Defense

February 14, 2014

Tunnel 9 engineers conduct boundary layer transition experiments at Mach 10

Air Force Project Engineer and Purdue University graduate student George Moraru examines the illuminated temperature sensitive paint coating on the large, 7-degree cone prior to the test program. The testing is performed under the Test Resource Management Center and the Air Force Office of Scientific Research funded Hypersonic Center of Testing Excellence.

The Arnold Engineering Development Complex Hypervelocity Wind Tunnel 9 is performing experiments on a large 7-degree cone test article at Mach 10 to improve the understanding of hypersonic boundary layer transition in testing and evaluation facilities.

The testing is made possible under the Test Resource Management Center and the Air Force Office of Scientific Research funded Hypersonic Center of Testing Excellence.

According to Dan Marren, White Oak site director, this program was modeled after the AFOSR lab programs and is the first of its kind for the test and evaluation commands.

“The first CoTE pilot program is in the area of hypersonics and is located at AEDC White Oak because of the unique hypervelocity test facility (Tunnel 9) located there. The University of Maryland was chosen as the anchor university,” he said. “The basic vision of this program is to bring together researchers, students and testing professionals working on U.S. Air Force priorities in a collaborative way that strengthens the vitality of both the research and the work force.”

The experimental program conducted during the Fall-Winter 2013-2014 consists of approximately 20 Mach 10 runs at units Reynolds number between two and 10 million per foot. High fidelity measurements are required to make better flight predictions based wind tunnel measurements and to validate boundary layer stability computations. Such computations offering predictive capabilities and their validation based on wind tunnel data are essential to reduce the developmental risks of new hypersonic systems. The ability to predict whether or not a vehicle will maintain a laminar boundary layer while travelling at hypersonic speed is critical. For turbulent boundary layers, the heating rate and viscous drag both increase sharply leading to more stringent thermal protection requirements, a reduced range and/or increased propulsion requirements.

Previous measurements in large scale hypersonic T&E facilities over the last 40 years were mostly limited to the determination of the transition location and provided limited insight into the physics of the transition process. Several cutting edge measurements techniques providing new insight are used during this test campaign including temperature sensitive paint, high frequency response pressure sensors and high speed Schlieren cinematography.

A new high speed data acquisition system is being deployed during this program to simultaneously acquire signals from 90 pressure transducers at sample rates between two and 10 MHz. Pressure measurements at frequencies up to 1 MHz are required to measure the acoustic instabilities (2nd mode wave) propagating and growing inside the laminar boundary layer and their breakdown which leads to turbulent flow. Such measurement can be achieved using high frequency response piezoelectric pressure transducers. The high density of pressure transducers on the cone test article allows the determination of the spatial amplification rates of instabilities which can be directly compared against boundary layer stability computations. In addition to point measurements, TSP provides global surface heat transfer to reveal detailed information about the shape of the transition front and transitional flow features such as crossflow vortices.

The 61-inch long cone model is fitted with interchangeable nose tips with radii between 0.06 and two inches in order to assess the effect of nose bluntness on boundary layer transition. Previous measurements have revealed that transition moves downstream with an increase in nose diameter. Such measurements have been corroborated with stability computations showing reduced second mode growth with increased bluntness. However, high fidelity instability measurements have yet to be performed for a wide range of nose bluntness.

The effect of small angle of attacks on the shape of the transition front is not well understood. This effect is investigated by using the unique Tunnel 9 hydraulic pitch system which allows measurement over a wide angle of attack range during a single wind tunnel run. Measurements at angles of attack up to 10 degrees are performed to quantify the effect of crossflow instability on the transition process.




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