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NASA NTRS1994

Computational Analysis in Support of the SSTO Flowpath Test

A synergistic approach of combining computational methods and experimental measurements is used in the analysis of a hypersonic inlet. There are four major focal points within this study which examine the boundary layer growth on a compression ramp upstream of the cowl lip of a scramjet inlet. Initially, the boundary layer growth on the NASP Concept Demonstrator Engine (CDE) is examined. The follow-up study determines the optimum diverter height required by the SSTO Flowpath test to best duplicate the CDE results. These flow field computations are then compared to the experimental measurements and the mass average Mach number is determined for this inlet.

Duncan, Beverly S.↗
NASA NTRS2007

Aerosol Source Plume Physical Characteristics from Space-based Multiangle Imaging

Models that assess aerosol effects on regional air quality and global climate parameterize aerosol sources in terms of amount, type, and injection height. The multiangle imaging spectroradiometer (MISR) aboard NASA's Terra satellite retrieves total column aerosol optical thickness (AOT), and aerosol type over cloud-free land and water. A stereo-matching algorithm automatically retrieves reflecting-layer altitude wherever clouds or aerosol plumes have discernable spatial contrast, with about 500-m accuracy, at 1.1-km horizontal resolution. Near-source biomass burning smoke, volcanic effluent, and desert dust plumes are observed routinely, providing information about aerosol amount, particle type, and injection height useful for modeling applications. Compared to background aerosols, the plumes sampled have higher AOT, contain particles having expected differences in Angstrom exponent, size, single-scattering albedo, and for volcanic plume and dust cloud cases, particle shape. As basic thermodynamics predicts, thin aerosol plumes lifted only by regional winds or less intense heat sources are confined to the boundary layer. However, when sources have sufficient buoyancy, the representative plumes studied tend to concentrate within discrete, high-elevation layers of local stability; the aerosol is not uniformly distributed up to a peak altitude, as is sometimes assumed in modeling. MISR-derived plume heights, along with meteorological profile data from other sources, make it possible to relate radiant energy flux observed by the moderate resolution imaging spectroradiometer (MODIS), also aboard the Terra spacecraft, to convective heat flux that plays a major role in buoyant plume dynamics. A MISR climatology of plume behavior based on these results is being developed.

aerosol plumes↗
NASA NTRS1994

Experimental studies of shock-wave/wall-jet interaction in hypersonic flow, part A

Experimental studies have been conducted to examine slot film cooling effectiveness and the interaction between the cooling film and an incident planar shock wave in turbulent hypersonic flow. The experimental studies were conducted in the 48-inch shock tunnel at Calspan at a freestream Mach number of close to 6.4 and at a Reynolds number of 35 x 10(exp 6) based on the length of the model at the injection point. The Mach 2.3 planar wall jet was generated from 40 transverse nozzles (with heights of both 0.080 inch and 0.120 inch), producing a film that extended the full width of the model. The nozzles were operated at pressures and velocities close to matching the freestream, as well as at conditions where the nozzle flows were over- and under-expanded. A two-dimensional shock generator was used to generate oblique shocks that deflected the flow through total turnings of 11, 16, and 21 degrees; the flows impinged downstream of the nozzle exits. Detailed measurements of heat transfer and pressure were made both ahead and downstream of the injection station, with the greatest concentration of measurements in the regions of shock-wave/boundary layer interaction. The major objectives of these experimental studies were to explore the effectiveness of film cooling in the presence of regions of shock-wave/boundary layer interaction and, more specifically, to determine how boundary layer separation and the large recompression heating rates were modified by film cooling. Detailed distributions of heat transfer and pressure were obtained in the incident-shock/wall-jet interaction region for a series of shock strengths and impingement positions for each of the two nozzle heights. Measurements were also made to examine the effects of nozzle lip thickness on cooling effectiveness. The major conclusion from these studies was that the effect of the cooling film could be readily dispersed by relatively weak incident shocks, so the peak heating in the recompression region was not significantly reduced by even the largest levels of film cooling. For the case studies in the absence of film cooling, the interaction regions were unseparated. However, adding film cooling resulted in regions of boundary layer separation induced in the film cooling layer, the size of which regions first increased and then decreased with increased film cooling. Surprisingly, the size of the separated regions and the magnitude of the recompression heating were not strongly influenced by the thickness of the cooling film, nor by the point of shock impingement relative to the exit plane of the nozzles. The lip thickness was found to have little effect on cooling effectiveness. Measurements with and in the absence of shock interaction were compared with the results of earlier experimental studies and correlated in terms of the major parameters controlling these flows.

Holden, Michael S.↗
NASA NTRS2019

Shockwave/Boundary-Layer Interaction Studies Performed in the NASA Langley 20-Inch Mach 6 Air Tunnel

This paper highlights results from a collaborative study performed by The University of Tennessee Space Institute (UTSI) and NASA Langley Research Center on the Shockwave/Boundary-Layer Interaction (SWBLI) generated by a cylindrical protuberance on a flat plate in a Mach 6 flow. The study was performed in the 20-Inch Mach 6 Air Tunnel at NASA Langley Research Center and consisted of two separate entries. In the first entry, simultaneous high-speed schlieren and high-speed pressure-sensitive paint (PSP) imaging – which was performed for the first time in the 20-Inch Mach 6 facility at NASA Langley – were performed as well as simultaneous high-speed schlieren and oil-flow imaging. In the second entry, the model configuration was modified to increase the size of the interaction region. High-speed schlieren and infrared thermography (IR) surface imaging were performed in this second entry. The goal of these tests was to characterize the SBLI in the presence of a laminar, transitional, and turbulent boundary layer using high-speed optical imaging techniques. AoA = sting angle-of-attack (°) dcylinder = cylinder diameter (mm) dtrip = cylindrical tripping element diameter (mm) Δshock = shock stand-off distance (mm) hcylinder = cylinder height (mm) htrip = cylindrical tripping element height (mm) HSS = high-speed schlieren M∞ = freestream Mach number PSP = pressure-sensitive paint Re∞ = freestream unit Reynolds number (m-1) SWBLI = shockwave/boundary-layer interaction θplate = model plate angle (°) Introduction his paper highlights two experimental entries performed in the 20-Inch Mach 6 Air Blowdown Tunnel at NASA Langley Research Center in collaboration with The University of Tennessee Space Institute (UTSI). The purpose of these entries was to characterize the dynamic shockwave/boundary-layer interaction (SWBLI) between a vertical cylinder on a flat plate and laminar, transitional (XSWBLI), and turbulent (SWTBLI) boundary layers with a freestream Mach number of 6 using non-intrusive optical diagnostics. Experiments performed by Murphree et al.1,2 were among the first to specifically characterize XSWBLI induced by a vertical cylinder on a flat plate geometry using several optical measurement techniques. Recent optical studies of XSWBLI phenomenon have been performed by UTSI at Mach 2 in their low-enthalpy blow wind tunnel3-8 and by Texas A&M University and UTSI at Mach numbers of 6 and 7 in their Adjustable Contour Expansion wind tunnel.9 The experiments described in this paper were intended to complement previous studies by expanding the freestream unit Reynolds number range, Re∞, over which the XSWBLI phenomena has been observed. Additionally these experiments, made possible under NASA’s new facility funding model under the Aeronautics Evaluation and Test Capabilities (AETC) project, promoted collaboration between university and NASA researchers. The initial entry in the 20-Inch Mach 6 Air Tunnel at NASA Langley occurred in December of 2016. Originally, testing was to occur in November of 2016 in the 31-Inch Mach 10 Air Tunnel at NASA Langley. This facility was chosen so that the XSWBLI phenomenon could be observed at much higher Mach numbers than had previously been attempted in ground test experiments. The model selected for this experiment, a 10° half-angle wedge with a sharp leading edge (described in detail in section II.B), had previously been used by Danehy et al. [10] for boundary layer transition studies using the nitric oxide planar laser-induced fluorescence (NO PLIF) flow visualization technique. In that work, it was determined that transition could be induced downstream of a single htrip = 1-mm tall, dtrip = 4-mm diameter cylindrical tripping element and that the streamwise location of the transition could be changed for a single Re∞ by changing the model angle-of-attack (AoA) (see Fig. A3 in Ref. [10] for more details). Based on the findings of that work, a decision was made to use the wedge model with the cylindrical tripping element to trip the boundary layer flow ahead of a cylindrical protuberance in order to achieve a XSWBLI. Unfortunately, the 31-Inch Mach 10 facility had been taken offline for repairs in October of 2016 and a decision was made to move the test to the 20-Inch Mach 6 facility. Since the behavior of the boundary layer with the chosen model configuration had not been studied before in that facility and the available test time was limited, the entry was considered to be exploratory and was used to collect spatially-resolved and time-resolved flow and surface visualization data that would be used to inform a second entry. Test techniques included simultaneous high-speed schlieren (HSS) captured at 160 kHz and high-speed pressure sensitive paint captured at 10 kHz as well as oil flow visualization, captured at 750 Hz. The second entry in the 20-Inch Mach 6 facility occurred in June and July of 2017. In this follow-on test, modifications to the wind tunnel model were made based on observations made during the first entry and included removing the cylindrical tripping element, increasing the size of the cylinder used to induce the SWBLI to increase the size of the interaction while simultaneously improving spatial resolution, and using a swept ramp array, similar to that described in Ref. [11], to trip the flow to turbulence. Simultaneous HSS (captured at 140 kHz, 100 kHz, and 40 kHz) and conventional IR thermography (captured at 30 Hz) imaging were performed simultaneously in this follow-on entry. This paper is intended to serve as a summary of the work performed during these two entries, to detail lessons learned from each entry, and to highlight some of the datasets acquired. Details on the experimental setup, model configuration, and techniques used are provided. Papers providing a more rigorous analysis of data acquired during the second entry, including statistical, spectral, and modal decomposition methods, can be found in Refs. [12,13]. An entry examining XSWBLI in the 31-Inch Mach 10 Blowdown Wind Tunnel facility is currently planned for mid-to-late calendar year 2019, pending the success of facility repairs. The work performed and described in this paper and the upcoming entry in the 31-Inch Mach 10 facility at NASA Langley have been made possible by NASA’s new facility funding model under the Aeronautics Evaluation and Test Capabilities (AETC) project. Wind Tunnel Facility All experiments discussed in this paper were performed in the 20-Inch Mach 6 Air Tunnel at NASA Langley Research Center. Specific details pertaining to this facility can be found in Refs. [14,15], with only a brief description of the facility provided here. For both entries, the nominal freestream unit Reynolds number was varied between 1.8×106 m-1 (0.5×106 ft-1) and 26.3×106 m-1 (8×106 ft-1). The nominal stagnation pressure was varied between 0.21 MPa and 3.33 MPa and the nominal stagnation temperature was varied between 480 K and 520 K to achieve the desired Re∞ condition. For all runs, the nominal freestream Mach number was 6. The nearly square test section is 520.7-mm (20.5-inches) wide by 508-mm (20-inches) high. Two 431.8-mm (17-inch) diameter windows made of Corning 7940, Grade 5F schlieren-quality glass serve as the side walls of the tunnel and provide optical access for the high-speed schlieren measurements. A rectangular window made of the same material as the side windows served as the top wall of the test section and provided optical access for the high-speed PSP and oil flow measurements. For the second entry, this top window was replaced with a Zinc Selenide (ZnSe) window with an anti-reflection coating capable of passing IR wavelengths between 8μm and 12μm with greater than 98% transmittance. The model was sting supported by a strut attached to a hydraulic system that allows for the model pitch angle to be adjusted between -5° to +55°. For the first entry, an initial pitch/pause sweep of the model AoA was performed to observe the resulting SWBLI. Ultimately, however, the sting pitch angle for this entry was fixed at +10.0° so that the angle of the top surface of the wedge relative to the streamwise axis of the tunnel (referred to herein as the plate angle, θplate), was θplate = 0°. For the second entry, θplate = 0° and θplate = -13.25° were initially tested with the swept ramp array (discussed in the following section) to determine which orientation produced conditions most favorable for XSWBLI to occur based on the heating signatures observed over the top surface of the model in the IR thermography images. Based on these initial tests, θplate = -13.25° was set for the remainder of the runs in the second entry. For both entries, any model changes were performed in a housing located beneath the closed test section. Prior to performing a run of the tunnel, the housing was sealed and the tunnel started. Once the appropriate freestream conditions were achieved, the model was injected into the test section using a hydraulic injection system. B. Model Geometry For all runs, a 10° half-angle (20° full-angle) wedge model with a sharp leading edge was used. The model is described in detail in Refs. [10,16]. The top surface of the sharp leading edge of the model extended 47.8 mm from its upstream-most edge to a junction with the upstream edge of a stainless steel top plate that then extended an (a) (c) (b) Fig. 1 (a) Schematic of top surface of wedge model with gas seeding insert, (b) perspective view of the model in the 20-Inch Mach 6 tunnel with centerline pressure orifices on sharp leading edge, and (c) a perspective view of the model with stainless steel (top) and SLA middle insert (bottom) during the first entry. Flow occurs from left to right.

Bathel, Brett F.↗
NASA NTRS1994

Low to middle tropospheric profiles and biosphere/troposphere fluxes of acidic gases in the summertime Canadian taiga

We report features of acidic gases in the troposphere from 9 to 5000 m altitude above ground over the Canadian taiga in the summer of 1990. The measurements were conducted at a 30-m meteorological tower and from the NASA Wallops Electra aircraft as part of the joint U.S.-Canadian Arctic Boundary Layer Expedition (ABLE) 3B Northern Wetland Studies (NOWES). We sampled air for acidic gases using the mist chamber collector coupled with subsequent analysis using ion chromatography. At the tower we collected samples at two heights during a 13-day period, including diurnal studies. Using eddy flux and profile data, we estimated the biosphere/troposphere fluxes of nitric, formic, and acetic acids and sulfur dioxide. For the organic acids, emissions from the taiga in the afternoon hours and deposition during the predawn morning hours were observed. The flux intensities alone were however not high enough to explain the observed changes in mixing ratios. The measured deposition fluxes of nitric acid were high enough to have a significant influence on its mixing ratio in the boundary layer. On three days we measured vertical profiles of nitric, formic, and acetic acids through the lower to midtroposphere. We found that the chemical composition of the troposphere was extremely heterogenous. Pronounced layers of polluted air were readily apparent from our measurements. Local photochemical production and episodic long-range transport of trace components, originating from biomass burning and possibly industrial emissions, appear to have a strong influence on the composition of the troposphere and biosphere/troposphere fluxes of acidic gases at this site.

Klemm, O.↗
NASA NTRS1959

Effects of Nose Radius and Extreme Cooling on Boundary-Layer Transition for Three Smooth 15°-Cone-Cylinders in Free Flight at Mach Numbers to 8.50

Three highly polished 15°- included- angle cone- cylinders with hemispherical tips of several diameters ( 2, 3, and 4 in.) have been flown in order to obtain boundary- layer transition data at very low wall to local stream temperature ratios, and heat- transfer data. All surfaces had a 2-microinch average roughness height. Laminar flow existed over the entire hemispherical nose of the 2- and 3-inch-tip- diameter models throughout the complete flight history. Extreme cooling to wall to local stream temperature ratios at the sonic point as low as 0.20 did not cause transition on the nose for diameters as large as 3 inches. However, extreme cooling did cause early transition on the 4-inch model where it appears probable that transition occurred forward of the 45° station at a wall to local stream temperature ratio of about 0.26. Variations in tip diameter influenced transition downstream of the nose under conditions of extreme cooling. The 2-inch- tip model was laminar at all cone- cylinder stations at temperature ratios as low as 0.32 whereas the 3- and 4-inch-tip models were turbulent at the same local flow conditions but at higher wall to local temperature ratios. Transition on the cone and cylinder of the 3- and 4-inch- tip bodies appeared to be sensitive to local Mach number, and occurred at higher local temperature ratios when values of local Mach number were higher. Increasing the nose diameter from 2 to 3 inches significantly changed the local flow conditions for which laminar flow existed on the cone- cylinder afterbody. However, a further increase in tip size t o a 4-inch diameter had no discernable effect on the local flow conditions at transition. The transition results of the 3- and 4-inch-nose-diameter smooth bodies are similar to those observed on a 7/8-inch-nose-diameter body with roughened surfaces. Turbulent boundary layers resulted in both cases at very low wall to local stream temperature ratios. Both laminar and turbulent heat-transfer data were in good agreement with theoretical Stanton numbers when heat-transfer reduction due to tip blunting was considered.

CYLINDRICAL AFTERBODY↗
NASA NTRS1990

Boundary layer roll circulations during FIRE

The probable mechanism underlying the development of boundary layer roll circulations are studied using wind and temperature profiles measured by the National Center for Atmospheric Research (NCAR) Electra during the stratocumulus phase of the First ISCCP Regional Experiment (FIRE). The expected, or preferred, roll orientations, horizontal wavelengths, and propagation periods are determined by finding the minimum values of the dynamic and thermodynamic forcing parameters, which here are the eddy Reynolds number (Re) and moist Rayleigh number (Ra sub m). These minimum values depend on the height z sub T of the capping temperature inversion and on the values of the Fourier coefficients of the background height-dependent vector wind profile. As input to our nonlinear spectral model, descent and ascent runs by the Electra provide for initial estimates of the inversion height and the wind profiles. In the first phase of the investigation presented here, a mechanism is said to be a probable contributor to the development of roll circulations within the stratocumulus-topped boundary layer if the modeled roll orientation and wavelengths agree with their observed values. Preliminary results using the 14-coefficient model of Haack-Hirschberg (1988) are discussed for the 7 July 1987 Electra Mission 188-A (Flight 5). This mission was flown across a sharp cloud boundary that was within a LANDSAT/SPOT scene. The stratocumulus deck was relatively solid in the eastern part of the scene, while there was a rapid decrease in cloud cover to scattered cumulus clouds aligned in streets to the west. These cloud streets were oriented nearly parallel to the mean wind direction in the layer, which was approximately 340 degrees. The hypothesis that roll circulations occurred in both the relatively clear and the cloudy regions is investigated using as model input a descent profile obtained in the relatively clear air and an ascent profile obtained in the cloudy air. Initial results for the clear air case are that the pure inflection point mode is not possible and the pure thermal mode was oriented 35 degrees to the right of the mean wind direction. The origin of this unacceptably large discrepancy between the observed and modeled results will be investigated further and the conclusions reported at the next FIRE workshop.

Shirer, Hampton N.↗
NASA NTRS1994

The boundary layer of Mars: Fluxes, stability, turbulent spectra, and growth of the mixed layer

Spectra of wind from high-frequency measurements in the Martian atmospheric surface layer, along with the diurnal variation of the height of the mixed surface layer, are calculated for the first time for Mars. Heat and momentum fluxes, stability, and z(sub 0) are estimated for early spring from a surface temperature model and from Viking Lander 2 temperatures and winds at 44 deg N, using Monin-Obukhov similarity theory. Flow distortion by the lander is also taken into account. Model spectra for two measuring heights and three surface roughnesses are calculated using the depth of the mixed layer and the surface-layer parameters. These experiments indicate that z(sub 0) probably lies between 1.0 and 3.0 cm, and most likely is closer to 1.0 cm. The spectra are adjusted to simulate aliasing and high-frequency rolloff, the latter caused by both the sensor response and the large Kolmogorov length on Mars. Since the spectral models depend on the surface parameters, including the estimated surface temperature, their agreement with the calculated spectra indicates that the surface-layer estimates are self-consistent. This agreement is especially noteworthy in that the inertial subrange is virtually absent in the Martian atmosphere at this height, due to the large Kolmogorov length scale. These analyses extend the range of applicability of terrestrial results and demonstrate that it is possible to estimate the effects of severe aliasing of wind measurements, to produce models that agree well with the measured spectra. The results show that similarity theory developed for Earth applies to Mars, and that the spectral models are universal.

Tillman, James E.↗
NASA NTRS2012

Reynolds-Averaged Turbulence Model Assessment for a Highly Back-Pressured Isolator Flowfield

The use of computational fluid dynamics in scramjet engine component development is widespread in the existing literature. Unfortunately, the quantification of model-form uncertainties is rarely addressed with anything other than sensitivity studies, requiring that the computational results be intimately tied to and calibrated against existing test data. This practice must be replaced with a formal uncertainty quantification process for computational fluid dynamics to play an expanded role in the system design, development, and flight certification process. Due to ground test facility limitations, this expanded role is believed to be a requirement by some in the test and evaluation community if scramjet engines are to be given serious consideration as a viable propulsion device. An effort has been initiated at the NASA Langley Research Center to validate several turbulence closure models used for Reynolds-averaged simulations of scramjet isolator flows. The turbulence models considered were the Menter BSL, Menter SST, Wilcox 1998, Wilcox 2006, and the Gatski-Speziale explicit algebraic Reynolds stress models. The simulations were carried out using the VULCAN computational fluid dynamics package developed at the NASA Langley Research Center. A procedure to quantify the numerical errors was developed to account for discretization errors in the validation process. This procedure utilized the grid convergence index defined by Roache as a bounding estimate for the numerical error. The validation data was collected from a mechanically back-pressured constant area (1 2 inch) isolator model with an isolator entrance Mach number of 2.5. As expected, the model-form uncertainty was substantial for the shock-dominated, massively separated flowfield within the isolator as evidenced by a 6 duct height variation in shock train length depending on the turbulence model employed. Generally speaking, the turbulence models that did not include an explicit stress limiter more closely matched the measured surface pressures. This observation is somewhat surprising, given that stress-limiting models have generally been developed to better predict shock-separated flows. All of the models considered also failed to properly predict the shape and extent of the separated flow region caused by the shock boundary layer interactions. However, the best performing models were able to predict the isolator shock train length (an important metric for isolator operability margin) to within 1 isolator duct height.

Baurle, Robert A.↗
NASA NTRS2018

A Global Analysis of Wildfire Smoke Injection Heights Derived from Space-Based Multi-Angle Imaging

We present an analysis of over 23,000 globally distributed wildfire smoke plume injection heights derived from Multi-angle Imaging SpectroRadiometer (MISR) space-based, multi-angle stereo imaging. Both pixel-weighted and aerosol optical depth (AOD)-weighted results are given, stratified by region, biome, and month or season. This offers an observational resource for assessing first-principle plume-rise modelling, and can provide some constraints on smoke dispersion modelling for climate and air quality applications. The main limitation is that the satellite is in a sun-synchronous orbit, crossing the equator at about 10:30 a.m. local time on the day side. Overall, plumes occur preferentially during the northern mid-latitude burning season, and the vast majority inject smoke near-surface. However, the heavily forested regions of North and South America, and Africa produce the most frequent elevated plumes and the highest AOD values; some smoke is injected to altitudes well above 2 km in nearly all regions and biomes. Planetary boundary layer (PBL) versus free troposphere injection is a critical factor affecting smoke dispersion and environmental impact, and is affected by both the smoke injection height and the PBL height; an example assessment is made here, but constraining the PBL height for this application warrants further work.

Val Martin, Maria↗
DOE OSTI2025

Savannah River Site Climatology for Research Activities Conducted Onsite and in P-Area (1994-2023)

This report presents a climatology of meteorological variables at the Savannah River Site (SRS) using measurements from the P-Area and Central Climatology meteorological towers to support research activities at the Savannah River Site. Key values included in this report are wind, surface pressure, temperature, relative humidity, soil moisture, cloud fraction, cloud base, lightning, atmospheric stability, and boundary layer parameters. The Central Climatology tower is useful in that it is instrumented at four levels, allowing for an examination of height dependence on some variables.

54 ENVIRONMENTAL SCIENCES↗
DOE OSTI2025

Savannah River Site Climatology for Research Activities Conducted Onsite and in P-Area (1994-2023)

Executive Summary This report presents a climatology of meteorological variables at the Savannah River Site (SRS) using measurements from the P-Area and Central Climatology meteorological towers to support research activities at the Savannah River Site. Key values included in this report are wind, surface pressure, temperature, relative humidity, soil moisture, cloud fraction, cloud base, lightning, atmospheric stability, and boundary layer parameters. The Central Climatology tower is useful in that it is instrumented at four levels, allowing for an examination of height dependence on some variables.

54 ENVIRONMENTAL SCIENCES↗
NASA NTRS1993

Synoptic/planetary-scale interactions and blocking over the North Atlantic Ocean

The central theme of this project has been the diagnosis of blocking anticyclogenesis and the corresponding interactions with synoptic-scale circulations. To that end an extensive investigation of the dynamics and energetics of a major blocking anticyclone and two upstream cyclones, all of which occurred over the North Atlantic Ocean and the United States in January 1979, was undertaken. Data for the study were provided by Goddard Laboratory for Atmospheres (GLA) 4 LAT by 5 LON FGGE analyses. The methodology has primarily focused on the diagnosis of circulation forcing mechanisms using the modified forms (referred to as the extended forms) of the height tendency and Zwack-Okossi equations developed by our research group. Calculations use routine second-order finite differencing with boundary layer fraction and sensible heating and latent heat release represented as parameterized quantities. Of particular interest are the latent heat release estimates, which combine convectional parameterized values with estimates derived from satellite IR data. The latter were obtained using an algorithm derived by Dr. Franklin R. Robertson of NASA's Marshall Space Flight Center. Results are contained in project reports, theses and publications identified in previous review summaries and reports, and publications listed at the end of this summary. Significant accomplishments in the past year are presented.

Smith, Phillip J.↗
NASA NTRS2016

Nonlinear Wakes Behind a Row of Elongated Roughness Elements

This paper is concerned with the high Reynolds number flow over a spanwise periodic array of roughness elements with inter-element spacing of the order of the local boundary-layer thickness. While earlier work by Goldstein, Sescu, Duck and Choudhari (2010) and Goldstein, Sescu, Duck and Choudhari (2011) was mainly concerned with smaller roughness heights that produced relatively weak distortions of the downstream flow, the focus here is on extending the analysis to larger roughness heights and streamwise elongated planform shapes that together produce a qualitatively different, nonlinear behavior of the downstream wakes. The roughness scale flow now has a novel triple-deck structure that is somewhat different from related studies that have previously appeared in the literature. The resulting flow is formally nonlinear in the intermediate wake region, where the streamwise distance is large compared to the roughness dimensions but small compared to the downstream distance from the leading edge, as well as in the far wake region where the streamwise length scale is of the order of the downstream distance from the leading edge. In contrast, the flow perturbations in both of these wake regions were strictly linear in the earlier work by Goldstein et al (2010, 2011). This is an important difference because the nonlinear wake flow in the present case provides an appropriate basic state for studying the secondary instability and eventual breakdown into turbulence.

Boundary layer control↗
NASA NTRSDate not supplied

Stability Analysis of the Flow Over a Swept Forward-Facing Step Using PIV Base Flows

Step excrescences are a type of surface imperfection encountered on swept wings of commercial aircraft, commonly because accessibility requirements prevent creating the wing’s surface from a single panel. If the step height is too large, super-critical, the flow can undergo an early transition to turbulence, which can be highly detrimental for the performance of the wing. In the case of a forward-facing step on a swept wing in a low-disturbance environment, stationary crossflow vortices can develop a significant amplitude upstream of the step and hence dominate the structure of the boundary-layer flow over the step. The main goal of the present investigation is to illuminate the path to transition supported by the fascinatingly complex flow field in the direct downstream vicinity of a step with a super critical height. The high-resolution, stereographic Particle Image Velocimetry (PIV) measurement dataset presently available for this flow field provides a complete description of the laminar flow for the execution of BiGlobal stability analysis in a plane parallel to the step. Although the notorious sensitivity of stability results to the description of the base flow demands a very careful uncertainty analysis of those results, it is argued that this very fact can be leveraged to produce new insight into the supported perturbation dynamics. In performing the analysis, several unsteady mode families are discovered that display the explosive perturbation expected for early transition to be induced. In considering domain widths equal to an integer-multiple of the incident crossflow-vortex wavelength and analyzing an extent of 5 crossflow-vortex wavelengths parallel to the step, it is found that the stability results converge while increasing the domain width. It is demonstrated, moreover, that the results for the wider domains can be approximated by appropriately averaging the results on neighboring single-crossflow-vortex-wavelength domains covering the same region. Besides being useful for computational purposes, this observed property suggests interpreting the instability mechanism as a distorted primary mechanism rather than a “proper” secondary mechanism. This follows in the context of the secondary in-stability analysis of three-dimensional boundary layers, because the secondary mechanism is usually characterized by being localized in a pocket of strong shear, while the distorted primary mechanism typically has an infinite support in the direction parallel to the step. Even though the growth rates are found to be sensitive to the interrogation-window size inherent to the PIV post-processing procedure, the spatial structure of the eigenfunctions is found to be relatively insensitive. Lastly, the spatial structure of the eigenfunctions corresponding to all velocity components are matched with the shape functions determined by computing the Spectral Proper Orthogonal Decomposition (SPOD) of a time-resolved measurement of the perturbation content.

forward-facing step↗
NASA NTRS2008

Boundary Layer Receptivity due to Roughness and Freestream Sound for Supersonic Flows over Axisymmetric Cones

The receptivity of supersonic boundary layers due to the interaction of a two-dimensional acoustic wave with a three-dimensional roughness is numerically investigated over a 5-degree sharp tipped straight cone at a free stream Mach number of 3.5 and a high Reynolds number of 10(exp 6)/inch. Both the steady and unsteady solutions are obtained by solving the full Navier-Stokes equations using the fifth-order accurate weighted essentially non-oscillatory (WENO) scheme for spatial discretization and using third-order total-variation-diminishing (TVD) Runge-Kutta scheme for temporal integration. The simulations are performed with different roughness heights and roughness locations. The steady computations with the roughness reveal that the perturbations induced by the roughness remain very close to the wall and do not penetrate to the outer part of the boundary layer. The effects of the roughness persist for about 10 boundary layer thicknesses downstream of each element. The unsteady simulation shows the generation of unstable three-dimensional first mode disturbances within a short distance from the roughness. However, the amplitudes of the instability waves are very small near the neutral point, on the order of 10(exp -4) compared to the free stream disturbances. The receptivity coefficients due to the roughness elements located very close to the tip are one order of magnitude smaller than those due to the roughness elements located closer to the neutral point.

Balakumar, P.↗
NASA NTRS2023

Experimental Investigation of Block-TPS Fence/Gap Roughness Effects on Transition Onset and Turbulent Heating at Mach 6

An experimental investigation of boundary-layer transition and convective surface heating on wind tunnel models that simulated aeroshells comprised of multiple blocks of thermal protection system (TPS) material has been performed. Two representative entry vehicle geometries, a sphere-cone aeroshell and a sphere-cone aeroshell, were considered. Multiple cast ceramic models of each geometry were fabricated with varying fence heights or gap depths to represent gap fillers that either protrude above, or are recessed below, the TPS blocks due to differential ablation during reentry. Wind tunnel testing was performed at Mach 6 over a range of Reynolds numbers sufficient to produce laminar, transitional, and turbulent flow. Convective surface heating and boundary-layer transition onset data were obtained using global phosphor thermography. The experimental heating data are presented herein, as are comparisons to laminar and turbulent smooth-wall heat transfer distributions from computational flow field simulations.

Aerothermodynamics↗