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Amanda Chou

Publications and source records attributed to Amanda Chou.

25 records · Page 2
All sourcesNASA NTRSDOE OSTI
NASA NTRSDate not supplied

Instabilities in the Wake of a Pseudorandom Roughness on a Supersonic Flat Plate

A roughness patch defined by a two-dimensional Fourier series was installed in a highly-polished flat plate model immersed in a supersonic quiet flow. The NASA Langley Research Center Supersonic Low Disturbance Tunnel is capable of providing a Mach 3.5 flow with low freestream acoustic noise and low turbulence levels. Measurements made with a hot-wire probe downstream of the roughness element showed the presence of instability modes that did not cause transition in the measurement region at a freestream unit Reynolds number of Re =12.6×10^6/m. The largest amplitude mode appeared to be similar to a symmetric mode. The same roughness patch immersed in a higher freestream acoustic noise environment showed little to no evidence of the instability modes. Instead, disturbances found in the boundary layer matched closely to the freestream spectra measured in the empty tunnel under the same conditions, indicating some evidence of freestream forcing in the boundary layer.

Roughness induced transition↗
NASA NTRSDate not supplied

Absorption Laser Differential Interferometry for Simultaneous Colinear Flow Property and Fluctuation Measurements

An absorption laser differential interferometer (A-LDI) was constructed and tested in an effort to simultaneously and colinearly measure flow properties (pressure, temperature, concentration) and flow fluctuations (changes in optical path length). Whereas typical LDI and focused LDI (FLDI) systems use fixed wavelength lasers as the light source (often either green at approximately 532 nm or red at approximately 630 nm), a tunable 760 nm diode laser was used to probe the A-band absorption spectrum of molecular oxygen. The A-LDI system was used to measure O2 absorption and flow fluctuations in a sub-atmospheric pressure-controlled test cell, with absorption profiles from direct absorption and wavelength modulation absorption matching simulated results, while simultaneously being sensitive to high frequency fluctuations from jets of either compressed air or 100% O2 from a gas bottle. The A-LDI system was also installed at the NASA Langley Research Center Probe Calibration Tunnel to assess the freestream flow state in a Mach 3.5 flow. LDI measurements of the freestream indicated increasing amplitudes of flow fluctuations with increasing stagnation pressure at a fixed stagnation temperature, while absorption results demonstrated sensitivity to freestream pressure changes.

Joshua M. Weisberger↗
NASA NTRSDate not supplied

Effect of Roughness on Transition on the BOLT-1a Geometry in Supersonic Flow

Infrared thermography measurements on an 8.8% scale BOLT-1a model were made in the NASA Langley Probe Calibration Tunnel for freestream unit Reynolds numbers from 7.7 × 10^ -6 /m to 30.7 × 10^ -6 /m, a freestream Mach number of 3.5, and nominal angle of attack of 0 degrees. The model was printed from polycarbonate and tested before and after fine-grit sanding to evaluate the effect of surface roughness on transition. Global surface roughness measurements were obtained using a 3D optical profilometer. The temperature distributions on the upper and lower experimental surfaces of the test article were simultaneously captured using two infrared cameras. Prior to testing, the facility’s freestream environment was characterized. At a freestream unit Reynolds number of approximately 17 × 10^ -6 /m, the nozzle-wall boundary layer began to transition non-uniformly, inducing transition on the model. The boundary-layer state was determined by analyzing Stanton number distributions computed from the surface temperature measurements. The present supersonic conditions that were examined on BOLT-1a were found to display both similarities and disparities in comparison to previous hypersonic experiments.

boundary layer transition↗
NASA NTRSDate not supplied

Joint AFOSR/ONR/NASA HTP Boundary Layer Transition Workshop Summary for RAeS

A workshop hosted by three US government entities (the Air Force Office of Scientific Research, the Office of Naval Research, and the National Aeronautics and Space Administration Hypersonic Technology Project) was held in June 2023. This workshop focused on high supersonic and hypersonic boundary layer transition. The purpose of the workshop was to gather subject matter experts in the nation to determine the current state of the art in the field, any gaps or areas that require more studies, and the future of the field. Participants of the workshop were drawn from industry, academia, and government agencies. Prior to the workshop they were provided a questionnaire to spark discussion during the workshop. Roadmaps for six different areas were created as one of the final end products of the workshop.

boundary layer transition↗
NASA NTRSDate not supplied

Preliminary Measurements on the BOLT Geometry in the Supersonic Low Disturbance Tunnel

Experiments were performed in the Mach 3.5 Supersonic Low Disturbance Tunnel on the BOundary-Layer Transition (BOLT) geometry. The goal of this campaign was to assess changes to the transition front by varying the freestream noise and the model surface quality. The model was printed of polycarbonate and is 30% scale of the flight geometry. It was tested under noisy and quiet conditions at different streamwise positions in the tunnel and was also tested before and after improvements were made to the model surface through sanding and gap reduction. Pitch and yaw angles were nominally zero. The model surface temperature was measured using infrared thermography and thermocouples as the unit Reynolds number was swept from 4.75 – 15.8 × 10 6 m −1 . Boundary-layer transition was observed near the midspan of the model for unit Reynolds numbers above 10×10 6 m −1 in quiet flow, while transition was already observed on the shoulders below 5×10 6 m −1 in noisy flow. The model surface enhancements improved the symmetry of the transition front for quiet flow but made only slight differences for noisy flow. In quiet flow, the impact of streamwise positioning on transition was more significant on the model shoulders compared to the central region. The difference in the shoulder heating was likely due to variation in noise radiated from the nozzle sidewalls. The total temperature and initial wall temperatures were near 300 K and 293 K, respectively, which generally resulted in negative convective heat-flux values into the model. Hot-wire anemometry was also utilized to characterize the freestream and to conduct a planar survey near the base of the model for the 7.92 × 10 6 m −1 quiet condition. The survey captured the thickening of the boundary layer at the centerline and the vortical nature of the flow outboard of it. Spectral analysis of the mass-flux fluctuations near the midspan demonstrates that the boundary layer is laminar and suggests that the thick boundary layer caused the relatively warm centerline observed in the infrared images.

boundary layer transition↗
NASA NTRSDate not supplied

Preliminary Measurements on the BOLT Geometry in the Supersonic Low Disturbance Tunnel

Experiments were performed in the Mach 3.5 Supersonic Low Disturbance Tunnel on the BOundary-Layer Transition (BOLT) geometry. The goal of this campaign was to assess changes to the transition front by varying the freestream noise and the model surface quality. The model was printed of polycarbonate and is 30% scale of the flight geometry. It was tested under noisy and quiet conditions at different streamwise positions in the tunnel and was also tested before and after improvements were made to the model surface through sanding and gap reduction. Pitch and yaw angles were nominally zero. The model surface temperature was measured using infrared thermography and thermocouples as the unit Reynolds number was swept from 4.75 -- 15.8 x 10^6 m^{-1}. Boundary-layer transition was observed near the midspan of the model for unit Reynolds numbers above 10 x 10^6 m^{-1} in quiet flow, while transition was already observed on the shoulders below 5 x 10^6 m^{-1} in noisy flow. The model surface enhancements improved the symmetry of the transition front for quiet flow but made only slight differences for noisy flow. In quiet flow, the impact of streamwise positioning on transition was more significant on the model shoulders compared to the central region. The difference in the shoulder heating was likely due to variation in noise radiated from the nozzle sidewalls. The total temperature and initial wall temperatures were near 300 K and 293 K, respectively, which generally resulted in negative convective heat-flux values into the model. Hot-wire anemometry was also utilized to characterize the freestream and to conduct a planar survey near the base of the model for the 7.92 x 10^6 m^{-1} quiet condition. The survey captured the thickening of the boundary layer at the centerline and the vortical nature of the flow outboard of it. Spectral analysis of the mass-flux fluctuations near the midspan demonstrates that the boundary layer is laminar and suggests that the thick boundary layer caused the relatively warm centerline observed in the infrared images.

boundary layer transition↗