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NASA NTRS1986

Calculations of cosmogenic nuclide production rates in the Earth's atmosphere and their inventories

The production rates of cosmogenic isotopes in the Earth's atmosphere and their resulting terrestrial abundances have been calculated, taking into account both geomagnetic and solar-modulatory effects. The local interstellar flux was assumed to be that of Garcia-Munoz, et al. Solar modulation was accounted for using the heliocentric potential model and expressed in terms of the Deep River neutron monitor count rates. The geomagnetic field was presented by vertical cutoffs calculated by Shea and Smart and the non-vertical cutoffs calculated using ANGRI. The local interstellar particle flux was first modulated using the heliocentric potential field. The modulated cosmic-ray fluxes reaching the earth's orbit then interacted with the geomagnetic field as though it were a high-pass filter. The interaction of the cosmic radiation with the Earth's atmosphere was calculated utilizing the Bolztmann transport equation. Spallation cross sections for isotope production were calculated using the formalism of Silberberg and Tsao and other cross sections were taken from standard sources. Inventories were calculated by accounting from the variation in solar modulation and geomagnetic field strength with time. Results for many isotope, including C-14, Be-7 and Be-10 are in generally good agreement with existing data. The C-14 inventory, for instance, amounts to 1.75/sq cm(e)/s, in excellent agreement with direct estimates.

Obrien, K.↗

NASA NTRS1986

An explicit finite-volume time-marching procedure for turbulent flow calculations

A method was developed which calculates two-dimensional, transonic, viscous flow in ducts. The finite-volume, time-marching formulation is used to obtain steady flow solutions of the Reynolds-averaged form of the Navier-Stokes equations. The entire calculation is performed in the physical domain. Control volumes are chosen so that smoothing of flow properties, typically required for stability, is not required. Different time steps are used in the different governing equations. A new pressure interpolation scheme is introduced which improves the shock capturing ability of the method. A multi-volume method for pressure changes in the boundary layer allows calculations which use very long and thin control volumes (length/height - 1000). The method is compared with two test cases. Essentially incompressible turbulent boundary layer flow in an adverse pressure gradient is calculated and the computed distributions of mean velocity and shear are in good agreement with the measurements. Transonic viscous flow in a converging diverging nozzle is calculated; the Mach number upstream of the shock is approximately 1.25. The agreement between the calculated and measured shock strength and total pressure losses is good.

Nicholson, Stephen↗

NASA NTRS1987

An explicit finite-volume time-marching procedure for turbulent flow calculations

Nicholson, Stephen↗

NASA NTRS1989

Measurement effects on the calculation of in-flight thrust for an F404 turbofan engine

A study has been performed that investigates parameter measurement effects on calculated in-flight thrust for the General Electric F404-GE-400 afterburning turbofan engine which powered the X-29A forward-swept wing research aircraft. Net-thrust uncertainty and influence coefficients were calculated and are presented. Six flight conditions were analyzed at five engine power settings each. Results were obtained using the mass flow-temperature and area-pressure thrust calculation methods, both based on the commonly used gas generator technique. Thrust uncertainty was determined using a common procedure based on the use of measurement uncertainty and influence coefficients. The effects of data nonlinearity on the uncertainty calculation procedure were studied and results are presented. The advantages and disadvantages of using this particular uncertainty procedure are discussed. A brief description of the thrust-calculation technique along with the uncertainty calculation procedure is included.

Conners, Timothy R.↗

NASA NTRS1989

Measurement effects on the calculation of in-flight thrust for an F404 turbofan engine

A study was performed that investigates parameter measurement effects on calculated in-flight thrust for the General Electric F404-GE-400 afterburning turbofan engine which powered the X-29A forward-swept wing research aircraft. Net-thrust uncertainty and influence coefficients were calculated and are presented. Six flight conditions were analyzed at five engine power settings each. Results were obtained using the mass flow-temperature and area-pressure thrust calculation methods, both based on the commonly used gas generator technique. Thrust uncertainty was determined using a common procedure based on the use of measurement uncertainty and influence coefficients. The effects of data nonlinearity on the uncertainty calculation procedure were studied and results are presented. The advantages and disadvantages of using this particular uncertainty procedure are discussed. A brief description of the thrust-calculation technique along with the uncertainty calculation procedure is included.

Conners, Timothy R.↗

NASA NTRS1990

Evaluation of various thrust calculation techniques on an F404 engine

In support of performance testing of the X-29A aircraft at the NASA-Ames, various thrust calculation techniques were developed and evaluated for use on the F404-GE-400 engine. The engine was thrust calibrated at NASA-Lewis. Results from these tests were used to correct the manufacturer's in-flight thrust program to more accurately calculate thrust for the specific test engine. Data from these tests were also used to develop an independent, simplified thrust calculation technique for real-time thrust calculation. Comparisons were also made to thrust values predicted by the engine specification model. Results indicate uninstalled gross thrust accuracies on the order of 1 to 4 percent for the various in-flight thrust methods. The various thrust calculations are described and their usage, uncertainty, and measured accuracies are explained. In addition, the advantages of a real-time thrust algorithm for flight test use and the importance of an accurate thrust calculation to the aircraft performance analysis are described. Finally, actual data obtained from flight test are presented.

Ray, Ronald J.↗

NASA NTRS1990

Real-time aerodynamic heating and surface temperature calculations for hypersonic flight simulation

A real-time heating algorithm was derived and installed on the Ames Research Center Dryden Flight Research Facility real-time flight simulator. This program can calculate two- and three-dimensional stagnation point surface heating rates and surface temperatures. The two-dimensional calculations can be made with or without leading-edge sweep. In addition, upper and lower surface heating rates and surface temperatures for flat plates, wedges, and cones can be calculated. Laminar or turbulent heating can be calculated, with boundary-layer transition made a function of free-stream Reynolds number and free-stream Mach number. Real-time heating rates and surface temperatures calculated for a generic hypersonic vehicle are presented and compared with more exact values computed by a batch aeroheating program. As these comparisons show, the heating algorithm used on the flight simulator calculates surface heating rates and temperatures well within the accuracy required to evaluate flight profiles for acceptable heating trajectories.

Quinn, Robert D.↗

NASA NTRS1991

Calculations on the rate of the ion-molecule reaction between NH3(+) and H2

The rate coefficient for the ion-molecule reaction NH3(+) + H2 yields NH4(+) + H has been calculated as a function of temperature with the use of the statistical phase space approach. The potential surface and reaction complex and transition state parameters used in the calculation have been taken from ab initio quantum chemical calculations. The calculated rate coefficient has been found to mimic the unusual temperature dependence measured in the laboratory, in which the rate coefficient decreases with decreasing temperature until 50-100 K and then increases at still lower temperatures. Quantitative agreement between experimental and theoretical rate coefficients is satisfactory given the uncertainties in the ab initio results and in the dynamics calculations. The rate coefficient for the unusual three-body process NH3(+) + H2 + He yields NH4(+) + H + He has also been calculated as a function of temperature and the result found to agree well with a previous laboratory determination.

Herbst, Eric↗

NASA NTRS1994

Three dimensional model calculations of the global dispersion of high speed aircraft exhaust and implications for stratospheric ozone loss

Two-dimensional (zonally averaged) photochemical models are commonly used for calculations of ozone changes due to various perturbations. These include calculating the ozone change expected as a result of change in the lower stratospheric composition due to the exhaust of a fleet of supersonic aircraft flying in the lower stratosphere. However, zonal asymmetries are anticipated to be important to this sort of calculation. The aircraft are expected to be restricted from flying over land at supersonic speed due to sonic booms, thus the pollutant source will not be zonally symmetric. There is loss of pollutant through stratosphere/troposphere exchange, but these processes are spatially and temporally inhomogeneous. Asymmetry in the pollutant distribution contributes to the uncertainty in the ozone changes calculated with two dimensional models. Pollutant distributions for integrations of at least 1 year of continuous pollutant emissions along flight corridors are calculated using a three dimensional chemistry and transport model. These distributions indicate the importance of asymmetry in the pollutant distributions to evaluation of the impact of stratospheric aircraft on ozone. The implications of such pollutant asymmetries to assessment calculations are discussed, considering both homogeneous and heterogeneous reactions.

Douglass, Anne R.↗

NASA NTRS1994

Errors induced by the neglect of polarization in radiance calculations for Rayleigh-scattering atmospheres

Although neglecting polarization and replacing the rigorous vector radiative transfer equation by its approximate scalar counterpart has no physical background, it is a widely used simplification when the incident light is unpolarized and only the intensity of the reflected light is to be computed. We employ accurate vector and scalar multiple-scattering calculations to perform a systematic study of the errors induced by the neglect of polarization in radiance calculations for a homogeneous, plane-parallel Rayleigh-scattering atmosphere (with and without depolarization) above a Lambertian surface. Specifically, we calculate percent errors in the reflected intensity for various directions of light incidence and reflection, optical thicknesses of the atmosphere, single-scattering albedos, depolarization factors, and surface albedos. The numerical data displayed can be used to decide whether or not the scalar approximation may be employed depending on the parameters of the problem. We show that the errors decrease with increasing depolarization factor and/or increasing surface albedo. For conservative or nearly conservative scattering and small surface albedos, the errors are maximum at optical thicknesses of about 1. The calculated errors may be too large for some practical applications, and, therefore, rigorous vector calculations should be employed whenever possible. However, if approximate scalar calculations are used, we recommend to avoid geometries involving phase angles equal or close to 0 deg and 90 deg, where the errors are especially significant. We propose a theoretical explanation of the large vector/scalar differences in the case of Rayleigh scattering. According to this explanation, the differences are caused by the particular structure of the Rayleigh scattering matrix and come from lower-order (except first-order) light scattering paths involving right scattering angles and right-angle rotations of the scattering plane.

Mishchenko, M. I.↗

NASA NTRS1997

Collection Efficiency and Ice Accretion Calculations for a Boeing 737-300 Inlet

Collection efficiency and ice accretion calculations have been made for a Boeing 737-300 inlet using a three-dimensional panel code, an adaptive grid code, the NASA Lewis LEWICE3D grid based ice accretion code. Flow solutions for the inlet were generated using the VSAERO panel code. Grids used in the ice accretion calculations were generated using the newly developed adaptive grid code ICEGRID3D. The LEWICE3D grid based ice accretion program was used to calculate impingement efficiency and ice shapes. Ice shapes typifying rime and mixed icing conditions were generated for a 30 minute hold condition. All calculations were performed on an SGI Power Challenge computer. The results have been compared to experimental flow and impingement data. In general, the calculated flow and collection efficiencies compared well with experiment, and the ice shapes looked reasonable and appeared representative of the rime and mixed icing conditions for which they were calculated.

Bidwell, Colin S.↗

NASA NTRS1995

Lagrangian Transport Calculations Using UARS Data: Ozone - Part 2

Trajectory calculations are used to examine ozone transport in the polar winter stratosphere during periods of the Upper Atmosphere Research Satellite (UARS) observations. The value of these calculations for determining mass transport was demonstrated previously using UARS observations of long-lived tracers, In the middle stratosphere, the overall ozone behavior observed by the Microwave Limb Sounder in the polar vortex is reproduced by this purely dynamical model. Calculations show the evolution of ozone in the lower stratosphere during early winter to be dominated by dynamics in December 1992 in the Arctic. Calculations for June 1992 in the Antarctic show evidence of chemical ozone destruction and indicate that approx. 50% of the chemical destruction may be masked by dynamical effects, mainly diabatic descent, which bring higher ozone into the lower-stratospheric vortex. Estimating differences between calculated and observed fields suggests that dynamical changes masked approx. 20% - 35% of chemical ozone loss during late February and early March 1993 in the Arctic. In the Antarctic late winter, in late August and early September 1992, below approx. 520 K, the evolution of vortex-averaged ozone is entirely dominated by chemical effects; above this level, however, chemical ozone depletion can be partially or completely masked by dynamical effects. Our calculations for 1992 showed that chemical loss was nearly completely compensated by increases due to diabatic descent at 655 K.

Manney, Gloria L.↗

NASA NTRS2002

On the Exit Boundary Condition for One-Dimensional Calculations of Pulsed Detonation Engine Performance

In one-dimensional calculations of pulsed detonation engine (PDE) performance, the exit boundary condition is frequently taken to be a constant static pressure. In reality, for an isolated detonation tube, after the detonation wave arrives at the exit plane, there will be a region of high pressure, which will gradually return to ambient pressure as an almost spherical shock wave expands away from the exit, and weakens. Initially, the flow is supersonic, unaffected by external pressure, but later becomes subsonic. Previous authors have accounted for this situation either by assuming the subsonic pressure decay to be a relaxation phenomenon, or by running a two-dimensional calculation first, including a domain external to the detonation tube, and using the resulting exit pressure temporal distribution as the boundary condition for one-dimensional calculations. These calculations show that the increased pressure does affect the PDE performance. In the present work, a simple model of the exit process is used to estimate the pressure decay time. The planar shock wave emerging from the tube is assumed to transform into a spherical shock wave. The initial strength of the spherical shock wave is determined from comparison with experimental results. Its subsequent propagation, and resulting pressure at the tube exit, is given by a numerical blast wave calculation. The model agrees reasonably well with other, limited, results. Finally, the model was used as the exit boundary condition for a one-dimensional calculation of PDE performance to obtain the thrust wall pressure for a hydrogen-air detonation in tubes of length to diameter ratio (L/D) of 4, and 10, as well as for the original, constant pressure boundary condition. The modified boundary condition had no performance impact for values of L/D > 10, and moderate impact for L/D = 4.

Wilson, Jack↗

NASA NTRS2004

Validation of NASCAP-2K Spacecraft-Environment Interactions Calculations

The recently released Nascap-2k, version 2.0, three-dimensional computer code models interactions between spacecraft surfaces and low-earth-orbit, geosynchronous, auroral, and interplanetary plasma environments. It replaces the earlier three-dimensional spacecraft interactions codes NASCAP/GEO, NASCAP/LEO, POLAR, and DynaPAC. Nascap-2k has improved numeric techniques, a modern user interface, and a simple, interactive satellite surface definition module (Object ToolKit). We establish the accuracy of Nascap-2k both by comparing computed currents and potentials with analytic results and by comparing Nascap-2k results with published calculations using the earlier codes. Nascap-2k predicts Langmuir-Blodgett or Parker-Murphy current collection for a nearly spherical (100 surfaces) satellite in a short Debye length plasma depending on the absence or presence of a magnetic field. A low fidelity (in geometry and time) Nascap-2k geosynchronous charging calculation gives the same results as the corresponding low fidelity NASCAP/GEO calculation. A high fidelity calculation (using the Nascap-2k improved geometry and time stepping capabilities) gives higher potentials, which are more consistent with typical observations. Nascap-2k predicts the same current as a function of applied potential as was observed and calculated by NASCAP/LEO for the SPEAR I rocket with a bipolar sheath. A Nascap-2k DMSP charging calculation gives results similar to those obtained using POLAR and consistent with observation.

Davis, V. A.↗

NASA NTRS2004

Martian Radiation Environment: Model Calculations and Recent Measurements with "MARIE"

The Galactic Cosmic Ray spectra in Mars orbit were generated with the recently expanded HZETRN (High Z and Energy Transport) and QMSFRG (Quantum Multiple-Scattering theory of nuclear Fragmentation) model calculations. These model calculations are compared with the first eighteen months of measured data from the MARIE (Martian Radiation Environment Experiment) instrument onboard the 2001 Mars Odyssey spacecraft that is currently in Martian orbit. The dose rates observed by the MARIE instrument are within 10% of the model calculated predictions. Model calculations are compared with the MARIE measurements of dose, dose-equivalent values, along with the available particle flux distribution. Model calculated particle flux includes GCR elemental composition of atomic number, Z = 1-28 and mass number, A = 1-58. Particle flux calculations specific for the current MARIE mapping period are reviewed and presented.

Saganti, P. B.↗

NASA NTRS2007

Core Physics and Kinetics Calculations for the Fissioning Plasma Core Reactor

Highly efficient, compact nuclear reactors would provide high specific impulse spacecraft propulsion. This analysis and numerical simulation effort has focused on the technical feasibility issues related to the nuclear design characteristics of a novel reactor design. The Fissioning Plasma Core Reactor (FPCR) is a shockwave-driven gaseous-core nuclear reactor, which uses Magneto Hydrodynamic effects to generate electric power to be used for propulsion. The nuclear design of the system depends on two major calculations: core physics calculations and kinetics calculations. Presently, core physics calculations have concentrated on the use of the MCNP4C code. However, initial results from other codes such as COMBINE/VENTURE and SCALE4a. are also shown. Several significant modifications were made to the ISR-developed QCALC1 kinetics analysis code. These modifications include testing the state of the core materials, an improvement to the calculation of the material properties of the core, the addition of an adiabatic core temperature model and improvement of the first order reactivity correction model. The accuracy of these modifications has been verified, and the accuracy of the point-core kinetics model used by the QCALC1 code has also been validated. Previously calculated kinetics results for the FPCR were described in the ISR report, "QCALC1: A code for FPCR Kinetics Model Feasibility Analysis" dated June 1, 2002.

Butler, C.↗

NASA NTRS2001

Calculation of the Aerodynamic Behavior of the Tilt Rotor Aeroacoustic Model (TRAM) in the DNW

Comparisons of measured and calculated aerodynamic behavior of a tiltrotor model are presented. The test of the Tilt Rotor Aeroacoustic Model (TRAM) with a single, 1/4-scale V- 22 rotor in the German-Dutch Wind Tunnel (DNW) provides an extensive set of aeroacoustic, performance, and structural loads data. The calculations were performed using the rotorcraft comprehensive analysis CAMRAD II. Presented are comparisons of measured and calculated performance and airloads for helicopter mode operation, as well as calculated induced and profile power. An aerodynamic and wake model and calculation procedure that reflects the unique geometry and phenomena of tiltrotors has been developed. There are major differences between this model and the corresponding aerodynamic and wake model that has been established for helicopter rotors. In general, good correlation between measured and calculated performance and airloads behavior has been shown. Two aspects of the analysis that clearly need improvement are the stall delay model and the trailed vortex formation model.

Johnson, Wayne↗

NASA NTRS2001

Airloads and Wake Geometry Calculations for an Isolated Tiltrotor Model in a Wind Tunnel

Comparisons of measured and calculated aerodynamic behavior of a tiltrotor model are presented. The test of the Tilt Rotor Aeroacoustic Model (TRAM) with a single, 0.25-scale V-22 rotor in the German-Dutch Wind Tunnel (DNW) provides an extensive set of aeroacoustic, performance, and structural loads data. The calculations were performed using the rotorcraft comprehensive analysis CAMRAD II. Presented are comparisons of measured and calculated performance for hover and helicopter mode operation, and airloads for helicopter mode. Calculated induced power, profile power, and wake geometry provide additional information about the aerodynamic behavior. An aerodynamic and wake model and calculation procedure that reflects the unique geometry and phenomena of tiltrotors has been developed. There are major differences between this model and the corresponding aerodynamic and wake model that has been established for helicopter rotors. In general, good correlation between measured and calculated performance and airloads behavior has been shown. Two aspects of the analysis that clearly need improvement are the stall delay model and the trailed vortex formation model.

Johnson, Wayne↗