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At least 433 records · Page 24

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NASA NTRS1991

Flowfield analysis of modern helicopter rotors in hover by Navier-Stokes method

The viscous, three-dimensional, flowfields of UH60 and BERP rotors are calculated for lifting hover configurations using a Navier-Stokes computational fluid dynamics method with a view to understand the importance of planform effects on the airloads. In this method, the induced effects of the wake, including the interaction of tip vortices with successive blades, are captured as a part of the overall flowfield solution without prescribing any wake models. Numerical results in the form of surface pressures, hover performance parameters, surface skin friction and tip vortex patterns, and vortex wake trajectory are presented at two thrust conditions for UH60 and BERP rotors. Comparison of results for the UH60 model rotor show good agreement with experiments at moderate thrust conditions. Comparison of results with equivalent rectangular UH60 blade and BERP blade indicates that the BERP blade, with an unconventional planform, gives more thrust at the cost of more power and a reduced figure of merit. The high thrust conditions considered produce severe shock-induced flow separation for UH60 blade, while the BERP blade develops more thrust and minimal separation. The BERP blade produces a tighter tip vortex structure compared with the UH60 blade. These results and the discussion presented bring out the similarities and differences between the two rotors.

Srinivasan, G. R.↗

NASA NTRS1993

Vortex generation and mixing in three-dimensional supersonic combustors

The generation and evolution of the flow vorticity established by instream injector ramps in a high Mach number/high enthalpy scramjet combustor flow-field are described in detail for a number of computational cases. Classical fluid dynamic circulation is presented for these cases in order to clarify the spatial distribution and convection of the vorticity. The ability of the simulations to accurately represent Stokes Law of circulation is discussed and shown. In addition, the conservation of swirl (effectively the moment-of-momentum theorem) is presented for these flows. The impact of both turbulent diffusion and the vortex/ramp non-uniformity on the downstream mixing rate is clearly illustrated. A correlation over the length of the combustor between fuel-air mixing and a parameter called the vortex stirring length is demonstrated. Finally, computational results for a representative ramp injector are compared with experimental data. Influence of the stream vorticity on the effective turbulent Prandtl number used in the simulation is discussed.

Riggins, D. W.↗

NASA NTRS1994

New constitutive equation for the volume viscosity in fluids

The traditional volume viscosity, Stokes' hypothesis, and acoustical relaxation are reviewed. The lossy Navier-Stokes Equation is applied to periodic (acoustic) flow, and it is shown that the traditional volume viscosity leads to a result which contradicts that describing acoustical relaxation. It is demonstrated that the addition of a second volume viscosity term to the constitutive equation, to account for pressure relaxation, resolves the conflict, and leads to a direct correspondence between the volume viscosity parameters and the acoustical relaxation parameters. The representation of volume viscosity is formulated for the case of multiple relaxations, as occur in air. Finally, an application of the new constitutive equation to a simple convective compressible flow, namely a linearly accelerating flow, demonstrates the impact of volume viscosity upon the flow and the physical conditions for which it is important.

Zuckerwar, Allan J.↗

NASA NTRS1996

Numerical Simulation of Mixing Enhancement in a Hot Supersonic Jet

Experimental observations show that the presence of small tabs on the edge of a hot, compressible jet exiting into a slower moving, colder ambient flow can increase the rate of spreading of the jet. This suggests that the rate of mixing of the jet and the ambient fluid is also increased. In order to elucidate the physical mechanism responsible for the increased spreading rate a set of calculations were carried out within the framework of the compressible three dimensional Navier-Stokes equations. A series of grid refinements were made to assess the accuracy of the results. The first simulated the flow without the tabs, obtaining reasonable agreement with experimental measurements of the velocity. We then simulated the flow, without tabs, over a range of values of the convective Mach number in order to determine the dependence of the mixing on this parameter. Simulations with modeled tabs were also carried out. In these calculations the effect of the tabs on the flow was modeled by pairs of counter rotating vortices. The results of these calculations indeed show that the presence of the tabs increase the spreading rate of the jet. The basic physical mechanism responsible for the enhanced spreading rate is discussed and qualitative comparisons with flow visualizations are made.

Grosch, C. E.↗

NASA NTRS1992

Analytical estimates of radial segregation in Bridgman growth from low-level steady and periodic accelerations

Approximate analytical solutions are developed that are mathematically tractable but provide the elements of CFD models for estimating convective flows subjected to low-level accelerations. The estimates are applicable to dilute systems that have horizontal temperature gradients in the vertical Bridgman configuration. The Navier-Stokes momentum equation is solved in 1D and subsequently in 2D by the first-order perturbation method. The analysis used is the case of plane-front solidification of a dilute system in the vertical thermally stable configuration. The model is found to predict the degree of radial segregation within a factor of two for a range of material and processing parameters. The method provides detailed information on the effects of processing on solute distribution in the grown crystal which are of interest in the orbital experiments emphasizing the control of radial segregation.

Naumann, Robert J.↗

NASA NTRS1971

The three-dimensional steady radial expansion of a viscous gas from a sonic source into a vacuum.

The three-dimensional steady radial expansion of a viscous, heat-conducting, compressible fluid from a spherical sonic source into a vacuum is analyzed using the Navier-Stokes equations as a basis. It is assumed that the model fluid is a perfect gas having constant specific heats, a constant Prandtl number of order unity, and viscosity coefficients varying as a power of the absolute temperature. Limiting forms for the flow variable solutions are studied for the Reynolds number based on the sonic source conditions, going to infinity and the Newtonian parameter both fixed and going to zero. For the case of the viscosity-temperature exponent between .5 and 1, it is shown that the velocity as well as the pressure approach zero as the radial distance goes to infinity. The formulations of the distinct regions that span the domain extending from the sonic source to the vacuum are presented.

Bush, W. B.↗

NASA NTRS1984

SSME thrust chamber simulation using Navier-Stokes equations

The capability of the PHOENICS fluid dynamics code in predicting two-dimensional, compressible, and reacting flow in the combustion chamber and nozzle of the space shuttle main engine (SSME) was evaluated. A non-orthogonal body fitted coordinate system was used to represent the nozzle geometry. The Navier-Stokes equations were solved for the entire nozzle with a turbulence model. The wall boundary conditions were calculated based on the wall functions which account for pressure gradients. Results of the demonstration test case reveal all expected features of the transonic nozzle flows. Of particular interest are the locations of normal and barrel shocks, and regions of highest temperature gradients. Calculated performance (global) parameters such as thrust chamber flow rate, thrust, and specific impulse are also in good agreement with available data.

Przekwas, A. J.↗

NASA NTRS1990

Numerical simulation of 3-D Benard convection with gravitational modulation

In this numerical study, randomly and sinusoidally modulated gravitational fields imposed on three-dimensional Rayleigh-Benard convection are investigated in an effort to understand the effects of vibration (G-Jitter) on fluid systems. The time-dependent, Navier-Stokes equations and the energy equation with Boussinesq approximations are solved by a semi-implicit, pseudospectral procedure. An analysis of energy balances indicates that with increasing modulation amplitude, transition from synchronous to relaxation oscillation goes through the subharmonic response. Random modulations are found to be less stabilizing than sinusoidal and are shown to impose three-dimensionality on the flow for some parameter ranges both at terrestrial and zero base gravity conditions.

Biringen, S.↗

NASA NTRS1993

Grid and design variables sensitivity analyses for NACA four-digit wing-sections

Two distinct parameterization procedures are developed for investigating the grid sensitivity with respect to design parameters of a wing-section example. The first procedure is based on traditional (physical) relations defining NACA four-digit wing-sections. The second is advocating a novel (geometrical) parameterization using spline functions such as NURBS (Non-Uniform Rational B-Splines) for defining the wing-section geometry. An interactive algebraic grid generation technique, known as Hermite Cubic Interpolation, is employed to generate C-type grids around wing-sections. The grid sensitivity of the domain with respect to design and grid parameters has been obtained by direct differentiation of the grid equations. A hybrid approach is proposed for more geometrically complex configurations. A comparison of the sensitivity coefficients with those obtained using a finite-difference approach has been made to verify the feasibility of the approach. The aerodynamic sensitivity coefficients are obtained using the compressible two-dimensional thin-layer Navier-Stokes equations.

Sadrehaghighi, Ideen↗

NASA NTRS1996

Investigation of shock-induced combustion past blunt projectiles

A numerical study is conducted to simulate shock-induced combustion in premixed hydrogen-air mixtures at various free-stream conditions and parameters. Two-dimensional axisymmetric, reacting viscous flow over blunt projectiles is computed to study shock-induced combustion at Mach 5.11 and Mach 6.46 in hydrogen-air mixture. A seven-species, seven reactions finite rate hydrogen-air chemical reaction mechanism is used combined with a finite-difference, shock-fitting method to solve the complete set of Navier-Stokes and species conservation equations. The study has allowed an improved understanding of the physics of shock-induced combustion over blunt projectiles and the numerical results can now be explained more readily with one-dimensional wave-interaction model.

Ahuja, J. K.↗

NASA NTRS2008

Computational Analyses in Support of Sub-scale Diffuser Testing for the A-3 Facility: Aero-Acoustic Analyses and Experimental Validation - Part 3

A unique assessment of acoustic similarity scaling laws and acoustic analogy methodologies in predicting the far-field acoustic signature from a sub-scale altitude rocket test facility at the NASA Stennis Space Center was performed. A directional, point-source similarity analysis was implemented for predicting the acoustic far-field. In this approach, experimental acoustic data obtained from "similar" rocket engine tests were appropriately scaled using key geometric and dynamic parameters. The accuracy of this engineering-level method is discussed by comparing the predictions with acoustic far-field measurements obtained. In addition, a CFD solver was coupled with a Lilley's acoustic analogy formulation to determine the improvement of using a physics-based methodology over an experimental correlation approach. In the current work, steady-state Reynolds-averaged Navier-Stokes calculations were used to model the internal flow of the rocket engine and altitude diffuser. These internal flow simulations provided the necessary realistic input conditions for external plume simulations. The CFD plume simulations were then used to provide the spatial turbulent noise source distributions in the acoustic analogy calculations. Preliminary findings of these studies will be discussed.

Allgood, Daniel C.↗

NASA NTRS1981

A parametric study of staged fuel injector configurations for scramjet applications

A parametric study of staged (multiple) perpendicular fuel injector configurations has been conducted using a computer code which solves the two-dimensional elliptic Navier-Stokes equations. The program computes the turbulent mixing and reaction of hydrogen fuel and air and allows the study of separated regions of the flow immediately preceding and following the injectors. The validity of the code is demonstrated in a cold flow helium injection study with a single injector. Results are presented that describe the flow field near opposing staged injectors over a range of parameters. Parameters that are varied include injector size, fuel split, and distance between injectors. Comparisons of the configurations are made to assess their mixing and potential flame-holding qualities.

Weidner, E. H.↗

NASA NTRS1986

Numerical Solution of the Incompressible Navier-Stokes Equations in Three-dimensional Generalized Curvilinear Coordinates

Numerically solving the incompressible Navier-Stokes equations is known to be time consuming and expensive. Testing of the INS3D computers code, which solves these equations with the use of the pseudocompressibility method, shows this method to be an efficient way to obtain the steady state solution. The effects of the waves introduced by the pseudocompressibility method are analyzed and criteria are set and tested for the choice of the pseudocompressibility parameter which governs the artificial sound speed. The code is tested using laminar flow over a two dimensional backward-facing step, and laminar flow over a two dimensional circular cylinder. The results of the computations over the backward-facing step are in excellent agreement with experimental results. The transient solution of the flow over the cylinder impulsively started from rest is in good agreement with experimental results. However, the computed frequency of periodic shedding of vortices behind the cylinder is not in agreement with the experimental value. For a three dimensional test case, computations were conducted for a cylinder end wall junction. The saddle point separation and horseshoe vortex system appear in the computed field. The solution also shows secondary vortex filaments which wrap around the cylinder and spiral up in the wake.

Rogers, S. E.↗

NASA NTRS2024

Intrastep, Stage-Value Predictors for Diagonally-Implicit Runge-Kutta Methods

To better identify the necessary attributes of good stage-value predictors (SVPs), numerous SVPs are designed for an existing: ESDIRK4(3)7L[2]SA [26] and a new: ESDIRK4(3)8L[2]SA scheme.1 Both are stifflyaccurate, stage-order two, explicit, singly-diagonally implicit Runge–Kutta (ESDIRK) schemes. Tradeoffs are studied in the parameter spaces enforcing the constraints on accuracy, linear stability, nonlinear stability and coefficient size to determine which objectives correlate with effective predictors. The SVPs are tested on three challenging external aerodynamics problems [107 − 108 degrees of freedom (DoFs)], each with a different level of stiffness. The problems include two 3D airfoils simulations and one canonical turbulence simulation. All simulations use the compressible Navier-Stokes equations (CNSE). An entropy stable spectral collocation formulation is used for discretizing the spatial terms in the equations. Simulations are performed at a wide variety of temporal error tolerances. Problems that are sufficiently stiff (e.g., lax temporal error tolerances) benefit from SVPs designed with second-order accuracy and stability properties: A-stability, and L-stability, rather than high accuracy constraints. Simulations with modest stiffness (e.g., strict error tolerances) are better suited for SVPs designed using high accuracy constraints. Designing SVPs with enhanced stability properties is tedious but worthwhile. Simulation times are reduced with optimal SVPs by as much as 100% on some stages, with combined stepwise improvements of between 50 − 100% for both methods. A comparative study is performed with the two aforementioned methods as well as four other ESDIRKs. The newly designed ESDIRK4(3)8L[2]SA with γ ≈ 1/10, proves to be the most efficient of the six tested ESDIRK schemes simulating the CNSE.

Diagonally-Implicit Runge-Kutta↗

NASA NTRS1993

Averaging techniques for steady and unsteady calculations of a transonic fan stage

It is often desirable to characterize a turbomachinery flow field with a few lumped parameters such as total pressure ratio or stage efficiency. Various averaging schemes may be used to compute these parameters. The momentum, energy, and area averaging schemes are described and compared. The schemes were compared for two computed solutions of the midspan section of a transonic fan stage: a steady averaging-plane solution in which average rotor outflow conditions were used as stator inflow conditions, and an unsteady rotor-stator interaction solution. The solutions were computed on identical grids using similar Navier-Stokes codes and an algebraic turbulence model. The unsteady solution is described, some unsteady flow phenomena are discussed, and the steady pressure distributions are compared. Despite large unsteady pressure fluctuations on the stator surface, the steady pressure distribution matched the average unsteady distribution almost exactly. Stator wake profiles, stator loss coefficient, and stage efficiency were computed for the two solutions with the three averaging schemes and are compared. In general, the energy averaging scheme gave good agreement between the averaging-plane solution and the time-averaged unsteady solution, even though certain phenomena due to unsteady wake migration were neglected.

Wyss, M. L.↗

NASA NTRS1993

A comparison of two incompressible Navier-Stokes algorithms for unsteady internal flow

A comparative study of two different incompressible Navier-Stokes algorithms for solving an unsteady, incompressible, internal flow problem is performed. The first algorithm uses an artificial compressibility method coupled with upwind differencing and a line relaxation scheme. The second algorithm uses a fractional step method with a staggered grid, finite volume approach. Unsteady, viscous, incompressible, internal flow through a channel with a constriction is computed using the first algorithm. A grid resolution study and parameter studies on the artificial compressibility coefficient and the maximum allowable residual of the continuity equation are performed. The periodicity of the solution is examined and several periodic data sets are generated using the first algorithm. These computational results are compared with previously published results computed using the second algorithm and experimental data.

Wiltberger, N. Lyn↗

NASA NTRS2022

Identification of Best Practices for Predicting Inlet Performance Using FUN3D Part 2: Installed Inlets.

A series of studies were performed to assess the impacts of boundary condition type and placement, grid reﬁnement, and modeling parameters such as turbulence model and ﬂux limiter on the predicted inlet performance for installed inlet conﬁgurations using the FUN3D ﬂow solver. Two conﬁgurations were considered for the study; a wall-mounted Boundary Layer Ingestion (BLI) inlet and the C607 propulsion model tested in the 8x6 Supersonic Wind Tunnel at the NASA Glenn Research Center. The results of the studies were to be used to recommend best practices, as well as to assess the accuracy of FUN3D for inlet predictions. The results of BLI inlet stud-ies showed a minimal impact of grid reﬁnement on the predicted inlet performance for a constant mass ﬂow rate through the inlet. For the C607 propulsion model, the results showed that while the FUN3D predictions at the Aerodynamic Inter-face Plane (AIP) qualitatively agree with the experimental data, FUN3D showed a tendency to overpredict the circumferential distortion metric (IDCmax) and un-derpredict both the radial distortion metric (IDRmax) and the pressure recovery at the AIP (PRAIP ), with the diﬀerences between FUN3D and the experimental data increasing with increasing grid reﬁnement and Mach number. Additionally, the outﬂow boundary location studies performed for both geometries showed that the solution at the AIP was not signiﬁcantly impacted by the outﬂow boundary location as long as it was not placed at the location of the AIP. The modeling parameter studies did not indicate a path forward for improved predictions for either inlet con-ﬁguration. Finally, comparisons between the mass ﬂow plug and outﬂow geometry versions of the C607 propulsion model illustrated favorable agreement, which indi-cates that the diﬀerences observed are not caused by the outﬂow boundary model for this problem. This problem poses signiﬁcant challenges to Reynolds-averaged Navier-Stokes (RANS) solvers due to the presence of shocks and ﬂow separation in the inlet.

FUN3D↗

NASA NTRS1979

Direct control of the grid point distribution in meshes generated by elliptic equations

The generation of computational grids suitable for obtaining accurate numerical solutions to the three-dimensional Navier-Stokes equations is the subject of intensive research. For a wide class of nozzle configurations, a three-dimensional grid can be constructed by a sequence of two-dimensional grids in successive cross-sectional planes. The present paper is concerned with numerical generation of two-dimensional grids. An effective method of interior grid control is presented based on a modified elliptic system containing free parameters. For a simply connected region, the free parameters are computed from the Dirichlet boundary values. The resulting interior grid point distribution is controlled entirely by a priori selection of the grid point distribution along the boundaries of the section.

Middlecoff, J. F.↗