CORNERSTONE RESEARCH GROUP INC — Department of Defense SBIR Phase I: N202-141

Description

Increasing global threats require the U.S. to establish hypersonic military proficiency and excellence. The U.S. military are prioritizing the development of hypersonic weapons as part of an effort to acquire the capability of launching attacks against targets around the world in under an hour. Hypersonic flight vehicles could also be used for future space flight applications such as emergency satellite repairs and providing a cheaper way to launch unmanned/manned payloads into Low Earth Orbit (LEO). Moreover, the primary hinderance to present-day production of hypersonic flight vehicles is laminar to turbulent boundary layer transition control. The prediction and control of hypersonic boundary layer transition is necessary in order to prevent excess surface heating that is caused by turbulent flow. Designing hypersonic vehicles that can endure the high heat loads from turbulence will significantly add to the structure's weight, impacting flight performance (payload, range, and maneuverability). It is expected that most near-term hypersonic vehicles will require a lifting body to sustain flight, indicating that if the body has a swept feature, then the turbulent transition on the lifting body will be dominated by a mean cross-flow velocity component that is subject to a cross-flow instability. Previous research demonstrated the application of discrete patterned roughness to control boundary layer laminar-turbulence transition on a swept-lift-generating aerodynamic body in a Mach 6 wind tunnel. Patterned discrete roughness in the form of micron-sized dimples was designed to excite selected wavelengths of stationary cross-flow modes, which are highly receptive to surface roughness. As a result, with the appropriate discrete patterned roughness, the transition Reynolds number on the swept-lifting body was increased by 25%; turbulent aero-heating occurred 25% farther downstream on the test article. The hypersonic flow control research discussed was conducted in the Mach 6 Ludwieg Tube Wind Tunnel at the United States Air Force Academy. The stagnation temperature for these wind tunnel runs was 480 K, and tunnel run time was 0.080 seconds, which prevented extreme heating or ablation to be achieved on the test article. Additionally, the test articles were highly polished (