Publications

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Two novel and challenging applications of high-frequency pressure-sensitive paint were attempted for ground testing at Sandia National Labs. Blast tube testing, typically used to assess the response of a system to an incident blast wave, was the first application. The paint was tested to show feasibility for supplementing traditional pressure instrumentation in the harsh outdoor environment. The primary challenge was the background illumination from sunlight and time-varying light contamination from the associated explosion. Optimal results were obtained in pre-dawn hours when sunlight contamination was absent; additional corrections must be made for the intensity of the explosive illumination. A separate application of the paint for acoustic testing was also explored to provide the spatial distribution of loading on systems that do not contain pressure instrumentation. In that case, the challenge was the extremely low level of pressure variations that the paint must resolve (120 dB). Initial testing indicated the paint technique merits further development for a larger scale reverberant chamber test with higher loading levels near 140 dB.

Femtosecond Laser Electronic Excitation Tagging (FLEET) is used to measure velocity flowfields in the wake of a sharp 7◦ half-angle cone in nitrogen at Mach 8, over freestream Reynolds numbers from 4.3∗106 /m to 13.8∗106 /m. Flow tagging reveals expected wake features such as the separation shear layer and two-dimensional velocity components. Frequency-tripled FLEET has a longer lifetime and is more energy efficient by tenfold compared to 800 nm FLEET. Additionally, FLEET lines written with 267 nm are three times longer and 25% thinner than that written with 800 nm at a 1 µs delay. Two gated detection systems are compared. While the PIMAX 3 ICCD offers variable gating and fewer imaging artifacts than a LaVision IRO coupled to a Photron SA-Z, its slow readout speed renders it ineffective for capturing hypersonic velocity fluctuations. FLEET can be detected to 25 µs following excitation within 10 mm downstream of the model base, but delays greater than 4 µs have deteriorated signal-to-noise and line fit uncertainties greater than 10%. In a hypersonic nitrogen flow, exposures of just several hundred nanoseconds are long enough to produce saturated signals and/or increase the line thickness, thereby adding to measurement uncertainty. Velocity calculated between the first two delays offer the lowest uncertainty (less than 3% of the mean velocity).

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A numerical study of the response of a conical structure to periodic turbulent spot loading at Mach 6 is conducted and compared with experimental results. First, a deterministic model which describes the birthing of turbulent spots established by a defined forcing frequency as well as the evolution of the spots is derived. The model is then used to apply turbulent spot loading to a calibrated finite element model of a slender cone structure. The numerical solution yielded acceleration response data for the cone structure. These data are compared to experimental measurement. Similar damping times and acceleration amplitudes are observed for isolated spots. At higher frequencies of turbulent spot generation, the panel response corresponds to the structural natural mode shape being forced; however, only qualitative agreement is observed. Finally, the convection velocity for two cases is varied. It is shown that marginal deviations in the convection velocity of turbulent spots yields little change in the resulting response of a structure. This result illustrates that the time between spot events provides the dominant determination of which structural modes are excited.

This paper details a joint numerical and experimental investigation of transition-delaying roughness. A numerical simulation was undertaken to design a surface roughness configuration that would suppress Mack’s 2nd mode instability in order to maintain laminar flow over a Mach 8 hypersonic blunt cone. Following the design process the roughness configuration was implemented on a hypersonic cone test article. Multiple experimental runs at the Mach 8 condition with different Reynolds numbers were run, as well as an off-design Mach 5 condition. The roughness did appear to delay transition in the Mach 8 case as intended, but did not appear to delay transition in the Mach 5 case. Concurrently, simulations of the roughness configuration were also computed for both Mach cases utilizing the experimental conditions. Linear stability theory was applied to the simulations in order to determine their boundary layer stability characteristics. This investigation of multiple cases helps to validate the numerical code with real experimental results as well as provide physical evidence for the transition-delaying roughness phenomenon.

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Prediction of boundary-layer transition is a critical part of the design of hypersonic vehicles because of the large increase in skin-friction drag and surface heating associated with the onset of transition. Testing in conventional (noisy) wind tunnels has been an important means of characterizing and understanding the boundary-layer transition (BLT) behavior of hypersonic vehicles. Because the existing low disturbance, i.e., quiet, facilities operate only at Mach 6, moderate Reynolds numbers, fairly small sizes, and low freestream enthalpy, conventional facilities will continue to be employed for testing and evaluation of hypersonic vehicles, especially for ground testing involving other Mach numbers, higher freestream enthalpies, and larger models. To enable better use of transition data from conventional facilities and more accurate extrapolation of wind-tunnel results to flight, one needs an in-depth knowledge of the broadband disturbance environment in those facilities as well as of the interaction between the freestream disturbances with laminar boundary layers.

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Fluid-structure interactions were studies on a 7° half-angle cone in the Sandia Hypersonic Wind Tunnel at Mach 5 and 8 and in the Purdue Boeing/AFOSR Mach 6 Quiet Tunnel. A thin composite panel was integrated into the cone and the response to boundary-layer disturbances was characterized by accelerometers on the backside of the panel. Here, under quiet-flow conditions at Mach 6, the cone boundary layer remained laminar. Artificially generated turbulent spots excited a directionally dependent panel response which would last much longer than the spot duration.

While low disturbance (“quiet”) hypersonic wind tunnels are believed to provide more reliable extrapolation of boundary layer transition behavior from ground to flight, the presently available quiet facilities are limited to Mach 6, moderate Reynolds numbers, low freestream enthalpy, and subscale models. As a result, only conventional (“noisy”) wind tunnels can reproduce both Reynolds numbers and enthalpies of hypersonic flight configurations, and must therefore be used for flight vehicle test and evaluation involving high Mach number, high enthalpy, and larger models. This article outlines the recent progress and achievements in the characterization of tunnel noise that have resulted from the coordinated effort within the AVT-240 specialists group on hypersonic boundary layer transition prediction. New Direct Numerical Simulation (DNS) datasets elucidate the physics of noise generation inside the turbulent nozzle wall boundary layer, characterize the spatiotemporal structure of the freestream noise, and account for the propagation and transfer of the freestream disturbances to a pitot-mounted sensor. The new experimental measurements cover a range of conventional wind tunnels with different sizes and Mach numbers from 6 to 14 and extend the database of freestream fluctuations within the spectral range of boundary layer instability waves over commonly tested models. Prospects for applying the computational and measurement datasets for developing mechanism-based transition prediction models are discussed.

The spanwise variation of resonance dynamics in the Mach 0.94 flow over a finite-span cavity of variable length-to-width ratio was explored using time-resolved particle image velocimetry (TR-PIV) in a planform plane above the cavity and time-resolved pressure sensitive paint (TR-PSP) on the floor and adjacent exterior surface. The TR-PIV showed a significant variation in resonant fluctuations to occur across the span of the cavity, which appears to arise from spillage vortices stemming from finite width effects. Thus, the spanwise variation was a strong function of the cavity aspect ratio and was only weakly dependent on the cavity mode number. Modal streamwise velocity fluctuations in the spillage vortices showed large peaks at modes one through three, indicating that resonance dynamics, and not just broadband turbulence effects, are prevalent near the sidewalls. Large peaks in modal pressures were also present on the walls just outside of the cavity. Interestingly, prominent peaks at the mode frequencies were observed in the spanwise velocity spectra as well. These peaks were strongest near the cavity sidewalls suggesting a coupling between the resonance mechanism and the spillage vortices.

Fluid-structure interactions were studied on a 7◦ half-angle cone in the Sandia Hypersonic Wind Tunnel at Mach 5 and 8 and in the Purdue Boeing/AFOSR Mach 6 Quiet Tunnel. A thin composite panel was integrated into the cone and the response to boundary-layer disturbances was characterized by accelerometers on the backside of the panel. Under quiet-flow conditions at Mach 6, the cone boundary layer remained laminar. Artificially generated turbulent spots excited a directionally dependent panel response which would last much longer than the spot duration. When the spot generation frequency matched a structural natural frequency of the panel, resonance would occur and responses over 200 times greater than under a laminar boundary layer were obtained. At Mach 5 and 8 under noisy flow conditions, natural transition driven by the wind-tunnel acoustic noise dominated the panel response. An elevated vibrational response was observed during transition at frequencies corresponding to the distribution of turbulent spots in the transitional flow. Once turbulent flow developed, the structural response dropped because the intermittent forcing from the spots no longer drove panel vibration.

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