ALBERT

Beschreibung: We use geodetic data from Very Long Baseline Interferometry (VLBI) to determine the pre- and postseismic velocities of two sites. We then place limits on variations in interseismic strain buildup. The 1987 and 1988 Gulf of Alaska earthquakes (each Ms = 7.6) broke the Pacific plate interior. During the earthquakes the Cape Yakataga site moved 78 mm toward southwest. During the 1989 Loma Prieta earthquake (Ms = 7.1) the Fort Ord site moved 48 mm toward north. Baselines (a) from Fairbanks to Cape Yakataga and (b) from Mojave to Fort Ord change at nearly the same rate before and after the earthquakes. Postseismic transients, which we determine from differences between post- and preseismic rates, are minor: at Cape Yakataga the transient is 3 +/- 4 mm in a postseismic interval of 23 months, and at Fort Ord the transient is 6 +/- 5 mm in 21 months. The slip beneath the Loma Prieta rupture needed to generate the Fort Ord transient is 0.22 +/- 0.19 m, one-tenth the coseismic slip (2 m). We analyze elastic lithosphere-viscous asthenosphere models to determine that the characteristic time describing exponential decay in deep fault slip is longer than 6 years. The VLBI measurements are consistent with uniform interseismic strain buildup. They disagree with fast postseismic rates caused by an asthenosphere with very low viscosity.

Beschreibung: The AEPI (Atmospheric Emissions Photometric Imager) experiment on the ATLAS-1 shuttle mission (launched March 24, 1992) imaged the earth night airglow emission in O2 Atmospheric (0,0) bands, at 762.0 nm. Earthward views of O2 A bands show structure from gravity waves which exhibit extended horizontal structure with horizontal wavelenghts on the order of 50-100 km. These observations of the O2 A (0,0) bands are particularly interesting since in this wavelength the lower atmosphere absorbs all the earth-reflected emissions and most of the spectrally diffuse backgrounds. Herein we present observations of gravity waves using a topside airglow imaging technique.

Beschreibung: Fujii et al. (1994) obtained characteristics of the electrodynamic parameters, that is, field-aligned currents, electric fields, and electron precipitation, which are associted with auroral substorm events in the nighttime sector, through a unique analysis that places the ionospheric measurements of these parameters into the context of a generic substorm determined from global auroral images. In this paper we investigate in considerably more detail the characteristics of the field-aligned currents using data from the same set of passes as the previuos study. We show for the first time that the net upward field-aligned currents throughout the surge and surge horn are sufficient to account for most if not all of the converging currents of the auroral electrojets. Current densities are largest in the surge and surge horn. Current region continuity does not appear to exist across the substorm bulge region. Much of the auroral substorm field-aligned current is composed of filamentary currents and finite current segments at large angles to each other. The westward electrojet may contain large gradients in intensity both in local time and latitude due to sets of localized field-aligned currents. The net downward current for several hours to the west of the surge is insufficient to account for the eastward electrojet, consistent with the concept that this electrojet originates primarily on the dayside. Our pattern of field-aligned currents associated with the surge has common features and also differs significantly from the patterns previously derived from data from radars and ground-based magnetometer arrays. Our pattern is considerably more complex, probably due to the much higher resolution in latitude of the satellite data. It is also larger in area, since our average substorm is much larger than those pertaining to the previous patterns, giving a substorm wedge considerably wider than that obtained from the radar and array data.

Beschreibung: An analytical model of the bidirectional reflectance for optically semi-infinite plant canopies has been extended to describe the reflectance of finite depth canopies contributions from the underlying soil. The model depends on 10 independent parameters describing vegetation and soil optical and structural properties. The model is inverted with a nonlinear minimization routine using directional reflectance data for lawn (leaf area index (LAI) is equal to 9.9), soybeans (LAI, 2.9) and simulated reflectance data (LAI, 1.0) from a numerical bidirectional reflectance distribution function (BRDF) model (Myneni et al., 1988). While the ten-parameter model results in relatively low rms differences for the BRDF, most of the retrieved parameters exhibit poor stability. The most stable parameter was the single-scattering albedo of the vegetation. Canopy albedo could be derived with an accuracy of less than 5% relative error in the visible and less than 1% in the near-infrared. Sensitivity were performed to determine which of the 10 parameters were most important and to assess the effects of Gaussian noise on the parameter retrievals. Out of the 10 parameters, three were identified which described most of the BRDF variability. At low LAI values the most influential parameters were the single-scattering albedos (both soil and vegetation) and LAI, while at higher LAI values (greater than 2.5) these shifted to the two scattering phase function parameters for vegetation and the single-scattering albedo of the vegetation. The three-parameter model, formed by fixing the seven least significant parameters, gave higher rms values but was less sensitive to noise in the BRDF than the full ten-parameter model. A full hemispherical reflectance data set for lawn was then interpolated to yield BRDF values corresponding to advanced very high resolution radiometer (AVHRR) scan geometries collected over a period of nine days. The resulting parameters and BRDFs are similar to those for the full sampling geometry, suggesting that the limited geometry of AVHRR measurements might be used to reliably retrieve BRDF and canopy albedo with this model.

Beschreibung: ARC2D is a computational fluid dynamics program developed at the NASA Ames Research Center specifically for airfoil computations. The program uses implicit finite-difference techniques to solve two-dimensional Euler equations and thin layer Navier-Stokes equations. It is based on the Beam and Warming implicit approximate factorization algorithm in generalized coordinates. The methods are either time accurate or accelerated non-time accurate steady state schemes. The evolution of the solution through time is physically realistic; good solution accuracy is dependent on mesh spacing and boundary conditions. The mathematical development of ARC2D begins with the strong conservation law form of the two-dimensional Navier-Stokes equations in Cartesian coordinates, which admits shock capturing. The Navier-Stokes equations can be transformed from Cartesian coordinates to generalized curvilinear coordinates in a manner that permits one computational code to serve a wide variety of physical geometries and grid systems. ARC2D includes an algebraic mixing length model to approximate the effect of turbulence. In cases of high Reynolds number viscous flows, thin layer approximation can be applied. ARC2D allows for a variety of solutions to stability boundaries, such as those encountered in flows with shocks. The user has considerable flexibility in assigning geometry and developing grid patterns, as well as in assigning boundary conditions. However, the ARC2D model is most appropriate for attached and mildly separated boundary layers; no attempt is made to model wake regions and widely separated flows. The techniques have been successfully used for a variety of inviscid and viscous flowfield calculations. The Cray version of ARC2D is written in FORTRAN 77 for use on Cray series computers and requires approximately 5Mb memory. The program is fully vectorized. The tape includes variations for the COS and UNICOS operating systems. Also included is a sample routine for CONVEX computers to emulate Cray system time calls, which should be easy to modify for other machines as well. The standard distribution media for this version is a 9-track 1600 BPI ASCII Card Image format magnetic tape. The Cray version was developed in 1987. The IBM ES/3090 version is an IBM port of the Cray version. It is written in IBM VS FORTRAN and has the capability of executing in both vector and parallel modes on the MVS/XA operating system and in vector mode on the VM/XA operating system. Various options of the IBM VS FORTRAN compiler provide new features for the ES/3090 version, including 64-bit arithmetic and up to 2 GB of virtual addressability. The IBM ES/3090 version is available only as a 9-track, 1600 BPI IBM IEBCOPY format magnetic tape. The IBM ES/3090 version was developed in 1989. The DEC RISC ULTRIX version is a DEC port of the Cray version. It is written in FORTRAN 77 for RISC-based Digital Equipment platforms. The memory requirement is approximately 7Mb of main memory. It is available in UNIX tar format on TK50 tape cartridge. The port to DEC RISC ULTRIX was done in 1990. COS and UNICOS are trademarks and Cray is a registered trademark of Cray Research, Inc. IBM, ES/3090, VS FORTRAN, MVS/XA, and VM/XA are registered trademarks of International Business Machines. DEC and ULTRIX are registered trademarks of Digital Equipment Corporation.

Beschreibung: Panel method computer programs are software tools of moderate cost used for solving a wide range of engineering problems. The panel code PMARC_12 (Panel Method Ames Research Center, version 12) can compute the potential flow field around complex three-dimensional bodies such as complete aircraft models. PMARC_12 is a well-documented, highly structured code with an open architecture that facilitates modifications and the addition of new features. Adjustable arrays are used throughout the code, with dimensioning controlled by a set of parameter statements contained in an include file; thus, the size of the code (i.e. the number of panels that it can handle) can be changed very quickly. This allows the user to tailor PMARC_12 to specific problems and computer hardware constraints. In addition, PMARC_12 can be configured (through one of the parameter statements in the include file) so that the code's iterative matrix solver is run entirely in RAM, rather than reading a large matrix from disk at each iteration. This significantly increases the execution speed of the code, but it requires a large amount of RAM memory. PMARC_12 contains several advanced features, including internal flow modeling, a time-stepping wake model for simulating either steady or unsteady (including oscillatory) motions, a Trefftz plane induced drag computation, off-body and on-body streamline computations, and computation of boundary layer parameters using a two-dimensional integral boundary layer method along surface streamlines. In a panel method, the surface of the body over which the flow field is to be computed is represented by a set of panels. Singularities are distributed on the panels to perturb the flow field around the body surfaces. PMARC_12 uses constant strength source and doublet distributions over each panel, thus making it a low order panel method. Higher order panel methods allow the singularity strength to vary linearly or quadratically across each panel. Experience has shown that low order panel methods can provide nearly the same accuracy as higher order methods over a wide range of cases with significantly reduced computation times; hence, the low order formulation was adopted for PMARC_12. The flow problem is solved by modeling the body as a closed surface dividing space into two regions: the region external to the surface in which an unknown velocity potential exists representing the flow field of interest, and the region internal to the surface in which a known velocity potential (representing a fictitious flow) is prescribed as a boundary condition. Both velocity potentials are required to satisfy Laplace's equation. A surface integral equation for the unknown potential external to the surface can be written by applying Green's Theorem to the external region. Using the internal potential and zero flow through the surface as boundary conditions, the unknown potential external to the surface can be solved for. When the internal flow option, which allows the analysis of closed ducts, wind tunnels, and similar internal flow problems, is selected, the geometry is modeled such that the flow field of interest is inside the geometry and the fictitious flow is outside the geometry. Items such as wings, struts, or aircraft models can be included in the internal flow problem. The time-stepping wake model gives PMARC_12 the ability to model both steady and unsteady flow problems. The wake is convected downstream from the wake-separation line by the local velocity field. With each time step, a new row of wake panels is added to the wake at the wake-separation line. Time stepping can start from time t=0 (no initial wake) or from time t=t0 (an initial wake is specified). A wide range of motions can be prescribed, including constant rates of translation, constant rate of rotation about an arbitrary axis, oscillatory translation, and oscillatory rotation about any of the three coordinate axes. Investigators interested in a visual representation of the phenomenon they are studying with PMARC_12 may want to consider obtaining the program GVS (ARC-13361), the General Visualization System. GVS is a Silicon Graphics IRIS program which was created for the purpose of supporting the scientific visualization needs of PMARC_12. GVS is available separately from COSMIC. PMARC_12 is written in standard FORTRAN 77, with the exception of the NAMELIST extension used for input. This makes the code fairly machine independent. A compiler which supports the NAMELIST extension is required. The amount of free disk space and RAM memory required for PMARC_12 will vary depending on how the code is dimensioned using the parameter statements in the include file. The recommended minimum requirements are 20Mb of free disk space and 4Mb of RAM. PMARC_12 has been successfully implemented on a Macintosh II running System 6.0.7 or 7.0 (using MPW/Language Systems Fortran 3.0), a Sun SLC running SunOS 4.1.1, an HP 720 running HP-UX 8.07, an SGI IRIS running IRIX 4.0 (it will not run under IRIX 3.x.x without modifications), an IBM RS/6000 running AIX, a DECstation 3100 running ULTRIX, and a CRAY-YMP running UNICOS 6.0 or later. Due to its memory requirements, this program does not readily lend itself to implementation on MS-DOS based machines. The standard distribution medium for PMARC_12 is a set of three 3.5 inch 800K Macintosh format diskettes and one 3.5 inch 1.44Mb Macintosh format diskette which contains an electronic copy of the documentation in MS Word 5.0 format for the Macintosh. Alternate distribution media and formats are available upon request, but these will not include the electronic version of the document. No executables are included on the distribution media. This program is an update to PMARC version 11, which was released in 1989. PMARC_12 was released in 1993. It is available only for use by United States citizens.

Beschreibung: This program determines the supersonic flowfield surrounding three-dimensional wing-body configurations of a delta wing. It was designed to provide the numerical computation of three dimensional inviscid, flowfields of either perfect or real gases about supersonic or hypersonic airplanes. The governing equations in conservation law form are solved by a finite difference method using a second order noncentered algorithm between the body and the outermost shock wave, which is treated as a sharp discontinuity. Secondary shocks which form between these boundaries are captured automatically. The flowfield between the body and outermost shock is treated in a shock capturing fashion and therefore allows for the correct formation of secondary internal shocks . The program operates in batch mode, is in CDC update format, has been implemented on the CDC 7600, and requires more than 140K (octal) word locations.

Beschreibung: MSIS is an empirical model of the thermosphere based on Mass Spectrometer and Incoherent Scatter data. It provides a description of atmospheric temperature, density, and composition for altitudes higher than 85 kilometers. There are coefficients in MSIS to account for yearly and daily variations, geodetic latitude and longitude, altitude, and solar activity. Variations due to magnetic storms are represented by three-hour magnetic ap indices. The MSIS model enables a more timely prediction of aeronomic densities for specific events, such as rocket flights. The database for this model is a comprehensive summary of rocket flight, satellite, incoherent scatter radar, grenade, and falling sphere measurements. Subsets of data were formed by random selection after sorting on altitude, latitude, time of day, etc. Curve fitting was done with four to five thousand data points at a time. The resulting coefficients are presented in subroutines which calculate thermospheric composition and temperature for a user-supplied position and time. MSIS is written in FORTRAN 77 for use with batch or interactive programs and has been implemented on a DEC VAX series computer operating under VMS 4.3 with a central memory requirement of approximately 25K of 8 bit bytes. MSIS is based on a 1977 thermosphere model and was last updated in 1987 to reflect the CIRA 1986 Neutral Thermosphere Model.

Beschreibung: This theoretical aerodynamics program, TAD, was developed to predict the aerodynamic characteristics of vehicles with sounding rocket configurations. These slender, axisymmetric finned vehicle configurations have a wide range of aeronautical applications from rockets to high speed armament. Over a given range of Mach numbers, TAD will compute the normal force coefficient derivative, the center-of-pressure, the roll forcing moment coefficient derivative, the roll damping moment coefficient derivative, and the pitch damping moment coefficient derivative of a sounding rocket configured vehicle. The vehicle may consist of a sharp pointed nose of cone or tangent ogive shape, up to nine other body divisions of conical shoulder, conical boattail, or circular cylinder shape, and fins of trapezoid planform shape with constant cross section and either three or four fins per fin set. The characteristics computed by TAD have been shown to be accurate to within ten percent of experimental data in the supersonic region. The TAD program calculates the characteristics of separate portions of the vehicle, calculates the interference between separate portions of the vehicle, and then combines the results to form a total vehicle solution. Also, TAD can be used to calculate the characteristics of the body or fins separately as an aid in the design process. Input to the TAD program consists of simple descriptions of the body and fin geometries and the Mach range of interest. Output includes the aerodynamic characteristics of the total vehicle, or user-selected portions, at specified points over the mach range. The TAD program is written in FORTRAN IV for batch execution and has been implemented on an IBM 360 computer with a central memory requirement of approximately 123K of 8 bit bytes. The TAD program was originally developed in 1967 and last updated in 1972.

Beschreibung: This program, which is called 'AOFA', determines the complete viscous and inviscid flow around a body of revolution at a given angle of attack and traveling at supersonic speeds. The viscous calculations from this program agree with experimental values for surface and pitot pressures and with surface heating rates. At high speeds, lee-side flows are important because the local heating is difficult to correlate and because the shed vortices can interact with vehicle components such as a canopy or a vertical tail. This program should find application in the design analysis of any high speed vehicle. Lee-side flows are difficult to calculate because thin-boundary-layer theory is not applicable and the concept of matching inviscid and viscous flow is questionable. This program uses the parabolic approximation to the compressible Navier-Stokes equations and solves for the complete inviscid and viscous regions of flow, including the pressure. The parabolic approximation results from the assumption that the stress derivatives in the streamwise direction are small in comparison with derivatives in the normal and circumferential directions. This assumption permits the equation to be solved by an implicit finite difference marching technique which proceeds downstream from the initial data point, provided the inviscid portion of flow is supersonic. The viscous cross-flow separation is also determined as part of the solution. To use this method it is necessary to first determine an initial data point in a region where the inviscid portion of the flow is supersonic. Input to this program consists of two parts. Problem description is conveyed to the program by namelist input. Initial data is acquired by the program as formatted data. Because of the large amount of run time this program can consume the program includes a restart capability. Output is in printed format and magnetic tape for further processing. This program is written in FORTRAN IV and has been implemented on a CDC 7600 with a central memory requirement of approximately 35K (octal) of 60 bit words.

Beschreibung: The Comprehensive Analytical Model of Rotorcraft Aerodynamics, CAMRAD, program is designed to calculate rotor performance, loads, and noise; helicopter vibration and gust response; flight dynamics and handling qualities; and system aeroelastic stability. The analysis is a consistent combination of structural, inertial, and aerodynamic models applicable to a wide range of problems and a wide class of vehicles. The CAMRAD analysis can be applied to articulated, hingeless, gimballed, and teetering rotors with an arbitrary number of blades. The rotor degrees of freedom included are blade/flap bending, rigid pitch and elastic torsion, and optionally gimbal or teeter motion. General two-rotor aircrafts can be modeled. Single main-rotor and tandem helicopter and sideby-side tilting proprotor aircraft configurations can be considered. The case of a rotor or helicopter in a wind tunnel can also be modeled. The aircraft degrees of freedom included are the six rigid body motion, elastic airframe motions, and the rotor/engine speed perturbations. CAMRAD calculates the load and motion of helicopters and airframes in two stages. First the trim solution is obtained; then the flutter, flight dynamics, and/or transient behavior can be calculated. The trim operating conditions considered include level flight, steady climb or descent, and steady turns. The analysis of the rotor includes nonlinear inertial and aerodynamic models, applicable to large blade angles and a high inflow ratio, The rotor aerodynamic model is based on two-dimensional steady airfoil characteristics with corrections for three-dimensional and unsteady flow effects, including a dynamic stall model. In the flutter analysis, the matrices are constructed that describe the linear differential equations of motion, and the equations are analyzed. In the flight dynamics analysis, the stability derivatives are calculated and the matrices are constructed that describe the linear differential equations of motion. These equations are analyzed. In the transient analysis, the rigid body equations of motion are numerically integrated, for a prescribed transient gust or control input. The CAMRAD program product is available by license for a period of ten years to domestic U.S. licensees. The licensed program product includes the CAMRAD source code, command procedures, sample applications, and one set of supporting documentation. Copies of the documentation may be purchased separately at the price indicated below. CAMRAD is written in FORTRAN 77 for the DEC VAX under VMS 4.6 with a recommended core memory of 4.04 megabytes. The DISSPLA package is necessary for graphical output. CAMRAD was developed in 1980.

Beschreibung: ARC2D is a computational fluid dynamics program developed at the NASA Ames Research Center specifically for airfoil computations. The program uses implicit finite-difference techniques to solve two-dimensional Euler equations and thin layer Navier-Stokes equations. It is based on the Beam and Warming implicit approximate factorization algorithm in generalized coordinates. The methods are either time accurate or accelerated non-time accurate steady state schemes. The evolution of the solution through time is physically realistic; good solution accuracy is dependent on mesh spacing and boundary conditions. The mathematical development of ARC2D begins with the strong conservation law form of the two-dimensional Navier-Stokes equations in Cartesian coordinates, which admits shock capturing. The Navier-Stokes equations can be transformed from Cartesian coordinates to generalized curvilinear coordinates in a manner that permits one computational code to serve a wide variety of physical geometries and grid systems. ARC2D includes an algebraic mixing length model to approximate the effect of turbulence. In cases of high Reynolds number viscous flows, thin layer approximation can be applied. ARC2D allows for a variety of solutions to stability boundaries, such as those encountered in flows with shocks. The user has considerable flexibility in assigning geometry and developing grid patterns, as well as in assigning boundary conditions. However, the ARC2D model is most appropriate for attached and mildly separated boundary layers; no attempt is made to model wake regions and widely separated flows. The techniques have been successfully used for a variety of inviscid and viscous flowfield calculations. The Cray version of ARC2D is written in FORTRAN 77 for use on Cray series computers and requires approximately 5Mb memory. The program is fully vectorized. The tape includes variations for the COS and UNICOS operating systems. Also included is a sample routine for CONVEX computers to emulate Cray system time calls, which should be easy to modify for other machines as well. The standard distribution media for this version is a 9-track 1600 BPI ASCII Card Image format magnetic tape. The Cray version was developed in 1987. The IBM ES/3090 version is an IBM port of the Cray version. It is written in IBM VS FORTRAN and has the capability of executing in both vector and parallel modes on the MVS/XA operating system and in vector mode on the VM/XA operating system. Various options of the IBM VS FORTRAN compiler provide new features for the ES/3090 version, including 64-bit arithmetic and up to 2 GB of virtual addressability. The IBM ES/3090 version is available only as a 9-track, 1600 BPI IBM IEBCOPY format magnetic tape. The IBM ES/3090 version was developed in 1989. The DEC RISC ULTRIX version is a DEC port of the Cray version. It is written in FORTRAN 77 for RISC-based Digital Equipment platforms. The memory requirement is approximately 7Mb of main memory. It is available in UNIX tar format on TK50 tape cartridge. The port to DEC RISC ULTRIX was done in 1990. COS and UNICOS are trademarks and Cray is a registered trademark of Cray Research, Inc. IBM, ES/3090, VS FORTRAN, MVS/XA, and VM/XA are registered trademarks of International Business Machines. DEC and ULTRIX are registered trademarks of Digital Equipment Corporation.

Beschreibung: Panel methods are moderate cost tools for solving a wide range of engineering problems. PMARC (Panel Method Ames Research Center) is a potential flow panel code that numerically predicts flow fields around complex three-dimensional geometries. PMARC's predecessor was a panel code named VSAERO which was developed for NASA by Analytical Methods, Inc. PMARC is a new program with many additional subroutines and a well-documented code suitable for powered-lift aerodynamic predictions. The program's open architecture facilitates modifications or additions of new features. Another improvement is the adjustable size code which allows for an optimum match between the computer hardware available to the user and the size of the problem being solved. PMARC can be resized (the maximum number of panels can be changed) in a matter of minutes. Several other state-of-the-art PMARC features include internal flow modeling for ducts and wind tunnel test sections, simple jet plume modeling essential for the analysis and design of powered-lift aircraft, and a time-stepping wake model which allows the study of both steady and unsteady motions. PMARC is a low-order panel method, which means the singularities are distributed with constant strength over each panel. In many cases low-order methods can provide nearly the same accuracy as higher order methods (where the singularities are allowed to vary linearly or quadratically over each panel). Low-order methods have the advantage of a shorter computation time and do not require exact matching between panels. The flow problem is solved by assuming that the body is at rest in a moving flow field. The body is modeled as a closed surface which divides space into two regions -- one region contains the flow field of interest and the other contains a fictitious flow. External flow problems, such as a wing in a uniform stream, have the external region as the flow field of interest and the internal flow as the fictitious flow. This arrangement is reversed for internal flow problems where the internal region contains the flow field of interest and the external flow field is fictitious. In either case it is assumed that the velocity potentials in both regions satisfy Laplace's equation. PMARC has extensive geometry modeling capabilities for handling complex, three-dimensional surfaces. As with all panel methods, the geometry must be modeled by a set of panels. For convenience, the geometry is usually subdivided into several pieces and modeled with sets of panels called patches. A patch may be folded over on itself so that opposing sides of the patch form a common line. For example, wings are normally modeled with a folded patch to form the trailing edge of the wing. PMARC also has the capability to automatically generate a closing tip patch. In the case of a wing, a tip patch could be generated to close off the wing's third side. PMARC has a simple jet model for simulating a jet plume in a crossflow. The jet plume shape, trajectory, and entrainment velocities are computed using the Adler/Baron jet in crossflow code. This information is then passed back to PMARC. The wake model in PMARC is a time-stepping wake model. The wake is convected downstream from the wake separation line by the local velocity flowfield. With each time step, a new row of wake panels is added to the wake at the wake separation line. PMARC also allows an initial wake to be specified if desired, or, as a third option, no wakes need be modeled. The effective presentation of results for aerodynamics problems requires the generation of report-quality graphics. PMAPP (ARC-12751), the Panel Method Aerodynamic Plotting Program, (Sterling Software), was written for scientists at NASA's Ames Research Center to plot the aerodynamic analysis results (flow data) from PMARC. PMAPP is an interactive, color-capable graphics program for the DEC VAX or MicroVAX running VMS. It was designed to work with a variety of terminal types and hardcopy devices. PMAPP is available separately from COSMIC. PMARC was written in standard FORTRAN77 using adjustable size arrays throughout the code. Redimensioning PMARC will change the amount of disk space and memory the code requires to be able to run; however, due to its memory requirements, this program does not readily lend itself to implementation on MS-DOS based machines. The program was implemented on an Apple Macintosh (using 2.5 MB of memory) and tested on a VAX/VMS computer. The program is available on a 3.5 inch Macintosh format diskette (standard media) or in VAX BACKUP format on TK50 tape cartridge or 9-track magnetic tape. PMARC was developed in 1989.

Beschreibung: The Transonic Airfoil analysis computer code, TAIR, was developed to employ a fast, fully implicit algorithm to solve the conservative full-potential equation for the steady transonic flow field about an arbitrary airfoil immersed in a subsonic free stream. The full-potential formulation is considered exact under the assumptions of irrotational, isentropic, and inviscid flow. These assumptions are valid for a wide range of practical transonic flows typical of modern aircraft cruise conditions. The primary features of TAIR include: a new fully implicit iteration scheme which is typically many times faster than classical successive line overrelaxation algorithms; a new, reliable artifical density spatial differencing scheme treating the conservative form of the full-potential equation; and a numerical mapping procedure capable of generating curvilinear, body-fitted finite-difference grids about arbitrary airfoil geometries. Three aspects emphasized during the development of the TAIR code were reliability, simplicity, and speed. The reliability of TAIR comes from two sources: the new algorithm employed and the implementation of effective convergence monitoring logic. TAIR achieves ease of use by employing a "default mode" that greatly simplifies code operation, especially by inexperienced users, and many useful options including: several airfoil-geometry input options, flexible user controls over program output, and a multiple solution capability. The speed of the TAIR code is attributed to the new algorithm and the manner in which it has been implemented. Input to the TAIR program consists of airfoil coordinates, aerodynamic and flow-field convergence parameters, and geometric and grid convergence parameters. The airfoil coordinates for many airfoil shapes can be generated in TAIR from just a few input parameters. Most of the other input parameters have default values which allow the user to run an analysis in the default mode by specifing only a few input parameters. Output from TAIR may include aerodynamic coefficients, the airfoil surface solution, convergence histories, and printer plots of Mach number and density contour maps. The TAIR program is written in FORTRAN IV for batch execution and has been implemented on a CDC 7600 computer with a central memory requirement of approximately 155K (octal) of 60 bit words. The TAIR program was developed in 1981.

Beschreibung: The TAWFIVE program calculates transonic flow over a transport-type wing and fuselage. Although more complex Euler and Navier-Stokes methods are available, TAWFIVE combines a multi-grid acceleration technique in the iterative solution of the potential equation with the use of integral-form boundary-layer equations to provide a computationally efficient and sufficiently accurate design tool. TAWFIVE simplifies the solution process by breaking the problem into a loosely coupled set of modified equations. The inviscid method, using standard inviscid equations (nonlinear full potential), is valid in the "outer" region away from the wing, whereas the boundary-layer equations are valid in the thin region near the solid surface of the wing. The two types of equations are coupled by a technique of modifying surface boundary conditions for the inviscid equations. This interaction process starts with a solution of the outer flow field. Pressures are computed at the wing surface and are used to calculate the boundary layer. The boundary-layer and wake properties are then computed using a three-dimensional integral method, and the computed displacement thickness is added to the surface of the "hard" geometry. This new displaced wing surface is then regridded and the inviscid flowfield is recomputed. New values of the inviscid pressures are then used by the boundary-layer method to predict a new displacement thickness distribution. An under-relaxed update of the previously predicted displacement thickness is then made to obtain a new displacement thickness correction that is added to the "hard" geometry. These global iterations are continued until suitable convergence is obtained. Input to TAWFIVE is limited to geometric definition of the configuration, free-stream flow quantities, and iteration control parameters. The geometric input consists of the definition of a series of airfoil sections to define the wing and a series of fuselage cross sections to model the fuselage. High-aspect-ratio wings are modeled more accurately than low-aspect-ratio wings since no special provisions are made to accurately model the wing-fuselage juncture or the wingtip region. The user can specify the solution either in terms of lift or in terms of angle of attack. TAWFIVE can produce tabular output and input files for PLOT3D (COSMIC program number ARC-12779). TAWFIVE is written in FORTRAN 77 for CRAY series computers running UNICOS. The main memory requirement is 2.7Mb for execution. This program is available on a 9-track 1600 BPI UNIX tar format magnetic tape. TAWFIVE was under development from 1979 to 1989 and first released by COSMIC in 1991. CRAY and UNICOS are registered trademarks of Cray Research, Inc.

Beschreibung: A summary is presented of vortex control applications and current techniques for the control of longitudinal vortices produced by bodies, leading edges, tips and intersections. Vortex control has up till now been performed by many approaches in an empirical fashion, assisted by the essentially inviscid nature of much of longitudinal vortex behavior. Attention is given to Reynolds number sensitivities, vortex breakdown and interactions, vortex control on highly swept wings, and vortex control in juncture flows.

Beschreibung: On an hourly time-scale the different roles of the solar wind and interplanetary magnetic field (IMF) parameters on ground micropulsation activity can be better investigated than at longer time-scales. A long-term comparison between ground measurements made at L'Aquila and IMP 8 observations confirms the solar wind speed as the key parameter for the onset of pulsations even at low latitudes, although additional control of the energy transfer from the interplanetary medium to the earth's magnetosphere is clearly exerted by the cone angle. Above about 20 mHz the frequency of pulsations is confirmed to be closely related to the IMF magnitude while, in agreement with model predictions, the IMF magnitude is related to the amplitude of the local fundamental resonant mode. We provide an interesting example in which high resolution measurements simultaneously obtained in the foreshock region and on the ground show that external transversal fluctuations do not penetrate deep into the low latitude magnetosphere.

Beschreibung: This paper discusses in situ measurements of O3, aerosol number density, and aerosol size obtained during NASA Arctic Boundary Layer Expedition aircraft flights over the Alaskan Arctic region. The major source of summer O3 for the troposphere in the intrusion of stratospheric air and subsequent transport to lower altitudes. Photochemistry of mixed layer emissions and O3 transported from high northern latitude urban/industrialized areas do not appear to play major roles as sources of O3 for the Alaska region. O3 gradients reflect the loss at the surface and supply from the stratosphere. Free tropospheric O3 averaged 74 ppbv compared to 32 ppbv for the mixed layer. O3 loss mechanisms are a combination of the destruction via photochemistry, chemical reaction with surface emissions, and direct loss through deposition to the surface. The boreal forest in the most efficient of the O3 sinks and has the largest increase in aerosol number density relative to the free troposphere.

Beschreibung: The flow-field within an axial flow turbomachine, such as a turbine or compressor, is extremely complex because of three-dimensional features such as hub-corner stall, tip-leakage flows, and airfoil wakes. These flow features interact with each other and with rotor and stator airfoils inducing time-varying forces on the airfoils. These complicated rotor-stator interactions must be understood in order to design turbomachines that are light and compact as well as reliable and efficient. Two codes, STAGE-2 and STAGE-3, have been developed to compute these unsteady rotor-stator interaction flows in multistage turbomachines. An implicit, thin-layer Euler/Navier-Stokes zonal algorithm is used to compute the unsteady flow-field within both turbine and compressor configurations. Results include surface pressures and wake profiles for two-dimensional turbine and compressor configurations and surface pressures for a three-dimensional single-stage turbine configuration. The results compare well with experimental data and other unsteady computations.

Beschreibung: The most important investigations leading to the International Reference Ionosphere 1990 (IRI-90) are overviewed, and the latest version of the model is described. The shortcomings and limitations of the IRI-90 are pointed out, together with the ways of overcoming them. The list of studies that the IRI group has yet to carry out includes the investigations of magnetic storm effects as the highest priority. This paper discusses determinations of and the available data on the electron density, plasma temperatures, ion composition, and ion drift in the ionosphere, together with future improvements needed on these parameters.

Beschreibung: In the aftermath of the eruption of Mt. Pinatubo, multiwavelength stratospheric aerosol extinction measurements by the satellite-borne Stratospheric Aerosol and Gas Experiment (SAGE II) revealed the presence of a previously unobserved mode of aerosol that exhibited high extinction but a small inferred particle size. This mode may represent a transitional phase between the very small aerosol created by gas-to-particle conversion and a quasi-steady state, post-volcanic aerosol that exhibits both large extinction and large particle size. The presence of a transitional small aerosol mode may have a significant impact on chemical and radiative processes in the stratosphere.

Beschreibung: CO2 exchange rates were measured at selected tundra sites near Bethel, Alaska using portable, climate-controlled, instrumented enclosures. The empirically modeled exchange rate for a representative area of vegetated tundra was 1.2 +/- 1.2 g/sq m/d, compared to a tower-measured exchange over the same time period of 1.1 +.0- 1.2 g/sq m/d. Net exchange in response to varying light levels was compared to wet meadow and dry upland tundra, and to the net exchange measured by the micrometeoroidal tower technique. The multispectral reflectance properties of the sites were measured and related to exchange rates in order to provide a quantitative foundation for the use of satellite remote sensing to monitor biosphere/atmosphere CO2 exchange in the tundra biome.

Beschreibung: Enhanced concentrations of CH4 in the unpolluted atmospheric mixed layer over both Arctic and subarctic tundra landscapes are documented here using data from the NASA Arctic Boundary Layer Expedition (ABLE 3A). The CH4 concentration gradients were determined mainly by interactions of biogenic emission from wet tundra and turbulent mixing proceses. The gradient were most frequently associated with intrusion of upper tropospheric or stratospheric air into the midtroposphere, emissions from forest and tundra fires, and long-range transport of enhanced concentration of these gases from unidentified sources. Summertime haze layers exhibited midtropospheric enhancements of CH4 similar to those measured in winter Arctic events. The observations confirm the importance of Arctic and Subarctic wetland environments as a regional source of global atmospheric CH4.

Beschreibung: The oxidation of the earth's crust and the increase in atmospheric oxygen early in earth history have been linked to the accumulation of reduced carbon in sedimentary rocks. Trends in the carbon isotope composition of sedimentary organic carbon and carbonate show that during the Proterozoic aeon (2.5-0.54 Gyr ago) the organic carbon reservoir grew in size, relative to the carbonate reservoir. This increase, and the concomitant release of oxidizing power in the environment, occurred mostly during episodes of global rifting and orogeny.

Beschreibung: Recently available spectral O3 line parameters are summarized and progress in remote sensing based on these parameters is discussed. Particular attention is given to covering line positions, line intensities and linewidths for both the main isotopic species (O-16)3 and the isotopic variants (O-16)(O-18)(O-16) and (O-16)(O-16)(O-18).

Beschreibung: Jet noise and jet-induced structural loads have become key issues in the design of commercial and military aircraft. Computational Fluid Dynamics (CFD) can be of use in predicting the underlying jet shear-layer instabilities and, in conjunction with classical acoustic theory, jet noise. The computational issues involved in the resolution of high Reynolds number unsteady jet flows are addressed in this paper. Once these jet flows can be accurately resolved, it should be possible to use acoustic theory to extract, for example, the far-field jet noise. An assessment of future work and computational resources required for directly computing far-field jet noise is also presented.

Beschreibung: The paper presents Alouette 2/ISIS 1 data which furnish evidence for the stimulation of high-order D(n) resonances (n greater than 4) by topside sounders, confirming the prediction of Osherovich (1990) based on an analogy with naturally occurring narrowband magnetospheric emissions. The results indicate that observations of stimulated ionospheric emissions and naturally occurring magnetospheric emissions can be used as complementary data sets to address such fundamental questions as the nature of the excitation mechanism for these emissions and the nature of the waves (i.e., whether the waves predominantly electrostatic or with a significant magnetic component present), and the question of the dominance among the Dn, Dn(+), and Dn(-) resonances and the conditions required for all three to be present at the same time.

Beschreibung: Ground magnetic field perturbations recorded by the CANOPUS magnetometer network in the 7 to 13 MLT sector are used to examine how reconfigurations of the dayside polar ionospheric flow take place in response to north-south changes of the IMF. During the 6-h interval in question, IMF Bz oscillates between +/- 7 nT with about a 1-h period. Corresponding variations in the ground magnetic disturbance are observed which we infer are due to changes in ionospheric flow. Cross correlation of the data obtained from two ground stations at 73.5 deg magnetic latitude, but separated by about 2 hours in MLT, shows that changes in the flow are initiated in the prenoon sector (about 10 MLT) and then spread outward toward dawn and dusk with a phase speed of about 5 km/s over the longitude range about 8 to 12 MLT, slowing to about 2 km/s outside this range. Cross correlating the data from these ground stations with IMP 8 IMF Bz records produces a MLT variation in the ground response delay relative to the IMF which is compatible with these deduced phase speeds.

Beschreibung: The particle scattering and current sheet stability features in the geomagnetic tail during the phase of substorm growth were investigated using Tsyganenko's (1989) magnetic field model. In a study of four substorm events which were observed both in the high-altitude nightside tail and in the auroral ionosphere, the model magnetic field was adjusted to each case so as to represent the global field development during the growth phase of the substorms. The model results suggest that the auroral brightenings are connected with processes taking place in the near-earth region inside about 15 earth radii. The results also suggest that there is a connection between the chaotization of the electrons and the auroral brightenings at substorm onset.

Beschreibung: Alternative geometries for the present-day configuration of plate boundaries in northeast Asia and Japan are tested using NUVEL-1 and 256 horizontal earthquake slip vectors from the Japan and northern Kuril trenches. Statistical analysis of the slip vectors is used to determine whether the North American, Eurasian, or Okhotsk plate overlies the trench. Along the northern Kuril trench, slip vectors are well-fit by the NUVEL-1 Pacific-North America Euler pole, but are poorly fit by the Pacific-Eurasia Euler pole. Results for the Japan trench are less conclusive, but suggest that much of Honshu and Hokkaido are also part of the North American plate. The simplest geometry consistent with the trench slip vectors is a geometry in which the North American plate extends south to 41 deg N, and possibly includes northern Honshu and southern Hokkaido. Although these results imply that the diffuse seismicity that connects the Lena River delta to Sakhalin Island and the eastern Sea of Japan records motion between Eurasia and North America, onshore geologic and seismic data define an additional belt of seismicity in Siberia that cannot be explained with this geometry. Assuming that these two seismic belts constitute evidence for an Okhotsk block, two published kinematic models for motion of the Okhotsk block are tested. The first model, which predicts motion of up to 15 mm/yr relative to North America, is rejected because Kuril and Japan trench slip vectors are fit more poorly than for the simpler geometry described above. The second model gives a good fit to the trench slip vectors, but only if Okhotsk-North America motion is slower than 5 mm/yr.

Beschreibung: This computer program is designed to calculate the flow fields in two-dimensional and three-dimensional axisymmetric supersonic inlets. The method of characteristics is used to compute arrays of points in the flow field. At each point the total pressure, local Mach number, local flow angle, and static pressure are calculated. This program can be used to design and analyze supersonic inlets by determining the surface compression rates and throat flow properties. The program employs the method of characteristics for a perfect gas. The basic equation used in the program is the compatibility equation which relates the change in stream angle to the change in entropy and the change in velocity. In order to facilitate the computation, the flow field behind the bow shock wave is broken into regions bounded by shock waves. In each region successive rays are computed from a surface to a shock wave until the shock wave intersects a surface or falls outside the cowl lip. As soon as the intersection occurs a new region is started and the previous region continued only in the area in which it is needed, thus eliminating unnecessary calculations. The maximum number of regions possible in the program is ten, which allows for the simultaneous calculations of up to nine shock waves. Input to this program consists of surface contours, free-stream Mach number, and various calculation control parameters. Output consists of printed and/or plotted results. For plotted results an SC-4020 or similar plotting device is required. This program is written in FORTRAN IV to be executed in the batch mode and has been implemented on a CDC 7600 with a central memory requirement of approximately 27k (octal) of 60 bit words.

Beschreibung: This program was developed to predict turbine stage performance taking into account the effects of complex passage geometries. The method uses a quasi-3D inviscid-flow analysis iteratively coupled to calculated losses so that changes in losses result in changes in the flow distribution. In this manner the effects of both the geometry on the flow distribution and the flow distribution on losses are accounted for. The flow may be subsonic or shock-free transonic. The blade row may be fixed or rotating, and the blades may be twisted and leaned. This program has been applied to axial and radial turbines, and is helpful in the analysis of mixed flow machines. This program is a combination of the flow analysis programs MERIDL and TSONIC coupled to the boundary layer program BLAYER. The subsonic flow solution is obtained by a finite difference, stream function analysis. Transonic blade-to-blade solutions are obtained using information from the finite difference, stream function solution with a reduced flow factor. Upstream and downstream flow variables may vary from hub to shroud and provision is made to correct for loss of stagnation pressure. Boundary layer analyses are made to determine profile and end-wall friction losses. Empirical loss models are used to account for incidence, secondary flow, disc windage, and clearance losses. The total losses are then used to calculate stator, rotor, and stage efficiency. This program is written in FORTRAN IV for batch execution and has been implemented on an IBM 370/3033 under TSS with a central memory requirement of approximately 4.5 Megs of 8 bit bytes. This program was developed in 1985.

Beschreibung: Turbomachinery components are often connected by ducts, which are usually annular. The configurations and aerodynamic characteristics of these ducts are crucial to the optimum performance of the turbomachinery blade rows. The ANDUCT computer program was developed to calculate the velocity distribution along an arbitrary line between the inner and outer walls of an annular duct with axisymmetric swirling flow. Although other programs are available for duct analysis, the use of the velocity gradient method makes the ANDUCT program fast and convenient while requiring only modest computer resources. A fast and easy method of analyzing the flow through a duct with axisymmetric flow is the velocity gradient method, also known as the stream filament or streamline curvature method. This method has been used extensively for blade passages but has not been widely used for ducts, except for the radial equilibrium equation. In ANDUCT, a velocity gradient equation derived from the momentum equation is used to determine the velocity variation along an arbitrary straight line between the inner and outer wall of an annular duct. The velocity gradient equation is used with an assumed variation of meridional streamline curvature. Upstream flow conditions may vary between the inner and outer walls, and an assumed total pressure distribution may be specified. ANDUCT works best for well-guided passages and where the curvature of the walls is small as compared to the width of the passage. The ANDUCT program is written in FORTRAN IV for batch execution and has been implemented on an IBM 370 series computer with a central memory requirement of approximately 60K of 8 bit bytes. The ANDUCT program was developed in 1982.

Beschreibung: The Panel Code for Planar Cascades was developed as an aid for the designer of turbomachinery blade rows. The effective design of turbomachinery blade rows relies on the use of computer codes to model the flow on blade-to-blade surfaces. Most of the currently used codes model the flow as inviscid, irrotational, and compressible with solutions being obtained by finite difference or finite element numerical techniques. While these codes can yield very accurate solutions, they usually require an experienced user to manipulate input data and control parameters. Also, they often limit a designer in the types of blade geometries, cascade configurations, and flow conditions that can be considered. The Panel Code for Planar Cascades accelerates the design process and gives the designer more freedom in developing blade shapes by offering a simple blade-to-blade flow code. Panel, or integral equation, solution techniques have been used for several years by external aerodynamicists who have developed and refined them into a primary design tool of the aircraft industry. The Panel Code for Planar Cascades adapts these same techniques to provide a versatile, stable, and efficient calculation scheme for internal flow. The code calculates the compressible, inviscid, irrotational flow through a planar cascade of arbitrary blade shapes. Since the panel solution technique is for incompressible flow, a compressibility correction is introduced to account for compressible flow effects. The analysis is limited to flow conditions in the subsonic and shock-free transonic range. Input to the code consists of inlet flow conditions, blade geometry data, and simple control parameters. Output includes flow parameters at selected control points. This program is written in FORTRAN IV for batch execution and has been implemented on an IBM 370 series computer with a central memory requirement of approximately 590K of 8 bit bytes. This program was developed in 1982.

Beschreibung: An exact, full-potential-equation model for the steady, irrotational, homoentropic, and homoenergetic flow of a compressible, inviscid fluid through a two-dimensional planar cascade together with its appropriate boundary conditions has been derived. The CAS2D computer program numerically solves an artificially time-dependent form of the actual full-potential-equation, providing a nonrotating blade-to-blade, steady, potential transonic cascade flow analysis code. Comparisons of results with test data and theoretical solutions indicate very good agreement. In CAS2D, the governing equation is discretized by using type-dependent, rotated finite differencing and the finite area technique. The flow field is discretized by providing a boundary-fitted, nonuniform computational mesh. This mesh is generated by using a sequence of conformal mapping, nonorthogonal coordinate stretching, and local, isoparametric, bilinear mapping functions. The discretized form of the full-potential equation is solved iteratively by using successive line over relaxation. Possible isentropic shocks are captured by the explicit addition of an artificial viscosity in a conservative form. In addition, a four-level, consecutive, mesh refinement feature makes CAS2D a reliable and fast algorithm for the analysis of transonic, two-dimensional cascade flows. The results from CAS2D are not directly applicable to three-dimensional, potential, rotating flows through a cascade of blades because CAS2D does not consider the effects of the Coriolis force that would be present in the three-dimensional case. This program is written in FORTRAN IV for batch execution and has been implemented on an IBM 370 series computer with a central memory requirement of approximately 200K of 8 bit bytes. The CAS2D program was developed in 1980.

Beschreibung: A computer program, QSONIC, has been developed for calculating the full potential, transonic quasi-three-dimensional flow through a rotating turbomachinery blade row. The need for lighter, more efficient turbomachinery components has led to the consideration of machines with fewer stages, each with blades capable of higher speeds and higher loading. As speeds increase, the numerical problems inherent in the transonic regime have to be resolved. These problems include the calculation of imbedded shock discontinuities and the dual nature of the governing equations, which are elliptic in the subcritical flow regions but become hyperbolic for supersonic zones. QSONIC provides the flow analyst with a fast and reliable means of obtaining the transonic potential flow distribution on a blade-to-blade stream surface of a stationary or rotating turbomachine blade row. QSONIC combines several promising transonic analysis techniques. The full potential equation in conservative form is discretized at each point on a body-fitted period mesh. A mass balance is calculated through the finite volume surrounding each point. Each local volume is corrected in the third dimension for any change in stream-tube thickness along the stream tube. The nonlinear equations for all volumes are of mixed type (elliptic or hyperbolic) depending on the local Mach number. The final result is a block-tridiagonal matrix formulation involving potential corrections at each grid point as the unknowns. The residual of each system of equations is solved along each grid line. At points where the Mach number exceeds unity, the density at the forward (sweeping) edge of the volume is replaced by an artificial density. This method calculates the flow field about a cascade of arbitrary two-dimensional airfoils. Three-dimensional flow is approximated in a turbomachinery blade row by correcting for stream-tube convergence and radius change in the through flow direction. Several significant assumptions were made in developing the QSONIC program, including: (1) the flow is inviscid and adiabatic, (2) the flow relative to the blade is steady, (3) the fluid is a perfect gas with constant specific heat, (4) the flow is isentropic and any discontinuities (shocks) are weak enough to be approximated as isentropic jumps, (5) there is no velocity component normal to the stream surface, and (6) the flow relative to a fixed frame in space (absolute velocity) is completely irrotational. These assumptions place some limitations on the application of QSONIC. Sharp leading edges at high incidence and high-Mach-number turbine blade trailing edges with substantial deviation will both cause large velocity peaks on the blade. In addition, the program may have difficulty converging if the passage is nearly choked. Input to QSONIC consists of case control parameters, a geometry description, upstream boundary conditions, and a rotor description. Output includes solution scheme parameters and flow field parameters. A data file is also output which contains data on the solution mesh, surface Mach numbers, surface static pressures, isomachs, and the velocity vector field. This data may be used for further processing or for plotting. The QSONIC is written in FORTRAN IV for batch execution and has been implemented on an IBM 370 series computer with a central memory requirement of approximately 500K of 8 bit bytes. QSONIC was developed in 1982.

Beschreibung: This computer program, WIND, was developed to numerically solve the exact, full-potential equation for three-dimensional, steady, inviscid flow through an isolated wind turbine rotor. The program automatically generates a three-dimensional, boundary-conforming grid and iteratively solves the full-potential equation while fully accounting for both the rotating and Coriolis effects. WIND is capable of numerically analyzing the flow field about a given blade shape of the horizontal-axis type wind turbine. The rotor hub is assumed representable by a doubly infinite circular cylinder. An arbitrary number of blades may be attached to the hub and these blades may have arbitrary spanwise distributions of taper and of the twist, sweep, and dihedral angles. An arbitrary number of different airfoil section shapes may be used along the span as long as the spanwise variation of all the geometeric parameters is reasonably smooth. The numerical techniques employed in WIND involve rotated, type-dependent finite differencing, a finite volume method, artificial viscosity in conservative form, and a successive overrelaxation combined with the sequential grid refinement procedure to accelerate the iterative convergence rate. Consequently, WIND is cabable of accurately analyzing incompressible and compressible flows, including those that are locally transonic and terminated by weak shocks. Along with the three-dimensional results, WIND provides the results of the two-dimensional calculations to aid the user in locating areas of possible improvement in the aerodynamic design of the blade. Output from WIND includes the chordwise distribution of the coefficient of pressure, the Mach number, the density, and the relative velocity components at spanwise stations along the blade. In addition, the results specify local values of the lift coefficient and the tangent and axial aerodynamic force components. These are also given in integrated form expressing the total torque and the total axial force acting on the shaft. WIND can also be used to analyze the flow around isolated aircraft propellers and helicopter rotors in hover as long as the relative oncoming flow is subsonic. The WIND program is written in FORTRAN IV for batch execution and has been implemented on an IBM 370 series computer with a central memory requirement of approximately 253K of 8 bit bytes. WIND was developed in 1980.

Beschreibung: This computer program calculates the flow field in the supersonic portion of a mixed-compression aircraft inlet at non-zero angle of attack. This approach is based on the method of characteristics for steady three-dimensional flow. The results of this program agree with those produced by the two-dimensional method of characteristics when axisymmetric flow fields are calculated. Except in regions of high viscous interaction and boundary layer removal, the results agree well with experimental data obtained for threedimensional flow fields. The flow field in a variety of axisymmetric mixed compression inlets can be calculated using this program. The bow shock wave and the internal shock wave system are calculated using a discrete shock wave fitting procedure. The internal flow field can be calculated either with or without the discrete fitting of the internal shock wave system. The influence of molecular transport can be included in the calculation of the external flow about the forebody and in the calculation of the internal flow when internal shock waves are not discretely fitted. The viscous and thermal diffussion effects are included by treating them as correction terms in the method of characteristics procedure. Dynamic viscosity is represented by Sutherland's law and thermal conductivity is represented as a quadratic function of temperature. The thermodynamic model used is that of a thermally and calorically perfect gas. The program assumes that the cowl lip is contained in a constant plane and that the centerbody contour and cowl contour are smooth and have continuous first partial derivatives. This program cannot calculate subsonic flow, the external flow field if the bow shock wave does not exist entirely around the forebody, or the internal flow field if the bow flow field is injected into the annulus. Input to the program consists of parameters to control execution, to define the geometry, and the vehicle orientation. Output consists of a list of parameters used, solution planes, and a description of the shock waves. This program is written in FORTRAN IV for batch execution and has been implemented on a CDC 6000 series machine with a central memory requirement of 110K (octal) of 60 bit words when it is overlayed. This flow analysis program was developed in 1978.

Beschreibung: This computer program was developed for calculating the subsonic or transonic flow on the hub-shroud mid-channel stream surface of a single blade row of a turbomachine. The design and analysis of blades for compressors and turbines ideally requires methods for analyzing unsteady, three-dimensional, turbulent viscous flow through a turbomachine. Since an exact solution is impossible at present, solutions on two-dimensional surfaces are calculated to obtain a quasi-three dimensional solution. When three-dimensional effects are important, significant information can be obtained from a solution on a cross-sectional surface of the passage normal to the flow. With this program, a solution to the equations of flow on the meridional surface can be carried out. This solution is chosen when the turbomachine under consideration has significant variation in flow properties in the hubshroud direction, especially when input is needed for use in blade-to-blade calculations. The program can also perform flow calculations for annular ducts without blades. This program should prove very useful in the design and analysis of any turbomachine. This program calculates a solution for two-dimensional, adiabatic shockfree flow. The flow must be essentially subsonic, but there may be local areas of supersonic flow. To obtain the solution, this program uses both the finite difference and the quasi-orthogonal (velocity gradient) methods combined in a way that takes maximum advantage of both. The finite-difference method solves a finite-difference equation along the meridional stream surface in a very efficient manner but is limited to subsonic velocities. This approach must be used in cases where the blade aspect ratios are above one, cases where the passage is curved, and cases with low hub-tip-ratio blades. The quasi-orthogonal method solves the velocity gradient equation on the meridional surface and is used if it is necessary to extend the range of solutions into the transonic regime. In general the blade row may be fixed or rotating and the blades may be twisted and leaned. The flow may be axial, radial, or mixed. The upstream and downstream flow conditions can vary from hub to shroud with provisions made for an approximate correction for loss of stagnation pressure. Also, viscous forces are neglected along solution mesh lines running from hub to tip. The capabilities of this program include handling of nonaxial flows without restriction, annular ducts without blades, and specified streamwise loss distributions. This program is written in FORTRAN IV for batch execution and has been implemented on an IBM 360 computer with a central memory requirement of approximately 700K of 8 bit bytes. This core requirement can be reduced depending on the size of the problem and the desired solution accuracy. This program was developed in 1977.

Beschreibung: A computer program has been developed for the design of supersonic rotor blades where losses are accounted for by correcting the ideal blade geometry for boundary layer displacement thickness. The ideal blade passage is designed by the method of characteristics and is based on establishing vortex flow within the passage. Boundary-layer parameters (displacement and momentum thicknesses) are calculated for the ideal passage, and the final blade geometry is obtained by adding the displacement thicknesses to the ideal nozzle coordinates. The boundary-layer parameters are also used to calculate the aftermixing conditions downstream of the rotor blades assuming the flow mixes to a uniform state. The computer program input consists essentially of the rotor inlet and outlet Mach numbers, upper- and lower-surface Mach numbers, inlet flow angle, specific heat ratio, and total flow conditions. The program gas properties are set up for air. Additional gases require changes to be made to the program. The computer output consists of the corrected rotor blade coordinates, the principal boundary-layer parameters, and the aftermixing conditions. This program is written in FORTRAN IV for batch execution and has been implemented on an IBM 7094. This program was developed in 1971.

Beschreibung: This program obtains a transonic flow solution on a blade-to-blade surface between blades of a turbomachine. The flow must be essentially subsonic, but there may be locally supersonic flow. The solution is two-dimensional, isentropic, and shock free. The blades may be fixed or rotating. The flow may be axial, radial, or mixed, and there may be a change in stream-channel thickness in the through-flow direction. A loss in relative stagnation pressure may be accounted for. The program input consists of blade and stream-channel geometry, stagnation flow conditions, inlet and outlet flow angles, and blade-to-blade stream-channel weight flow. The output includes blade surface velocities, velocity magnitude and direction at all interior mesh points in the blade-to-blade passage, and streamline coordinates throughout the passage. The transonic solution is obtained by a combination of a finite-difference, stream-function solution and a velocity-gradient solution. The finite-difference solution at a reduced weight flow provides information needed to obtain a velocity-gradient solution. This program is written in FORTRAN IV for batch execution and has been implemented on the IBM 360 computer with a central memory requirement of approximately 36K of 8 bit bytes. This program was developed in 1969 and last updated in 1979.

Beschreibung: This program is a revision of an existing program for blade-to-blade aerodynamic analysis of turbomachine blades and it is a simpler program while consistent with related programs. The analysis is for two-dimensional, subsonic, compressible (or incompressible), nonviscous flow in a circular or straight infinite cascade of blades, which may be fixed or rotating. The flow may be axial, radial, or mixed, and the stream channel thickness may change in the through-flow direction. The program input consists of blade and stream channel geometry, total flow conditions, inlet and outlet flow angles, and blade-to-blade stream channel weight flow. The output includes blade surface velocities, velocity magnitude and direction at all interior mesh points in the blade-to-blade passage, and streamline coordinates throughout the passage. This program was developed on an IBM 7094/7044 DCS.

Beschreibung: This computer program gives the blade-to-blade solution of the two-dimensional, subsonic, compressible (or incompressible), nonviscous flow problem for a circular or straight infinite cascade of tandem or slotted turbomachine blades. The blades may be fixed or rotating. The flow may be axial, radial , or mixed. The method of solution is based on the stream function using an iterative solution of nonlinear finite-difference equations. These equations are solved using two major levels of iteration. The inner iteration consists of the solution of simultaneous linear equations by successive over-relaxation, using an estimated optimum over-relaxation factor. The outer iteration then changes the coefficients of the simultaneous equations to correct for compressibility. The program input consists of the basic blade geometry, the meridional stream channel coordinates, fluid stagnation conditions, weight flow and flow split through the slot, and inlet and outlet flow angles. The output includes blade surface velocities, velocity magnitude and direction throughout the passage, and the streamline coordinates.

Beschreibung: This FORTRAN IV computer program which incorporates the method of characteristics was written to assist in the design of supersonic inlets. There were two objectives: (1) to study a greater variety of supersonic inlet configurations and (2) to reduce the time required for trial-and-error procedures to arrive at optimum inlet design. The computer program was written with the intention of being able to construct a variety of inlet configurations by interchanging specific subroutines. In this manner, greater flexibility of choice was attained, and the time required to program a specific inlet configuration was greatly reduced. The second objective was accomplished by a reformulation of the boundary value problem for hyperbolic equations. By this reformulation of the boundary data, the engineering design quantities, throat Mach number and flow angle, were introduced as direct input quantities to the computer program. As a consequence of introducing the engineering parameters as input, the computer program will calculate the surface contours required to satisfy the specific throat conditions. Inviscid flow is assumed and the method used to calculate the inlet contour results in minimum distortion to the flow in the throat. This program was developed on an IBM 7094.

Beschreibung: This program represents a subsonic aerodynamic method for determining the mean camber surface of trimmed noncoplaner planforms with minimum vortex drag. With this program, multiple surfaces can be designed together to yield a trimmed configuration with minimum induced drag at some specified lift coefficient. The method uses a vortex-lattice and overcomes previous difficulties with chord loading specification. A Trefftz plane analysis is used to determine the optimum span loading for minimum drag. The program then solves for the mean camber surface of the wing associated with this loading. Pitching-moment or root-bending-moment constraints can be employed at the design lift coefficient. Sensitivity studies of vortex-lattice arrangements have been made with this program and comparisons with other theories show generally good agreement. The program is very versatile and has been applied to isolated wings, wing-canard configurations, a tandem wing, and a wing-winglet configuration. The design problem solved with this code is essentially an optimization one. A subsonic vortex-lattice is used to determine the span load distribution(s) on bent lifting line(s) in the Trefftz plane. A Lagrange multiplier technique determines the required loading which is used to calculate the mean camber slopes, which are then integrated to yield the local elevation surface. The problem of determining the necessary circulation matrix is simplified by having the chordwise shape of the bound circulation remain unchanged across each span, though the chordwise shape may vary from one planform to another. The circulation matrix is obtained by calculating the spanwise scaling of the chordwise shapes. A chordwise summation of the lift and pitching-moment is utilized in the Trefftz plane solution on the assumption that the trailing wake does not roll up and that the general configuration has specifiable chord loading shapes. VLMD is written in FORTRAN for IBM PC series and compatible computers running MS-DOS. This program requires 360K of RAM for execution. The Ryan McFarland FORTRAN compiler and PLINK86 are required to recompile the source code; however, a sample executable is provided on the diskette. The standard distribution medium for VLMD is a 5.25 inch 360K MS-DOS format diskette. VLMD was originally developed for use on CDC 6000 series computers in 1976. It was originally ported to the IBM PC in 1986, and, after minor modifications, the IBM PC port was released in 1993.

Beschreibung: This code was developed to aid design engineers in the selection and evaluation of aerodynamically efficient wing-canard and wing-horizontal-tail configurations that may employ simple hinged-flap systems. Rapid estimates of the longitudinal aerodynamic characteristics of conceptual airplane lifting surface arrangements are provided. The method is particularly well suited to configurations which, because of high speed flight requirements, must employ thin wings with highly swept leading edges. The code is applicable to wings with either sharp or rounded leading edges. The code provides theoretical pressure distributions over the wing, the canard or horizontal tail, and the deflected flap surfaces as well as estimates of the wing lift, drag, and pitching moments which account for attainable leading edge thrust and leading edge separation vortex forces. The wing planform information is specified by a series of leading edge and trailing edge breakpoints for a right hand wing panel. Up to 21 pairs of coordinates may be used to describe both the leading edge and the trailing edge. The code has been written to accommodate 2000 right hand panel elements, but can easily be modified to accommodate a larger or smaller number of elements depending on the capacity of the target computer platform. The code provides solutions for wing surfaces composed of all possible combinations of leading edge and trailing edge flap settings provided by the original deflection multipliers and by the flap deflection multipliers. Up to 25 pairs of leading edge and trailing edge flap deflection schedules may thus be treated simultaneously. The code also provides for an improved accounting of hinge-line singularities in determination of wing forces and moments. To determine lifting surface perturbation velocity distributions, the code provides for a maximum of 70 iterations. The program is constructed so that successive runs may be made with a given code entry. To make additional runs, it is necessary only to add an identification record and the namelist data that are to be changed from the previous run. This code was originally developed in 1989 in FORTRAN V on a CDC 6000 computer system, and was later ported to an MS-DOS environment. Both versions are available from COSMIC. There are only a few differences between the PC version (LAR-14458) and CDC version (LAR-14178) of AERO2S distributed by COSMIC. The CDC version has one main source code file while the PC version has two files which are easier to edit and compile on a PC. The PC version does not require a FORTRAN compiler which supports NAMELIST because a special INPUT subroutine has been added. The CDC version includes two MODIFY decks which can be used to improve the code and prevent the possibility of some infrequently occurring errors while PC-version users will have to make these code changes manually. The PC version includes an executable which was generated with the Ryan McFarland/FORTRAN compiler and requires 253K RAM and an 80x87 math co-processor. Using this executable, the sample case requires about four hours to execute on an 8MHz AT-class microcomputer with a co-processor. The source code conforms to the FORTRAN 77 standard except that it uses variables longer than six characters. With two minor modifications, the PC version should be portable to any computer with a FORTRAN compiler and sufficient memory. The CDC version of AERO2S is available in CDC NOS Internal format on a 9-track 1600 BPI magnetic tape. The PC version is available on a set of two 5.25 inch 360K MS-DOS format diskettes. IBM AT is a registered trademark of International Business Machines. MS-DOS is a registered trademark of Microsoft Corporation. CDC is a registered trademark of Control Data Corporation. NOS is a trademark of Control Data Corporation.

Beschreibung: This program provides a wing design algorithm based on modified linear theory which takes into account the effects of attainable leading-edge thrust. A primary objective of the WINGDES2 approach is the generation of a camber surface as mild as possible to produce drag levels comparable to those attainable with full theoretical leading-edge thrust. WINGDES2 provides both an analysis and a design capability and is applicable to both subsonic and supersonic flow. The optimization can be carried out for designated wing portions such as leading and trailing edge areas for the design of mission-adaptive surfaces, or for an entire planform such as a supersonic transport wing. This program replaces an earlier wing design code, LAR-13315, designated WINGDES. WINGDES2 incorporates modifications to improve numerical accuracy and provides additional capabilities. A means of accounting for the presence of interference pressure fields from airplane components other than the wing and a direct process for selection of flap surfaces to approach the performance levels of the optimized wing surfaces are included. An increased storage capacity allows better numerical representation of those configurations that have small chord leading-edge or trailing-edge design areas. WINGDES2 determines an optimum combination of a series of candidate surfaces rather than the more commonly used candidate loadings. The objective of the design is the recovery of unrealized theoretical leading-edge thrust of the input flat surface by shaping of the design surface to create a distributed thrust and thus minimize drag. The input consists of airfoil section thickness data, leading and trailing edge planform geometry, and operational parameters such as Mach number, Reynolds number, and design lift coefficient. Output includes optimized camber surface ordinates, pressure coefficient distributions, and theoretical aerodynamic characteristics. WINGDES2 is written in FORTRAN V for batch execution and has been implemented on a CDC CYBER computer operating under NOS 2.7.1 with a central memory requirement of approximately 344K (octal) of 60 bit words. This program was developed in 1984, and last updated in 1990. CDC and CYBER are trademarks of Control Data Corporation.

Beschreibung: Newell and Meng (1992) present maps of the occurrence probability of various classifications of particle precipitation as seen in the dayside topside ionosphere. It is argued that these are maps of the magnetospheric regions, a contention with which their critics disagree. The latter conclude that, because of convection, any one population of particles seen at low altitudes will have originated from a wide variety of locations, and particle characteristics cannot be mapped back to those in the magnetosphere without detailed knowledge of both the convection and magnetic field. Steplike boundaries between the regions will arise from nonsteady-state conditions and cannot be envisaged as steady-state magnetospheric boundaries between two plasma populations. In their reply Newell and Meng contend that convection does not move plasma from the LLBL into the cusp. Most of the LLBL plasma comes from the magnetosheath, so the direction of plasma transfer is in the other direction.

Beschreibung: Unsteady flow behavior and load characteristics of a 2D VR-7 airfoil with and without a leading-edge slat were studied in the water tunnel of the Aeroflightdynamics Directorate, NASA Ames Research Center. Both airfoils were oscillated sinusoidally between 5 and 25 deg at Re = 200,000 to obtain the unsteady lift, drag, and pitching moment data. A fluorescent dye was released from an orifice located at the leading edge of the airfoil for the purpose of visualizing the boundary layer and wake flow. The flowfield and load predictions of an incompressible Navier-Stokes code based on a velocity-vorticity formulation were compared with the test data. The test and predictions both confirm that the slatted VR-7 airfoil delays both static and dynamic stall as compared to the VR-7 airfoil alone.

Beschreibung: To monitor future global temperature trends, it would be extremely useful if parameters nonlinearly related to surface temperature could be found, thereby amplifying any warming signal that may exist. Evidence that global thunderstorm activity is nonlinearly related to diurnal, seasonal and interannual temperature variations is presented. Since global thunderstorm activity is also well correlated with the earth's ionospheric potential, it appears that variations of ionospheric potential, that can be measured at a single location, may be able to supply valuable information regarding global surface temperature fluctuations. The observations presented enable a prediction that a 1 percent increase in global surface temperatures may result in a 20 percent increase in ionospheric potential.

Beschreibung: The largest known effusive eruptions during the Cenozoic and Mesozoic Eras, the voluminous flood basalts, have long been suspected as being associated with major extinctions of biotic species. Despite the possible errors attached to the dates in both time series of events, the significance level of the suspected correlation is found here to be 1 percent to 4 percent. Statistically, extinctions lag eruptions by a mean time interval that is indistinguishable from zero, being much less than the average residual derived from the correlation analysis. Oceanic flood basalts, however, must have had a different biological impact, which is still uncertain owing to the small number of known examples and differing physical factors. Although not all continental flood basalts can have produced major extinction events, the noncorrelating eruptions may have led to smaller marine extinction events that terminated at least some of the less catastrophically ending geologic stages. Consequently, the 26 Myr quasi-periodicity seen in major marine extinctions may be only a sampling effect, rather than a manifestation of underlying periodicity.

Beschreibung: The monthly and daily samples of the Ap index for the interval from 1932 through 1982 were studied using the power spectrum technique. Results obtained for Bartel's period (about 27 days), the semiannual period, the dual-peak solar cycle distribution of geomagnetic storms, and certain other medium-scale periodicities are examined in detail. In addition, results on the cumulative occurrence number of storms per decade as a function of the Ap and Dst indices for the storm are presented.

Beschreibung: The circulation of polar wind ions from the high-latitude ionosphere to the plasma sheet is investigated. Single-particle trajectory simulations for the geomagnetic tail show, in addition to the breaking of adiabaticity, the possible creation of new high-altitude mirror points. This trajectory feature results from an intense parallel deceleration imparted by the magnetic field rotation during fast ExB transport. This centrifugal deceleration yields a critical parallel velocity which depends on the magnitude of the convection electric field and below which ions remain trapped inside the neutral sheet.

Beschreibung: The problem of the hypersonic double ellipse in rarefied flow is treated by a particle method using the collision model first described by McDonald (1988). In the approach used here, the computational overhead is reduced by using simple cubic cells. The problem of the definition of complex geometries is addressed by developing an algorithm to define the relation of a body surface to the network of cells.

Beschreibung: Program LAURA (Langley Aerothermodynamic Upwind Relaxation Algorithm) is an upwind-biased, point-implicit relaxation algorithm for obtaining the numerical solution to the governing equations for 3D viscous hypersonic flows in chemical and thermal nonequilibrium. The algorithm is derived using a finite-volume formulation in which the inviscid components of flux across cell walls are described with a modified Roe's averaging and with second-order corrections based on Yee's Symmetric Total Variation Diminishing scheme. The code has been applied to Problem 8.2 of this workshop for the case of thermochemical nonequilibrium flow through a nozzle. Chemical reaction rates are defined with the model of Park (1987). Thermal nonequilibrium is modeled using a two-temperature approximation in which the vibrational energies of all molecules are assumed to be in equilibrium at a single temperature which is generally different from the translational-rotational temperature. Two grids were used to define the flow for the original problem, with a stagnation temperature of 6500 K. A third case with a stagnation temperature of 10,000 K is also presented. The solution domain includes the converging nozzle, subsonic flow domain in which the gas is substantially in thermochemical equilibrium and the diverging nozzle, hypersonic flow domain in which the gas is substantially in thermochemical nonequilibrium.

Beschreibung: Solutions have been computed and results are presented for Problem 1, the case of Mach 9 transitional flow past a 7 deg half-angle cone at zero incidence. The solutions were computed using a code developed for the integration of the parabolized Navier-Stokes equations. The algorithm employed in the code is based on a Roe-type flux-difference-splitting scheme applied following a finite-volume approach. The basic algorithm has been modified to make it implicit and second-order accurate in the crossflow directions. Results are presented in terms of surface pressure and heat transfer as well as boundary layer profiles of pitot pressure, Mach number, and tangential velocity. The case was recalculated several times in an effort to determine sensitivities to such parameters as grid density, wall temperature, turbulence model parameters, as well as freestream expansion. Comparisons with the experimental data are presented and discussed.

Beschreibung: A development status evaluation is presented for the theoretical understanding and design conceptualization of boundary layer control (BLC) systems applicable to supersonic transports, such as the currently envisioned NASA High Speed Civil Transport. By reducing fuel burned, supersonic BLC techniques could expand ranges to Pacific-crossing scales, while lowering sonic boom effects and upper-atmosphere pollution and even reducing skin friction temperature. The critical consideration for supersonic BLC is the presence of wave effects.

Beschreibung: Attention is drawn to the influence of preexisting finite-amplitude instabilities on the growth of other disturbances; current design tools for LFC take no notice of this kind of interaction. When a rational accounting is accomplished for the evolution of incoming disturbances in finite-amplitude solutions of the equations of motion, future transition-prediction methods will need to take these wave interactions into account. Attention is given here to interactions in the presence of crossflow vortices and interactions involving Goertler vortices.

Beschreibung: A development history and a development-trends evaluation are presented for laminar flow controlled airfoil technologies and design concepts, including the search for 'natural' laminar flow and actively controlled flow via suction through small pores on the airfoil surface. While most NASA activities in this field have been concerned with subsonic aircraft, it has been projected that the control of boundary layer turbulence may be even more critical to the aerodynamic efficiency of supersonic aircraft. Developmental programs for these techniques have been conducted with several modified conventional aircraft.

Beschreibung: An account is given of the development history of natural laminar-flow (NLF) airfoil profiles under guidance of an experimentally well-verified theoretical method for the design of airfoils suited to virtually all subcritical applications. This method, the Eppler Airfoil Design and Analysis Program, contains a conformal-mapping method for airfoils having prescribed velocity-distribution characteristics, as well as a panel method for the analysis of potential flow about given airfoils and a boundary-layer method. Several of the NLF airfoils thus obtained are discussed.

Beschreibung: High resolution in situ measurements of reactive nitrogen (NO(y)) and O3 were made in the upper troposphere and lower stratosphere at a variety of latitudes and seasons. In the lower stratosphere, NO(y) and O3 are very highly positively correlated at all times and spatial scales sampled. The ratio NO(y)/NO3 is much less variable than either species measured separately. The ratio has a much weaker gradient with altitude than the mixing ratios of O3 or NO(y). The ratio is smaller and decreases more rapidly with altitude in the tropics than at midlatitudes. In the upper troposphere NO(y) and O3 are only weakly correlated. Their ratio in the tropical upper troposphere is about 0.005-0.025 and the ratio in the midlatitude upper troposphere is about 0.004-0.010. The NO(y) in the upper troposhere is probably partly due to lightning.

Beschreibung: The chemistry and dynamics of a convective system observed on February 2, 1987 in the EMEX and STEP campaigns are analyzed. Chemical and thermodynamic profiles in undisturbed air near the EMEX 9 system indicate that the troposphere was well mixed by previous convection. As a consequence there was little direct transport from the boundary layer to the upper troposphere and air transported upward was detrained throughout the middle and upper troposphere. There was minimal effect on O3 production. Other convective complexes located 800-900 km upstream produced greater perturbations on trace gas profile immediately below the tropopause. Ozone production was reduced by about 0.25 ppbv/d at these altitudes, representing a reduction in P(O3) of 15-20 percent over the column from 14.5 to 17 km. P(O3) from 12 to 17 km in a region distant from active convection and subject to lightning was 2-3 times higher than it would have been without lightning.

Beschreibung: Lateral variations of the temperature field in the lower mantle have been reconstructed using new results in mineral physics and seismic tomographic data. We show that, with the application of high-pressure experimental values of thermal expansivity and of sound velocities, the slow seismic anomalies in the lower mantle under the Pacific and Africa can be converted into realistic-looking plume structures with large dimensions of 0(1000 km). The outer fringes of the plumes have an excess temperature of around 400 K. In the core of the plumes are found tonguelike structures with extremely high thermal anomalies. These values can exceed 1200 K and are too high to be explained on the basis of thermal anomalies alone. We suggest that these major plumes in the deep mantle may be driven by both thermal and chemical buoyancies or that enhanced conductive heat-transfer may be important there.

Beschreibung: The largest known Cenozoic impact craters with the most accurately measured ages are found to correlate very closely with geologic stage boundaries. The level of confidence in this result is 98-99 percent even under the most pessimistic assumptions concerning dating errors. One or more large impacts may have led, in at least some cases, to the extinctions and first appearances of biotic species that mark many of the geologic stage boundaries.

Beschreibung: Ozone profiles obtained with the Differential Absorption Lidar (DIAL) system at the National Institute for Environmental Studies (NIES) (Tsukuba, Japan) were compared with data provided by the satellite sensor SAGE II. The SAGE II data were selected based on criteria of spatial and temporal differences between the DIAL and the SAGE II measurements: five degrees in latitude and 15 degrees in longitude, within a latitudinal band from 31 deg to 41 deg N, and within one, three and five days after or before the DIAL measurements. Results show very good agreement for the individual and the zonal-mean profiles. The average mean difference between the DIAL and the SAGE II measurements over the altitudes 15-50 km was about 10 percent.

Beschreibung: The BATSE on-board burst system has been triggered by over 400 terrestrial electron precipitation events. These are the single largest cause of false triggers. Bremsstrahlung is generated as the precipitating electrons interact in the Earth's atmosphere, or in the spacecraft, and this radiation is detected by the Large Area Detectors, often triggering the instrument into burst mode. Several examples of such events are presented here, and the different classes of events are described. A possible correlation of events to strong magnetospheric activity is presented, and the association of a sub-set of events to a powerful VLF transmitter on the western coast of Australia is described.

Beschreibung: The Tsyganenko models contain 'modules' for two external current systems, the tail and ring currents; the former is the more detailed. No specific modules, however, cover the magnetopause and the Birkeland current system. The effects of these currents are represented by an all-purpose 'polynomial', and the resulting formulas involve about 30 parameters that specify the model. The dependence on magnetospheric indices is also simplified.

Beschreibung: A quantitative model is developed of the magnetic field produced by the electric current system of region 2 Birkeland currents, closed via the partial ring current. The distribution of j-perpendicular is computed from a given axially asymmetric spatial distribution of hot isotropic magnetospheric plasma over an infinitely thin L shell in an axisymmetric purely dipolar geomagnetic field, while the field-aligned current density is found from the continuity of the net electric current. The magnetic field distribution is derived by a Biot-Savart integral over the electric current system. An assumed cosine dependence of the plasma pressure on local time makes it possible to reduce the problem of analytical representation of the B field to two dimensions. The obtained numerical fits for the partial ring current/region 2 Birkeland current magnetic field are relatively simple, continuous, and valid throughout the whole extraterrestrial space from ionospheric heights up to tens of earth radii.

Beschreibung: The effects of multiple phase transitions on mantle convection are investigated by numerical simulations that are based on three-dimensional models. These simulations show that cold sheets of mantle material collide at junctions, merge, and form a strong downflow that is stopped temporarily by the transition zone. The accumulated cold material gives rise to a strong gravitational instability that causes the cold mass to sink rapidly into the lower mantle. This process promotes a massive exchange between the lower and upper mantles and triggers a global instability in the adjacent plume system. This mechanism may be cyclic in nature and may be linked to the generation of superplumes.

Beschreibung: Airloads measured on a two-bladed helicopter rotor in flight, from the Tip Aerodynamic and Acoustic Test, are compared with calculations from a comprehensive helicopter analysis (CAMRAD/JA), and the pressures compared with calculations from a full-potential rotor code (FPR). The flight test results cover an advance ratio range from 0.19 to 0.38. The lowest speed case is characterized by the presence of significant blade-vortex interactions. Good correlation of peak-to-peak vortex-induced loads and the corresponding pressures is obtained. The results of the correlation for this two-bladed rotor are substantially similar to the results for three- and four-bladed rotors, concerning the tip vortex core size for best correlation, calculation of the peak-to-peak loads on the retreating side, and calculation of vortex-induced loads on inboard radial stations.

Beschreibung: A new CFD potential code, FPX (eXtended Full-Potential), has been developed for application to both helicopters and tilt-rotors. The code solves the unsteady, three-dimensional full potential equation and is an extension of the rotor code, FPR. Both entropy and viscosity corrections are included to enhance the physical modeling capabilities. A number of efficiency related modifications have yielded a factor of two speed-up in the code. An axial flow capability has been added to treat tilt-rotor in forward flight (cruise mode). In order to employ streamwise periodicity and accurately solve for the propagation of acoustic signals in the tip region, an H-H topology has been added to the basic O-H grid system. Computations are performed for the XV-15 Standard and ATB blades at high-speed conditions. Comparisons are made for the blade aerodynamics and the induced fuselage cabin pressure for a range of Mach numbers. Grid generation, wake treatment, and far-field wall treatment are identified as problem areas with recommendations for future research.

Beschreibung: A preliminary test/theory correlation evaluation is conducted for wake measurement test results obtained by LDV for a B360 helicopter rotor, at conditions critical to the understanding of wake-rollup and blade-vortex interaction phenomena. The LDV data were complemented by acoustic, blade pressure, rotor performance, and blade/control load measurements.

Beschreibung: Analytic expressions have been derived for estimating the nonablating laminar and turbulent boundary-layer convective heating rates on inclined flat surfaces for the Martian atmosphere in thermochemical equilibrium. The equations are valid in the speed and altitude regime where aerobraking would occur at Mars. Comparisons with limited experimental measurements and calculations for CO2 (the Martian atmosphere is 95.6 percent CO2) yielded reasonably good agreement, especially for the ratios of heating rates in CO2 to those in air at the same conditions. In the aerobraking speed regime, the laminar flat surface boundary layer heating rates are 15-25 percent greater at Mars than in air. The differences between the turbulent heating rates are even more pronounced. The turbulent heating rates can be over 50 percent greater at Mars than in air at the same flight conditions.

Beschreibung: A new method has been developed to provide a direct test of the error calibrations of gravity models based on actual satellite observations. The basic approach projects the error estimates of the gravity model parameters onto satellite observations, and the results of these projections are then compared with data residual computed from the orbital fits. To allow specific testing of the gravity error calibrations, subset solutions are computed based on the data set and data weighting of the gravity model. The approach is demonstrated using GEM-T3 to show that the gravity error estimates are well calibrated and that reliable predictions of orbit accuracies can be achieved for independent orbits.

Beschreibung: This code was developed to aid design engineers in the selection and evaluation of aerodynamically efficient wing-canard and wing-horizontal-tail configurations that may employ simple hinged-flap systems. Rapid estimates of the longitudinal aerodynamic characteristics of conceptual airplane lifting surface arrangements are provided. The method is particularly well suited to configurations which, because of high speed flight requirements, must employ thin wings with highly swept leading edges. The code is applicable to wings with either sharp or rounded leading edges. The code provides theoretical pressure distributions over the wing, the canard or horizontal tail, and the deflected flap surfaces as well as estimates of the wing lift, drag, and pitching moments which account for attainable leading edge thrust and leading edge separation vortex forces. The wing planform information is specified by a series of leading edge and trailing edge breakpoints for a right hand wing panel. Up to 21 pairs of coordinates may be used to describe both the leading edge and the trailing edge. The code has been written to accommodate 2000 right hand panel elements, but can easily be modified to accommodate a larger or smaller number of elements depending on the capacity of the target computer platform. The code provides solutions for wing surfaces composed of all possible combinations of leading edge and trailing edge flap settings provided by the original deflection multipliers and by the flap deflection multipliers. Up to 25 pairs of leading edge and trailing edge flap deflection schedules may thus be treated simultaneously. The code also provides for an improved accounting of hinge-line singularities in determination of wing forces and moments. To determine lifting surface perturbation velocity distributions, the code provides for a maximum of 70 iterations. The program is constructed so that successive runs may be made with a given code entry. To make additional runs, it is necessary only to add an identification record and the namelist data that are to be changed from the previous run. This code was originally developed in 1989 in FORTRAN V on a CDC 6000 computer system, and was later ported to an MS-DOS environment. Both versions are available from COSMIC. There are only a few differences between the PC version (LAR-14458) and CDC version (LAR-14178) of AERO2S distributed by COSMIC. The CDC version has one main source code file while the PC version has two files which are easier to edit and compile on a PC. The PC version does not require a FORTRAN compiler which supports NAMELIST because a special INPUT subroutine has been added. The CDC version includes two MODIFY decks which can be used to improve the code and prevent the possibility of some infrequently occurring errors while PC-version users will have to make these code changes manually. The PC version includes an executable which was generated with the Ryan McFarland/FORTRAN compiler and requires 253K RAM and an 80x87 math co-processor. Using this executable, the sample case requires about four hours to execute on an 8MHz AT-class microcomputer with a co-processor. The source code conforms to the FORTRAN 77 standard except that it uses variables longer than six characters. With two minor modifications, the PC version should be portable to any computer with a FORTRAN compiler and sufficient memory. The CDC version of AERO2S is available in CDC NOS Internal format on a 9-track 1600 BPI magnetic tape. The PC version is available on a set of two 5.25 inch 360K MS-DOS format diskettes. IBM AT is a registered trademark of International Business Machines. MS-DOS is a registered trademark of Microsoft Corporation. CDC is a registered trademark of Control Data Corporation. NOS is a trademark of Control Data Corporation.

Beschreibung: Since its early beginnings, NASA has been actively involved in the design and testing of airfoil sections for a wide variety of applications. Recently a set of programs has been developed to smooth and scale arbitrary airfoil coordinates. The smoothing program, AFSMO, utilizes both least-squares polynomial and least-squares cubic-spline techniques to iteratively smooth the second derivatives of the y-axis airfoil coordinates with respect to a transformed x-axis system which unwraps the airfoil and stretches the nose and trailing-edge regions. The corresponding smooth airfoil coordinates are then determined by solving a tridiagonal matrix of simultaneous cubic-spline equations relating the y-axis coordinates and their corresponding second derivatives. The camber and thickness distribution of the smooth airfoil are also computed. The scaling program, AFSCL, may then be used to scale the thickness distribution generated by the smoothing program to a specified maximum thickness. Once the thickness distribution has been scaled, it is combined with the camber distribution to obtain the final scaled airfoil contour. The airfoil smoothing and scaling programs are written in FORTRAN IV for batch execution and have been implemented on a CDC CYBER 170 series computer with a central memory requirement of approximately 70K (octal) of 60 bit words. Both programs generate plotted output via CALCOMP type plotting calls. These programs were developed in 1983.

Beschreibung: The Supersonic Wing Nonlinear Aerodynamics computer program, LTSTAR, was developed to provide for the estimation of the nonlinear aerodynamic characteristics of a wing at supersonic speeds. This corrected linearized-theory method accounts for nonlinearities in the variation of basic pressure loadings with local surface slopes, predicts the degree of attainment of theoretical leading-edge thrust forces, and provides an estimate of detached leading-edge vortex loadings that result when the theoretical thrust forces are not fully realized. Comparisons of LTSTAR computations with experimental results show significant improvements in detailed wing pressure distributions, particularly for large angles of attack and for regions of the wing where the flow is highly three-dimensional. The program provides generally improved predictions of the wing overall force and moment coefficients. LTSTAR could be useful in design studies aimed at aerodynamic performance optimization and for providing more realistic trade-off information for selection of wing planform geometry and airfoil section parameters. Input to the LTSTAR program includes wing planform data, freestream conditions, wing camber, wing thickness, scaling options, and output options. Output includes pressure coefficients along each chord, section normal and axial force coefficients, and the spanwise distribution of section force coefficients. With the chordwise distributions and section coefficients at each angle of attack, three sets of polars are output. The first set is for linearized theory with and without full leading-edge thrust, the second set includes nonlinear corrections, and the third includes estimates of attainable leading-edge thrust and vortex increments along with the nonlinear corrections. The LTSTAR program is written in FORTRAN IV for batch execution and has been implemented on a CDC 6000 series computer with a central memory requirement of approximately 150K (octal) of 60 bit words. The LTSTAR program was developed in 1980.

Beschreibung: The nozzle afterbody is one of the main drag-producing components of an aircraft propulsion system. Thus, considerable effort has been devoted to developing techniques for predicting the afterbody flow field and drag. The RAXJET computer program was developed to predict the transonic, axisymmetric flow over nozzle afterbodies with supersonic jet exhausts and includes the effects of boundary-layer displacement, separation, jet entrainment, and inviscid jet plume blockage. RAXJET iteratively combines the South-Jameson relaxation procedure, the Reshotko-Tucker boundary-layer solution, the Presz separation model, the Dash-Pergament mixing model, and the Dash-Thorpe inviscid plume model into a single, comprehensive model. The approach taken in the RAXJET program requires considerably less computational time than the Navier-Stokes solutions and generally yields results of comparable accuracy. In RAXJET, the viscous-inviscid interaction model is constructed by dividing the afterbody flow field into six separate computational regions: (1) The inviscid external flow solution is based on the relaxation procedure of South and Jameson for solving the exact nonlinear potential flow equation in nonconservative form. (2) The flow field in the inviscid jet exhaust is solved by explicit spatial marching of the conservative finite-difference form of the inviscid flow equations for a uniform composition gas mixture. (3) The properties in the attached boundary-layer region are solved by a modified version of the Reshotko-Tucker integral method for turbulent flows. (4) The analysis of the separated flow region consists of predicting the separation location and calculating the discriminating streamline shape. (5) The jet wake region is determined by either a simple extrapolation model or by an integral method that accounts for entrainment effects. (6) The displacement-thickness distribution arising from entrainment into the jet mixing layer is calculated by the overlaid mixing model. The inviscid external flow solution and inviscid jet exhaust solution provide the necessary flow conditions to calculate the flow in the viscous regions. The viscous and inviscid flow fields are iteratively solved until a final solution is obtained. Input to the RAXJET program consists of body geometry data, free-stream conditions, main logic control parameters, and condition and control parameters for each of the six computational flow regions. Output from RAXJET includes detailed flow results and aerodynamic coefficients. The RAXJET program is written in FORTRAN IV for batch execution and has been implemented on a CDC CYBER 170 series computer with a central memory requirement of approximately 60K(octal) of 60 bit words. The RAXJET program was developed in 1982.

Beschreibung: The computer program SALLY was developed to compute the incompressible linear stability characteristics and integrate the amplification rates of boundary layer disturbances on swept and tapered wings. For some wing designs, boundary layer disturbance can significantly alter the wing performance characteristics. This is particularly true for swept and tapered laminar flow control wings which incorporate suction to prevent boundary layer separation. SALLY should prove to be a useful tool in the analysis of these wing performance characteristics. The first step in calculating the disturbance amplification rates is to numerically solve the compressible laminar boundary-layer equation with suction for the swept and tapered wing. A two-point finite-difference method is used to solve the governing continuity, momentum, and energy equations. A similarity transformation is used to remove the wall normal velocity as a boundary condition and place it into the governing equations as a parameter. Thus the awkward nonlinear boundary condition is avoided. The resulting compressible boundary layer data is used by SALLY to compute the incompressible linear stability characteristics. The local disturbance growth is obtained from temporal stability theory and converted into a local growth rate for integration. The direction of the local group velocity is taken as the direction of integration. The amplification rate, or logarithmic disturbance amplitude ratio, is obtained by integration of the local disturbance growth over distance. The amplification rate serves as a measure of the growth of linear disturbances within the boundary layer and can serve as a guide in transition prediction. This program is written in FORTRAN IV and ASSEMBLER for batch execution and has been implemented on a CDC CYBER 70 series computer with a central memory requirement of approximately 67K (octal) of 60 bit words. SALLY was developed in 1979.

Beschreibung: An incompressible three-dimensional laminar boundary-layer flow over a swept wing is used as a model to study both the wall-curvature and streamline-curvature effects on the stationary crossflow instability. The basic state is obtained by solving the full Navier-Stokes (N-S) equations numerically. The linear disturbance equations are cast on a fixed, body-intrinsic, curvilinear coordinate system. Those nonparallel terms which contribute mainly to the streamline-curvature effect are retained in the formulation of the disturbance equations and approximated by their local finite difference values. The resulting eigenvalue problem is solved by a Chebyshev collocation method. The present results indicate that the convex wall curvature has a stabilizing effect, whereas the streamline curvature has a destabilizing effect. A validation of these effects with an N-S solution for the linear disturbance flow is provided.

Beschreibung: The paper presents a general approach to constructing mean velocity profiles for compressible turbulent boundary layers with isothermal or adiabatic walls. The theory is based on a density-weighted transformation that allows the extension of the incompressible similarity laws of the wall to the compressible regions. The velocity profile family is compared to a range of experimental data, and excellent agreement is obtained. A self-consistent skin friction law, which satisfies the proposed velocity profile family, is derived and compared with the well-known Van Driest II theory for boundary layers in zero pressure gradient. The results are found to be at least as good as those obtained by using the Van Driest II transformation.

Beschreibung: The stability of TIROS-N Microwave Sounding Unit (MSU) channel 4, which monitors the deep-layer averaged temperature of the lower stratosphere, is tested by intercalibrating MSU channel 4 data from the NIROS-N series of NOAA satellites during 1979-1991. The monthly gridpoint anomalies are validated with 10 years of radiosonde data during 1979-1988. The results demonstrated that the satellite sensors are very stable in their calibration and that the previously reported uncertainties in the stratospheric temperature information produced by NOAA are not the result of calibration changes in the MSUs. It was found that the largest globally averaged temperature variations during 1979-1991 occur after the El Chichon (1982) and Pinatubo (1991) volcanic eruptions. These warm events are superimposed upon a net downward trend in temperatures during the period. The cooling trend is strongest in polar regions and the Northern Hemisphere middle latitudes.

Beschreibung: Remote sensing of aerosol optical thickness from space is difficult over continental surfaces. There are two retrieval algorithms, one based on the use of dark targets and a second based on contrast reduction between selected pixels. Improvements in the contrast reduction method are reported. A procedure is developed for using the satellite image to evaluate whether conditions for applying the structure method are met. The theoretical background is discussed, and the usefulness of the structure functions is demonstrated. The method is applied to NOAA Advanced Very High Resolution Radiometer (AVHRR) imagery where simultaneous ground measurements are used for validation.

Beschreibung: An MHD simulation was performed to obtain a self-consistent model of magnetic field and plasma density near the X point reconnection region. The MHD model was used to perform extensive ray tracing calculations in order to clarify the propagation characteristics of the plasma waves near the X point reconnection region. The dynamic wave spectra possibly observed by the Geotail spacecraft during a typical cross-tail trajectory are reconstructed. By comparing the extensive ray tracing calculations with the plasma wave data from Geotail, it is possible to perform a kind of 'remote sensing' to identify the location and structure of potential X point reconnection regions.

Beschreibung: A number of chemical-kinetic problems related to phenomena occurring behind a shock wave surrounding an object flying in the earth atmosphere are discussed, including the nonequilibrium thermochemical relaxation phenomena occurring behind a shock wave surrounding the flying object, problems related to aerobraking maneuver, the radiation phenomena for shock velocities of up to 12 km/sec, and the determination of rate coefficients for ionization reactions and associated electron-impact ionization reactions. Results of experiments are presented in form of graphs and tables, giving data on the reaction rate coefficients for air, the ionization distances, thermodynamic properties behind a shock wave, radiative heat flux calculations, Damkoehler numbers for the ablation-product layer, together with conclusions.

Beschreibung: Computations were made for those test cases of Problem 3 which were designated as laminar flows, viz., test cases 3.1, 3.2, 3.4, and 3.5. These test cases corresponded to flows over a flat plate and a compression ramp at high Mach number and at high Reynolds number. The computations over the compression ramps indicate a substantial streamwise extent of separation. Based on previous experience with separated laminar flows at high Mach numbers which indicated a substantial effect with spatial grid refinement, a series of computations with different grid sizes were performed. Also, for the flat plate, comparisons of the results for two different algorithms were made.

Beschreibung: A development status evaluation is presented for CFD methods applicable to fuselage-integrated scramjet powerplant incorporating hypersonic vehicles; these methods are critically important due to the unavailability of experimental facilities for such elevated Mach number/high-enthalphy conditions. Advancements are required in algorithm robustness and speed, geometric flexibility, and the inclusion of more complete flow physics. The most serious deficiencies lie in turbulence modeling, the lack of complete transition-prediction methods, and combustion modeling.

Beschreibung: An overview is given of research activity on the application of computational fluid dynamics (CDF) for hypersonic propulsion systems. After the initial consideration of the highly integrated nature of air-breathing hypersonic engines and airframe, attention is directed toward computations carried out for the components of the engine. A generic inlet configuration is considered in order to demonstrate the highly three dimensional viscous flow behavior occurring within rectangular inlets. Reacting flow computations for simple jet injection as well as for more complex combustion chambers are then discussed in order to show the capability of viscous finite rate chemical reaction computer simulations. Finally, the nozzle flow fields are demonstrated, showing the existence of complex shear layers and shock structure in the exhaust plume. The general issues associated with code validation as well as the specific issue associated with the use of CFD for design are discussed. A prognosis for the success of CFD in the design of future propulsion systems is offered.