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Surface Pressure Fluctuations Induced by a Hypersonic Turbulent Boundary Layer on a Sharp Cone at Angle of Attack
High-fidelity simulations are performed to characterize the turbulence-induced wall pressure fluctuations on a sharp cone at a 5.5° angle-of-attack in a Mach 8 flow. Wall-resolved large-eddy simulation (LES) and wall-modeled large-eddy simulation (WMLES) results are compared to measurements at several locations on the cone body, where comparisons with high-frequency PCB sensors are good, while comparison with low-frequency kulite sensors varies depending on the location. For a given streamwise location, simulation results show significant azimuthal variation in the wall pressure fluctuation statistics. Comparisons between LES and WMLES results indicate that WMLES can accurately reproduce the low-frequency component of the autospectra quite well, while slight sensitivity is apparent in the coherence. A grid sensitivity study for WMLES further demonstrated sensitivity in the wall pressure fluctuation coherence, but not in the autospectra.
Hypersonic Thermal Boundary Layer Profiles via Femtosecond Coherent Anti-Stokes Raman Spectroscopy
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FLEET Velocimetry to Probe Hypersonic Transitional Boundary Layers
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Hypersonic Boundary-Layer Receptivity to Oblique Vorticity Waves
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Hypersonic turbulent boundary layer measurements using an electron beam Final report, Oct. 1966 - May 1968
Direct measurement of density in turbulent boundary layer on shock tunnel nozzle wall using electron beam probe
Hypersonic boundary-layer transition in the presence of wind-tunnel noise.
Description of additional results indicating that a unique relationship exists between rms sound pressure level and transition Reynolds number, provided all other parameters are held nearly invariant. The results illustrate the different levels of the transition Reynolds numbers obtained in six facilities with nearly invariant laminar boundary layers on the different models (at the same local unit Reynolds number) and indicate that some facility characteristic has a strong influence on the transition Reynolds numbers.
Numerical solution of the interaction of a strong shock wave with a hypersonic turbulent boundary layer
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Some topics in hypersonic turbulent boundary layer, heat transfer, and skin friction
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Real gas effects on hypersonic boundary-layer stability
High-temperature effects alter the physical and transport properties of a gas, air in particular, due to vibrational excitation and gas dissociation, and thus the chemical reactions have to be considered in order to compute the flow field. Linear stability of high-temperature boundary layers is investigated under the assumption of chemical equilibrium and this gas model is labeled here as real gas model. In this model, the system of stability equations remains of the same order as for the perfect gas and the effect of chemical reactions is introduced only through mean flow and gas property variations. Calculations are performed for Mach 10 and 15 boundary layers and the results indicate that real gas effects cause the first mode instability to stabilize while the second mode is made more unstable. It is also found that the second mode instability shifts to lower frequencies. There is a slight destabilizing influence of real gas on the Goertler instability as compared to the perfect gas results.
Wall Cooling Effects on Hypersonic Transitional/Turbulent Boundary Layers at High Reynolds Numbers
A 4 degree wedge was used to produce a thick turbulent boundary layer with an edge Mach number of 11. By using a two-dimensional model, the boundary layer was nearly free from upstream history effects associated with nozzle wall turbulent boundary layers. Heat-transfer distributions were used to define regions of laminar, transitional, and turbulent flow at several values of T(sub w)/T(sub t) for an edge unit Reynolds number of 0.47 x lot per cm. Pitot and total temperature profiles and skin-friction measurements were obtained at selected stations along the model. Turbulence parameters (mixing length/sigma and epsilon) were derived from the fully turbulent profiles and used to more completely define the "low Reynolds number" effect. Turbulent Prandtl number distributions are also presented.
Hypersonic Laminar Boundary Layer Velocimetry with Discrete Roughness on a Flat Plate
Laminar boundary layer velocity measurements are made on a 10-degree half-angle wedge in a Mach 10 flow. Two types of discrete boundary layer trips were used to perturb the boundary layer gas. The first was a 2-mm tall, 4-mm diameter cylindrical trip. The second was a scaled version of the Orbiter Boundary Layer Transition (BLT) Detailed Test Objective (DTO) trip. Both 1-mm and 2.5-mm tall BLT DTO trips were tested. Additionally, side-view and plan-view axial boundary layer velocity measurements were made in the absence of these tripping devices. The free-stream unit Reynolds numbers tested for the cylindrical trips were 1.7x10(exp 6)/m and 3.3x10(exp 6)/m. The free-stream unit Reynolds number tested for the BLT DTO trips was 1.7x10(exp 6)/m. The angle of attack was kept at approximately 5-degrees for most of the tests resulting in a Mach number of approximately 8.3. These combinations of unit Reynolds numbers and angle of attack resulted in laminar flowfields. To study the precision of the measurement technique, the angle of attack was varied during one run. Nitric-oxide (NO) molecular tagging velocimetry (MTV) was used to obtain averaged axial velocity values and associated uncertainties. These uncertainties are as low as 20 m/s. An interline, progressive scan CCD camera was used to obtain separate images of the initial reference and shifted NO molecules that had been tagged by the laser. The CCD configuration allowed for sub-microsecond sequential acquisition of both images. The maximum planar spatial resolution achieved for the side-view velocity measurements was 0.07-mm in the wall-normal direction by 1.45-mm in the streamwise direction with a spatial depth of 0.5-mm. For the plan-view measurements, the maximum planar spatial resolution in the spanwise and streamwise directions was 0.69-mm by 1.28-mm, respectively, with a spatial depth of 0.5-mm. Temperature sensitive paint (TSP) measurements are provided to compliment the velocity data and to provide further insight into the behavior of the boundary layers. The experiments were performed at the NASA Langley Research Center 31-Inch Mach 10 Air tunnel.
Simulation and Modeling of Hypersonic Turbulent Boundary Layers Subject to Pressure Gradient and Wall Cooling (Year 1)
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Hypersonic Boundary-Layer Transition on Blunted Cones at Angle of Attack
Experimental studies of cones at several high-speed facilities have demonstrated that, for small nosetip bluntness, transition onset over a circular cone moves upstream along the leeward side and downstream along the windward side, but this trend may be reversed at large bluntness values, where transition onset moves downstream along the leeward side and upstream along the windward side. A theoretical and numerical investigation is performed to characterize the effects of nose bluntness on disturbance amplification over the circular cone for several angles of attack, with the goal of understanding the potential physical mechanisms behind the experimental observations. The three-dimensional laminar basic states over a 1.5 m long, 7-degree half-angle cone with 9.525 mm nosetip radius are computed for selected angles of attack values and freestream conditions that are selected to match the Mach 10 experiments conducted within the Hypervelocity Wind Tunnel 9 at the Arnold Engineering Development Complex (AEDC). The solutions at a freestream unit Reynolds number of 17.1 million per meter are used to perform detailed instability analyses for angles of attack equal to 0, 1, 3, and 5 degrees. Results indicate that the linear amplification of stationary crossflow waves along inflection lines may begin to influence transition along the acreage of the cone for angles of attack equal to or larger than 5 degrees. The measured trend in transition front with respect to increasing angle of attack is found to be consistent with the predicted increase in the amplification factors for Mack mode disturbances along the streamline trajectories. The increase in Mack mode amplification along the windward ray for higher angles of attack is shown to be the result of a progressively earlier entropy-layer swallowing. Computations also indicate that the transition amplification factor along the windward ray is not constant and increases with the angle of attack and that the transition #-factors along the leeward ray are rather small. The nonmodal analysis for zero degrees angle of attack shows that entropy-layer disturbances with appreciably strong energy growth can coexist with Mack mode instabilities at the measured transition location.
Hypersonic Boundary-Layer Transition on Blunted Cones at Angle of Attack
Experimental studies of cones at several high-speed facilities have demonstrated that, for small nosetip bluntness, transition onset over a circular cone moves upstream along the leeward side and downstream along the windward side, but this trend may be reversed at large bluntness values, where transition onset moves downstream along the leeward side and upstream along the windward side. A theoretical and numerical investigation is performed to characterize the effects of nose bluntness on disturbance amplification over the circular cone for several angles of attack, with the goal of understanding the potential physical mechanisms behind the experimental observations. The three-dimensional laminar basic states over a 1.5 m long, 7-degree half-angle cone with 9.525 mm nosetip radius are computed for selected angles of attack values and freestream conditions that are selected to match the Mach 10 experiments conducted within the Hypervelocity Wind Tunnel 9 at the Arnold Engineering Development Complex (AEDC). The solutions at a freestream unit Reynolds number of 17.1 million per meter are used to perform detailed instability analyses for angles of attack equal to 0, 1, 3, and 5 degrees. Results indicate that the linear amplification of stationary crossflow waves along inflection lines may begin to influence transition along the acreage of the cone for angles of attack equal to or larger than 5 degrees. The measured trend in transition front with respect to increasing angle of attack is found to be consistent with the predicted increase in the amplification factors for Mack mode disturbances along the streamline trajectories. The increase in Mack mode amplification along the windward ray for higher angles of attack is shown to be the result of a progressively earlier entropy-layer swallowing. Computations also indicate that the transition amplification factor along the windward ray is not constant and increases with the angle of attack and that the transition #-factors along the leeward ray are rather small. The nonmodal analysis for zero degrees angle of attack shows that entropy-layer disturbances with appreciably strong energy growth can coexist with Mack mode instabilities at the measured transition location.
Hypersonic Boundary-Layer Instabilities over Cone-Flare Models
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Hypersonic Boundary-Layer Instabilities over Ogive-Cylinder Models
Computational investigations of an ogive-cylinder geometry with varying nosetips at zero degrees angle of attack are presented. The model geometry and conditions are selected to match experiments conducted in the Air Force Research Laboratory (AFRL) Mach 6 Ludwieg Tube. Five nosetips of interest were selected for the computational studies herein: two sharp ogives, two blunt ogives, and one hemispherical nosetip. Computations are performed at a freestream Reynolds number of 7.01 × 10 6 m -1 . Each sharp and blunt tip ogives had a 14 and 28 degree version for the tip angles. A cylindrical section follows the nosetip, resulting in a meter long model such as in the experiments. The laminar flow solutions are analyzed. The boundary-layer-edge properties and the velocity and temperature profiles are compared across streamwise locations aft of the ogive-cylinder junction. The blunt nosetips induce an entropy layer that envelopes the boundary-layer profiles. Modal stability analysis identifies most amplified frequencies corresponding to Mack’s second modes that agree with experimental results for the sharp tips. Similar to the experimental measurements based on wall-mounted pressure sensors, no unstable modes are found for the blunt models. Nonmodal analysis revealed a broadband set of disturbances present for the blunter tips, in agreement with experimental observations. Flow perturbation contours of most amplified planar and oblique disturbances are shown to qualitatively match wind tunnel Schlieren images, with a switch from rope-like to elongated structures, i.e., from high frequency Mack’s second modes to low frequency Mack’s first modes, as the nosetip angle is increased for the sharp tip, and from boundary-layer to entropy-layer disturbances as the bluntness is increased.
Thermomechanical Relaxation of Vibrationally Excited NO in a Hypersonic Turbulent Boundary Layer
The effects of thermal nonequilibrium are studied using a DC glow discharge plasma to selectively modify the vibrational temperature of an otherwise cold flow. These experiments were performed at 𝑴=5.7and 𝑹𝒆=6×10 6 /m in the Actively Controlled Expansion tunnel at Texas A&M University on a canonical 2.75° half-angle wedge test article. A∼47 W plasma was generated across spanwise electrodes, and removable trips produced either a laminar or turbulent boundary layer. NO planar laser-induced fluorescence (PLIF) quantified the extent of the NO 𝑿 2 𝚷 (𝒗=1) vibrational perturbation due to the plasma and its subsequent interplay with the boundary layer, the first such data of which the authors are aware. NOPLIF measurements were taken at three locations along the test article, and quasi-temperature measurements are reported at each. Despite not quantifiably perturbing the boundary layer, the plasma was shown to alter the NO vibrational temperature profile therein. Independent analysis of the thermodynamics and fluid mechanics of the system suggests each set of physics could explain the results observed near and far from the wall, respectively.