Navier-Stokes Code Research Articles

Hydrodynamic Design of Pump Diffuser Using Inverse Design Method and CFD

A new approach to optimizing a pump diffuser is presented, based on a three-dimensional inverse design method and a Computational Fluid Dynamics (CFD) technique. The blade shape of the diffuser was designed for a specified distribution of circulation and a given meridional geometry at a low specific speed of 0.109 (non-dimensional) or 280 (m3/min, m, rpm). To optimize the three-dimensional pressure fields and the secondary flow behavior inside the flow passage, the diffuser blade was more fore-loaded at the hub side as compared with the casing side. Numerical calculations, using a stage version of Dawes three-dimensional Navier-Stokes code, showed that such a loading distribution can suppress flow separation at the corner region between the hub and the blade suction surface, which was commonly observed with conventional designs having a compact bowl size (small outer diameter). The improvements in stage efficiency were confirmed experimentally over the corresponding conventional pump stage. The application of multi-color oil-film flow visualization confirmed that the large area of the corner separation was completely eliminated in the inverse design diffuser.

A THEORETICAL STUDY OF PAPER FLUTTER

Two different methods of analysis were developed in order to clarify the phenomenon of paper flutter. One of these is a flutter simulation using a Navier–Stokes code (N–S simulation). N–S simulation was used to determine the unsteady lift force, the amplitude of flutter, and the air-flow around a paper sheet, by means of a time-marching scheme. The other form of analysis is based on a potential-flow analysis of an oscillating thin airfoil via an eigenvalue analysis for determining stability. The flutter speeds and flutter modes obtained by each method are consistent. Aspects of the behavior of paper flutter that had not been clarified in the experimental analysis were clarified by potential flow analysis. From the results of this study, it was shown that potential flow analysis is very convenient and adequate for a parametric study of this problem.

DEVELOPMENT OF A THREE-DIMENSIONAL VISCOUS AEROELASTIC SOLVER FOR NONLINEAR PANEL FLUTTER

A new three-dimensional (3-D) viscous aeroelastic solver for nonlinear panel flutter is developed in this paper. A well-validated full Navier–Stokes code is coupled with a finite-difference procedure for the von Karman plate equations. A subiteration strategy is employed to eliminate lagging errors between the fluid and structural solvers. This approach eliminates the need for the development of a specialized, tightly coupled algorithm for the fluid/structure interaction problem. The new computational scheme is applied to the solution of inviscid two-dimensional panel flutter problems for subsonic and supersonic Mach numbers. Supersonic results are shown to be consistent with the work of previous researchers. Multiple solutions at subsonic Mach numbers are discussed. Viscous effects are shown to raise the flutter dynamic pressure for the supersonic case. For the subsonic viscous case, a different type of flutter behavior occurs for the downward deflected solution with oscillations occurring about a mean deflected position of the panel. This flutter phenomenon results from a true fluid/structure interaction between the flexible panel and the viscous flow above the surface. Initial computations have also been performed for inviscid, 3-D panel flutter for both supersonic and subsonic Mach numbers.

Numerical simulation of flow in a circular duct fitted with air‐jet vortex generators

AbstractMost of the fundamental studies of the use of air‐jet vortex generators (AJVGs) have concentrated on their potential ability to inhibit boundary layer separation on aerofoils. However, AJVGs may be of use in controlling or enhancing certain features of internal duct flows. For example, they may be of use in controlling the boundary layer at the entrance to engine air intakes, or as a means of increasing mixing and heat transfer. The objective of this paper is to analyse the flow field in the proximity of an air‐jet vortex generator array in a duct by using two local numerical models, i.e. a simple flat plate model and a more geometrically faithful sector model. The sector model mirrors the circular nature of the duct's cross‐section and the centre line conditions on the upper boundary. The flow was assumed fully turbulent and was solved using the finite volume, Navier–Stokes Code CFX 4 (CFDS, AEA Technology, Harwell) on a non‐orthogonal, body‐fitted, grid using the k–ε turbulence model and standard wall functions. Streamwise, vertical and cross‐stream velocity profiles, circulation and peak vorticity decay, peak vorticity paths in cross‐stream and streamwise direction, cross‐stream vorticity profiles and cross‐stream wall shear stress distributions were predicted. Negligible difference in results was observed between the flat plate and the sector model, since the produced vortices were small relative to the duct diameter and close to the surface. The flow field was most enhanced, i.e. maximum thinning of the boundary layer, with a configuration of 30° pitch and 75° skew angle. No significant difference in results could be observed between co‐ and counter‐rotating vortex arrays. Copyright © 2002 John Wiley & Sons, Ltd.

Numerical investigation on the effect of blade sweep on the performance of Wells turbine

A Wells turbine is one of the simplest and promising self-rectifying air turbines which is basic to the needs of the near future and likely to be economically viable. With the recent development in computer hardware and software, it has now become practicable to conduct a reasonable computation of three-dimensional turbulent flows through complex geometry. To investigate the effect of blade sweep on the performance of the Wells turbine, the numerical investigation was carried out under steady flow condition with a fully 3-D Navier–Stokes code for two kinds of blades, NACA0020 and CA9. As a result, it was found that the performance of the Wells turbine is considerably influenced by the blade sweep. The optimum blade sweep ratio ( f=0.35) for the NACA0020 was found. This value is just the same as one obtained experimentally by the authors in the past. It was also found that the overall turbine performance for the NACA0020 is better than that for the CA9. It was shown that the numerical method is able quite well to predict the effect of blade sweep of the Wells turbine. The detailed flow patterns for several blade sweeps were also shown and discussed in this paper.

Hybrid Unstructured Navier-Stokes Simulation of Military Aircraft Configuration Including Inlet Airflow and Jet Exhaust Effects.

A CFD-based design system, CASPER (Computational Aerodynamics System for Performance Evaluation and Research), was developed at TRDI-JDA for advanced aircraft design at the next generation. In this paper, hybrid unstructured Reynolds-Averaged Navier-Stokes computations around a military aircraft, F-16 A, powered configurations including inlet airflow and jet exhaust effects of propulsion system were carried out as a practical example by using the CASPER. In order to validate the present CFD code, the present computed results are well compared to the wind-tunnel testing data with respect to total pressure recovery ratio on engine compressor face and aft-body pressure distributions at steady state. Thus, the present hybrid unstructured Navier-Stokes code is expected to be effective as one of design tools of advanced aircraft propulsion system.

Control of Pseudo-Shock by Slot Injection

It is well known that the boundary layer separation due to the adverse pressure gradient of the pseudo-shock affects the dynamic feature of the pseudo-shock such as the shape of the first shock, the pressure recovery and the self-induced oscillation. The suppression of the separation is effective to control the interaction between the shock wave and the boundary layer. In this research, the boundary layer control with the slot injection is studied experimentally and numerically. The experiment is carried out in a blow-down supersonic wind tunnel at free stream Mach number of 2.0. The numerical simulation is made with the two-dimensional compressible Navier-Stokes code. In solving the equations, the second-order accurate Harten-Yee's upwind TVD scheme and k-e turbulence model modified for the compressible flow are used. The results show that the separation is appreciably suppressed and the shape of the first shock is changed, resulting the wall pressure fluctuation to be greatly reduced. Furthermore the total pressure loss is decreased with the low slot height.

A new technique for a parallel dealiased pseudospectral Navier–Stokes code

A novel aspect of a parallel procedure for the numerical simulation of the solution of the Navier–Stokes equations through the Fourier–Galerkin pseudospectral method is presented. It consists of a dealiased (“3/2” rule) transposition of the data that organizes the computations in the distributed direction in such a way that whenever a Fast Fourier Transform must be calculated, the algorithm will employ data stored solely on the proper memory of the processor which is computing it. This provide for the employment of standard routines for the computations of the Fourier transform. The aliasing removal procedure has been directly inserted into the transposition algorithm. The code is written for distributed memory computers, but not specifically for a peculiar architecture. The use on a variety of machines is allowed by the adoption of the Message Passing Interface library. The portability of the code is demonstrated by the similar performances, in particular the high efficiency, that all the machines tested show up to a number of parallel processors equal to 1/2 the truncation parameter N/2. Explicit time integration is used. The present code organization is relevant to physical and mathematical problems which require a three dimensional spectral treatment.

Two-equation turbulence models for prediction of heat transfer on a transonic turbine blade

Two versions of the two-equation k– ω model and a shear stress transport (SST) model are used in a three-dimensional, multi-block, Navier–Stokes code to compare the detailed heat transfer measurements on a transonic turbine blade. It is found that the SST model resolves the passage vortex better on the suction side of the blade, thus yielding a better comparison with the experimental data than either of the k– ω models. However, the comparison is still deficient on the suction side of the blade. Use of the SST model does require the computation of distance from a wall, which for a multi-block grid, such as in the present case, can be complicated. However, a relatively easy fix for this problem was devised. Also addressed are issues such as (1) computation of the production term in the turbulence equations for aerodynamic applications, and (2) the relation between the computational and experimental values for the turbulence length scale, and its influence on the passage vortex on the suction side of the turbine blade.

Aerodynamic performance improvement by divergent trailing edge modification to a supercritical airfoil

A computational study has been performed to determine the effects of divergent trailing edge (DTE) modification to a supercritical airfoil in transonic flow field. For this, the computational result with the original DLBA 186 supercritical airfoil was compared to that of the modified DLBA 283. A Navier-Stokes code, Fluent 5. 1, was used with Spalart-Allmaras’s one-equation turbulence model. Results in this study showed that the reduction in drag due to the DTE modification is associated with weakened shock and delayed shock appearance. The decrease in drag due to the DTE modification is greater than the increase in base drag. The effect of the recirculating flow region on lift increase was also observed. An airfoil with DTE modification achieved the same lift coefficient at a lower angle of attack while giving a lower drag coefficient. Thus, the lift-to-drag ratio increases in transonic flow conditions compared to the original airfoil. The lift coefficient increases considerably whereas the lift slope increases just a little due to DTE modification.

Underexpanded Jet for Testing Laser-Based Combustion Diagnostics

The design, construction, and operation of an underexpanded jet facility developed as a test  owŽ eld for laserbased supersonic combustion diagnostics are described. The continuous-running device accelerates hot gases from a 160-kPa, 2000-K, stoichiometric H2 /Air combustion chamber through a 3.15-mm i.d. alumina tube into a 11-kPa chamber. The resulting underexpanded jet of  ame species provides a range of  owŽ eld parameters (M = 1–4, T = 1700–500 K, P = 72.8–2.7 kPa, and 1⁄2 = 0.12–0.004 kg/m3 ) typical of practical supersonic combustion experiments. The modeling of the jet  owŽ eld using an axisymmetric Navier–Stokes code is also described.

Actuator Duct Model of Turbomachinery Components for Powered-Nacelle Navier-Stokes Calculations

An actuator duct model was installed into the WIND Navier-Stokes code. With this model, flow through fan rotors and fan exit guide vanes can be simulated without blade geometry. Flow turning, total pressure and total temperature changes across blade rows are controlled by body forces added to the Navier-Stokes equations. This model provides an affordable and relatively accurate prediction of a three-dimensional flow field through rotor and stator blade rows. A NASA Stage 35 compressor and a NASA 22 test rig flow were analyzed to quantify the ability of the actuator duct theory to represent blade rows. Favorable comparisons between predicted and measured total pressure, total temperature and swirl profiles indicated that the actuator duct logic was correctly installed in the WIND code. Comparisons of inlet separation angle-of-attack were made between test data and predictions using the actuator duct theory. The WIND code with the actuator duct model predicted higher separation angle-of-attack by an average of 2° for different fan speeds. The effect of the rotor on inlet flow separation was also determined by comparing the predictions of a flow-through nacelle with the predictions using the actuator duct to simulate the rotating fan. The comparison shows that the presence of the rotor increased separation free angle-of-attack over the flow-through nacelle which is consistent with the observation from the experiment. NOMENCLATURE

Transport Coefficients for High-Temperature Nonequilibrium Air Flows

A new transport coefficient model based on the first approximation of the Chapman-Cowling theory is implemented into the URANUS nonequilibrium NavierStokes Code in order to predict re-entry flows more accurately. The new model is compared to the widely used, simplified model of Gupta and Yos. The simplifications used to obtain the latter model are described in detail. The transport coefficients computed with both methods differ by up to a factor of two in the case of partially ionized air under equilibrium conditions. Therefore, the new model allows a more accurate description of the thermo-chemical relaxation in the shock region of significantly ionized, high-temperature flows. The differences in surface heat fluxes between the new model and the model of Yos and Gupta are smaller than 4%.

Modeling film-coolant flow characteristics at the exit of shower-head holes

Abstract The coolant flow characteristics at the hole exits of a film-cooled blade are derived from an earlier study where the hole pipes and coolant plenum were also discretized. The blade chosen is the VKI rotor with three staggered rows of shower-head holes. The present analysis applies these flow characteristics at the shower-head hole exits. A multi-block three-dimensional Navier–Stokes code with Wilcox's k – ω model is used to compute the heat transfer coefficient on the film-cooled turbine blade. A reasonably good comparison with the experimental data as well as with the more complete earlier analysis where the hole pipes and coolant plenum were also gridded is obtained. If the 1/7th power law is assumed for the coolant flow velocity and temperature distributions at the hole exits, considerable differences in the heat transfer coefficient on the blade surface, especially in the leading-edge region, are observed even though the span-averaged values of h match well with the experimental data. This calls for spatially-resolved experimental data near film-cooling holes on a blade for better validation of the code.

MEGAFLOW: Parallel complete aircraft CFD

As a consequence of the worldwide competition in the aircraft market requirements for the accurate prediction of aerodynamic performance and optimization of configurations have increased very much. More sophisticated wind tunnel as well as high quality CFD techniques have become necessary and essential tools for aircraft industry aerodynamic development groups. This requires an ongoing struggle for efficiency improvements where parallel computing is one of the major issues. This paper considers the MEGAFLOW activities in the area of parallel flow solvers including their application and use in the industrial framework. It describes the parallelization principles for the structured multi-block Navier–Stokes code FLOWer and the unstructured Navier–Stokes code TAU. Principle results for industrial applications are given with respect to efficiency, speed-up, load balancing, etc. Computational examples demonstrate the quality of the solvers for 3D flow over wing/body configurations as well as complete transport aircraft.

A Computational Method for Eu's Generalized Hydrodynamic Equations of Rarefied and Microscale Gasdynamics

Generalized hydrodynamic equations have been proposed by Eu (Kinetic Theory and Irreversible Thermodynamics, 1992) for modeling the motion of gases far removed from equilibrium. His generalized hydrodynamic equations are consistent with the laws of thermodynamics. In this paper, a computational method of solving Eu's generalized hydrodynamic equations is presented. It has been shown that the new equations are applicable to all Mach numbers and indeed satisfy the second law of thermodynamics at all Knudsen numbers and to every order of approximation. The computational method of the generalized hydrodynamic equations is based on the finite-volume formulation and is exactly the same as the compressible Navier–Stokes codes, except for an additional routine for calculating the shear stress and the heat flux from the given conserved variables and thermodynamic forces. To check its validity and potential for hydrodynamics applications, the method is tested for the structure of one-dimensional shock wave and for a two-dimensional flat plate flow problem. The numerical results show that the new computational model yields the shock solutions for Mach numbers tested, up to M=30, and removes the singularity near the leading edge of a flat plate that is ill-defined in the case of the Navier–Stokes theory.

Parallel processing with the Wind CFD code at Boeing

Wind is a general purpose CFD code capable of either Euler or Navier–Stokes analysis (R.H. Bush, A Three Dimensional Zonal Navier–Stokes Code for Subsonic through Hypersonic Propulsion Flowfields, AIAA 88-2380, 11 July 1998; R.H. Bush, G.D. Power, The Wind Production Flow Solver For the NPARC Alliance, AIAA 98-0935, 12 January 1998). The Wind code and its predecessor NASTD have been used for CFD applications, using its parallel mode, at Boeing since 1993. The Wind code is currently developed and distributed by the NPARC alliance. For more information see the NPARC web site at http://www.lerc.nasa.gov/www/winddocs/.

Computation of ship transom stern flow using a multiblock grid method

The incompressible Reynolds Averaged Navier-Stokes (RANS) code using the multiblock grid method has been developed for simulating the steady free surface flows over ship having complex geometries. “Horizontal Mode”, “Vertical Mode” and “BJ-SGS” which are methods of updating boundary conditions on block interfaces are incorporated in the present code and investigated in the computation of free surface flows about Wigley hull. The new code is applied to computation of transom stern flow about a container ship and a high speed vessel and those numerical results are compared with the experimental results.

Aerothermal predictions with transition models for high-pressure turbine blades

Turbine aerothermal analysis is a key issue of the entire engine design as higher performance and longer component life must always be reached. For this purpose, a substantial effort is dedicated to aerothermal methods at Snecma-Moteur. This paper deals with the problem of heat transfer prediction on uncooled turbine blades. The main parameter investigated is the transition process, but consideration is also given to the turbulence model, effect of free stream turbulence on laminar boundary layer and mesh refinement. All the models are implemented in a three-dimensional Navier-Stokes code. Quasi-three-dimensional computations are used to evaluate all models, and then fully three-dimensional computations are performed. Aerothermal predictions are compared with static pressure and heat transfer measurements from the IACA and TATEF Brite-EuRam programmes carried out in the test facilities of DERA, Pyestock and the von Karman Institute.

Numerical Analysis of Three Dimensional Dynamic Stall of an Oscillating Wing

Three-dimensional dynamic stall over an oscillating wing has been analyzed by using a compressible Navier-Stokes code. The code solved the thin-layer Navirer-Stokes equations with a second-order time accuracy for a semispan wing with 0.3048m chord, a NACA 0015 airfoil section, and zero twist Computations were made for a freestream Mach number of 0.29, a chord Reynolds number of 1.9510(sup)6 and a reduced frequency equal to 0.1. Numerical results were compared with experimental data which include the hysteresis of lift, drag and moment at various wing span. The comparison reveals the quantitative as well as qualitative nature of the three-dimensional dynamic stall.