|M.Sc Student||Eli Yakirevich|
|Subject||Design of Continuous Closed-Loop Transonic Linear Cascade|
for Aero-Thermal Performance Studies in
|Department||Department of Aerospace Engineering||Supervisor||Dr. Cukurel Beni|
The present research focused on the design process of a new continuous closed-loop hot transonic linear cascade. This work was done to fulfill the world-wide need to conduct aero-thermal research in micro-turbomachinery that cannot be conducted using existing state-of-art cascades. The facility features fully modular design which is intended to serve as a test bench for axial micro-turbomachinery components in independently varying Mach and Reynolds numbers ranges of 0 - 1.3 and 2⋅104 - 6⋅105 respectively. Moreover, for preserving heat transfer characteristics of the hot gas section, the gas to solid temperature ratio (up to 2) is retained. Mach - Reynolds - temperature ratio independency is required to accommodate the requirements of the similarity analysis that is described in this work, in order to enable experiments for different scales and features of turbomachinery components. This operational environment has not been sufficiently addressed in prior art, although it is critical for the future development of ultra-efficient high power or thrust devices.
In order to alleviate the dimension specific challenges associated with micro-turbomachinery, the facility is designed in a highly versatile manner, and can easily accommodate different geometric configurations (pitch, ±20° stagger angle, ±20° incidence angle), absent of any alterations to the test section. Owing to the quick swap design, the vane geometry can be easily replaced without manufacturing or re-assembly of other components. Flow periodicity is achieved by multiple components that were specially designed for this cascade. A settling chamber directly upstream of the cascade is designed to generate uniform flow into the cascade inlet. Controlled boundary layer suction was designed to further enhance uniformity at the entrance to the tested section and independently adjustable tailboard mechanisms are included to contour the flowlines at the exit of the tested sections. All these components combined result in enhanced stability and periodicity while suppressing shockwave reflections.
The cascade setup enables utilization of test equipment that would measure integral loss towards quantifying overall profile performance. Moreover, total and static pressures will be mapped upstream and downstream of the cascade. In addition to that, heat transfer measurements would server towards film cooling effectiveness evaluating. Overall, as aero-thermal performance of various advanced geometries will be assessed in engine relevant environments, construction of this facility would enable test-aided design capability for national and world-wide micro gas turbine manufacturers.