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At least 1,693 records · Page 94
All sourcesNASA NTRSDOE OSTI
NASA NTRS1992

An inverse method for the calculation of through-thickness fatigue crack closure behavior

An inverse technique was used to calculate through-thickness fatigue crack closure behavior. The through-thickness variation in crack opening stress-intensity factor was calculated by considering the variation in the three-dimensional stress-intensity factor, the variation in crack growth rate along the crack front, and a relationship between the crack growth rate and effective stress-intensity factor range (da/dN-Delta-K(eff)). The three-dimensional stress-intensity factor variation was obtained from an elastic finite element analysis of specific crack-front profiles observed experimentally. The variation in crack growth rate along the crack front was obtained experimentally from comparison of observed crack front changes. The da/dN-Delta-K(eff) relationship was estimated from high stress ratio, constant load amplitude, and fatigue crack growth tests. The through-thickness crack opening stress-intensity factor results agreed with crack opening measurements obtained from fatigue striations, near-tip strain gages, and remote strain and displacement gages.

Dawicke, D. S.↗
NASA NTRS1993

Numerical calculations of the linear response of a gaseous disk to a protoplanet

Numerical calculations of the linear response of a 2D gaseous disk to the perturbations induced by a protoplanet and the corresponding torque are presented. When the pressure gradient is taken into account, torques are increased in disks with gradients in either surface density and sound speed, the effect of the latter being much greater for the same-sized gradient as measured by the power law index. The torques in turn may be used to calculate timescales for orbital migration of protoplanets.

Korycansky, D. G.↗
NASA NTRS1991

Calculations of separated 3-D flows with a pressure-staggered Navier-Stokes equations solver

A Navier-Stokes equations solver based on a pressure correction method with a pressure-staggered mesh and calculations of separated three-dimensional flows are presented. It is shown that the velocity pressure decoupling, which occurs when various pressure correction algorithms are used for pressure-staggered meshes, is caused by the ill-conditioned discrete pressure correction equation. The use of a partial differential equation for the incremental pressure eliminates the velocity pressure decoupling mechanism by itself and yields accurate numerical results. Example flows considered are a three-dimensional lid driven cavity flow and a laminar flow through a 90 degree bend square duct. For the lid driven cavity flow, the present numerical results compare more favorably with the measured data than those obtained using a formally third order accurate quadratic upwind interpolation scheme. For the curved duct flow, the present numerical method yields a grid independent solution with a very small number of grid points. The calculated velocity profiles are in good agreement with the measured data.

Kim, S.-W.↗
NASA NTRS1993

Experiment-specific cosmic microwave background calculations made easier - Approximation formula for smoothed delta T/T windows

Simple and easy to implement elementary function approximations are introduced to the spectral window functions needed in calculations of model predictions of the cosmic microwave backgrond (CMB) anisotropy. These approximations allow the investigator to obtain model delta T/T predictions in terms of single integrals over the power spectrum of cosmological perturbations and to avoid the necessity of performing the additional integrations. The high accuracy of these approximations is demonstrated here for the CDM theory-based calculations of the expected delta T/T signal in several experiments searching for the CMB anisotropy.

Gorski, Krzysztof M.↗
NASA NTRS1993

Time-accurate Navier-Stokes calculations with multigrid acceleration

An efficient method for calculating unsteady flows is presented, with emphasis on a modified version of the thin-layer Navier-Stokes equations. Fourier stability analysis is used to illustrate the effect of treating the source term implicitly instead of explicity, as well as to illustrate other algorithmic choices. A 2D circular cylinder (with a Reynolds number of 1200 and a Mach number of 0.3) is calculated. The present scheme requires only about 10 percent of the computer time required by global minimum time stepping.

Melson, N. D.↗
NASA NTRS1993

An improved far field drag calculation method for nonlinear CFD codes

An improved method to calculate drag based on far field conditions is presented and demonstrated. The method is illustrated using two CFD codes, PARC and CFL3D, and examining their ability to preserve lift and drag coefficients along grid line defined integration paths. The flow fields were generated by solving the Euler equations for a NACA 0012 airfoil at a free stream Mach number of 0.8 and angle of attack of 1.25 degrees. In comparison to force coefficients obtained by surface pressure integration, neither code acceptably preserved both force coefficients throughout the near and far fields. An investigation into the relationship between numerical prediction error and calculated force coefficients revealed a direct connection between solution mass conservation error and force coefficients error. A method to correct the predicted force coefficients based on integrated mass conservation error is described and demonstrated. The corrected force coefficients for both codes are shown to be more accurate than the uncorrected values. The correction method is applicable to both two and three dimensions and is independent of the algorithm used to generate the flow field.

Sirbaugh, James R.↗
NASA NTRS1993

Installed F/A-18 inlet flow calculations at 60 deg angle-of-attack and 10 deg side slip

This paper presents the results of PARC3D numerical calculations on a 19.78 percent scale forebody/inlet model of the F/A-18 at a Mach number of 0.20, an angle-of-attack of 60 deg, and a side-slip angle of 10 deg. The main purpose of these calculations is to support an upcoming wind-tunnel test program in the prediction of engine inlet compressor face total pressure recovery and flow distortion. The GRIDGEN system was used to generate a grid which includes the inlet and lip, and other aircraft components which are considered to be important to inlet performance, such as the ramp/splitter plate, the diverter and slot, and the deflected leading edge flap. PARC3D shows complex flow patterns on the fuselage surfaces below the leading edge extensions, on the ramp/splitter plate, inlet lip, and inside the inlet. PARC3D tends to underpredict total pressure recovery and overpredict the flow distortion at the inlet compressor face.

Podleski, S. D.↗
NASA NTRS1993

A simplified reaction mechanism for calculation of emissions in hydrocarbon (Jet-A) combustion

The paper presents a simplified reaction mechanism developed for use in calculations of NO(x) emissions in Jet-A combustion. The rate for the N2 + O2 reaction was selected to match available experimental results, and the reaction rates for other global reactions in the Jet-A are empirical reaction rates adjusted to match species concentrations using the detailed mechanism of Miller and Bowman (1989). The mechanism was validated by comparing the emission of NO(x) calculated using propane as fuel.

Kundu, K. P.↗
NASA NTRS1993

Backscattering by nonspherical hydrometeors as calculated by the coupled-dipole method - An application in radar meteorology

The severest test of a theory of scattering by particles is how well it calculates scattering in the backward direction. The coupled-dipole method can be used for accurately calculating backscattering at 94 GHz by hexagonal ice crystals. Backscattering by columns is markedly different from that by plates, which indicates that it might be possible to infer size and shape distributions of ice crystals using recently developed millimeter wave radar.

Dungey, Clifton E.↗
NASA NTRS1993

Reynolds stress calculations of homogeneous turbulent shear flow with bounded energy states

Reynolds stress calculations of homogeneous turbulent shear flow are conducted with a second-order closure model modified to account for nonequilibrium vortex stretching in the dissipation rate transport equation as recently proposed by Bernard and Speziale (1992). As with the earlier reported K-epsilon model calculations incorporating this vortex stretching effect, a production-equals-dissipation equilibrium is obtained with bounded turbulent kinetic energy and dissipation. However, this equilibrium is now not achieved until the dimensionless time St greater than 60 - an elapsed time that is at least twice as large as any of those considered in previous numerical and physical experiments on homogeneous shear flow. Direct quantitative comparisons between the model predictions and the results of experiments are quite favorable. In particular, it is shown that the inclusion of this nonequilibrium vortex stretching effect has the capability of explaining the significant range of production to dissipation ratios observed in experiments.

Speziale, Charles G.↗
NASA NTRS1934

A method of calculating the performance of controllable propellers with sample computations

This paper contains a series of calculations showing how the performance of controllable propellers may be derived from data on fixed-pitch propellers given in N.A.C.A. Technical Report No. 350, or from similar data. Sample calculations are given which compare the performance of airplanes with fixed-pitch and with controllable propellers. The gain in performance with controllable propellers is shown to be largely due to the increased power available, rather than to an increase in efficiency. Controllable propellers are of particular advantage when used with geared and with supercharged engines. A controllable propeller reduces the take-off run, increases the rate of climb and the ceiling, but does not increase the high speed, except when operating above the design altitude of the previously used fixed-pitch propeller or when that propeller was designed for other than high speed.

Hartman, Edwin P↗
NASA NTRS1939

A simplified method for the calculation of airfoil pressure distribution

A method is presented for the rapid calculation of the pressure distribution over an airfoil section when the normal-force distribution and the pressure distribution over the "base profile" (i.e., the profile of the same airfoil were the camber line straight and the resulting airfoil at zero angle of attack) are known. This note is intended as a supplement to N.A.C.A. Report Nos. 631 and 634 wherein methods are presented for the calculation of the normal-force distribution over plain and flapped airfoils, respectively, but not of the pressures on the individual surfaces. Base-profile pressure-coefficient distributions for the usual N.A.C.A. family of airfoils, which are also suitable for several other commonly employed airfoils, are included in tabular form. With these tabulated base-profile pressures and the computed normal-force distributions, pressure distributions adequate for most engineering purposes can be obtained.

Allen, H Julian↗
NASA NTRS1941

The calculation of span LCAD distributions of swept-back wings

Span load distributions of swept-back wings have been calculated. The method used was to replace the wing with a bound vortex at the quarter-chord line and to calculate the downwash due to the system of bound and trailing vortices to conform at the three-quarter-chord line to the slope of the flat-plate wing surface. Results are given for constant-chord and 5:1 tapered plan forms, for sweep-back angles of 0 degrees, 30 degrees, and 45 degrees, and for aspect ratios of 3, 6, and 9. Some comments on the stalling of swept-back wings are included.

Mutterperl, William↗
NASA NTRS1947

Experimental and calculated characteristics of three wings of NACA 64-210 and 65-210 airfoil sections with and without 2 degree washout

Report presents the results of an investigation conducted to determine some of the effects of airfoil section and washout on the experimental and calculated characteristics of 10-percent-thick wings. Three wings of aspect ratio 9 and ratio of root chord to tip chord 2.5 were tested. One wing had NACA 64-210 sections and 2 degree washout, the second had NACA 65-210 sections and 2 degree washout, and the third had NACA 65-210 sections and 0 degree washout. It was found that the experimental characteristics of the wings could be satisfactorily predicted from calculations based upon two-dimensional data when the airfoil contours of the wings conformed to the true airfoil sections with the same high degree of accuracy as the two-dimensional models.

AIRFOILS - NACA 65-210 - CHAR & MOMENTS LOW-DRAG↗
NASA NTRS1947

Calculation of Compressible Flows past Aerodynamic Shapes by Use of the Streamline Curvature

A simple approximate method is given for the calculation of isentropic irrotational flows past symmetrical airfoils, including mixed subsonic-supersonic flows. The method is based on the choice of suitable values for the streamline curvature in the flow field and the subsequent integration of the equations of motion. The method yields limiting solutions for potential flow. The effect of circulation is considered. A comparison of derived velocity distributions with existing results that are based on calculation to the third order in the thickness ratio indicated satisfactory agreement. The results are also presented in the form of a set of compressibility correction rules that lie between the Prandtl-Glauert rule and the von Karman-Tsien rule (approximately). The different rules correspond to different values of the local shape parameter square root sign YC sub a, in which Y is the ordinate and C sub a is the curvature at a point on an airfoil. Bodies of revolution, completely supersonic flows, and the significance of the limiting solutions for potential flow are also briefly discussed.

Perl, W↗
NASA NTRS1953

Method for calculating the rolling and yawing moments due to rolling for unswept wings with or without flaps or ailerons by use of nonlinear section lift data

The methods of NACA Reports 865 and 1090 have been applied to the calculation of the rolling- and yawing-moment coefficients due to rolling for unswept wings with or without flaps or ailerons. The methods allow the use of nonlinear section lift data together with lifting-line theory. Two calculated examples are presented in simplified computing forms in order to illustrate the procedures involved.

CONTROLS - COMPLETE WINGS↗
NASA NTRS1956

Calculation of the forces and moments on a slender fuselage and vertical fin penetrating lateral gusts

A theory is presented for calculating the variation with frequency of the lateral-force and yawing-moment coefficients due to sinusoidal side gusts passing over the profile of a simple fuselage-vertical-fin combination. The analysis is based on slender-body theory. The method considers the penetration effect of both fuselage and vertical tail in calculating side force and yawing moment due to side gusts, as opposed to a simple lag concept which considers the flow angle to be uniform over the configuration.

STABILITY, LATERAL AND DIRECTIONAL - DYNAMIC↗
NASA NTRS1956

A method for calculation of free-space sound pressures near a propeller in flight including considerations of the chordwise blade loading

This report presents tabulated values of certain definite integral that are involved in the calculation of near-field propeller noise when the chordwise forces are assumed to be either uniform or of a Dirac delta type. The tabulations are over a wide range of operating conditions and are useful for estimating propeller noise when either the concept of an effective radius or radial distributions of forces are considered. Use of the tabulations is illustrated by several examples of calculated results for some specific propellers.

NOISE↗