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NASA NTRS1980

Application of optical methods to the study of jet noise and turbulence

Optical methods are generally applied in fluid mechanics for either visualization or measurement. The use of a laser anemometer to study flow velocity in a jet and its relaton to the sound radiated is described. The same acoustic emission phenomenon is also measued by combining the signals from four Schlieren systems with that from an interferometer. The use of an optical Fourier transformation approach with real time analysis to determine the spatio-temporal structure of a field of mass volume such as a waveguide or free jet is also examined.

Sava, P. G.↗

NASA NTRS2000

Photogrammetry Applied to Wind Tunnel Testing

In image-based measurements, quantitative image data must be mapped to three-dimensional object space. Analytical photogrammetric methods, which may be used to accomplish this task, are discussed from the viewpoint of experimental fluid dynamicists. The Direct Linear Transformation (DLT) for camera calibration, used in pressure sensitive paint, is summarized. An optimization method for camera calibration is developed that can be used to determine the camera calibration parameters, including those describing lens distortion, from a single image. Combined with the DLT method, this method allows a rapid and comprehensive in-situ camera calibration and therefore is particularly useful for quantitative flow visualization and other measurements such as model attitude and deformation in production wind tunnels. The paper also includes a brief description of typical photogrammetric applications to temperature- and pressure-sensitive paint measurements and model deformation measurements in wind tunnels.

Liu, Tian-Shu↗

NASA NTRS1971

System distortion errors in radiorespirometry curves.

Description of a method for evaluating the maximum error producible in the system distortion of a carrier gas type radiorespirometer. The maximum producible error is evaluated on the basis of a second order. The maximum producible decrease in peak height output is calculated as a function of the system's time constant or flow rate by using the system's transfer function and the Laplace transformation equation (Gardner and Barnes, 1961).

Neville, E. D.↗

NASA NTRS1981

Simulation of three-dimensional incompressible flows with a vortex-in-cell method

A new method for the numerical simulation of three-dimensional incompressible flows is described. The vortex-in-cell (VIC) method presented traces the motion of the vortex filaments in the velocity field which these filaments create. The velocity field is not calculated directly by the Biot-Savart law of interaction but by creating a mesh-record of the vorticity field, then integrating a Poisson's equation via the fast Fourier transform to generate a mesh-record of the velocity field. The computed scales of motion are assumed to be essentially inviscid. Viscous of subgrid-scale effects are incorporated into a filtering procedure in wave vector space. Results of tracing a periodic array of single vortex rings are compared with a Green's function calculation. The agreement is very good.

Couet, B.↗

NASA NTRS1990

Computation of three-dimensional compressible boundary layers to fourth-order accuracy on wings and fuselages

A solution method, fourth-order accurate in the body-normal direction and second-order accurate in the stream surface directions, to solve the compressible 3-D boundary layer equations is presented. The transformation used, the discretization details, and the solution procedure are described. Ten validation cases of varying complexity are presented and results of calculation given. The results range from subsonic flow to supersonic flow and involve 2-D or 3-D geometries. Applications to laminar flow past wing and fuselage-type bodies are discussed. An interface procedure is used to solve the surface Euler equations with the inviscid flow pressure field as the input to assure accurate boundary conditions at the boundary layer edge. Complete details of the computer program used and information necessary to run each of the test cases are given in the Appendix.

Iyer, Venkit↗

NASA NTRSDate not supplied

A Design Exploration of Natural Laminar Flow Applications for the SUSAN Electrofan Concept

A trade study supported by the NASA Convergent Aeronautics Solutions (CAS) Project is presently underway to explore the desirability, feasibility, and viability of a new vehicle concept. The vehicle under development is the SUbsonic Single Aft eNgine (SUSAN) Electrofan configuration, which is a subsonic regional jet with the transformative concept of combining wing-mounted distributed electrified aircraft propulsion with a single engine design. The multidisciplinary team is investigating several potential technologies that would help the SUSAN Electrofan configuration achieve its performance goals, one of which is the application of Natural Laminar Flow (NLF) to the main wing. The present computational study utilizes the Crossflow Attenuated Natural Laminar Flow (CATNLF) design method to reshape the wing airfoils to obtain significant extents of NLF at the cruise condition. Fully-turbulent and laminar designs were performed to quantify the aerodynamic performance potential of NLF on the SUSAN Electrofan configuration. The laminar design supported laminar flow on53% of the area of the wing upper surface, resulting in a 19count (8.8%) drag reduction for the wing-fuselage configuration. The near-cruise off-design characteristics are also studied and illustrate that the robust laminar design sustains an NLF performance benefit through perturbations in Mach and angle of attack. A discussion is also included on multidisciplinary implications of utilizing NLF for the configuration, including design, manufacturing, and operational considerations required to support NLF on the main wing, as well as the interaction between the potential wing-mounted boundary layer ingestion technology that is also being explored for the SUSAN Electrofan configuration.

Natural Laminar Flow↗

NASA NTRS1979

Applications of Laplace transform methods to airfoil motion and stability calculations

This paper reviews the development of generalized unsteady aerodynamic theory and presents a derivation of the generalized Possio integral equation. Numerical calculations resolve questions concerning subsonic indicial lift functions and demonstrate the generation of Kutta waves at high values of reduced frequency, subsonic Mach number, or both. The use of rational function approximations of unsteady aerodynamic loads in aeroelastic stability calculations is reviewed, and a reformulation of the matrix Pade approximation technique is given. Numerical examples of flutter boundary calculations for a wing which is to be flight tested are given. Finally, a simplified aerodynamic model of transonic flow is used to study the stability of an airfoil exposed to supersonic and subsonic flow regions.

Edwards, J. W.↗

NASA NTRS1979

Measurement of vortex frequencies in a lean, premixed prevaporized combustor

A modified laser schlieren and fast Fourier transform analysis system has been developed and applied to the study of vortex frequencies in a lean, premixed prevaporized combustor. The system is sensitive to the passage of vortex structures which dominate the flow in the combustor studied. The frequency spectra of the passage of vortices show distinct frequency peaks which are postulated to be characteristic of the processes triggering the vortex shedding. The method provides additional information which should aid in the characterization and understanding of shear flows.

Parker, L. J.↗

NASA NTRS1991

Compressible flows with periodic vortical disturbances around lifting airfoils

A numerical method is developed for solving periodic, three-dimensional, vortical flows around lifting airfoils in subsonic flow. The first-order method that is presented fully accounts for the distortion effects of the nonuniform mean flow on the convected upstream vortical disturbances. The unsteady velocity is split into a vortical component which is a known function of the upstream flow conditions and the Lagrangian coordinates of the mean flow, and an irrotational field whose potential satisfies a nonconstant-coefficient, inhomogeneous, convective wave equation. Using an elliptic coordinate transformation, the unsteady boundary value problem is solved in the frequency domain on grids which are determined as a function of the Mach number and reduced frequency. The numerical scheme is validated through extensive comparisons with known solutions to unsteady vortical flow problems. In general, it is seen that the agreement between the numerical and analytical results is very good for reduced frequencies ranging from 0 to 4, and for Mach numbers ranging from .1 to .8. Numerical results are also presented for a wide variety of flow configurations for the purpose of determining the effects of airfoil thickness, angle of attack, camber, and Mach number on the unsteady lift and moment of airfoils subjected to periodic vortical gusts. It is seen that each of these parameters can have a significant effect on the unsteady airfoil response to the incident disturbances, and that the effect depends strongly upon the reduced frequency and the dimensionality of the gust. For a one-dimensional (transverse) or two-dimensional (transverse and longitudinal) gust, the results indicate that airfoil thickness increases the unsteady lift and moment at the low reduced frequencies but decreases it at the high reduced frequencies. The results show that an increase in airfoil Mach number leads to a significant increase in the unsteady lift and moment for the low reduced frequencies, but a significant decrease for the high reduced frequencies.

Scott, James R.↗

NASA NTRS1991

A finite-difference, frequency-domain numerical scheme for the solution of the linearized unsteady Euler equations

A numerical method is developed for solving periodic, three-dimensional, vortical flows around lifting airfoils in subsonic flow. The first-order method, that is presented, fully accounts for the distortion effects of the nonuniform mean flow on the convected upstream vortical disturbances. The unsteady velocity is split into a vortical component which is a known function of the upstream flow conditions and the Lagrangian coordinates of the mean flow, and an irrotational field whose potential satisfies a nonconstant-coefficient, inhomogeneous, convective wave equation. Using an elliptic coordinate transformation, the unsteady boundary value problem is solved in the frequency domain on grids which are determined as a function of the Mach number and reduced frequency. Extensive comparisons are made with known solutions to unsteady vortical flow problems, and it is seen that the agreement is generally very good for reduced frequencies ranging from 0 up to 4.

Scott, James R.↗

NASA NTRS1985

A tomographic technique for aerodynamics at transonic speeds

Computer aided tomography (CAT) provides a means of noninvasively measuring the air density distribution around an aerodynamic model. This technique is global in that a large portion of the flow field can be measured. A test of the applicability of CAT to transonic velocities was studied. A hemispherical-nose cylinder afterbody model was tested at a Mach number of 0.8 with a new laser holographic interferometer at the 2- by 2-Foot Transonic Wind Tunnel. Holograms of the flow field were taken and were reconstructed into interferograms. The fringe distribution (a measure of the local densities) was digitized for subsequent data reduction. A computer program based on the Fourier-transform technique was developed to convert the fringe distribution into three-dimensional densities around the model. Theoretical aerodynamic densities were calculated for evaluating and assessing the accuracy of the data obtained from the tomographic method.

Lee, G.↗

NASA NTRS1985

A tomographic technique for aerodynamics at transonic speeds

Lee, G.↗

NASA NTRS1974

A modified Newton-Raphson analysis of flight measurements of the trailing vortices of a heavy jet transport

A modified Newton-Raphson method has been used to determine the parameters in equations describing the vortex flow to obtain a best match with flight measurements of the flow behind the C-5A airplane. The flight measurements were made using a specially instrumented T-33 airplane which passed as closely as possible to the centers of the trailing vortices at several distances behind the C-5A airplane. The flight measurements were transformed to flow velocity relative to an inertial frame of reference. The assumed form of the flow consisted of the superposition of two counterrotating, finite core vortices. The positions of the vortex centers, their total circulation, the effective eddy viscosity and measurement bias were the parameters adjusted. The assumed form of vortex flow fit well the measured velocities for the numerous sets of data, both flaps up and down for the C-5A airplane. The resulting values of total circulation, however, were about two-thirds that expected of a wing with an elliptical loading. A partial explanation of the less than expected circulation is a dip in the spanwise lift distribution at the airplane's center line. The distance between the trailing vortices at the smallest times encountered is somewhat less than that expected for an elliptical wing loading.

Taylor, L. W., Jr.↗

NASA NTRS1978

Numerical solution of the Navier-Stokes equations for super-sonic flows with strong shocks

The numerical solution of the full Navier-Stokes Equations for viscous flows with high Mach numbers and a strong detached bow shock was obtained. Two dimensional flows around a circular cylinder, and a circular cylinder with an aft-body in the form of a fairing, were considered. The solution of the compressible N.S. equations was accomplished by the method of finite differences. An implicit scheme of solution, the S.O.R., was used with the optimum acceleration parameters determined by trial and error. The tensor notation was used in writing the N-S Equations transformed into general curvilinear coordinates. The equations for the generation of the coordinate system were solved, followed by the solution of the N.S. equations, at the end of a set of given number of time steps. "Wiggles", constituted the one major problem that needed to be overcome. These oscillations give rise to quantities such as negative temperatures, which ultimately caused the computational program to break down. Certain dissipative finite-difference schemes damped these oscillations.

Devarayalu, K.↗

NASA NTRS2012

Battery system and method for sensing and balancing the charge state of battery cells

A battery system utilizes a plurality of transformers interconnected with the battery cells. The transformers each have at least one transformer core operable for magnetization in at least a first magnetic state with a magnetic flux in a first direction and a second magnetic state with a magnetic flux in a second direction. The transformer cores retain the first magnetic state and the second magnetic state without current flow through said plurality of transformers. Circuitry is utilized for switching a selected transformer core between the first and second magnetic states to sense voltage and/or balance particular cells or particular banks of cells.

Davies, Francis J.↗

NASA NTRS1987

Transition and mixing in axisymmetric jets and vortex rings

A class of impulsively started, axisymmetric, laminar jets produced by a time dependent joint source of momentum are considered. These jets are different flows, each initially at rest in an unbounded fluid. The study is conducted at three levels of detail. First, a generalized set of analytic creeping flow solutions are derived with a method of flow classification. Second, from this set, three specific creeping flow solutions are studied in detail: the vortex ring, the round jet, and the ramp jet. This study involves derivation of vorticity, stream function, entrainment diagrams, and evolution of time lines through computer animation. From entrainment diagrams, critical points are derived and analyzed. The flow geometry is dictated by the properties and location of critical points which undergo bifurcation and topological transformation (a form of transition) with changing Reynolds number. Transition Reynolds numbers were calculated. A state space trajectory was derived describing the topological behavior of these critical points. This state space derivation yielded three states of motion which are universal for all axisymmetric jets. Third, the axisymmetric round jet is solved numerically using the unsteady laminar Navier Stokes equations. These equations were shown to be self similar for the round jet. Numerical calculations were performed up to a Reynolds number of 30 for a 60x60 point mesh. Animations generated from numerical solution showed each of the three states of motion for the round jet, including the Re = 30 case.

Allen, G. A., Jr.↗

NASA NTRS1978

The practical application of a finite difference method for analyzing transonic flow over oscillating airfoils and wings

Separating the velocity potential into steady and unsteady parts and linearizing the resulting unsteady equations for small disturbances was performed. The steady velocity potential was obtained first from the well known nonlinear equation for steady transonic flow. The unsteady velocity potential was then obtained from a linear differential equation in complex form with spatially varying coefficients. Since sinusoidal motion is assumed, the unsteady equation is independent of time. The results of an investigation into the relaxation-solution-instability problem was discussed. Concepts examined include variations in outer boundary conditions, a coordinate transformation so that the boundary condition at infinity may be applied to the outer boundaries of the finite difference region, and overlapping subregions. The general conclusion was that only a full direct solution in which all unknowns are obtained at the same time will avoid the solution instabilities of relaxation. An analysis of the one-dimensional form of the unsteady transonic equation was studied to evaluate errors between exact and finite difference solutions. Pressure distributions were presented for a low-aspect-ratio clipped delta wing at Mach number of 0.9 and for a moderate-aspect-ratio rectangular wing at a Mach number of 0.875.

Weatherill, W. H.↗