificial temperature sensitive structure to the gas turbine engine component. 
     
     
         19 . The method of  claim 18 , wherein the affixing further includes thermally attaching the sacrificial temperature sensitive structure to a wall defining an airflow conduit. 
     
     
         20 . The method of  claim 18 , wherein the affixing further includes placing the sacrificial temperature sensitive structure in a thermal communication path with a portion of an actuation device. 
     
     
         21 . The method of  claim 16 , wherein the thermodynamically coupling further includes at least partially filling a cavity of the gas turbine component with the sacrificial temperature sensitive structure, the gas turbine engine component in the form of a turbine vane. 
     
     
         22 . The method of  claim 16 , wherein the determining a thickness includes estimating the duration of an expendable aircraft mission. 
     
     
         23 . The method of  claim 16 , wherein the determining a thickness includes estimating the duration of a gas turbine engine exigent condition.