Hypersonic vehicle design relies heavily on computational fluid dynamics simulations. Thus, these simulations must be able to accurately predict and model the various phenomena observed for hypersonic flows such as shock wave/boundary layer interaction (SBLI). Vehicle designers rely on Reynolds-Averaged Navier-Stokes simulations, which have proven to be a poor predictor of SBLI, but can still be useful as a tool for parametric studies. The Arnold Engineering Development Center (AEDC) Hypervelocity Tunnel 9 has built a hollow cylinder flare geometry to conduct SBLI experiments. The Johns Hopkins University Applied Physics Laboratory supported AEDC Hypervelocity Tunnel 9’s goal of designing future experiments by conducting a series of parametric studies varying CFD code, turbulence models, flare angles, and boundary-layer transition location. Additionally, stability analyses were conducted to inform AEDC Hypervelocity Tunnel 9 on what ideal leading-edge radii might be to enable transition earlier, enabling a fully turbulent boundary layer prior to the flare.

The boundary-layer transition and stability characteristics of sharp cones at angle-of-attack are investigated with measurements at Mach 10 in the Arnold Engineering Development Complex (AEDC) Hypervelocity Wind Tunnel 9 on a 1.5-m long, 7-deg cone at unit Reynolds numbers between 1.8 and 15 million per meter. The transition location is determined with coaxial thermocouples and temperature sensitive paint, and stability measurements are obtained using high-frequency response pressure sensors. The measurements are used to validate the STABL-3D linear stability theory (LST) code at angles-of-attack up to 6-deg. The computations are found to reproduce the experimental trends regarding the effect of angle-of-attack on the growth of 2 mode waves. The amplitude of the 2 mode waves near breakdown on the leeward and windward meridians scale linearly with edge Mach number. The initial amplitudes estimated using linear stability computations are found to scale with Pitot noise in the unstable 2 mode frequency band. Linear stability computations along with the ability to correlate initial 2 mode amplitudes to tunnel noise and to correlate maximum 2 mode amplitudes to edge Mach number enables the use of Mack’s amplitude method to predict 2 mode transition. This methodology is investigated to accurately predict sharp cone boundary layer transition at 0-deg AoA. Extension to sharp cones at angle-of-attack is expected to be straightforward.

Boundary-layer transition and stability data were obtained at Mach 10 in the Arnold Engineering Development Complex (AEDC) Hypervelocity Wind Tunnel 9 on a 1.5-m long, 7-deg cone at unit Reynolds numbers between 1.8 and 31 million per meter. A total of 24 runs were performed at angles-of-attack between 0 and 10-deg on sharp and blunted cones with nose radii between 5.1 and 50.8-mm. The transition location was determined with coaxial thermocouples and temperature sensitive paint while stability measurements were obtained using high-frequency response pressure sensors. Mean flow and boundary layer-stability computations were also conducted and compared with the experiment. The effect of angle-of-attack and bluntness on the transition location displays similar trends compared to historical hypersonic wind tunnel data at similar Mach and Reynolds numbers. The N factor at start of transition on sharp cones increases with unit Reynolds number. Values between 4 and 7 were observed. The N factor at start of transition significantly decreases as bluntness increases and is successfully correlated with the ratio of transition location to entropy layer swallowing length. Good agreement between the computed and measured spatial amplification rates and most amplified 2 mode frequencies are obtained for sharp and moderately blunted cones. For large bluntness, where the ratio of transition to entropy swallowing length is below 0.1, 2 mode waves were not observed before the start of transition on the frustum.

Laminar stagnation region heating augmentation is investigated in the AEDC Tunnel 9 at Mach 10 by performing high frequency surface pressure and heat transfer measurements on the Orion CEV capsule at zero degree angle-of-attack for unit Reynolds numbers between 0.5 and 15 million per foot. Heating augmentation increases with Reynolds number, but is also model size dependent as it is absent on a 1.25-inch diameter model at Reynolds numbers where it reaches up to 15% on a 7-inch model. Heat transfer space-time correlations on the 7-inch model show that disturbances convect at the boundary layer edge velocity and that the streamwise integral scale increases with distance. Therefore, vorticity amplification due to stretching and piling-up in the stagnation region appears to be responsible for the stagnation point heating augmentation on the larger model. This assumption is reinforced by the f(exp -11/3) dependence of the surface pressure spectrum compared to the f(exp -1) dependence in the free stream. Vorticity amplification does not occur on the 1.25- inch model because the disturbances are too large. Improved free stream fluctuation measurements will be required to determine if significant vorticity is present upstream or mostly generated behind the bow shock.

The Arnold Engineering Development Center (AEDC) Hypervelocity Wind Tunnel No. 9 facility has played a key role in the development of hypersonic vehicles for over 30 years, providing high-quality aerodynamic and aerothermal test data covering high Mach number and high Reynolds number flight simulations. Although Tunnel 9 can achieve flight-level Reynolds numbers and naturally transitioning boundary layers on most test articles, the presence of small pressure fluctuations can complicate the understanding of the boundarylayer transition phenomenon. In an attempt to better characterize the freestream disturbances described as “tunnel noise,” a set of data was collected using flush-mounted pitot acoustic probes. The data quantify the relative pitot acoustic noise of the freestream flow across an extensive bandwidth of 0 to 1000 kHz for the Mach 10 nozzle at a nominal Reynolds number of 2×10/ft in AEDC Tunnel 9.

The Arnold Engineering Development Center (AEDC) Hypervelocity Wind Tunnel No. 9 facility has played a key role in the development of hypersonic vehicles for over 30 years, providing high-quality aerodynamic and aerothermal test data covering high Mach number and high Reynolds number flight simulations. Although Tunnel 9 can achieve flight level Reynolds numbers and naturally transitioning boundary layers on most test articles, the presence of “tunnel noise” can complicate the understanding of the boundary-layer transition phenomenon. In an attempt to better characterize the freestream disturbances described as “tunnel noise” a set of data was collected using a flush-mounted Pitot acoustic probe. The data quantify the relative Pitot acoustic noise of the freestream flow for the Mach 8, 10, and 14 nozzles at AEDC Tunnel 9. The percent noise level for each nozzle varied on the basis of Reynolds number from approximately 2 to 3.5 percent at Mach 8, 2.5 to 4 percent at Mach 10, and 3.75 to 6.25 percent at Mach 14.

* The research reported herein was performed by the Arnold Engineering Development Center (AEDC), Air Force Materiel Command. Work and analysis for this research were performed by Sverdrup Technology, Inc., AEDC Group, technical services contractor for AEDC; SY Technology, Inc., Huntsville, Alabama; and WaveFront Sciences, Albuquerque, New Mexico. Further reproduction is authorized to satisfy needs of the U. S. Government. # Associate Fellow, AIAA + Senior Member, AIAA ABSTRACT A series of aero-optics tests have been carried out at Mach 7 in the Hypervelocity Wind Tunnel 9 (Tunnel 9) at the Arnold Engineering Development Center (AEDC). The test-bed used for the measurements were two flat plates which had sapphire windows mounted in titanium frames. Aero-optic measurements included near-field phase and intensity measurements made with two wavefront sensors, farfield point spread functions made with an imaging camera, and high frequency optical tilts (bore sight error) made with an X-Y Detector. Ancillary measurements of pressure and heat transfer on the testbed plates were also made. The aero-optic measurements coupled with a variety of computations resulted in phase and intensity maps, bore sight errors, contained energy diameters (CED’s) and point spread functions (PSF’s). Comparisons between the various measurements are made to ascertain aerodynamic effects, instrument errors, facility-induced errors and measurement uncertainties.

Hypervelocity Wind Tunnel No. 9, located at the White Oak, MD site of the Arnold Engineering Development Complex (AEDC), has long been recognized as a unique world class ground-test facility. The facility was developed in the early 1970s to provide critical low-altitude, high Mach number data in support of the Navy's reentry development programs. Since its inception, Tunnel 9 has maintained a leading role in hypersonic ground testing by continually expanding its operational capabilities to match the needs of current and projected programs, maintaining data quality, and understanding customer requirements. Tunnel 9 started with a unique design built around a state-of-the-art supply heater that provided a clean, high-pressure, high-temperature nitrogen supply. Initial operation of Tunnel 9 realized a Mach 10 and 14 aerodynamic simulation capability. Additional Mach 7 and 8 high Reynolds number capabilities were subsequently developed. Each upgrade to Tunnel 9 during the past 40 years of operation has been in response to various sponsors or hypersonic basic research requirements. These capability enhancements have helped maintain Tunnel 9's position as a core DoD hypersonic test and evaluation (T&E) ground-test facility, which has been identified as a leading facility in all major hypersonic facility studies. Recent improvements and modernization over the past 10 years have focused on test article measurements and have significantly changed the types, quantity, and quality of test data that are readily acquired in a Tunnel 9 test entry. Recent advancements include major system changes such as a completely new control room and high-speed data system to entirely new measurement capabilities such as global heattransfer measurements using Temperature Sensitive paint technology. These along with other incremental improvements have allowed Tunnel 9 to provide new insights into the physics associated with complex hypersonic flows.