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At least 667 records · Page 37
All sourcesNASA NTRSDOE OSTI
NASA NTRS1994

Measurements of trapped protons and cosmic rays from recent shuttle flights

We have flown two new charged particle detectors in five recent Shuttle flights. In this paper we report on the dose rate, equivalent dose rate, and radiation quality factor for trapped protons and cosmic radiation separately. A comparison of the integral Linear Energy Transfer (LET) spectra with recent transport code calculations show significant disagreement. Using the calculated dose rate from the omni-directional AP8MAX model with IGRF reference magnetic field epoch 1970, and observed dose rate as a function of (averaged over all geographic latitude) and longitude, we have determined the westward drift of the South Atlantic anomaly. We have also studied the east-west effect, and observed a 'second' radiation belt. A comparison of the galactic cosmic radiation lineal energy transfer spectra with model calculations shows disagreement comparable to those of the trapped protons.

Badhwar, G. D.↗
NASA NTRS1979

Initial comparison of single cylinder Stirling engine computer model predictions with test results

A NASA developed digital computer code for a Stirling engine, modelling the performance of a single cylinder rhombic drive ground performance unit (GPU), is presented and its predictions are compared to test results. The GPU engine incorporates eight regenerator/cooler units and the engine working space is modelled by thirteen control volumes. The model calculates indicated power and efficiency for a given engine speed, mean pressure, heater and expansion space metal temperatures and cooler water inlet temperature and flow rate. Comparison of predicted and observed powers implies that the reference pressure drop calculations underestimate actual pressure drop, possibly due to oil contamination in the regenerator/cooler units, methane contamination in the working gas or the underestimation of mechanical loss. For a working gas of hydrogen, the predicted values of brake power are from 0 to 6% higher than experimental values, and brake efficiency is 6 to 16% higher, while for helium the predicted brake power and efficiency are 2 to 15% higher than the experimental.

Tew, R. C., Jr.↗
NASA NTRS1981

Linear/nonlinear behavior in unsteady transonic aerodynamics

The accurate calculation of the aerodynamic forces in unsteady transonic flow requires the solution of the nonlinear flow equations. The aeroelastician, on the other hand, seeks to treat his problems (flutter, for example) by means of linear equations whenever possible. He may do this, even when the underlying flow is nonlinear, if the perturbation forces are linear over some (perhaps small) range of unsteady amplitude of motion. This paper assesses the range of parameters over which linear behavior occurs. In particular calculations are made for an NACA 64A006 airfoil oscillating in pitch over a range of amplitudes, frequencies, and Mach numbers. The primary aerodynamic method used is the well known LTRAN2 code of Ballhaus and Goorjian that provides a finite-difference solution to the low frequency, small disturbance, two-dimensional potential flow equation. Comparisons are made with linear subsonic theory, local linearization, and, for steady flow, with the full potential equation code of Bauer, Garabedian, and Korn.

Dowell, E. H.↗
NASA NTRS1982

Prediction of sound radiation from different practical jet engine inlets

Computer codes which were based upon a special integral representation of the external solutions of the Helmholtz equation were upgraded so that they would yield accurate results for the acoustic radiation patterns in the field surrounding an axisymmetric body for nondimensional wave numbers, based on duct radius, of up to twenty. The accuracy of these computer programs was checked by the use of the point source method for the generation of exact solutions and then by comparison with the results of other experimental and theoretical studies. These computer codes were used in a parametric study of the dependence of the radiated sound field on input modal distribution, wave number, and inlet lip shape. The results of this study show that: (1) as the wave number (i.e., cut off ratio) is increased for a given input modal distribution that the acoustic radiation peak moves towards the inlet centerline and becomes more compact (i.e., narrows); (2) as the input mode number is increased for a given cut off ratio the acoustic radiation peak moves away from the inlet centerline and becomes more compact.

Meyer, W. L.↗
NASA NTRS1985

Non-isentropic unsteady transonic small disturbance theory

Modifications to transonic small disturbance theory (TSD) which more accurately model the Euler equations and seem to remove the problem of nonunique potential flow solutions are presented. The modifications are implemented in the two-dimensional computer code XTRAN2L, and steady and unsteady flow calculations made for the NACA 0012, NLR 7301, and NACA 64A010A airfoils. Comparisons are made with unmodified and modified TSD, Euler, and full potential theories and with experimental data. The modified theory requires only minor coding changes in existing algorithms for calculating small disturbance flows, and results in relatively small increases in computational cost.

Fuglsang, D. F.↗
NASA NTRS1978

Prediction of nearfield jet entrainment by an interactive mixing/afterburning model

The development of a computational model (BOAT) for calculating nearfield jet entrainment, and its application to the prediction of nozzle boattail pressures, is discussed. BOAT accounts for the detailed turbulence and thermochemical processes occurring in the nearfield shear layers of jet engine (and rocket) exhaust plumes while interfacing with the inviscid exhaust and external flowfield regions in an overlaid, interactive manner. The ability of the model to analyze simple free shear flows is assessed by detailed comparisons with fundamental laboratory data. The overlaid methodology and the entrainment correction employed to yield the effective plume boundary conditions are assessed via application of BOAT in conjunction with the codes comprising the NASA/LRC patched viscous/inviscid model for determining nozzle boattail drag for subsonic/transonic external flows. Comparisons between the predictions and data on underexpanded laboratory cold air jets are presented.

Dash, S. M.↗
NASA NTRS1997

A Three-Dimensional Linearized Unsteady Euler Analysis for Turbomachinery Blade Rows

A three-dimensional, linearized, Euler analysis is being developed to provide an efficient unsteady aerodynamic analysis that can be used to predict the aeroelastic and aeroacoustic responses of axial-flow turbo-machinery blading.The field equations and boundary conditions needed to describe nonlinear and linearized inviscid unsteady flows through a blade row operating within a cylindrical annular duct are presented. A numerical model for linearized inviscid unsteady flows, which couples a near-field, implicit, wave-split, finite volume analysis to a far-field eigenanalysis, is also described. The linearized aerodynamic and numerical models have been implemented into a three-dimensional linearized unsteady flow code, called LINFLUX. This code has been applied to selected, benchmark, unsteady, subsonic flows to establish its accuracy and to demonstrate its current capabilities. The unsteady flows considered, have been chosen to allow convenient comparisons between the LINFLUX results and those of well-known, two-dimensional, unsteady flow codes. Detailed numerical results for a helical fan and a three-dimensional version of the 10th Standard Cascade indicate that important progress has been made towards the development of a reliable and useful, three-dimensional, prediction capability that can be used in aeroelastic and aeroacoustic design studies.

Montgomery, Matthew D.↗
NASA NTRS2002

Low-Pressure Turbine Separation Control: Comparison With Experimental Data

The present work details a computational study, using the Glenn HT code, that analyzes the use of vortex generator jets (VGJs) to control separation on a low-pressure turbine (LPT) blade at low Reynolds numbers. The computational results are also compared with the experimental data for steady VGJs. It is found that the code determines the proper location of the separation point on the suction surface of the baseline blade (without any VGJ) for Reynolds numbers of 50,000 or less. Also, the code finds that the separated region on the suction surface of the blade vanishes with the use of VGJs. However, the separated region and the wake characteristics are not well predicted. The wake width is generally over-predicted while the wake depth is under-predicted.

Garg, Vijay K.↗
NASA NTRS2003

Comparison of the AUSM(+) and H-CUSP Schemes for Turbomachinery Applications

Many turbomachinery CFD codes use second-order central-difference (C-D) schemes with artificial viscosity to control point decoupling and to capture shocks. While C-D schemes generally give accurate results, they can also exhibit minor numerical problems including overshoots at shocks and at the edges of viscous layers, and smearing of shocks and other flow features. In an effort to improve predictive capability for turbomachinery problems, two C-D codes developed by Chima, RVCQ3D and Swift, were modified by the addition of two upwind schemes: the AUSM+ scheme developed by Liou, et al., and the H-CUSP scheme developed by Tatsumi, et al. Details of the C-D scheme and the two upwind schemes are described, and results of three test cases are shown. Results for a 2-D transonic turbine vane showed that the upwind schemes eliminated viscous layer overshoots. Results for a 3-D turbine vane showed that the upwind schemes gave improved predictions of exit flow angles and losses, although the HCUSP scheme predicted slightly higher losses than the other schemes. Results for a 3-D supersonic compressor (NASA rotor 37) showed that the AUSM+ scheme predicted exit distributions of total pressure and temperature that are not generally captured by C-D codes. All schemes showed similar convergence rates, but the upwind schemes required considerably more CPU time per iteration.

Chima, Rodrick V.↗
NASA NTRS1982

Comparison of viscous-shock-layer solutions by time-asymptotic and steady-state methods

Two flow-field codes employing the time- and space-marching numerical techniques were evaluated. Both methods were used to analyze the flow field around a massively blown Jupiter entry probe under perfect-gas conditions. In order to obtain a direct point-by-point comparison, the computations were made by using identical grids and turbulence models. For the same degree of accuracy, the space-marching scheme takes much less time as compared to the time-marching method and would appear to provide accurate results for the problems with nonequilibrium chemistry, free from the effect of local differences in time on the final solution which is inherent in time-marching methods. With the time-marching method, however, the solutions are obtainable for the realistic entry probe shapes with massive or uniform surface blowing rates; whereas, with the space-marching technique, it is difficult to obtain converged solutions for such flow conditions. The choice of the numerical method is, therefore, problem dependent. Both methods give equally good results for the cases where results are compared with experimental data.

Gupta, R. N.↗
NASA NTRS1993

A detailed study of mean-flow solutions for stability analysis of transitional flows

A finite-volume upwind-difference parabolized Navier-Stokes code is utilized to obtain laminar mean-flow solutions at Mach 3.5 on a half-angle cone of 5 deg at an angle-of-attack of 2 deg. A detailed study is conducted on this configuration; the main focus is the velocity profiles in the leeward and windward symmetry planes at various axial locations. Comparisons of the solution profiles are made with both a central-difference code that incorporates scalar and matrix dissipation models and another state-of-the-art upwind-difference finitevolume code. The results obtained emphasize the importance of using matrix dissipation models for schemes that require explicit artificial dissipation. These results also illustrate the accuracy and efficiency of the planeby-plane marching procedure for computing mean-flow solutions for predicting the onset of transition with linear instability.

Ramakrishnan, R.↗
NASA NTRS1992

Analysis of film cooling in rocket nozzles

Computational Fluid Dynamics (CFD) programs are customarily used to compute details of a flow field, such as velocity fields or species concentrations. Generally they are not used to determine the resulting conditions at a solid boundary such as wall shear stress or heat flux. However, determination of this information should be within the capability of a CFD code, as the code supposedly contains appropriate models for these wall conditions. Before such predictions from CFD analyses can be accepted, the credibility of the CFD codes upon which they are based must be established. This report details the progress made in constructing a CFD model to predict the heat transfer to the wall in a film cooled rocket nozzle. Specifically, the objective of this work is to use the NASA code FDNS to predict the heat transfer which will occur during the upcoming hot-firing of the Pratt & Whitney 40K subscale nozzle (1Q93). Toward this end, an M = 3 wall jet is considered, and the resulting heat transfer to the wall is computed. The values are compared against experimental data available in Reference 1. Also, FDNS's ability to compute heat flux in a reacting flow will be determined by comparing the code's predictions against calorimeter data from the hot firing of a 40K combustor. The process of modeling the flow of combusting gases through the Pratt & Whitney 40K subscale combustor and nozzle is outlined. What follows in this report is a brief description of the FDNS code, with special emphasis on how it handles solid wall boundary conditions. The test cases and some FDNS solution are presented next, along with comparison to experimental data. The process of modeling the flow through a chamber and a nozzle using the FDNS code will also be outlined.

Woodbury, Keith A.↗
NASA NTRS1993

Comparison Between Predicted and Experimentally Measured Flow Fields at the Exit of the SSME HPFTP Impeller

Validation of CFD codes is a critical first step in the process of developing CFD design capability. The MSFC Pump Technology Team has recognized the importance of validation and has thus funded several experimental programs designed to obtain CFD quality validation data. The first data set to become available is for the SSME High Pressure Fuel Turbopump Impeller. LDV Data was taken at the impeller inlet (to obtain a reliable inlet boundary condition) and three radial positions at the impeller discharge. Our CFD code, TASCflow, is used within the Propulsion and Commercial Pump industry as a tool for pump design. The objective of this work, therefore, is to further validate TASCflow for application in pump design. TASCflow was used to predict flow at the impeller discharge for flowrates of 80, 100 and 115 percent of design flow. Comparison to data has been made with encouraging results.

Bache, George↗
NASA NTRS1994

Fracture mechanics life analytical methods verification testing

The objective was to evaluate NASCRAC (trademark) version 2.0, a second generation fracture analysis code, for verification and validity. NASCRAC was evaluated using a combination of comparisons to the literature, closed-form solutions, numerical analyses, and tests. Several limitations and minor errors were detected. Additionally, a number of major flaws were discovered. These major flaws were generally due to application of a specific method or theory, not due to programming logic. Results are presented for the following program capabilities: K versus a, J versus a, crack opening area, life calculation due to fatigue crack growth, tolerable crack size, proof test logic, tearing instability, creep crack growth, crack transitioning, crack retardation due to overloads, and elastic-plastic stress redistribution. It is concluded that the code is an acceptable fracture tool for K solutions of simplified geometries, for a limited number of J and crack opening area solutions, and for fatigue crack propagation with the Paris equation and constant amplitude loads when the Paris equation is applicable.

Favenesi, J. A.↗
NASA NTRS1996

Performance Analysis of Two Early NACA High Speed Propellers with Application to Civil Tiltrotor Configurations

The helicopter industry is vigorously pursuing development of civil tiltrotors. One key to efficient high speed performance of this rotorcraft is prop-rotor performance. Of equal, if not greater, importance is assurance that the flight envelope is free of aeroelastic instabilities well beyond currently envisioned cuise speeds. This later condition requires study at helical tip Match numbers well in excess of 1.0. Two 1940's 'supersonic' propeller experiments conducted by NACA have provided an immensely valuable data bank with which to study prop-rotor behavior at transonic and supersonic helical tip Mach numbers. Very accurate 'blades alone' data were obtained by using nearly an infinite hub. Tabulated data were recreated from the many thrust and power figures and are included in two Appendices to this report. This data set is exceptionally well suited to re-evaluating classical blade element theories as well as evolving computational fluid dynamic (CFD) analyses. A limited comparison of one propeller's experimental results to a modem rotorcraft CFD code is made. This code, referred to as TURNS, gives very encouraging results. Detailed analysis of the performance data from both propellers is provided in Appendix A. This appendix quantifies the minimum power required to produce usable prop-rotor thrust. The dependence of minimum profile power on Reynolds number is quantified. First order compressibility power losses are quantified as well and a first approximation to design air-foil thickness ratio to avoid compressibility losses is provided. Appendix A's results are applied to study high speed civil tiltrotor cruise performance. Predicted tiltrotor performance is compared to two turboprop commercial transports. The comparison shows that there is no fundamental aerodynamic reason why the rotorcraft industry could not develop civil tiltrotor aircraft which have competitive cruise performance with today's regional, turboprop airlines. Recommendations for future study that will insure efficient prop-rotor performance to well beyond 400 knots are given.

Harris, Franklin D.↗
NASA NTRS1983

An experimental investigation of the subcritical and supercritical flow about a swept semispan wing

An experimental investigation of the turbulent, subcritical and supercritical flow over a swept, semispan wing in a solid wall wind tunnel is described. The program was conducted over a range of Mach numbers, Reynolds numbers, and angles of attack to provide a variety of test cases for assessment of wing computer codes and tunnel wall interference effects. Wing flows both without and with three dimensional flow separation are included. Data include mean surface pressures for both the wing and tunnel walls; surface oil flow patterns on the wing; and mean velocity, flow field surveys. The results are given in tabular form and presented graphically to illustrate some of the effects of the test parameters. Comparisons of the wing pressure data with the results from two inviscid wing codes are also shown to assess the importance of viscous flow and tunnel wall effects.

Lockman, W. K.↗
NASA NTRS1986

Comparison of experimental surface pressures with theoretical predictions on twin two-dimensional convergent-divergent nozzles

A three-dimensional subsonic aerodynamic panel code (VSAERO) was used to predict the effects of upper and lower external nozzle flap geometry on the external afterbody/nozzle pressure coefficient distributions and external nozzle drag of nonaxisymmetric convergent-divergent exhaust nozzles having parallel external sidewalls installed on a generic twin-engine high performance aircraft model. Nozzle static pressure coefficient distributions along the upper and lower surfaces near the model centerline and near the outer edges (corner) of the two surfaces were calculated, and nozzle drag was predicted using these surface pressure distributions. A comparison between the theoretical predictions and experimental wind tunnel data is made to evaluate the utility of the code in calculating the flow about these types of non-axisymmetric afterbody configurations. For free-stream Mach numbers of 0.60 and 0.90, the conditions where the flows were attached on the boattails yielded the best comparison between the theoretical predictions and the experimental data. For the Boattail terminal angles of greater than 15 deg., the experimental data for M = 0.60 and 0.90 indicated areas of separated flow, so the theoretical predictions failed to match the experimental data. Even though calculations of regions of separated flows are within the capabilities of the theoretical method, acceptable solutions were not obtained.

Carlson, J. R.↗
NASA NTRS1987

Analytical determination of propeller performance degradation due to ice accretion

A computer code capable of computing the propeller performance for clean, glaze, or rime ice propeller configurations to determine the performance degradation resulting from a given icing encounter has been developed. The inviscid, incompressible flowfield at each specified propeller radial location is first computed using the Theodorsen method. A droplet trajectory computation then calculates the droplet impingement points and airfoil collection efficiency for each radial location. User-selectable empirical correlations are available for determining the aerodynamic penalties due to ice accretion. Propeller performance is finally computed using strip analysis for either the clean or iced propeller. In the iced mode, the thrust and torque coefficient equations are modified by the drag and lift coefficient increments due to ice to obtain the appropriate iced values. Comparison with available experimental propeller icing data shows generally good agreement. The code's capability of properly predicting the thrust coefficient, power coefficient, and propeller efficiency of an iced propeller is shown to be dependent on the choice of empirical correlation employed as well as on the proper specification of the radial icing extent and propeller blade angle.

Miller, T. L.↗