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Ultra-High Temperature (UHT) Sensor Technology for Application in the Austere Environment of Gas Turbine Engines


Current blade health monitoring sensors are capable of operating at 1100°F continuously uncooled, and have been demonstrated to work up to 1800°F with cooling. Use of compressor air for sensor cooling would adversely impact the cycle efficiency and potentially produce case distortion, and hence, a need exists to develop uncooled sensors that can operate in a +2500°F environment in the aft end of the turbo machinery. Design progression also alludes to a high-temperature need in the compression system. Current developmental and future engines for Department of Defense (DoD) propulsion systems will need to operate at high efficiencies to meet the mission-weighted fuel burn (MWFB) requirements. This means that engines operate with tighter running clearances, increased blade loading and at substantially higher temperatures. As these designs have matured, high cycle fatigue (HCF) failures have become a more common problem. Engine structural integrity program (ENSIP) criteria (MIL-HDBK-1783B, A.4.13.3) was modified to state that components shall have a minimum HCF life of 10^9 cycles (up from 10^7) and must be designed to 40% of the endurance limit (down from 60%); however, these contemporary design standards are still lacking in their ability to mitigate HCF failures. As a result, broader design margins are required to mitigate safety risks. Mission requirements for enhanced propulsion system performance and reduced fuel burn present significant design challenges for original engine manufacturers (OEM’s). An effective method to increase performance, while decreasing weight and specific fuel consumption (SFC), is to dramatically increase compression ratio and use fewer, more efficient stages. These increased compression ratios directly translate to increased temperatures which exacerbate the HCF issues. There is a need to understand the temperature environment in which these airfoils operate to better assess the impact to dwindling HCF margins due to thermal effects on fatigue and endurance capabilities. Current developmental and future engines are expected to see increased creep incidents due to rising engine stresses and temperatures. With larger stresses present inside engines, creep can occur at lower temperatures (cold creep) and may appear more prevalent in the cold sections of engines, although creep has traditionally been classified as a hot section issue. There exists a need to understand the impact of degradation mechanisms such as HCF, cold creep and others on material capabilities during engine operation. Important aspects of a sensor system for this environment to consider are footprint, overall ruggedness, and practicality of system-engine interface. An onboard diagnostic system must be of compact geometry and maintain a minimum weight. Should a system utilize line of sight probes requiring ports through the engine casing, said ports should also be minimized in size and number. An onboard system will be subjected to engine vibration, debris in gas path and possible perturbations from bypass air (on a turbofan engine). The vibration experienced by the engine case and the rotor is not necessarily in unison so the system must be able to discern blade vibration due to foreign or domestic object damage (FOD or DOD) or HCF from other vibratory noise such as that from relative motion between case and rotor. Ideally, the system must be structurally ruggedized to withstand the vibrations and loads of an operating aircraft, debris that may be in the flow path and the high-speed airflow. Currently we work with probes that are 1/2 to 1/4 inches in diameter, and just over 1 inch in length. There is no current upper bound on weight, however one must consider that weight is a premium in aircraft design and the more light-weight a design is the more attractive it is as a solution. These probes penetrate engine cases and today we would like to see innovation in decreasing these penetrant sizes. Of course, an innovative design proposal that challenges our concepts of probe construction and functionality is welcomed. It is highly recommended, though not required, that the collaboration with original equipment manufacturers be maintained throughout this development. PHASE I: Design and demonstrate the feasibility of an UHT in-situ engine sensor system at temperatures in the region of 2500°F and as cited in the Description section. Proof of concept demonstrating the system’s ability to identify and report airfoil high cycle fatigue and FOD/DOD events in real time, at elevated temperatures, is necessary. Operating in austere engine environments should be given serious consideration in this phase PHASE II: Based on Phase I effort, further develop and test in-situ engine sensor system prototype. Rig testing should be conducted to elevate technology readiness level (TRL) to an appropriate level for a representative demonstration (a minimum TRL level of 4 is expected). The system should ultimately demonstrate robustness in design and function for successful operation in a representative environment. Conceptual design for the final system configuration should also be completed. PHASE III: Ruggedize the design and perform required operational testing. Commercialize and transition the developed technology to appropriate Navy platforms.
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