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At least 1,513 records · Page 84
All sourcesNASA NTRSDOE OSTI
NASA NTRS1986

Unsteady separated boundary layer in a transonic diffuser flow with self-excited oscillations

A numerical investigation of two-dimensional unsteady boundary layer in a transonic diffuser flow with self-excited oscillations and strong flow separation by solving the compressible, Reynolds-averaged, thin-layer Navier-Stokes equations with two-equations turbulence model is described. Three different meshes with constant streamwise mesh distribution and varying vertical mesh distribution were used. Results obtained indicate that a refinement of mesh studied here has minimal effect on the mean boundary layer flow but significantly increases the amplitude of oscillation of all flow variables. Comparisons of unsteady wall pressure, velocity profile, terminal shock, and separation pocket among computations and with experiment are presented.

Hsieh, T.↗
NASA NTRS2010

Analysis of Nozzle Jet Plume Effects on Sonic Boom Signature

An axisymmetric full Navier-Stokes computational fluid dynamics (CFD) study was conducted to examine nozzle exhaust jet plume effects on the sonic boom signature of a supersonic aircraft. A simplified axisymmetric nozzle geometry, representative of the nozzle on the NASA Dryden NF-15B Lift and Nozzle Change Effects on Tail Shock (LaNCETS) research airplane, was considered. The highly underexpanded nozzle flow is found to provide significantly more reduction in the tail shock strength in the sonic boom N-wave pressure signature than perfectly expanded and overexpanded nozzle flows. A tail shock train in the sonic boom signature, similar to what was observed in the LaNCETS flight data, is observed for the highly underexpanded nozzle flow. The CFD results provide a detailed description of the nozzle flow physics involved in the LaNCETS nozzle at different nozzle expansion conditions and help in interpreting LaNCETS flight data as well as in the eventual CFD analysis of a full LaNCETS aircraft. The current study also provided important information on proper modeling of the LaNCETS aircraft nozzle. The primary objective of the current CFD research effort was to support the LaNCETS flight research data analysis effort by studying the detailed nozzle exhaust jet plume s imperfect expansion effects on the sonic boom signature of a supersonic aircraft. Figure 1 illustrates the primary flow physics present in the interaction between the exhaust jet plume shock and the sonic boom coming off of an axisymmetric body in supersonic flight. The steeper tail shock from highly expanded jet plume reduces the dip of the sonic boom N-wave signature. A structured finite-volume compressible full Navier-Stokes CFD code was used in the current study. This approach is not limited by the simplifying assumptions inherent in previous sonic boom analysis efforts. Also, this study was the first known jet plume sonic boom CFD study in which the full viscous nozzle flow field was modeled, without coupling to a sonic boom propagation analysis code, from the stagnation chamber of the nozzle to the far field external flow, taking into account all nonisentropic effects in the shocks, boundary layers, and free shear layers, and their interactions at distances up to 30 times the nozzle exit diameter from the jet centerline. A CFD solution is shown in Figure 2. The flow field is very complicated and multi-dimensional, with shock-shock and shockplume interactions. At the time of this reporting, a full three-dimensional CFD study was being conducted to evaluate the effects of nozzle vectoring on the aircraft tail shock strength.

Bui, Trong↗
NASA NTRS1994

Optimized design of a hypersonic nozzle

Conventional procedures for designing nozzles involve the design of an inviscid contour (using the method of characteristics) that is corrected with a displacement thickness calculated from boundary-layer theory. However, nozzles designed using this classical procedure have been shown to exhibit poor flow quality at Mach numbers characteristic of hypersonic applications. The nozzle to be designed will be a part of the NASA HYPULSE facility which is being used for hypervelocity flight research. Thus, the flow quality of the nozzle is a critical question that needs to be addressed. Design of nozzles for hypersonic applications requires a proper assessment of the effects of the thick boundary layer on the inviscid flowfield. Since the flow field is largely supersonic, the parabolized form of the Navier-Stokes (PNS) equations can be used. The requirement of a uniform flow at the exit plane of the nozzle can be used to define an objective function as part of an optimization procedure. The design procedure used in this study involves the coupling of a nonlinear (least-squares) optimization algorithm with an efficient, explicit PNS solver. The thick boundary layers growing on the walls of the nozzle limit the extent of the usable core region (region with uniform flow) for testing models (especially rectangular). In order to maximize the region of uniform flow, it was decided to have the exit plane of this nozzle to be (nearly) rectangular. Thus, an additional constraint on the nozzle shape resulted, namely the nozzle will have a shape transitioning from a circular one at the inlet to that of a rectangle at the exit. In order to provide for a smooth shape transition, the cross sectional contour of the nozzle is defined by a superellipse. The nozzle is taken to be a meter in length. The axial variations of the major and minor radii of the superellipse are governed by cubic splines. The design parameters are the coefficients of the splines associated with the local nozzle wall slopes. Extensive calculations have been made (with a three-dimensional Euler code) to understand the effects of various parameters such as location of the knot points of the spline function, different ways of characterizing the uniformity of the flow in the exit plane, as well as the effect of constraining the area of the nozzle to be invariant. Turbulent flow (measurements indicate that the flow at the nozzle inlet is turbulent) calculations are now being performed (with the inviscidly designed nozzle contours) to assess the flow quality.

Krishnamurthy, Ramesh↗
DOE OSTI2025

An infrared on-shell action and its implications for soft charge fluctuations in asymptotically flat spacetimes

We study the infrared on-shell action of Einstein gravity in asymptotically flat spacetimes (AFSs), obtaining an effective, gauge-invariant boundary action for memory and shockwave spacetimes. We show that the phase space is in both cases parameterized by the leading soft variables in AFSs, thereby extending the equivalence between shockwave and soft commutators to spacetimes with non-vanishing Bondi mass. We then demonstrate that our on-shell action is equal to three quantities studied separately in the literature: (i) the soft supertranslation charge; (ii) the shockwave effective action, or equivalently the modular Hamiltonian; and (iii) the soft effective action. Finally, we compute the quantum fluctuations in the soft supertranslation charge and, assuming the supertranslation parameter may be promoted to an operator, we obtain an area law, consistent with earlier results showing that the modular Hamiltonian has such fluctuations.

asymptotically flat spacetimes↗
NASA NTRS1987

Prediction and control of transition in hypersonic boundary layers

In this paper, the role of compressible linear stability theory in prediction of boundary layer transition at supersonic and hypersonic speeds is investigated. Computations for sharp cones, using the e exp N method with N = 10, show that the first oblique Tollmien-Schlichting mode is responsible for transition at adiabatic wall conditions for freestream Mach numbers up to 7. For cold walls, the two-dimensional second mode dominates the transition process at lower hypersonic Mach numbers due to the well-known destabilizing effect of cooling on the second mode. It is shown that pressure gradient and suction may be used to stabilize this mode. Some results on the real gas effects on hypersonic boundary-layer stability are presented.

Malik, Mujeeb R.↗
NASA NTRS2009

Numerical Study of Outlet Boundary Conditions for Unsteady Turbulent Internal Flows Using the NCC

This paper presents the results of studies on the outlet boundary conditions for turbulent internal flow simulations. Several outlet boundary conditions have been investigated by applying the National Combustion Code (NCC) to the configuration of a LM6000 single injector flame tube. First of all, very large eddy simulations (VLES) have been performed using the partially resolved numerical simulation (PRNS) approach, in which both the nonlinear and linear dynamic subscale models were employed. Secondly, unsteady Reynolds averaged Navier- Stokes (URANS) simulations have also been performed for the same configuration to investigate the effects of different outlet boundary conditions in the context of URANS. Thirdly, the possible role of the initial condition is inspected by using three different initial flow fields for both the PRNS/VLES simulation and the URANS simulation. The same grid is used for all the simulations and the number of mesh element is about 0.5 million. The main purpose of this study is to examine the long-time behavior of the solution as determined by the imposed outlet boundary conditions. For a particular simulation to be considered as successful under the given initial and boundary conditions, the solution must be sustainable in a physically meaningful manner over a sufficiently long period of time. The commonly used outlet boundary condition for steady Reynolds averaged Navier-Stokes (RANS) simulation is a fixed pressure at the outlet with all the other dependent variables being extrapolated from the interior. The results of the present study suggest that this is also workable for the URANS simulation of the LM6000 injector flame tube. However, it does not work for the PRNS/VLES simulation due to the unphysical reflections of the pressure disturbances at the outlet boundary. This undesirable situation can be practically alleviated by applying a simple unsteady convection equation for the pressure disturbances at the outlet boundary. The numerical results presented in this paper suggest that this unsteady convection of pressure disturbances at the outlet works very well for all the unsteady simulations (both PRNS/VLES and URANS) of the LM6000 single injector flame tube.

Liu, Nan-Suey↗
NASA NTRS1986

Flow-separation patterns on symmetric forebodies

Flow-visualization studies of ogival, parabolic, and conical forebodies were made in a comprehensive investigation of the various types of flow patterns. Schlieren, vapor-screen, oil-flow, and sublimation flow-visualization tests were conducted over an angle-of-attack range from 0 deg. to 88 deg., over a Reynolds-number range from 0.3X10(6) to 2.0X10(6) (based on base diameter), and over a Mach number range from 0.1 to 2. The principal effects of angle of attack, Reynolds number, and Mach number on the occurrence of vortices, the position of vortex shedding, the principal surface-flow-separation patterns, the magnitude of surface-flow angles, and the extent of laminar and turbulent flow for symmetric, asymmetric, and wake-like flow-separation regimes are presented. It was found that the two-dimensional cylinder analogy was helpful in a qualitative sense in analyzing both the surface-flow patterns and the external flow field. The oil-flow studies showed three types of primary separation patterns at the higher Reynolds numbers owing to the influence of boundary-layer transition. The effect of angle of attack and Reynolds number is to change the axial location of the onset and extent of the primary transitional and turbulent separation regions. Crossflow inflectional-instability vortices were observed on the windward surface at angles of attack from 5 deg. to 55 deg. Their effect is to promote early transition. At low angles of attack, near 10 deg., an unexpected laminar-separation bubble occurs over the forward half of the forebody. At high angles of attack, at which vortex asymmetry occurs, the results support the proposition that the principal cause of vortex asymmetry is the hydrodynamic instability of the inviscid flow field. On the other hand, boundary-layer asymmetries also occur, especially at transitional Reynolds numbers. The position of asymmetric vortex shedding moves forward with increasing angle of attack and with increasing Reynolds number, and moves rearward with increasing Mach number.

Keener, Earl R.↗
NASA NTRS1978

Methods and results of boundary layer measurements on a glider

Boundary layer measurements were carried out on a glider under natural conditions. Two effects are investigated: the effect of inconstancy of the development of static pressure within the boundary layer and the effect of the negative pressure difference in a sublaminar boundary layer. The results obtained by means of an ion probe in parallel connection confirm those results obtained by means of a pressure probe. Additional effects which have occurred during these measurements are briefly dealt with.

Nes, W. V.↗
DOE OSTI2021

Size quantization of an exciton: A toy model of the “dead layer”

In this paper, size-quantization levels of an exciton in large nanocrystals is studied theoretically. For the nanocrystal size, L , much bigger than the Bohr radius, a B , the level positions do not depend on a B . The correction to the levels in a small parameter a B / L depends on the reflection phase of the exciton from the boundary. Calculation of this phase constitutes a three-body problem: electron, hole, and the boundary. This calculation can be performed analytically in the limit when the hole is much heavier than the electron. Physically, a slow motion of the hole towards the boundary takes place in the effective potential created by the fast motion of the electron orbiting the hole and touching the boundary. As a result, the hole is reflected before reaching the boundary. The distance of the closest approach of the hole to the boundary (the dead layer) exceeds a B parametrically .

71 CLASSICAL AND QUANTUM MECHANICS, GENERAL PHYSIC↗
DOE OSTI2020

A phase-field model for non-isothermal phase transformation and plasticity in polycrystalline yttria-stabilized tetragonal zirconia

Here we propose an elastoplastic phase-field (PF) model to investigate the mechanics of tetragonal-to-monoclinic phase transformation (TMPT) and elastoplastic deformation of polycrystalline yttria-stabilized tetragonal zirconia (YSTZ). A Landau polynomial with non-vanishing chemical energy at the equilibrium temperature is introduced to account for the actual formation energies of the phases. The effects of different grain orientations, latent heat, and temperature on TMPT and deformation mechanisms are considered. The suppressive transformation effects of the grain boundaries (GBs) is modeled using an inhomogeneous kinetic coefficient in the bulk and GBs. The simulation results for single crystals demonstrate the capability of the model to reproduce the orientation-dependent compressive deformation of YSTZ similar to atomistic simulations and micropillar experiments. The single crystal with [100] crystallographic orientation along the loading direction (SC[100]) displays both TMPT and plasticity, SC[101] experiences only phase transformation, while SC[001] undergoes only plastic yielding. The TMPT induced by compressive loading exhibits shape memory effect (SME) below the equilibrium transformation temperature and pseudoelasticity above it, while the critical transformation stress increases with increasing loading temperature. The irrecoverable plastic strain is found to trap a part of the monoclinic phase, which prevents a complete reverse transformation. The polycrystalline cases also display SME and PE at low and high temperatures, respectively. Due to the orientation differences between grains and the stress concentrations at geometric nonlinearities, plastic deformation occurs in polycrystalline YSTZ for an applied load less than the yield stress. The results suggest a possible limitation of plasticity and an improvement of the shape recovery of YSTZ if one can control the orientation of the grains and/or increase the density of stacking faults at the GBs during material processing.

36 MATERIALS SCIENCE↗
NASA NTRS1977

Shock-tube chemistry. I - The laminar-to-turbulent boundary layer transition

A model is proposed for calculating laminar-turbulent transition in the boundary layer of flows in chemical shock tubes. It is shown that the time to turbulent transition may be treated as the time required by the shocked gas to move over one shock-tube diameter. The model is shown to provide a means of predicting the turbulent transition time in any shock tube experiment, provided the condition of limiting flow is met. Knowing the transition time, it becomes possible to obtain the density and temperature profiles and, hence, to determine precisely the effect of shock tube boundary layers on the measurement of chemical rate constants.

Bander, J. A.↗
NASA NTRS1992

An investigation of shock/turbulent boundary layer bleed interactions

A numerical investigation was conducted to determine the effect of bleed on oblique shock wave/turbulent boundary layer interactions. The numerical solution to the compressible Navier-Stokes equations reveal the flow details throughout the interaction zone and inside the normal bleed slot. Results are presented for an incident oblique shock of sufficient strength to cause boundary layer separation in the absence of bleed. Bleed is applied across the shock impingement location over a range of bleed mass flow rates corresponding to different values of plenum pressures. The results indicate a complex flow structure with large variations in both normal and tangential flow velocities across the bleed slot. The flow entrainment into the slot is accompanied by an expansion-compression wave system with a bow shock originating inside the bleed slot. Increasing the bleed mass flow by decreasing the plenum pressure caused an initial decrease then a later increase in the boundary layer momentum and displacement thickness downstream of the interaction.

Hamed, A.↗
NASA NTRS2010

Receptivity of Hypersonic Boundary Layers over Straight and Flared Cones

The effects of adverse pressure gradients on the receptivity and stability of hypersonic boundary layers were numerically investigated. Simulations were performed for boundary layer flows over a straight cone and two flared cones. The steady and the unsteady flow fields were obtained by solving the two-dimensional Navier-Stokes equations in axi-symmetric coordinates using the 5th order accurate weighted essentially non-oscillatory (WENO) scheme for space discretization and using third-order total-variation-diminishing (TVD) Runge-Kutta scheme for time integration. The mean boundary layer profiles were analyzed using local stability and non-local parabolized stability equations (PSE) methods. After the most amplified disturbances were identified, two-dimensional plane acoustic waves were introduced at the outer boundary of the computational domain and time accurate simulations were performed. The adverse pressure gradient was found to affect the boundary layer stability in two important ways. Firstly, the frequency of the most amplified second-mode disturbance was increased relative to the zero pressure gradient case. Secondly, the amplification of first- and second-mode disturbances was increased. Although an adverse pressure gradient enhances instability wave growth rates, small nose-tip bluntness was found to delay transition due to the low receptivity coefficient and the resulting weak initial amplitude of the instability waves. The computed and measured amplitude-frequency spectrums in all three cases agree very well in terms of frequency and the shape except for the amplitude.

Balakumar, Ponnampalam↗
NASA NTRS1979

Influence of corrugation-roughness shape on aerodynamic heating in thick turbulent boundary layers

The results of an experimental program to evaluate the effects of wave shape on heat transfer and pressure distributions of corrugation roughened flat plates in thick turbulent boundary layers are presented. The wave shaping effects studied were wavelength, wave amplitude, and wave radius. The experimental program consisted of tests in the sidewall boundary layer of the Langley Continuous Flow Hypersonic Tunnel at a Mach number of 10.3 and Reynolds numbers from 1.3 to 8.2 x 10 to the 6th per meter. The effects of wave shape are analyzed, correlated and compared with previous data obtained in both thick and thin boundary layers. The correlations include the effects of a large range of geometric, inviscid flow, internal boundary layer, and bulk boundary layer parameters in supersonic and hypersonic flow.

Brandon, H. J.↗
NASA NTRS1992

Modeling of near wall turbulence and modeling of bypass transition

The objectives for this project are as follows: (1) Modeling of the near wall turbulence: We aim to develop a second order closure for the near wall turbulence. As a first step of this project, we try to develop a kappa-epsilon model for near wall turbulence. We require the resulting model to be able to handle both near wall turbulence and turbulent flows away from the wall, computationally robust, and applicable for complex flow situations, flow with separation, for example, and (2) Modeling of the bypass transition: We aim to develop a bypass transition model which contains the effect of intermittency. Thus, the model can be used for both the transitional boundary layers and the turbulent boundary layers. We require the resulting model to give a good prediction of momentum and heat transfer within the transitional boundary and a good prediction of the effect of freestream turbulence on transitional boundary layers.

Yang, Z.↗
NASA NTRS1990

Polymer/riblet combination for hydrodynamic skin friction reduction

A process is disclosed for reducing skin friction and inhibiting the effects of liquid turbulence in a system involving the flow of a liquid along the surface of a body, e.g., a marine vehicle. This process includes injecting a drag reducing polymer into the valleys of adjacent, evenly spaced, longitudinal grooves extending along the length of the surface of the body, so that the rate of diffusion of the polymer from individual grooves into the liquid flow is predictably controlled by the groove dimensions. When the polymer has diffused over the tips of the grooves into the near wall region of the boundary layer, the polymer effectively reduces the turbulent skin friction. A substantial drag reducing effect is achieved with less polymer than must be used to lower skin friction when the surface of the body is smooth.

Bushnell, Dennis M.↗
NASA NTRS1995

Polymer/riblet combination for hydrodynamic skin friction reduction

A process is disclosed for reducing skin friction and inhibiting the effects of liquid turbulence in a system involving the flow of a liquid along the surface of a body, e.g. a marine vehicle. This process includes injecting a drag reducing polymer into the valleys of adjacent, evenly spaced, longitudinal grooves extending along the length of the surface of the body, so that the rate of diffusion of the polymer from individual grooves into the liquid flow is predictably controlled by the groove dimensions. When the polymer has diffused over the tips of the grooves into the near wall region of the boundary layer, the polymer effectively reduces the turbulent skin friction. A substantial drag reducing effect is achieved with less polymer than must be used to lower skin friction when the surface of the body is smooth.

Reed, Jason C.↗