Eric C. Marineau

Mach 10 Boundary Layer Transition Experiments on Sharp and Blunted Cones (Invited)

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.

Investigation of mach 10 boundary layer stability of sharp cones at angle-of-attack, part 1: Experiments

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.

Development of Non-Intrusive Velocity Measurement Capabilities at AEDC Tunnel 9

Instantaneous velocity measurements are necessary to improve the understanding of hypersonic turbulence and shock/turbulent boundary-layer interactions (STBLI). Such measurements are needed to improve the computational fluid dynamic codes required to reduce the developmental risks of new hypersonic systems. Particle image velocimetry (PIV) has been previously used in small-scale experimental facilities to provide global measurements of velocity fluctuations and Reynolds stresses in hypersonic turbulent boundary layers. However, PIV measurements in large-scale test & evaluation hypersonic facilities such as Tunnel 9 are challenging due to the facility scale, reduced test time, and seeding requirements. Based on an extensive literature review, a PIV development strategy has been developed for Tunnel 9. Using 0.25 micron liquid droplets injected at the wall, Stokes numbers of 0.4 and 0.02 are predicted upstream and downstream of the STBLI region produced by a 33-deg flare at Mach 10 for a free-stream Reynolds number of /ft. PIV development work will begin in the Mach 3 Calibration Laboratory before being transferred to Tunnel 9 for PIV measurements on a large scale hollow-cylinder-flare test article. Schlieren images obtained in the new Mach 3 Calibration Lab indicate that an 11 mm thick turbulent boundary layer can be obtained without tripping.

Development of Particle Image Velocimetry in a Mach 2.7 Wind Tunnel at AEDC White Oak

Particle image velocimetry (PIV) experiments have been performed in the AEDC White Oak Mach 3 Calibration Wind Tunnel to validate a new seeding methodology applicable to Tunnel 9 for hypersonic boundary layer measurements. This methodology involved tangential wall injection of a highly concentrated Polyalphaolefins (PAO-4) aerosol. Validation was performed in a well characterized 7.5 to 9 mm thick Mach 2.7 turbulent boundary layer with a 590 m/s edge velocity. The effect of mass injection on the velocity, as determined from Pitot surveys with and without wall injection and PIV measurements with global and local seeding is small, Δu≈0.7-1.3%. Velocity measurements derived from Pitot pressure and traditional globally seeded PIV agree with the locally seeded PIV measurements. Streamwise velocity fluctuations follow the same trend as DNS data and previous experiments. Image processing techniques were used to perform a quantitative comparison of the particle concentration across the boundary layer. The analysis shows that the local seeding methodology provides a good seeding uniformity over 80% of the boundary layer thickness as well as twice as much near wall particles.

Investigation of Hypersonic Laminar Heating Augmentation in the Stagnation Region

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.

PIV Measurements of Mach 2.7 Turbulent Boundary Layer with Varying Reynolds Numbers

A series of Particle image velocimetry (PIV) experiments have been performed in the AEDC White Oak Mach 3 Calibration Wind Tunnel to investigate the effect of the Reynolds number based on the momentum thickness, Reθ = 1673 − 7130, on mean streamwise and wall normal fluctuating velocity and Reynolds stress profiles in a turbulent boundary layer. High Reθ cases show the Morkovin transformed wall normal fluctuating velocity profile in good agreement with Klebanoff incompressible data as well as DNS above ~0. 4δ, and streamwise fluctuating velocity agreement above 0. 1δ. No conclusive fluctuating velocity magnitude dependence on Reθ is observed. The greater error at lower spatial resolution and higher Stokes number indicates that the spatial resolution and particle response have a greater influence on fluctuating velocity profiles. Estimates of the power spectra density from PIV measurements reveal truncation of energy at high wavenumbers due to particle lag and spatial filtering. The flatter spectra of the wall normal velocity, compared to the streamwise direction, leads to increased energy at high wavenumbers which could explain why the wall normal fluctuating velocity profile is more susceptible to measurement error from particle lag and spatial filtering.

Direct Measurements of Skin Friction at AEDC Hypervelocity Wind Tunnel 9

This investigation concerns the design, build, and testing of a new class of direct-measuring skin friction sensors capable of performing favorably in sustained hypervelocity flow conditions and detecting boundary layer transition effects. A rigorous, multi-step approach was developed to systematically test the sensors through various bench test setups and wind tunnel facilities. First, the sensors underwent NIST-traceable force calibrations to ensure that any measured wall shear values taken in testing were obtained and reported with well-documented experimental uncertainties. The calibration process was conducted through a series of tests to characterize the new sensors static, thermal, pressure, and dynamic responses. Validation testing was then conducted in the medium-scale, blow-down Supersonic Tunnel at Virginia Tech under Mach 4.0 flow conditions. Once completed, the final capstone investigations were conducted in AEDC Wind Tunnel 9, a unique hypervelocity blowdown wind tunnel that uses a nitrogen gas test medium. One of the skin friction sensors was integrated into a steel 155.6 cm-(61.27 in.-) long, 7-degree half-angle, cone model. Flow was nominally maintained at Mach 10 and a stagnation temperature of 1000 K (1800 °R). The stagnation pressure and unit Reynolds number were varied over the range of 2.3 to 43.4 MPa (330 to 6300 psia) and 1.6 to 30.3x10/m (0.5 to 9.24x10/ft), respectively. Skin friction was measured over several boundary layer states including near-laminar, transitional, and turbulent flows. Wall shear measurements ranged between 0.92 and 340 Pa (0.02 and 7.1 psf), while the skin friction coefficients ranged between 0.0003 and 0.0060. The total uncertainty of the skin friction sensor remained at ±9.2% of the measurement for a 95% confidence level. The magnitude of skin friction depended on the state of the boundary layer, the Reynolds number of the flow, the nose tip radius of the cone model, and angle-of-attach under pitching runs.

The Hypervelocity Wind Tunnel No. 9 - Continued excellence through improvement and modernization

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 Navys 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 9s 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.

Compilation and Analysis of Second Mode Amplitudes on Sharp Cones in Hypersonic Wind Tunnels

This research effort coordinated by the NATO AVT-240 specialists’ group compiles and analyzes second-mode amplitudes on sharp slender cones at 0 degrees angle of attack. The analysis focuses on pressure fluctuations measured with piezoelectric sensors in 11 hypersonic wind tunnels operated by 9 organizations located in 3 NATO countries (Belgium, Germany, and USA) and Japan. The measurements are at freestream Mach numbers between 5 and 14, unit Reynolds numbers Re/m between 1.5 and 16 million per meter, and wall-to total temperature ratios between 0.1 and 0.8. The study shows that second-mode growth rates can be predicted with Parabolized Stability Equations (PSE) over the wide range of conditions. The maximum second-mode amplitudes vary weakly at edge Mach number Me greater than ~5.8, but significantly decrease at lower Me. The maximum N factor envelope from PSE and the measured amplitudes are used to estimate the initial amplitudes A0. At each Mach number, A0 varies approximately as Re/m^(-1). This leads to transition N factors that increase with Re/m. This behavior is consistent with the results from Marineau (AIAA Journal, Vol. 55, No. 2, 2017).

Direct Skin Friction Measurements at Mach 10 in a Hypervelocity Wind Tunnel

This investigation concerns the design, build, and testing of direct-measuring skin friction sensors capable of performing in sustained hypersonic flow and detecting transition. A multistep approac...