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At least 271 records · Page 15

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NASA NTRS1986

Unsteady pressure measurements on a biconvex airfoil in a transonic oscillating cascade

Flush-mounted dynamic pressure transducers were installed on the center airfoil of a transonic oscillating cascade to measure the unsteady aerodynamic response as nine airfols were simultaneously driven to provide 1.2 deg of pitching motion about the midchord. Initial tests were performed at an incidence and angle of 0 deg and A Mach number of 0.65 in order to obtain results in a shock-free compressible flowfield. Subsequent tests were performed at an incidence angle of 7 deg and Mach number of 0.8 in order to observe the surface pressures with an oscillating shock near the leading edge of the airfoil. Results are presented for interblade phase angles of 90 and -90 deg and at blade oscillatory frequencies of 200 and 500 Hz (semi-chord reduced frequencies up to about 0.5 at a Mach number of 0.8). Results from the zero-incidence cascade are compared with a clasical usnteady flat-plate analysis. Flow visualization results depicting the shock motion on the airfoils in the high-incidence cascade are discussed. The airfoil pressure data are tabulated.

Shaw, L. M.↗

NASA NTRS1992

Three-dimensional shock-wave/boundary-layer interactions with bleed through a circular hole

Computations were performed to study three-dimensional (3-D), shock-wave/boundary-layer interactions on a flat plate in which fluid in the boundary layer was bled through a circular hole into a plenum to control shock-wave induced separation. Results are presented which show the details of the 3-D flowfield about the bleed hole and how bleed-hole placement relative to shock-wave impingement affect upstream, spanwise, and downstream influence lengths. This study revealed an underlying mechanisms by which bleed holes can affect shock-wave/boundary-layer interactions. This investigation is based on the ensemble-averaged, "full-compressible" Navier-Stokes equations closed by the Baldwin-Lomax turbulence model. Solutions to these equations were obtained by an implicit finite-volume method based on the partially-split, two-factored algorithm of Steger.

Rimlinger, M. J.↗

NASA NTRS1985

Upwind relaxation algorithms for the Navier Stokes equations

The development of upwind relaxation algorithms for obtaining efficient steady-state solutions to the compressible Navier-Stokes equations is described. The method is second-order accurate spatially and naturally disipative, using third-order flux splitting of the pressure and convective terms and second-order central differencing for shear and heat flux terms. A line Gauss-Seidel relaxation approach, shown to be unconditionally stable for model convection and diffusion equations, is used. The algorithm is demonstrated for several flows using the thin-layer form of the equations, including the problem of shock-induced separation over a flat plate.

Thomas, J. L.↗

NASA NTRS1976

An efficient numerical method for solving the time-dependent compressible Navier-Stokes equations at high Reynolds number

A fine-mesh method incorporating two new operators, which drastically reduces the computation time, has been developed for solving the time-dependent Navier-Stokes equations at flight Reynolds numbers. The approach time-splits the equations into a hyperbolic part and a parabolic part, solves the hyperbolic part by a new explicit numerical method based on characteristics theory, and solves the parabolic part by a new efficient implicit parabolic method. The method has reduced the computation time by one and two orders of magnitude from that required previously to solve for the interaction of a shock wave with a boundary layer on a flat plate.

Maccormack, R. W.↗

NASA NTRS1992

The Proteus Navier-Stokes code

An effort is currently underway at NASA Lewis to develop two and three dimensional Navier-Stokes codes, called Proteus, for aerospace propulsion applications. Proteus solves the Reynolds-averaged, unsteady, compressible Navier-Stokes equations in strong conservation law form. Turbulence is modeled using a Baldwin-Lomax based algebraic eddy viscosity model. In addition, options are available to solve thin layer or Euler equations, and to eliminate the energy equation by assuming constant stagnation enthalpy. An extensive series of validation cases have been run, primarily using the two dimensional planar/axisymmetric version of the code. Several flows were computed that have exact solution such as: fully developed channel and pipe flow; Couette flow with and without pressure gradients; unsteady Couette flow formation; flow near a suddenly accelerated flat plate; flow between concentric rotating cylinders; and flow near a rotating disk. The two dimensional version of the Proteus code has been released, and the three dimensional code is scheduled for release in late 1991.

Towne, Charles E.↗

NASA NTRS1995

The Physics of Turbulence in the Boundary Layer

The geometry of the velocity field in a numerically simulated incompressible turbulent boundary layer over a flat plate at Re theta=670 has been studied using the invariants of the velocity gradient tensor. These invariants are computed at every grid point in the flow and used to form the discriminant. Of primary interest are those regions in the flow where the discriminant is positive; regions where, according to the characteristic equation, the eigenvalues of the velocity gradient tensor are complex. An observer moving with a frame of reference which is attached to a fluid particle lying within such a region would see a local flow pattern of the type stable-focus-stretching or unstable-focus-compressing. When the flow is visualized this way, continuous, connected, large-scale structures are revealed that extend from the point just below the buffer layer out to the beginning of the wake region. These structures are aligned with the mean shear close to the wall and arch in the cross-stream direction away from the wall. In some cases the structures observed are very similar to to the hairpin eddy vision of boundary layer structure proposed by Theodorsen. That the structure of the flow is revealed more effectively by the discriminant rather than by the vorticity is important and adds support to recent observations of the discriminant in a channel flow simulation. Of particular importance is the fact that the procedure does not require the use of an arbitrary threshold in the discriminant. Further analysis using computer flow visualization shows a high degree of spatial correlation between regions of positive discriminant, extreme negative pressure fluctuations and large instantaneous values of Reynolds shear stress.

Kline, Stephen↗

NASA NTRS1971

Calculation of compressible nonadiabatic boundary layers in laminar, transitional and turbulent flow by the method of integral relations

A computer program was developed to do the calculations for two-dimensional or axisymmetric configurations from low speeds to hypersonic speeds with arbitrary streamwise pressure, temperature, and Mach number distributions. Options are provided for obtaining initial conditions either from experimental information or from a theoretical similarity solution. The transition region can be described either by an arbitrary distribution of intermittency or by a function based on Emmons' probability theory. Correlations were developed for use in estimating the parameters of the theoretical intermittency function. Correlations obtained from other sources are used for estimating the transition point. Comparisons were made between calculated and measured boundary layer quantities for laminar, transitional, and turbulent flows on flat plates, cones, cone flares, and a waisted body of revolution. Excellent agreement was obtained between the present theory and two other theories based on the method of finite differences. The intermittency required to reproduce some experimental heat transfer results in hypersonic flow was found to be quite different from the theoretical function. It is suggested that the simple probability theory of Emmons may not be valid for representing the intermittency of hypersonic transitional boundary layers and that the program could be useful as a tool for detailed study of the intermittency of the transition region.

Kuhn, G. D.↗

DOE OSTI2022

Verification and Validation Activities for the Multi-Fidelity Toolkit

The Multi-Fidelity Toolkit (MFTK) is a simulation tool being developed at Sandia National Laboratories for aerodynamic predictions of compressible flows over a range of physics fidelities and computational speeds. These models include the Reynolds-Averaged-Navier-Stokes (RANS) equations, the Euler equations, and modified Newtonian aerodynamics (MNA) equations, and they can be invoked independently or coupled with hierarchical Kriging to interpolate between high-fidelity simulations using lower-fidelity data. However, as with any new simulation capability, verification and validation are necessary to gather credibility evidence. This work describes formal code- and solution-verification activities as well as model validation with uncertainty considerations. Code verification is performed on the MNA model by comparing with an analytical solution for flat-plate and inclined-plate geometries. Solution-verification activities include grid-refinement studies of HIFiRE-1 wind tunnel measurements, which are used for validation, for all model fidelities. A thorough treatment of the validation comparison with prediction error and validation uncertainty is also presented.

42 ENGINEERING↗

NASA NTRS1972

A theoretical and experimental investigation of the hypersonic turbulent boundary layer subject to normal and longitudinal pressure gradients and cross flow along a windward plane of symmetry.

A theoretical and experimental investigation for the hypersonic turbulent boundary layer undergoing both normal and longitudinal pressure gradients and cross flow along the plane of symmetry on the windward side of a body of revolution is presented. The purpose of this investigation is to provide a method for calculating boundary layers that occur on the centerline of symmetry of an inlet. The three-dimensional compressible integral equations are written along the symmetry plane and integrated numerically, and the effect of the cross flow on the behavior of the boundary layer is also studied. The experiments were performed at Mach number 6 and Reynolds number of 44,000,000/ft for the axisymmetric case and for the case with cross flow. The velocity profiles and the boundary layer parameters agree relatively well with theoretical values with and without cross flow effects. The presence of the cross flow alters significantly the value of the profiles. Total temperature-velocity correlations deviate from the linear Crocco distribution. Heat transfer rates agree with theoretical values and values obtained by the flat plate reference enthalpy method. Cross flow influence on the heat transfer rates is of minor relevance.

Zakkay, V.↗

NASA NTRS1987

Navier-Stokes cascade analysis with a stiff Kappa-Epsilon turbulence solver

The two dimensional, compressible, thin layer Navier-Stokes equations with the Baldwin-Lomax turbulence model and the kinetic energy-energy dissipation (k-epsilon) model are solved numerically to simulate the flow through a cascade. The governing equations are solved for the entire flow domain, without the boundary layer assumptions. The stiffness of the k-epsilon equations is discussed. A semi-implicit, Runge-Kutta, time-marching scheme is developed to solve the k-epsilon equations. The impact of the k-epsilon solver on the explicit Runge-Kutta Navier-Stokes solver is discussed. Numerical solutions are presented for two dimensional turbulent flow over a flat plate and a double circular arc cascade and compared with experimental data.

Liu, Jong-Shang↗

NASA NTRS1988

Navier-Stokes cascade analysis with a stiff k-epsilon turbulence solver

Liu, Jong-Shang↗

NASA NTRS1990

A finite element solver for 3-D compressible viscous flows

Computation of the flow field inside a space shuttle main engine (SSME) requires the application of state of the art computational fluid dynamic (CFD) technology. Several computer codes are under development to solve 3-D flow through the hot gas manifold. Some algorithms were designed to solve the unsteady compressible Navier-Stokes equations, either by implicit or explicit factorization methods, using several hundred or thousands of time steps to reach a steady state solution. A new iterative algorithm is being developed for the solution of the implicit finite element equations without assembling global matrices. It is an efficient iteration scheme based on a modified nonlinear Gauss-Seidel iteration with symmetric sweeps. The algorithm is analyzed for a model equation and is shown to be unconditionally stable. Results from a series of test problems are presented. The finite element code was tested for couette flow, which is flow under a pressure gradient between two parallel plates in relative motion. Another problem that was solved is viscous laminar flow over a flat plate. The general 3-D finite element code was used to compute the flow in an axisymmetric turnaround duct at low Mach numbers.

Reddy, K. C.↗

DOE OSTI2022

Direct numerical simulation of hypersonic turbulent boundary layers: effect of spatial evolution and Reynolds number

Direct numerical simulations (DNS) are performed to investigate the spatial evolution of flat-plate zero-pressure-gradient turbulent boundary layers over long streamwise domains ( >300δ i , with δ i the inflow boundary-layer thickness) at three different Mach numbers, 2.5, 4.9 and 10.9, with the surface temperatures ranging from quasiadiabatic to highly cooled conditions. The settlement of turbulence statistics into a fully developed equilibrium state of the turbulent boundary layer has been carefully monitored, either based on the satisfaction of the von Kármán integral equation or by comparing runs with different inflow turbulence generation techniques. The generated DNS database is used to characterize the streamwise evolution of multiple important variables in the high-Mach-number, cold-wall regime, including the skin friction, the Reynolds analogy factor, the shape factor, the Reynolds stresses, and the fluctuating wall quantities. The data confirm the validity of many classic and newer compressibility transformations at moderately high Reynolds numbers (up to friction Reynolds number Re τ ≈1200 ) and show that, with proper scaling, the sizes of the near-wall streaks and superstructures are insensitive to the Mach number and wall cooling conditions. The strong wall cooling in the hypersonic cold-wall case is found to cause a significant increase in the size of the near-wall turbulence eddies (relative to the boundary-layer thickness), which leads to a reduced-scale separation between the large and small turbulence scales, and in turn to a lack of an outer peak in the spanwise spectra of the streamwise velocity in the logarithmic region.

71 CLASSICAL AND QUANTUM MECHANICS, GENERAL PHYSIC↗

NASA NTRS1959

Use of a Stanton Tube for Skin-Friction Measurements

A small total-pressure tube resting against a flat-plate surface was used as a Stanton tube and calibrated as a skin-friction meter at various subsonic and supersonic speeds. Laminar flow was maintained for the supersonic runs at a Mach number M(sub infinity) of 2. At speeds between M(sub infinity) = 1.33 and M(sub infinity) = 1.87, the calibrations were carried-out in a turbulent boundary layer. The subsonic flows were found to be in transition. The skin-friction readings of a floating-element type of balance served as the reference values against which the Stanton tube was calibrated. A theoretical model was developed which, for moderate values of the shear parameter tau, accurately predicts the performance of the Stanton tube in subsonic and supersonic flows. A "shear correction factor" was found to explain the deviations from the basic model when T became too large. Compressibility effects were important only in the case of turbulent supersonic flows, and they did not alter the form of the calibration curve. The test Reynolds numbers, based on the distance from the leading edge and free-stream conditions, ranged from 70,000 to 875,000. The turbulent-boundary-layer Reynolds numbers, based on momentum thickness, varied between 650 and 2,300. Both laminar and turbulent velocity profiles were taken and the effect of pressure gradient on the calibration was investigated.

Abarbanel, S. S.↗

NASA NTRS1992

A unified approach for numerical simulation of viscous compressible and incompressible flows over adiabatic and isothermal walls

A new formulation (including the choice of variables, their non-dimensionalization, and the form of the artificial viscosity) is proposed for the numerical solution of the full Navier-Stokes equations for compressible and incompressible flows with heat transfer. With the present approach, the same code can be used for constant as well as variable density flows. The changes of the density due to pressure and temperature variations are identified and it is shown that the low Mach number approximation is a special case. At zero Mach number, the density changes due to the temperature variation are accounted for, mainly through a body force term in the momentum equation. It is also shown that the Boussinesq approximation of the buoyancy effects in an incompressible flow is a special case. To demonstrate the new capability, three examples are tested. Flows in driven cavities with adiabatic and isothermal walls are simulated with the same code as well as incompressible and supersonic flows over a wall with and without a groove. Finally, viscous flow simulations of an oblique shock reflection from a flat plate are shown to be in good agreement with the solutions available in literature.

Hafez, M.↗

DOE OSTI2021

RapidClave® Technology Demonstrations - I

This demonstration sought to show a dramatic reduction in tool cost for RapidClave® tooling, taking advantage of recent modifications made on the RapidClave® system. Instead of building tools with complex internal heating, simple aluminum tools would be produced and heated externally through the recently added hot plate and air blower inside the RapidClave® system. Baselining against this tooling is the legacy RapidClave® tooling, as well as compression tooling, which would be the incumbent processing technique that the RapidClave® is displacing. Both an automotive and aerospace application were used to evaluate the tooling and process. Specifically, a four-piece Volkswagen hood structure was selected as an appropriate automotive geometry, while a wing structure representative of a current wing entering limited production was provided by Cornerstone Research Group for evaluating aerospace applications. In total, six tools were fabricated, and costs were compared against compression tooling quotes of the same tool geometry. A cost savings of over 80% was shown on all the tools. This demonstration also produced parts in the RapidClave® using the abovementioned tools to validate the low-cost tooling and the machine modifications. Solvay recommended two prepreg systems, CYCOM® EP2750 and SolvaLite™ 712, as appropriate for aerospace and automotive applications, respectively. Flat panels of each material were fabricated to verify cure parameters before proceeding with actual part fabrication. On-tool cycle times of approximately 30 minutes was achieved, with modifications identified to lower the time further.

36 MATERIALS SCIENCE↗

NASA NTRS1994

Innovative design of composite structures: Design, manufacturing, and testing of plates utilizing curvilinear fiber trajectories

As a means of improving structural design, the concept of fabricating flat plates containing holes by incorporating curvilinear fiber trajectories to transmit loads around the hole is studied. In the present discussion this concept is viewed from a structural level, where access holes, windows, doors, and other openings are of significant size. This is opposed to holes sized for mechanical fasteners. Instead of cutting the important load-bearing fibers at the hole edge, as a conventional straightline design does, the curvilinear design preserves the load-bearing fibers by orienting them in smooth trajectories around the holes, their loading not ending abruptly at the hole edge. Though the concept of curvilinear fiber trajectories has been studied before, attempts to manufacture and test such plates have been limited. This report describes a cooperative effort between Cincinnati Milacron Inc., NASA Langley Research Center, and Virginia Polytechnic Institute and State University to design, manufacture, and test plates using the curvilinear fiber trajectory concept. The paper discusses details of the plate design, details of the manufacturing, and a summary of results from testing the plates with inplane compressive buckling loads and tensile loads. Comparisons between the curvilinear and conventional straightline fiber designs based on measurements and observation are made. Failure modes, failure loads, strains, deflections, and other key responses are compared.

Hyer, M. W.↗

NASA NTRS2005

Apparatus for Testing Flat Specimens of Thermal Insulation

An apparatus has been developed to implement an improved method of testing flat-plate specimens of thermal-insulation materials for cryogenic application. The method includes testing under realistic use conditions that could include vacuum and mechanical loading at a pressure up to 70 psi (=0.48 MPa). The apparatus can accommodate a rigid or flexible specimen having thickness up to 1.25 in. (=3.2 cm) and diameters between 6 and 10 in. (about 15.2 and 25.4 cm, respectively). Typical test conditions include boundary temperatures between 77 K and 373 K and vacuum/interstitial gas filling at a pressure between 10(exp -6) torr (=1.3 x 10(exp -4) Pa) and 760 torr (atmospheric pressure =0.1 MPa). The interstitial gas could be N2, He, CO2, or any other suitable gas to which the insulation is expected to be exposed in use. Relative to prior apparatuses and testing methods, this apparatus and the testing method that it implements offer advantages of relative simplicity and ease of use. The basic principle of operation of the apparatus is that of boil-off calorimetry, using liquid nitrogen or any other suitable liquid that boils at a desired temperature below ambient temperature. Comparative rates of flow of heat through the thicknesses of the specimens (heat-leak rates) and apparent-thermal-conductivity values are obtained from tests of specimens. Absolute values of heat-leak rates and apparent thermal conductivities are computed from a combination of (1) the aforementioned comparative values and (2) calibration factors obtained by testing reference specimens of materials that have known thermal-insulation properties. The apparatus includes a full complement of temperature sensors, a vacuum pump and chamber, a monitoring and control system, and tools and fixtures that enable rapid and reliable installation and removal of specimens. A specimen is installed at the bottom of the vacuum chamber, and a cold-mass assembly that includes a tank is lowered into position above and around the specimen (see figure). A spring-based compensating fixture helps to ensure adequate thermal contact with possibly irregular specimen surfaces. For a high-compression test, the springs can be replaced with spacers. A flat circular load cell at the bottom of the chamber measures the compressive load on the specimen. Once the desired compressive-load, temperature, and vacuum/gas-filling conditions are established, testing begins. During a test, all measurements are recorded by use of a portable data-acquisition system and a computer. The total heat-leak rate is measured and calculated as the boil-off flow rate multiplied by the latent heat of vaporization. The parasitic heat leak (to the side of the specimen and to the top and side of the cold-mass tank) is reduced to a small fraction of the total heat leak by use of a combination of multilayer-insulation (MLI) shield rings, reflective film, a fiberglass/epoxy centering ring, and a bulk fill of aerogel beads. This combination eliminates the need for a cryogenic guard chamber used in a typical prior apparatus to reduce the parasitic heat leak.

Fesmire, James E.↗