Critical evaluation of two-equation models for near-wall turbulence
Explore the source record for details and available documents.
Speziale, Charles G.↗
Engineering topics
Anderson, E. C.
Publications and source records attributed to Anderson, E. C..
Real gas effects on hypersonic boundary-layer stability
High-temperature effects alter the physical and transport properties of a gas, air in particular, due to vibrational excitation and gas dissociation, and thus the chemical reactions have to be considered in order to compute the flow field. Linear stability of high-temperature boundary layers is investigated under the assumption of chemical equilibrium and this gas model is labeled here as real gas model. In this model, the system of stability equations remains of the same order as for the perfect gas and the effect of chemical reactions is introduced only through mean flow and gas property variations. Calculations are performed for Mach 10 and 15 boundary layers and the results indicate that real gas effects cause the first mode instability to stabilize while the second mode is made more unstable. It is also found that the second mode instability shifts to lower frequencies. There is a slight destabilizing influence of real gas on the Goertler instability as compared to the perfect gas results.
Malik, M. R.↗
Assessment and application of Reynolds stress closure models to high-speed compressible flows
The paper presents results from the development of higher order closure models for the phenomological modeling of high-speed compressible flows. The work presented includes the introduction of an improved pressure-strain correlationi model applicable in both the low- and high-speed regime as well as modifications to the isotropic dissipation rate to account for dilatational effects. Finally, the question of stiffness commonly associated with the solution of two-equation and Reynolds stress transport equations in wall-bounded flows is examined and ways of relaxing these restrictions are discussed.
Gatski, T. B.↗
Aerodynamic design of the contoured wind-tunnel liner for the NASA supercritical, laminar-flow-control, swept-wing experiment
An overview is presented of the entire procedure developed for the aerodynamic design of the contoured wind tunnel liner for the NASA supercritical, laminar flow control (LFC), swept wing experiment. This numerical design procedure is based upon the simple idea of streamlining and incorporates several transonic and boundary layer analysis codes. The liner, presently installed in the Langley 8 Foot Transonic Pressure Tunnel, is about 54 ft long and extends from within the existing contraction cone, through the test section, and into the diffuser. LFC model testing has begun and preliminary results indicate that the liner is performing as intended. The liner design results presented in this paper, however, are examples of the calculated requirements and the hardware implementation of them.
Newman, P. A.↗
Numerical design of the contoured wind-tunnel liner for the NASA swept-wing LFC test
A contoured, nonporous, wind-tunnel liner has been designed in order to simulate a free-flight infinite yawed-wing, transonic-flow condition about a large-chord, supercritical-section, laminar-flow-control (LFC), swept-wing test panel. The numerical procedure developed for this aerodynamic liner design is based upon the simple idea of streamlining and incorporates several existing transonic and boundary-layer analysis codes. A summary of the entire procedure is presented to indicate: what was done and why, the sequence of steps, and the overall data flow. The liner is being installed in the NASA Langley 8-Foot Transonic Pressure Tunnel (TPT). Test results indicating the aerodynamic performance of the liner are not yet available; thus, the liner design results given here are examples of the calculated requirements and the hardware implementation.
Newman, P. A.↗
User guide for STRMLN: A boundary-layer program for contoured wind-tunnel liner design
A 2-D boundary layer computer code developed to process data for an arbitrary number of streamlines is presented. Provisions are included for the computer code to determine either mass transfer rates necessary for an effective boundary layer displacement of zero thickness or the effective displacement thickness for a specified mass transfer-rate distribution. The computer code was developed to be compatible with other computer codes which are being modified and/or developed at the NASA-Langley Research Center in order to design the three dimensional, contoured, wind tunnel liner used in transonic testing of a laminar flow control system installed on a supercritical airfoil section. A brief discription of the liner design procedure, representative liner calculations, adaptive-wall design for a two dimensional wind tunnel test, and other applications are reported.
Anderson, E. C.↗
The impact of turbulence on a radiating shock layer with coupled ablation injection
This paper provides a description of the potential impact of turbulence on a radiating flow field with large surface blowing. This is accomplished by calculating the forebody flow field with coupled carbon-phenolic mass injection for a probe entering the Jupiter atmosphere. Both laminar and turbulent flow conditions are assumed. For the no blowing solutions, turbulence is shown to have no impact on the surface radiative heating. However, with ablation injection, turbulence significantly increases the surface radiative heating rate. This occurs because the turbulence brings the high temperature gases closer to the surface, thus thinning the cool molecular gases at the wall which are responsible for blocking much of the radiation.
Moss, J. N.↗
Transitional Boundary-Layer Solutions Using a Mixing-Length and a Two-Equation Turbulence Model
Boundary-layer solutions were obtained using the conventional two-layer mixing-length turbulence model and the Wilcox-Traci two-equation model of turbulence. Both flatplate and blunt-body geometries were considered. The most significant result of the study is development of approximations for the two-equation model which permit streamwise stepsize comparable to that used in mixing-length computations. Additionally, a set of model-equation boundary conditions derived which apply equally well to both flat-plate and blunt-body geometries. Solutions obtained with the two-equations turbulence model are compared with experimental data and/or corresponding solutions obtained using the mixing-length model. Agreement is satisfactory for flat-plate boundary layers but not for blunt body boundary layers.
Anderson, E. C.↗
Numerical design of streamlined tunnel walls for a two-dimensional transonic test
An analytical procedure is discussed for designing wall shapes for streamlined, nonporous, two-dimensional, transonic wind tunnels. It is based upon currently available 2-D inviscid transonic and boundary layer analysis computer programs. Predicted wall shapes are compared with experimental data obtained from the NASA Langley 6 by 19 inch Transonic Tunnel where the slotted walls were replaced by flexible nonporous walls. Comparisons are presented for the empty tunnel operating at a Mach number of 0.9 and for a supercritical test of an NACA 0012 airfoil at zero lift. Satisfactory agreement is obtained between the analytically and experimentally determined wall shapes.
Newman, P. A.↗
Analytical design of a contoured wind-tunnel liner for supercritical testing
The present analytical design procedure is being developed in order to determine the shape of a contoured nonporous wind tunnel liner for use in the Ames 12-foot pressure wind tunnel test of a large chord, laminar flow control, swept wing panel which has a supercritical airfoil section. This procedure is applicable to the two-dimensional streamlined tunnel problem wall shape with that found experimentally.
Newman, P. A.↗
A study of the aerothermal environment for the Pioneer Venus multiprobe mission
In December 1978, four Pioneer Venus probe spacecraft are scheduled for almost simultaneous entry into the Venusian atmosphere at widely dispersed points about the planet. In this study, both detailed and approximate flow field analyses are used to define the entry aerothermal environment for the forebody of each of the four probes. The results show that approximate analyses can be used to predict inviscid radiative and laminar convective heating rates with acceptable accuracy. However, the radiative heating rates obtained with inviscid analyses are significantly greater than those obtained with a nonablating viscous-shock-layer (VSL) analysis, because the VSL analysis includes a strongly absorbing boundary layer. Also, the results show that the radiative heating is sensitive to small variations in atmospheric gas composition while the convective heating is not affected. With carbon-phenolic injection, the convective heating is reduced substantially while the overall radiative heating reduction is very small. Most of the radiative blockage occurs in the atomic line transitions which is significant only in the stagnation region.
Moss, J. N.↗
Vorticity interaction effects on blunt bodies
Numerical solutions of the viscous shock layer equations governing laminar and turbulent flows of a perfect gas and radiating and nonradiating mixtures of perfect gases in chemical equilibrium are presented for hypersonic flow over spherically blunted cones and hyperboloids. Turbulent properties are described in terms of the classical mixing length. Results are compared with boundary layer and inviscid flowfield solutions; agreement with inviscid flowfield data is satisfactory. Agreement with boundary layer solutions is good except in regions of strong vorticity interaction; in these flow regions, the viscous shock layer solutions appear to be more satisfactory than the boundary layer solutions. Boundary conditions suitable for hypersonic viscous shock layers are devised for an advanced turbulence theory.
Anderson, E. C.↗
Aerothermal environment for Jovian entry probes
Results are presented that describe the aerothermal environment encountered by the forebody of probes during Jovian entry. Two probe configurations are considered: a 55 deg half-angle spherically blunted cone and a 50 deg half-angle hyperboloid. The main emphasis is focused on defining the heating environment during entry and on determining the effectiveness of coupled ablation injection in reducing the surface heating rate. This is accomplished by calculating both the convective and radiative heating at several points along the entry trajectory both with and without coupled ablation injection. Results are presented for both laminar and turbulent calculations.
Moss, J. N.↗
Turbulent viscous-shock-layer solutions with strong vorticity interaction
Numerical solutions of the viscous-shock-layer equations governing laminar and turbulent flows of a perfect gas and radiating and nonradiating mixtures of perfect gases in chemical equilibrium are presented for hypersonic flow over spherically blunted cones and hyperboloids. The results are compared with boundary-layer and inviscid flow-field solutions. The agreement with the inviscid flow field data is satisfactory. The agreement with boundary-layer solutions is good except in regions of strong vorticity interaction. In these flow regions, the viscous-shock-layer solutions appear to be more satisfactory than the boundary-layer solutions.
Anderson, E. C.↗
Numerical solution of the steady-state Navier-Stokes equations for hypersonic flow about blunt axisymmetric bodies
The steady-state Navier-Stokes equations are solved for hypersonic flow about blunt axisymmetric bodies. The equations of motion are solved by successive approximations using an implicit finite-difference scheme. The results are compared with viscous shock-layer theory, experimental data, and time-dependent solutions of the Navier-Stokes equations. It is demonstrated that viscous shock-layer theory is sufficiently accurate for the range of flight conditions normally encountered by entry vehicles.
Anderson, E. C.↗
Viscous-shock-layer solutions for turbulent flow of radiating gas mixtures in chemical equilibrium
The viscous-shock-layer equations for hypersonic laminar and turbulent flows of radiating or nonradiating gas mixtures in chemical equilibrium are presented for two-dimensional and axially-symmetric flow fields. Solutions were obtained using an implicit finite-difference scheme and results are presented for hypersonic flow over spherically-blunted cone configurations at freestream conditions representative of entry into the atmosphere of Venus. These data are compared with solutions obtained using other methods of analysis.
Anderson, E. C.↗
Viscous shock layer solutions for turbulent flow of radiating gas mixtures in chemical equilibrium
The viscous shock layer equations for hypersonic laminar and turbulent flows of radiating or nonradiating gas mixtures in chemical equilibrium are presented for two-dimensional and axially symmetric flow fields. Solutions are obtained using an implicit finite difference scheme and results are presented for hypersonic flow over spherically blunted cone configurations at free stream conditions representative of entry into the atmosphere of Venus. These data are compared with solutions obtained using other methods of analysis.
Anderson, E. C.↗
Viscous-shock-layer solutions with radiation and ablation injection for Jovian entry
This paper presents results that describe the flow field and surface response conditions for a planetary probe entering the atmosphere of Jupiter. The results are obtained with a viscous-shock-layer analysis which includes mass injection, radiative transfer, diffusion, and viscous effects. Radiative transfer is calculated with an existing nongray radiation model that accounts for molecular band, atomic line, and continuum transitions. Particular emphasis is given to the effect of mass injection on the stagnation and downstream flow. Results for both specified and coupled mass injection are obtained for a carbon heat shield.
Moss, J. N.↗