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NASA NTRS2015

Urbanization Causes Increased Cloud Base Height and Decreased Fog in Coastal Southern California

Subtropical marine stratus clouds regulate coastal and global climate, but future trends in these clouds are uncertain. In coastal Southern California (CSCA), interannual variations in summer stratus cloud occurrence are spatially coherent across 24 airfields and dictated by positive relationships with stability above the marine boundary layer (MBL) and MBL height. Trends, however, have been spatially variable since records began in the mid-1900s due to differences in nighttime warming. Among CSCA airfields, differences in nighttime warming, but not daytime warming, are strongly and positively related to fraction of nearby urban cover, consistent with an urban heat island effect. Nighttime warming raises the near-surface dew point depression, which lifts the altitude of condensation and cloud base height, thereby reducing fog frequency. Continued urban warming, rising cloud base heights, and associated effects on energy and water balance would profoundly impact ecological and human systems in highly populated and ecologically diverse CSCA.

coasts↗

NASA NTRS2017

Computational Design and Analysis of a Transonic Natural Laminar Flow Wing for a Wind Tunnel Model

A natural laminar flow (NLF) wind tunnel model has been designed and analyzed for a wind tunnel test in the National Transonic Facility (NTF) at the NASA Langley Research Center. The NLF design method is built into the CDISC design module and uses a Navier-Stokes flow solver, a boundary layer profile solver, and stability analysis and transition prediction software. The NLF design method alters the pressure distribution to support laminar flow on the upper surface of wings with high sweep and flight Reynolds numbers. The method addresses transition due to attachment line contamination/transition, Gortler vortices, and crossflow and Tollmien-Schlichting modal instabilities. The design method is applied to the wing of the Common Research Model (CRM) at transonic flight conditions. Computational analysis predicts significant extents of laminar flow on the wing upper surface, which results in drag savings. A 5.2 percent scale semispan model of the CRM NLF wing will be built and tested in the NTF. This test will aim to validate the NLF design method, as well as characterize the laminar flow testing capabilities in the wind tunnel facility.

Lynde, Michelle N.↗

NASA NTRS2021

Assessment and Improvement of RANS-based Transition Models based on Experimental Data of the Common Research Model with Natural Laminar Flow

Transition models based on auxiliary transport equations augmenting the Reynolds-averaged Navier-Stokes (RANS) framework often rely upon the correlations that were derived from a limited number of low-speed experiments and do not account for all of the transition mechanisms and/or their variation with the significant flow parameters. Available data from a recent experiment in the National Transonic Facility at the NASA Langley Research Center are used to assess the current transition modeling capability in NASA's OVERFLOW 2.2o code for a swept wing configuration at transonic cruise conditions. Specifically, the OVERFLOW solutions are used together with detailed stability analysis of the boundary layer flow over the new Common Research Model with Natural Laminar Flow (CRM-NLF) to evaluate the accuracy and the robustness of the transport-equation-based transition models, with the goal of proposing improvements that would help to strengthen the physical basis of these models for the important class of flows involving the combined effects of crossflow and flow compressibility. Results highlight the significant underprediction of the laminar flow extent within the inboard region of the wing, wherein the onset of transition may be attributed to a gradual amplification of Tollmien-Schlichting instabilities.

Boundary layer transition↗

NASA NTRS2022

Defining the upper boundary of the Asian Tropopause Aerosol Layer (ATAL) using the static stability

The Asian Tropopause Aerosol Layer (ATAL) is located in the Upper Troposphere and Lower Stratosphere (UTLS) during the Asian Summer Monsoon. However, what dynamical feature separates the ATAL from the well-known stratospheric ‘Junge layer’ is not yet clear. In this study, using the in-situ (Radiosonde, Ozonesonde, backscatter sonde and cryogenic frost-point hygrometer) observations from multiple locations in India (Gadanki (13.45° N, 79.18° E), Hyderabad (17.47° N, 78.58° E) and Varanasi (25.27° N, 82.99° E)) and multi-satellite observations ((Cloud-Aerosol Lidar and Infrared Pathfinder Observation, (CALIPSO), Atmospheric Chemistry Experiment (ACE) Fourier Transform Spectrometer (FTS) and Constellation Observation System for Meteorology, Ionosphere and Climate (COSMIC) Global Position System (GPS) Radio Occultation (RO) (COSMIC GPS-RO)) we show that the ATAL can exist up to the layer of maximum stability (LmaxS), located a few kilometers above the tropopause, determined using the square of Brunt Väisäla frequency. These in-situ observations over Indian stations collected during the ISRO-NASA Balloon Measurement Campaigns of the Asian Tropopause Aerosol Layer (BATAL) show that the ATAL top can reach up to ∼442 K potential temperature level over the Indian region. The LmaxS delineated from COSMIC GPSRO observations over the Asian Summer Monsoon Anticyclone (ASMA) region indicates that the top of ATAL can reach up to 454 K potential temperature level, which is lower than the earlier Lagrangian transport model predicted 460 K. The temperature inversion at LmaxS acts as a lid and constrains the direct transport of aerosols to higher altitudes.

S.T. Akhil Raj↗

NASA NTRS1987

Stationary disturbances in three-dimensional boundary layers over concave surfaces

A two-dimensional boundary layer on a concave surface is known to be susceptible to centrifugal instability which manifests itself in the form of stationary streamwise counter-rotating vortices commonly known as Goertler vortices. In this paper, the problem of the stability of a three-dimensional boundary layer on a concave surface is considered. Linear stability equations, including streamline and surface curvature effects, are solved for an infinitely swept wing. The results indicate that the Goetler vortex structure begins to transform into co-rotating vortices when the crossflow Reynolds number is increased. This transition is almost complete when the crossflow Reynolds number is in excess of 45. It is shown that the centrifugal effects destabilize the crossflow vortices. Some comparisons with available experimental results are made.

Collier, F. S., Jr.↗

NASA NTRS1990

Design of laminar flow bodies in compressible flow

An optimization method was developed to design axisymmetric body shapes such as fuselages, nacelles, and external fuel tanks with increased transition Reynolds numbers in subsonic compressible flow. The method involves a constraint minimization procedure coupled with analysis of the inviscid and viscous flow regions, and linear stability analysis of the compressible boundary-layer. Boundary-layer transition is predicted by a 'hybrid' transition criterion based on Granville's transition criterion and a criterion using linear stability theory coupled with the e(sup n)-method. A tiptank of a business jet is used as an example to illustrate that the method can be utilized to design an axisymmetric body shape with extensive natural laminar flow. On the original tiptank boundary layer transition is predicted to occur at a transition Reynolds number of 6.04 x 10 6 on the original tiptank. On the designed body shape a transition Reynolds number of 7.22 x 10 6 is predicted using compressible linear stability theory coupled with e n -method.

Simha S. Dodbele↗

NASA NTRS1990

Viscous drag reduction in boundary layers

The present volume discusses the development status of stability theory for laminar flow control design, applied aspects of laminar-flow technology, transition delays using compliant walls, the application of CFD to skin friction drag-reduction, active-wave control of boundary-layer transitions, and such passive turbulent-drag reduction methods as outer-layer manipulators and complex-curvature concepts. Also treated are such active turbulent drag-reduction technique applications as those pertinent to MHD flow drag reduction, as well as drag reduction in liquid boundary layers by gas injection, drag reduction by means of polymers and surfactants, drag reduction by particle addition, viscous drag reduction via surface mass injection, and interactive wall-turbulence control.

Bushnell, Dennis M.↗

NASA NTRSDate not supplied

Eigenvalue Sensitivity Computations for Linear Stability Theory

To realize the drag reduction benefit of boundary-layer transition control strategies, it is crucial to integrate transition prediction into the vehicle design through an optimization process. The integration of transition prediction based on linear stability analysis into adjoint d design optimization requires coupling an adjoint enabled computational fluid dynamics (CFD) solver with an adjoint enabled linear stability code. In particular, the boundary-layer transition location is often predicted using the N-factor method based on linear stability theory (LST). Thus, sensitivity of the linear-stability eigenvalues constitute an essential building block for optimizing the laminar flow performance. The present paper describes an implementation of LST eigenvalue sensitivity analysis that can be easily coupled with a CFD solver. Specifically, we describe a discrete adjoint formulation for the transition location prediction based on the N-factor method. The verification of this formulation is carried out by comparing the adjoint-based sensitivity of the local growth rate of a given instability mode with respect to the disturbance frequency, and the adjoint-based sensitivity of the transition location with respect to spanwise wavenumber with those sensitivities computed using a finite-difference approximation. Finally, the adjoint LST formulation is applied to flat-plate boundary-layer flows at transonic, supersonic, and hypersonic conditions, to determine the behavior and sensitivities of the transition location with respect to a range of disturbance spanwise wavenumbers.

Boundary Layer Transition↗

NASA NTRS1990

Modulational stability of rotating-disk flow

The problem of modulational primary stability of rotating-disk boundary layer flow for stationary modes was studied by using the method of multiple scales. The system of equations for infinitesimal disturbances is solved by linear stability theory to determine the modulated solution which represents the time and azimuthal variations of the wave pattern. The solution for the amplitude of the most critical disturbance indicates that the disturbance spreads out as the flow time increases and grows exponentially at the center.

Riahi, Daniel N.↗

NASA NTRSDate not supplied

Laminar to Turbulence Transition in Boundary Layers due to Tripping Devices

Direct numerical simulations (DNS)and linear stability analysis were performed to identify the mechanisms by which two-and three-dimensional trips promote early transition in subsonic boundary layers over a flat plate. Investigations were conducted for two boundary layer transition trip configurations: one is a two-dimensional rectangular-shaped rod and the other is a three-dimensional zigzag tape. Reynolds numbers based on the trip heights are 683 and 300 for the two-and three-dimensional trips, respectively. Simulations with the two-dimensional trip showed that the flow separates at the top of the trip and forms a long recirculation zone downstream. However, the flow remains steady in the simulation without any external disturbances. Linear stability analysis showed that the inflectional profile in the recirculation zone is strongly unstable to frequencies in the range of 100-300 Hz. Simulations with the external disturbances forced at the frequency of 200 Hz showed that disturbances grow strongly inside the recirculation zone as observed in the Klebanoff1 experiment.

Transition↗

NASA NTRS2009

Stability Analysis of Roughness Array Wake in a High-Speed Boundary Layer

Computations are performed to examine the effects of both an isolated and spanwise periodic array of trip elements on a high-speed laminar boundary layer, so as to identify the potential physical mechanisms underlying an earlier transition to turbulence as a result of the trip(s). In the context of a 0.333 scale model of the Hyper-X forebody configuration, the time accurate solution for an array of ramp shaped trips asymptotes to a stationary field at large times, indicating the likely absence of a strong absolute instability in the mildly separated flow due to the trips. A prominent feature of the wake flow behind the trip array corresponds to streamwise streaks that are further amplified in passing through the compression corner. Stability analysis of the streaks using a spatial, 2D eigenvalue approach reveals the potential for a strong convective instability that might explain the earlier onset of turbulence within the array wake. The dominant modes of streak instability are primarily sustained by the spanwise gradients associated with the streaks and lead to integrated logarithmic amplification factors (N factors) approaching 7 over the first ramp of the scaled Hyper-X forebody, and substantially higher over the second ramp. Additional computations are presented to shed further light on the effects of both trip geometry and the presence of a compression corner on the evolution of the streaks.

Choudhari, Meelan↗

NASA NTRS2001

Transition In Supersonic Flows With Corners

It is proposed to investigate the effect of sudden change in the slope of the surface on the stability and transition onset point in supersonic and hypersonic boundary layers. Since we can not use the linear stability theory in flows where there are discontinuities we have to solve the full Navier-Stokes equations to determine the amplification of the disturbances across the oblique shock which exists at the corner. We will use the Navier-Stokes code which was developed by Dr. Harold Atkins in the aerodynamic and acoustic methods branch. The code is developed using the higher order ENO scheme and is currently set up for two-dimensional flows. In the first year of our investigation, we will investigate the evolution of small amplitude two-dimensional disturbances and determine the N-factors for the most amplified disturbances. We will perform the analysis for the parameter similar to X-31 model and flight experiment conditions. From these computation we will infer what is the maximum amplification rate possible in quite environments without any tripping devices. In the second part of the investigation, we will perform nonlinear computations to determine what is the minimum amplitudes necessary to cause the transition at a designed location. In the third part, we have to determine what kind of trip devices necessary to excite the required disturbances. The principal investigator of this project is Dr. P. Balakumar, Associate Professor of Department of Aerospace Engineering, Old Dominion University. We will actively collaborate with research scientists in the Aero-Thermodynamics Branch at NASA/Langley Research Center. A half-time graduate student will also participate in the project.

Balakumar, P.↗

NASA NTRS1978

An interaction solution algorithm for viscous aerodynamic performance prediction

A weak-interaction solution algorithm is established for aerodynamic flow field prediction about an isolated airfoil. It requires numerical solution of differential equations governing potential flow, viscous and turbulent boundary layer flow, and the turbulent wake flow downstream of a trailing edge. The algorithm accounts for computed viscous displacement effects on the potential flow. These in turn alter the viscous flow through imposed pressure gradients. Closure for turbulence is accomplished using a second order model. Numerical evaluations assess factors affecting solution accuracy, convergence and stability for the combined potential, boundary layer, and parabolic Navier-Stokes equation systems as solved using a finite element algorithm.

Baker, A. J.↗

NASA NTRS1998

Numerical Simulations of Unsteady Low-Reynolds-Number Flows Over the APEX Airfoil

Laminar and transitional separation bubbles are an important feature of low-Reynolds-number flows over airfoils. The separation bubbles are unsteady and have a significant impact on the aerodynamic properties of the airfoils. In this paper unsteady low-Reynolds-number separated flows over the APEX airfoil are calculated using a Navier-Stokes solver. The numerical results show the presence of unsteady separation bubbles in the flowfield. An analysis of the numerical results shows that flowfield disturbances are amplified significantly in the separation bubble, leading to periodic vortex shedding. A linear stability analysis of the separated boundary layer is performed and the results show that the dominant wavenumber and frequency in the numerical simulations agree with the most unstable wavenumber and frequency from the linear stability analysis. The numerical results also show the growth and interaction of disturbance waves in the separation bubble. For transonic flows over the APEX airfoil, the calculations show that the presence of shocks causes significant changes in the separation location and consequently, the overall flowfield.

Tatineni, Mahidhar↗

NASA NTRS2022

Stability Analysis of Streaks Induced by Optimized Vortex Generators

Numerical computations are performed to investigate the potential for transition control in an axisymmetric boundary layer via fully realizable, streamwise stationary streaks induced by an azimuthally periodic array of surface mounted vortex generators (VGs). Previous work has shown that suitable streaks of this type can significantly reduce the growth of Mack’s second mode instabilities, but large streak amplitudes can make the flow susceptible to previously absent streak instabilities that can become the leading cause of transition. Here, we use the adjoint capabilities of the SU2 flow solver to optimize the VG shape to maximize the reduction in the growth of second-mode disturbances while also preventing the streak amplitudes from reaching large enough values to precipitate an earlier onset of transition via streak instabilities. The geometry and the freestream flow conditions are selected to match a relevant trajectory lo-cation from the HIFiRE-1 flight experiment. Results show that the optimized VGs can increase the mean streak amplitude by 117% with respect to a manually developed baseline design. The stability of this optimized basic state is analyzed via the plane-marching parabolized stability equations, predicting a fully laminar flow over the entire cone, or equivalently, yielding transition delay of 130% versus the 17% for the baseline VGs.

Boundary layer transition↗

NASA NTRS2016

Measurements in a Transitioning Cone Boundary Layer at Freestream Mach 3.5

An experimental study was conducted in the Supersonic Low-Disturbance Tunnel to investigate naturally-occurring instabilities in a supersonic boundary layer on a 7 deg half- angle cone. All tests were conducted with a nominal freestream Mach number of M(sub infinity) = 3:5, total temperature of T(sub 0) = 299:8 K, and unit Reynolds numbers of Re(sub infinity) x 10(exp -6) = 9:89, 13.85, 21.77, and 25.73 m(exp -1). Instability measurements were acquired under noisy- ow and quiet- ow conditions. Measurements were made to document the freestream and the boundary-layer edge environment, to document the cone baseline flow, and to establish the stability characteristics of the transitioning flow. Pitot pressure and hot-wire boundary- layer measurements were obtained using a model-integrated traverse system. All hot- wire results were single-point measurements and were acquired with a sensor calibrated to mass ux. For the noisy-flow conditions, excellent agreement for the growth rates and mode shapes was achieved between the measured results and linear stability theory (LST). The corresponding N factor at transition from LST is N 3:9. The stability measurements for the quiet-flow conditions were limited to the aft end of the cone. The most unstable first-mode instabilities as predicted by LST were successfully measured, but this unstable first mode was not the dominant instability measured in the boundary layer. Instead, the dominant instabilities were found to be the less-amplified, low-frequency disturbances predicted by linear stability theory, and these instabilities grew according to linear theory. These low-frequency unstable disturbances were initiated by freestream acoustic disturbances through a receptivity process that is believed to occur near the branch I locations of the cone. Under quiet-flow conditions, the boundary layer remained laminar up to the last measurement station for the largest Re1, implying a transition N factor of N greater than 8:5.

King, Rudolph A.↗

NASA NTRS1984

Heat and mass transfer in combustion - Fundamental concepts and analytical techniques

Fundamental combustion phenomena and the associated flame structures in laminar gaseous flows are discussed on physical bases within the framework of the three nondimensional parameters of interest to heat and mass transfer in chemically-reacting flows, namely the Damkoehler number, the Lewis number, and the Arrhenius number which is the ratio of the reaction activation energy to the characteristic thermal energy. The model problems selected for illustration are droplet combustion, boundary layer combustion, and the propagation, flammability, and stability of premixed flames. Fundamental concepts discussed include the flame structures for large activation energy reactions, S-curve interpretation of the ignition and extinctin states, reaction-induced local-similarity and non-similarity in boundary layer flows, the origin and removal of the cold boundary difficulty in modeling flame propagation, and effects of flame stretch and preferential diffusion on flame extinction and stability. Analytical techniques introduced include the Shvab-Zeldovich formulation, the local Shvab-Zeldovich formulation, flame-sheet approximation and the associated jump formulation, and large activation energy matched asymptotic analysis. Potentially promising research areas are suggested.

Law, C. K.↗

NASA NTRS2019

Effect of 3D Roughness Patch on Instability Amplification in a Supersonic Boundary Layer

Surface roughness is known to have a substantial impact on the aerothermodynamic loading of high-speed vehicles, particularly via its influence on the laminar-turbulent transition process within the boundary layer. Numerical simulations are performed to investigate the effects of a distributed region of densely packed, sinusoidal shape roughness elements on a Mach 3.5 flat plate boundary layer for flow conditions corresponding to the planned conditions of an upcoming experiment in the Mach 3.5 Supersonic Low Disturbance Tunnel at the NASA Langley Research Center. Analysis of convective instabilities in the wake of the roughness patch was reported in a previous paper and the current work extends that analysis to instability amplification across the length of the roughness patch. Quasi-parallel stability analysis of the modified boundary layer flow over the patch indicates two dominant families of unstable disturbances, namely, a group of high frequency modes that amplify in localized regions along the roughness patch and another group of low frequency modes that have smaller peak amplification rates but amplify steadily both above the roughness patch and in the wake region behind it. The results suggest that the amplification factors associated with the high-frequency modes are sufficiently low, at least for the roughness patches considered in this paper, so that these modes are unlikely to have a major influence on the transition process. The amplification of the low-frequency modes within the region of the roughness patch is further quantified via direct numerical simulations. Results confirm the strongly destabilizing influence of the roughness patch on the first mode instabilities, yielding an N-factor increment of ΔN ≈ 3.6 for a roughness patch length of eight wavelengths.

Laminar turbulent transition↗