2020/01/28 by Harrison B. Yates, Eric H. Matlis, Eric Matlis +2 · 30 citations
Engineering · #Actuator #Aerospace engineering #Boundary layer #Computational Fluid Dynamics and Aerodynamics #Engineering #Flow control (data) #Fluid Dynamics and Turbulent Flows #Hypersonic speed #Mach number #Mach wave #Materials science #Mechanics #Physics #Plasma #Plasma actuator #Plasma and Flow Control in Aerodynamics #Reynolds number #Stanton number #Turbulence #Vortex
paper · open access · doi:10.2514/1.j058981
published in AIAA Journal 58(5), 2093-2108 (American Institute of Aeronautics and Astronautics)
openalex publication_date 2020/01/28 · openalex created_date 2025/10/10 · openalex updated_date 2026/07/25
The purpose of this research was to design, fabricate, and test a plasma-based active flow-control system to accelerate and delay crossflow-induced boundary-layer transition on a cone at an angle of attack at Mach 6 under quiet-flow conditions. A model with interchangeable nosetips was designed and fabricated from stainless steel, polyether ether ketone (PEEK), and Macor®. Transition on the model was characterized using infrared thermography and Kulite pressure transducers in the Boeing/U.S. Air Force Office of Scientific Research Mach 6 Quiet Tunnel at Purdue University. The flow controllers were assessed by their impact on the transition location and wave number of the largest-amplitude hot streaks. The transition location was accelerated by critical forcing (actuator wave number equals wave number of naturally largest-amplitude waves) and delayed by subcritical forcing (actuator wave number larger than natural waves). The disturbance wave number input of the plasma actuators was observed downstream on the model for many of the plasma-on runs, demonstrating that the plasma actuators introduced discrete forcing into the flow. The precise locations of the hot streaks arising from stationary crossflow vortices varied for different nosetips, presumably due to differences in their microscale roughness.