ALBERT

Description: The 3.0-micrometers water of hydration absorption feature observed in the IR photometry of many low-albedo and some medium-albedo asteroids strongly correlates with the 0.7-micrometers Fe(+2) to Fe(+3) oxidized iron absorption feature observed in narrowband spectrophotometry of these asteroids. Using this relationship, an empirical algorithm for predicting the presence of water of hydration in the surface material of a Solar System body using photometry obtained through the Eight-Color Asteroid Survey nu (0.550 micrometers), w (0.701 micrometers), and x (0.853 micrometers) filters was developed and applied to the ECAS photometry of asteroids and outer planet satellites. The percentage of objects in low-albedo, outer main-belt asteroid classes that test positively for water of hydration increases from P to B to C to G class and correlates linearly with the increasing mean albedos of those objects testing positively. The medium-albedo M-class asteroids do not test positively in large number using this algorithm. Aqueously altered asteroids dominate the Solar System population between heliocentric distances of 2.6 to 3.5 AU, bracketing the Solar System region where the aqueous alteration mechanism operated most strongly. One jovian satellite, J VI Himalia, and one saturnian satellite. Phoebe, tested positively for water of hydration, supporting the hypothesis that these may be captured C-class asteroids from a postaccretional dispersion. The proposed testing technique could be applied to an Earth-based survey of asteroids or a space-probe study of an asteroid's surface characteristic in order to identify a potential water source.

Description: The infrared transmission spectra and photochemical behavior of various organic compounds isolated in solid N2 ices, appropriate for applications to Triton ad Pluto, are presented. It is shown that excess absorption in the surface spectra of Triton and Pluto, i.e., absorption not explained by present models incorporating molecules already identified on these bodies (N2, CH4, CO, and CO2), that starts near 4450/cm (2.25 microns) and extends to lower frequencies, may be due to alkanes (C(n)H(2n+2)) and related molecules frozen in the nitrogen. Branched and linear alkanes may be responsible. Experiments in which the photochemstry of N2: CH4 and N2: CH4: CO ices was explored demonsrtrate that the surface ices of Triton and Pluto may contain a wide variety of additional species containing H, C, O, and N. Of these, the reactive molecule diazomethane, CH2N2, is particularly important since it may be largely responsible for the synthesis of larger alkanes from CH4 and other small alkanes. Diazomethane would also be expected to drive chemical reactions involving organics in the surface ices of Triton and Pluto toward saturation, i.e., to reduce multiple CC bonds. The positions and intrinsic strengths (A values) of many of the infrared absorption bands of N2 matrix-isolated molecules of relevance to Triton and Pluto have also been determined. These can be used to aid in their search and to place constraints on their abundances.

Description: The compressible dynamic stall flowfield over a NACA 0012 airfoil transiently pitching from 0 to 60 deg at a constant rate under compressible flow conditions has been studied using real-time interferometry. A quantitative description of the overall flowfield, including the finer details of dynamic stall vortex formation, growth, and the concomitant changes in the airfoil pressure distribution, has been provided by analyzing the interferograms. For Mach numbers above 0.4, small multiple shocks appear near the leading edge and are present through the initial stages of dynamic stall. Dynamic stall was found to occur coincidentally with the bursting of the separation bubble over the airfoil. Compressibility was found to confine the dynamic stall vortical structure closer to the airfoil surface. The measurements show that the peak suction pressure coefficient drops with increasing freestream Mach number, and also it lags the steady flow values at any given angle of attack. As the dynamic stall vortex is shed, an anti-clockwise vortex is induced near the trailing edge, which actively interacts with the post-stall flow.

Description: The effect of the porous leading edge of an airfoil on the blade-vortex interaction noise, which dominates the far-field acoustic spectrum of the helicopter, is investigated. The thin-layer Navier-Stokes equations are solved with a high-order upwind-biased scheme and a multizonal grid system. The Baldwin-Lomax turbulence model is modified for considering transpiration on the surface. The amplitudes of the propagating acoustic wave in the near field are calculated directly from the computation. The porosity effect on the surface is modeled in two ways: (1) imposition of prescribed transpiration velocity distribution and (2) calculation of transpiration velocity distribution by Darcy's law. Results show leading-edge transpiration can suppress pressure fluctuations at the leading edge during blade-vortex interaction and consequently reduce the amplitude of propagating noise by 30% at a maximum in the near field.

Description: A method has been developed for calculating the viscous flow about airfoils with and without deflected flaps at -90 deg incidence. This method provides for the solution of the unsteady incompressible Navier-Stokes equations by means of an implicit technique. The solution is calculated on a body-fitted computational mesh using a staggered-grid method. The vorticity is defined at the node points, and the velocity components are defined at the mesh-cell sides. The staggered-grid orientation provides for accurate representation of vorticity at the node points and the continuity equation at the mesh-cell centers. The method provides for the noniterative solution of the flowfield and satisfies the continuity equation to machine zero at each time step. The method is evaluated in terms of its stability to predict two-dimensional flow about an airfoil at -90-deg incidence for varying Reynolds number and laminar/turbulent models. The variations of the average loading and surface pressure distribution due to flap deflection, Reynolds number, and laminar or turbulent flow are presented and compared with experimental results. The comparisom indicate that the calculated drag and drag reduction caused by flap deflection and the calculated average surface pressure are in excellent agreement with the measured results at a similar Reynolds number.

Description: This report presents the most recent spherical harmonic topography model of Venus developed at Jet Propulsion Laboratory. It was produced by a spherical harmonic analysis of the most complete set of Magellan altimetry data, augmented by Pioneer Venus and Venera data. The harmonic coefficients of the topography were computed to degree and order 360. Compared to previous topography models, this one has the highest correlation with the gravity field of Venus.

Description: The 500-Myr average crater retention age for Venus has raised questions about the present-day level of tectonic activity. In this study we examine the relationship between the gravity and topography of four large volcanic swells, Beta, Atla, Bell, and Western Eistla Regiones, for clues about their stage evolution. The Magellan line-of-sight gravity data are inverted using a point mass model of the anomalous mass to solve for the local vertical gravity field. Spectral admittance calculated from both the local gravity inversions and a spherical harmonic model is compared to three models of compensation: local compensation, a 'flexural' model with local and regional compensation of surface and subsurface loads, and a 'hotspot' model of compensation that includes top loading by volcanoes and subsurface loading due to a deep, low density mass anomaly. The coherence is also calculated in each region, but yields an elastic thickness estimate only at Bell Regio. In all models, the long wavelengths are compensated locally. Our results may indicate a relatively old, possibly inactive plume.

Description: Rotor noise prediction codes predict the thickness and loading noise produced by a helicopter rotor, given the blade motion, rotor operating conditions, and fluctuating force distribution over the blade surface. However, the criticality of these various inputs, and their respective effects on the predicted acoustic field, have never been fully addressed. This paper examines the importance of these inputs, and the sensitivity of the acoustic predicitions to a variation of each parameter. The effects of collective and cyclic pitch, as well as coning and cyclic flapping, are presented. Blade loading inputs are examined to determine the necessary spatial and temporal resolution, as well as the importance of the chordwise distribution. The acoustic predictions show regions in the acoustic field where significant errors occur when simplified blade motions or blade loadings are used. An assessment of the variation in the predicted acoustic field is balanced by a consideration of Central Processing Unit (CPU) time necessary for the various approximations.

Description: Recent mapping studies west of Elysium Mons, Mars, have pinpointed subice features that suggest the existence of a frozen paleolake in Utopia Planitia as recently as 1.8 billion years ago. The subice features are interpreted to be hyaloclastic ridges and hills, table moutains, associated joekulhalaup deposits, and fluvial channels. Photoclinometric studies of these features and of a basal scarp around the northwest flank of Elysium Mons interpreted to have been an ice-sheet boundary indicate that the maximum thickness of ice within the basin may have been about 180 m. This thickness of ice during a relatively late stage of Martian geologic history would have important implications concerning the atmospheric, the climatic, and possibly the exobiologic history of the planet.

Description: The ion and electron momentum equations, along with Ampere's law, are solved for the ion and electron drift velocities and the electric field in the subsolar Venus ionosphere, assuming a partially ionized gas and a single ion species having the ion mean mass. All collision terms among the ions, electrons and neutral particles are retained in the equations. A general expression for the evolution of the magnetic field is derived and compared with earlier expressions. Subsolar region data in the altitude range 150-300 km from the Pioneer Venus Orbiter are used to calculate altitude profiles of the components of the current due to the electric field, gradients of pressure, and gravity. Altitude profiles of the ion and electron velocities as well as the electric field, electrodynamic heating, and the energy density are determined. Only orbits having a complete set of measured plasma temperatures and densities, neutral densities, and magnetic field were considered for analysis; the results are shown only for orbit 202. The vertical velocity at altitudes above 220 km is upgoing for orbit 202. This result is consistent with observations of molecular ions at high altitudes and of plasma flow to the nightside, both of which require upward velocity of ions from the dayside ionosphere. Above about 230 km the momentum equations are extremely sensitive to the altitude profiles of density, temperature, and magnetic field.

Description: We report here analyses of olivines and pyroxenes, and petrofabrics of 27 chondritic interplanetary dust particles (IDPs), comparing those from anhydrous and hydrous types. Approximately 40% of the hydrous particles contain diopside, a probable indicator of parent body thermal metamorphism, while this mineral is rarely present in the anhydrous particles. Based on this evidence, we find that hydrous and anhydrous IDPs are, in general, not directly related, and we conclude that olivine and pyroxene major-element compositions can be used to help discriminate between IDPs that are (1) predominantly nebular condensates, and lately resided in anhydrous or icy (no liquids) primitive parent bodies, and (2) those originating from more geochemically active parent bodies (probably hydrous and anhydrous asteroids).

Description: This paper reports on a magnetic field phenomenon, hereafter referred to as null fields, which were discovered during the inbound pass of the recent flyby of Jupiter by the Ulysses spacecraft. These null fields which were observed in the outer dayside magnetosphere are characterised by brief but sharp decreases of the field magnitude to values less than 1 nT. The nulls are distinguished from the current sheet signatures characteristic of the middle magnetosphere by the fact that the field does not reverse across the event. A field configuration is suggested that accounts for the observed features of the events.

Description: The U.S. National Aeronautics and Space Administration (NASA) Balloon Program has been highly successful since recovering from the catastrophic balloon failure problems of the early to mid 1980s. Balloons have continued to perform at unprecedented success rates. The comprehensive research and development (R&D) effort has continued with advances being made across the spectrum of balloon related disciplines. The long duration balloon project will be transitioning from a development effort to an operational capability this year. Recently, emphasis has been placed on the development and implementation of new support systems and facilities. A new permanent launch facility at Fort Sumner, New Mexico has been established. New ground station support equipment is being implemented, and a new heavy load launch vehicle is scheduled to be implemented in 1992. The progress, status and future plans for these and other aspects of the NASA program will be presented.

Description: The catastrophic balloon failure during the first half of the 1980's identified the need for a comprehensive and continuing balloon research and development (R&D) commitment by NASA. Technical understanding was lacking in many of the disciplines and processes associated with scientific ballooning. A comprehensive balloon R&D plan was developed in 1986 and implemented in 1987. The objectives were to develop the understanding of balloon system performance, limitations, and failure mechanisms. The program consisted of five major technical areas: structures, performance and analysis, materials, chemistry and processing, and quality control. Research activitites have been conducted at NASA/Goddard Space Flight Center (GSFC)-Wallops Flight Facility (WFF), other NASA centers and government facilities, universities, and the balloon manufacturers. Several new and increased capabilities and resources have resulted from this activity. The findings, capabilities, and plan of the balloon R&D program are presented.

Description: Caps have been used to structurally reinforce scientific research balloons since the late 1950's. The scientific research balloons used by the National Aeronautics and Space Administration (NASA) use internal caps. A NASA cap placement specification does not exist since no empirical information exisits concerning cap placement. To develop a cap placement specification, NASA has completed two in-hangar inflation tests comparing the structural contributions of internal caps and external caps. The tests used small scale test balloons designed to develop the highest possible stresses within the constraints of the hangar and balloon materials. An externally capped test balloon and an internally capped test balloon were designed, built, inflated and simulated to determine the structural contributions and benefits of each. The results of the tests and simulations are presented.

Description: The analysis of balloon envelopes by the finite element (FE) method is plagued by convergence problems. A pratical FE analysis approach is based on the fact that in thin shells with non-zero Gaussian curvature the membrane solution component is essentially decoupled from the bending solution component. A proxy-problem is solved by using a small artificial bending stiffness that assures convergence without significantly affecting the membrane solution component. This approach has been previously validated on slightly overpressurized balloon envelopes. Extensions of this approach to more difficult problems in the structural analysis of balloon envelopes are presented. The convergence forcing modelling measures are discussed. Implications of the findings of the analysis results to future balloon designs are also discussed.

Description: Geochemical profiles of surface units, impact, and volcanic features are studied in detail to determine the underlying structure in an area of extensive mare/highland interface, Sinus Amoris. This study region includes and surrounds the northeastern embayment of Mare Tranquillitatis. The concentrations of two major rock-forming elements (Mg and Al), which were derived from the Apollo 15 orbital geochemical measurements, were used in this study. Mapped units and deposits associated with craters in the northwestern part of the region tend to have correlated low Mg and Al concentrations, indicating the presence of Potassium (K)-Rare Earth Elements (REE)-Phosphorus (P) (KREEP)-enriched basalt. Found along the northeastern rim of Tranquillitatis were areas with correlated high Mg and Al concentration, indicating the presence of troctolite. Distinctive west/east and north/south trends were observed in the concentrations of Mg and Al, and, by implication, in the distribution of major rock components on the surface. Evidence for a systematic geochemical transition in highland or basin-forming units may be observed here in the form of distinctive differences in chemistry in otherwise similar units in the western and eastern portions of the study region.

Description: The purpose of this Note is to present results from an analytic/experimental study that investigated the potential for passively changing blade twist through the use of extension-twist coupling. A set of composite model rotor blades was manufactured from existing blade molds for a low-twist metal helicopter rotor blade, with a view toward establishing a preliminary proof concept for extension-twist-coupled rotor blades. Data were obtained in hover for both a ballasted and unballasted blade configuration in sea-level atmospheric conditions. Test data were compared with results obtained from a geometrically nonlinear analysis of a detailed finite element model of the rotor blade developed in MSC/NASTRAN.

Description: The paper considers the compressible Rayleigh equation as a model for the Mach wave emission mechanism associated with high-temperature supersonic jets. Solutions to the compressible Rayleigh equation reveal the existence of several families of supersonically convecting instability waves. These waves directly radiate noise to the jet far field. The predicted noise characteristics are compared to previously acquired experimental data for an axisymmetric Mach 2 fully pressure balanced jet operating over a range of jet total temperatures from ambient to 1370 K. The results of this comparison show that the first-order supersonic instability wave and the Kelvin-Hemlhlotz first-, second-, and third-order modes have directional radiation characteristics that are in agreement with observed data. The assumption of equal initial amplitudes for all of the waves leads to the conclusion that the flapping mode of instability dominates the noise radiatio process of supersonic jets. At a jet temperature of 1370 K, supersonic instability waves are predicted to dominate the noise radiated at high frequency at narrow angles to the jet axis.

Description: The objective of the present work is to study the mixing characteristics of a linear array of supersonic rectangular jets under conditions of screech synchronization. The screech synchronization at a fully expanded jet Mach number of 1.61 is achieved by a precise adjustment of the internozzle spacing. To our knowledge, such an experiment on the resonant mixing of screech synchronized multiple rectangular jets has not been reported before. The results are compared with the case where the screech was suppressed in the multijet configuration.

Description: The objective of the present investigation is to assess the effect of the spatial order of accuracy used for the evaluation of the inviscid fluxes on the resolution of higher order quantitites, such as velocity gradients. The viscous terms are computed as second-order accurate with central difference formulas, even though for the explicit part of the algorithm higher order approximations may be used. A viscous/inviscid method is used, and the outer part of the flowfield is computed with the inviscid flow equations. The viscous boundary-layer type flow region close to the body surface is computed with an algebraic eddy viscosity model. Results obtained with the conservative and nonconservative formulations and the viscous/inviscid approach are compared with available experimental data. The effect of grid refinement on the accuracy of the solution is also presented.

Description: The aerobraking orbital activities of Magelland during the gravity mapping of Venus are discussed. The goal of aerobraking was to circularize Magellan's orbit. By aerobraking the spacecraft into a nearly circula orbit, the Magellan team was able to provide scientists with a different data set to deepen their understanding of what is going on beneath Venus' surface. Before undertaking its gravity-mapping mission, Magellan completed three cycles of radar mapping. This repeated coverage allowed the spacecraft to see some of Venus' geologic features from different viewing angles. Various aspects of the mission are discussed, and maps of Venus are presented.

Description: Spatial correlation among densely packed particles can substantially change their single-scattering properties, thus making questionable the applicability of the independent scattering approximation in calculations of light scattering by planetary regoliths. The same problem arises in geophysics in light scattering computations for snow, frosts, and bare soil. In this paper, we use a dense-medium light-scattering theory based on the introduction of the static structure factor to calculate asymmetry parameters of the phase function for densely packed particles with real refractive indices 1.31 and 1.66, approximating water ice and soil particles, respectively, and imaginary refractive indices 0, 0.01, and 0.3. For sparsely distributed, independently scattering grains, the calculated asymmetry parameters are always positive and always larger than those for densely packed particles. For densely packed grains, the asymmetry parameters may be negative but only for radius-to-wavelength ratios from about 0.1 to about 0.4. With decreasing particle size, the calculated asymmetry parameters tend to zero independently of the compaction state. In the geometrical optics regime, the asymmetry parameters for densely packed scatterers are positive and very close to those for independently scattering grains. These results may have important implications for remote sensing of the Earth and solid planetary surfaces. In particular, it is demonstrated that negative asymmetry parameters derived with some approximate multiple-scattering theories may be physically irrelevant and can be the result of using an inaccurate bidirectional reflection function combined with the ill-conditionally of the inverse scattering problem.

Description: The first comprehensive discussion of the south seasonal polar cap spectra obtained by the Mariner 7 infrared spectrometer in the short-wavelength region (2-4 microns) is presented. The infrared spectra is correlated with images acquired by the wide-angle camera. Significant spectral variation is noted in the cap interior and regions of varying water frost abundance, CO2 ice/frost cover, and CO2-ice path length can be distinguished. Many of these spectral variations correlate with heterogeneity noted in the camera images, but certain significant infrared spectral variations are not discernible in the visible. Simple reflectance models are used to classify the observed spectral variations into four regions. Region I is at the cap edge, where there is enhanced absorption beyond 3 microns inferred to be caused by an increased abundance of water frost. The increase in water abundance over that in the interior is on the level of a few parts per thousand or less. Region II is the typical cap interior characterized by spectral features of CO2 ice at grain sizes of several millimeters to centimeters. These spectra also indicate the presence of water frost at the parts per thousand level. A third, unusual region (III), is defined by three spectra in which weak CO2 absorption features are as much as twice as strong as in the average cap spectra and are assumed to be caused by an increased path length in the CO2. Such large paths are inconsistent with the high reflectance in the visible and at 2.2 microns and suggest layered structures or deposition conditions that are not accounted for in current reflectance models. The final region (IV) is an area of thinning frost coverage or transparent ice well in the interior of the seasonal cap. These spectra are a combination of CO2 and ground signatures.

Description: We report new measurements of the sodium emission intensity seen in a line of sight just above the surface of the Moon. These data show a strong dependence on lunar phase. The emission intensity decreases from a maximum around first quarter (phase angle 90 deg) to very small values near full Moon (phase angle 0 deg). This suggests that the rate of sodium vapor production from the lunar surface is largest at the subsolar point and becomes small near the terminator. However, the sodium emission near full Moon falls below that which would be expected for solar photon-driven processes. Since the solar wind flux decreases substantially when the Moon enters the Earth's magnetotail near full Moon, while the global solar photon flux is undiminished, we suggest that solar wind sputtering is the dominant process for sodium production.

Description: Infrared diffuse reflectance spectra (2.53-25 microns) of some carbonaceous (C) chondrites were measured. The integrated intensity of the absorption bands near 3 microns caused by hydrous minerals were compared with the modal content of hydrous minerals for the meteorites. The CM and CI chondrites show larger values of the intergated intensity than those of the unique C chondrites Y82162, Y86720 and B7904, suggesting that the amount of hydrous minerals in the CM and CI chondrites is larger, which supports the contention that hydrous minerals were dehydrated by thermal metamorphism in the unique chondrites. Orgueil (CI) has the largest value of the integrated intensity among the C chondrites we measured and shows a sharp absorption band at 3685/cm (2.71 microns) that is not seen in the spectra of the CM chondrites. There is an excellent correlation between the observed hydrogen content in C chondrites and the integrated intensity. The CM chondrites show a wide variation in the strength of absorption bands at 1470/cm (6.8 microns), despite the similarity in absorption features near 3 micron for all CM chondites. The 1470/cm band could be due to the presence of some hydrocarbons but may also be a result of terrestrial alteration processes.

Description: Oxygen production from a lunar rock has been experimentally demonstrated for the first time. A 10 g sample of high-Ti basalt 70035 was reduced with hydrogen in seven experiments at temperatures of 900-1050 C and pressures of 14.7-150 psia. In all experiments, water evolution began almost immediately and was essentially complete in tens of minutes. Oxygen yields ranged from 2.93 to 4.61% of the starting sample weight, and showed weak dependence on temperature and pressure. Analysis of the solid samples demonstrated total reduction of Fe(2+) in ilmenite and small degrees of reduction in olivine and pyroxene. Ti O2 was also partially reduced to one or more suboxides. Data from these experiments provide a basis for predicting the yield of oxygen from lunar basalt as well as new constraints on natural reduction in the lunar regolith.

Description: 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.

Description: 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.

Description: 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.

Description: The ORBSIM program was developed for the accurate extraction of geophysical model parameters from Doppler radio tracking data acquired from orbiting planetary spacecraft. The model of the proposed planetary structure is used in a numerical integration of the spacecraft along simulated trajectories around the primary body. Using line of sight (LOS) Doppler residuals, ORBSIM applies fast and efficient modelling and optimization procedures which avoid the traditional complex dynamic reduction of data. ORBSIM produces quantitative geophysical results such as size, depth, and mass. ORBSIM has been used extensively to investigate topographic features on the Moon, Mars, and Venus. The program has proven particulary suitable for modelling gravitational anomalies and mascons. The basic observable for spacecraft-based gravity data is the Doppler frequency shift of a transponded radio signal. The time derivative of this signal carries information regarding the gravity field acting on the spacecraft in the LOS direction (the LOS direction being the path between the spacecraft and the receiving station, either Earth or another satellite). There are many dynamic factors taken into account: earth rotation, solar radiation, acceleration from planetary bodies, tracking station time and location adjustments, etc. The actual trajectories of the spacecraft are simulated using least squares fitted to conic motion. The theoretical Doppler readings from the simulated orbits are compared to actual Doppler observations and another least squares adjustment is made. ORBSIM has three modes of operation: trajectory simulation, optimization, and gravity modelling. In all cases, an initial gravity model of curved and/or flat disks, harmonics, and/or a force table are required input. ORBSIM is written in FORTRAN 77 for batch execution and has been implemented on a DEC VAX 11/780 computer operating under VMS. This program was released in 1985.

Description: The Composite Plate Buckling Analysis Program (COMPPAP) was written to help engineers determine buckling loads of orthotropic (or isotropic) irregularly shaped plates without requiring hand calculations from design curves or extensive finite element modeling. COMPPAP is a one element finite element program that utilizes high-order displacement functions. The high order of the displacement functions enables the user to produce results more accurate than traditional h-finite elements. This program uses these high-order displacement functions to perform a plane stress analysis of a general plate followed by a buckling calculation based on the stresses found in the plane stress solution. The current version assumes a flat plate (constant thickness) subject to a constant edge load (normal or shear) on one or more edges. COMPPAP uses the power method to find the eigenvalues of the buckling problem. The power method provides an efficient solution when only one eigenvalue is desired. Once the eigenvalue is found, the eigenvector, which corresponds to the plate buckling mode shape, results as a by-product. A positive feature of the power method is that the dominant eigenvalue is the first found, which is this case is the plate buckling load. The reported eigenvalue expresses a load factor to induce plate buckling. COMPPAP is written in ANSI FORTRAN 77. Two machine versions are available from COSMIC: a PC version (MSC-22428), which is for IBM PC 386 series and higher computers and compatibles running MS-DOS; and a UNIX version (MSC-22286). The distribution medium for both machine versions includes source code for both single and double precision versions of COMPPAP. The PC version includes source code which has been optimized for implementation within DOS memory constraints as well as sample executables for both the single and double precision versions of COMPPAP. The double precision versions of COMPPAP have been successfully implemented on an IBM PC 386 compatible running MS-DOS, a Sun4 series computer running SunOS, an HP-9000 series computer running HP-UX, and a CRAY X-MP series computer running UNICOS. COMPPAP requires 1Mb of RAM and the BLAS and LINPACK math libraries, which are included on the distribution medium. The COMPPAP documentation provides instructions for using the commercial post-processing package PATRAN for graphical interpretation of COMPPAP output. The UNIX version includes two electronic versions of the documentation: one in LaTex format and one in PostScript format. The standard distribution medium for the PC version (MSC-22428) is a 5.25 inch 1.2Mb MS-DOS format diskette. The standard distribution medium for the UNIX version (MSC-22286) is a .25 inch streaming magnetic tape cartridge (Sun QIC-24) in UNIX tar format. For the UNIX version, alternate distribution media and formats are available upon request. COMPPAP was developed in 1992.

Description: The Composite Plate Buckling Analysis Program (COMPPAP) was written to help engineers determine buckling loads of orthotropic (or isotropic) irregularly shaped plates without requiring hand calculations from design curves or extensive finite element modeling. COMPPAP is a one element finite element program that utilizes high-order displacement functions. The high order of the displacement functions enables the user to produce results more accurate than traditional h-finite elements. This program uses these high-order displacement functions to perform a plane stress analysis of a general plate followed by a buckling calculation based on the stresses found in the plane stress solution. The current version assumes a flat plate (constant thickness) subject to a constant edge load (normal or shear) on one or more edges. COMPPAP uses the power method to find the eigenvalues of the buckling problem. The power method provides an efficient solution when only one eigenvalue is desired. Once the eigenvalue is found, the eigenvector, which corresponds to the plate buckling mode shape, results as a by-product. A positive feature of the power method is that the dominant eigenvalue is the first found, which is this case is the plate buckling load. The reported eigenvalue expresses a load factor to induce plate buckling. COMPPAP is written in ANSI FORTRAN 77. Two machine versions are available from COSMIC: a PC version (MSC-22428), which is for IBM PC 386 series and higher computers and compatibles running MS-DOS; and a UNIX version (MSC-22286). The distribution medium for both machine versions includes source code for both single and double precision versions of COMPPAP. The PC version includes source code which has been optimized for implementation within DOS memory constraints as well as sample executables for both the single and double precision versions of COMPPAP. The double precision versions of COMPPAP have been successfully implemented on an IBM PC 386 compatible running MS-DOS, a Sun4 series computer running SunOS, an HP-9000 series computer running HP-UX, and a CRAY X-MP series computer running UNICOS. COMPPAP requires 1Mb of RAM and the BLAS and LINPACK math libraries, which are included on the distribution medium. The COMPPAP documentation provides instructions for using the commercial post-processing package PATRAN for graphical interpretation of COMPPAP output. The UNIX version includes two electronic versions of the documentation: one in LaTex format and one in PostScript format. The standard distribution medium for the PC version (MSC-22428) is a 5.25 inch 1.2Mb MS-DOS format diskette. The standard distribution medium for the UNIX version (MSC-22286) is a .25 inch streaming magnetic tape cartridge (Sun QIC-24) in UNIX tar format. For the UNIX version, alternate distribution media and formats are available upon request. COMPPAP was developed in 1992.

Description: Structural flaws and cracks may grow under fatigue inducing loads and, upon reaching a critical size, cause structural failure to occur. The growth of these flaws and cracks may occur at load levels well below the ultimate load bearing capability of the structure. The Fatigue Crack Growth Computer Program, NASA/FLAGRO, was developed as an aid in predicting the growth of pre-existing flaws and cracks in structural components of space systems. The earlier version of the program, FLAGRO4, was the primary analysis tool used by Rockwell International and the Shuttle subcontractors for fracture control analysis on the Space Shuttle. NASA/FLAGRO is an enhanced version of the program and incorporates state-of-the-art improvements in both fracture mechanics and computer technology. NASA/FLAGRO provides the fracture mechanics analyst with a computerized method of evaluating the "safe crack growth life" capabilities of structural components. NASA/FLAGRO could also be used to evaluate the damage tolerance aspects of a given structural design. The propagation of an existing crack is governed by the stress field in the vicinity of the crack tip. The stress intensity factor is defined in terms of the relationship between the stress field magnitude and the crack size. The propagation of the crack becomes catastrophic when the local stress intensity factor reaches the fracture toughness of the material. NASA/FLAGRO predicts crack growth using a two-dimensional model which predicts growth independently in two directions based on the calculation of stress intensity factors. The analyst can choose to use either a crack growth rate equation or a nonlinear interpolation routine based on tabular data. The growth rate equation is a modified Forman equation which can be converted to a Paris or Walker equation by substituting different values into the exponent. This equation provides accuracy and versatility and can be fit to data using standard least squares methods. Stress-intensity factor numerical values can be computed for making comparisons or checks of solutions. NASA/FLAGRO can check for failure of a part-through crack in the mode of a through crack when net ligament yielding occurs. NASA/FLAGRO has a number of special subroutines and files which provide enhanced capabilities and easy entry of data. These include crack case solutions, cyclic load spectrums, nondestructive examination initial flaw sizes, table interpolation, and material properties. The materials properties files are divided into two types, a user defined file and a fixed file. Data is entered and stored in the user defined file during program execution, while the fixed file contains already coded-in property value data for many different materials. Prompted input from CRT terminals consists of initial crack definition (which can be defined automatically), rate solution type, flaw type and geometry, material properties (if they are not in the built-in tables of material data), load spectrum data (if not included in the loads spectrum file), and design limit stress levels. NASA/FLAGRO output includes an echo of the input with any error or warning messages, the final crack size, whether or not critical crack size has been reached for the specified stress level, and a life history profile of the crack propagation. NASA/FLAGRO is modularly designed to facilitate revisions and operation on minicomputers. The program was implemented on a DEC VAX 11/780 with the VMS operating system. NASA/FLAGRO is written in FORTRAN77 and has a memory requirement of 1.4 MB. The program was developed in 1986.

Description: 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.

Description: Delaminations near the outer surface of a laminate are susceptible to local buckling and buckling-induced delamination propagation when the laminate is subjected to transverse impact loading. This results in a loss of stiffness and strength. TRBUCKL is an unique dynamic delamination buckling and delamination propagation analysis capability that can be incorporated into the structural analysis program, NASTRAN. This capability will aid engineers in the design of structures incorporating composite laminates. The capability consists of: (1) a modification of the direct time integration solution sequence which provides a new analysis algorithm that can be used to predict delamination buckling in a laminate subjected to dynamic loading; and (2) a new method of modeling the composite laminate using plate bending elements and multipoint constraints. The capability now exists to predict the time at which the onset of dynamic delamination buckling occurs, the dynamic buckling mode shape, and the dynamic delamination strain energy release rate. A procedure file for NASTRAN, TRBUCKL predicts both impact induced buckling in composite laminates with initial delaminations and the strain energy release rate due to extension of the delamination. In addition, the file is useful in calculating the dynamic delamination strain energy release rate for a composite laminate under impact loading. This procedure simplifies the simulation of progressive crack extension. TRBUCKL has been incorporated into COSMIC NASTRAN. TRBUCKL is a DMAP Alter for NASTRAN. It is intended for use only with the COSMIC NASTRAN Direct Transient Analysis (RF 9) solution sequence. The program is available as a listing only. TRBUCKL was developed in 1987.

Description: 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.

Description: Predictions of fatigue crack growth behavior can be made with the Fatigue Crack Growth Structural Analysis (FASTRAN II) computer program. As cyclic loads are applied to a selected crack configuration with an initial crack size, FASTRAN II predicts crack growth as a function of cyclic load history until either a desired crack size is reached or failure occurs. FASTRAN II is based on plasticity-induced crack-closure behavior of cracks in metallic materials and accounts for load-interaction effects, such as retardation and acceleration, under variable-amplitude loading. The closure model is based on the Dugdale model with modifications to allow plastically deformed material to be left along the crack surfaces as the crack grows. Plane stress and plane strain conditions, as well as conditions between these two, can be simulated in FASTRAN II by using a constraint factor on tensile yielding at the crack front to approximately account for three-dimensional stress states. FASTRAN II contains seventeen predefined crack configurations (standard laboratory fatigue crack growth rate specimens and many common crack configurations found in structures); and the user can define one additional crack configuration. The baseline crack growth rate properties (effective stress-intensity factor against crack growth rate) may be given in either equation or tabular form. For three-dimensional crack configurations, such as surface cracks or corner cracks at holes or notches, the fatigue crack growth rate properties may be different in the crack depth and crack length directions. Final failure of the cracked structure can be modelled with fracture toughness properties using either linear-elastic fracture mechanics (brittle materials), a two-parameter fracture criterion (brittle to ductile materials), or plastic collapse (extremely ductile materials). The crack configurations in FASTRAN II can be subjected to either constant-amplitude, variable-amplitude or spectrum loading. The applied loads may be either tensile or compressive. Several standardized aircraft flight-load histories, such as TWIST, Mini-TWIST, FALSTAFF, Inverted FALSTAFF, Felix and Gaussian, are included as options. FASTRAN II also includes two other methods that will help the user input spectrum load histories. The two methods are: (1) a list of stress points, and (2) a flight-by-flight history of stress points. Examples are provided in the user manual. Developed as a research program, FASTRAN II has successfully predicted crack growth in many metallic materials under various aircraft spectrum loading. A computer program DKEFF which is a part of the FASTRAN II package was also developed to analyze crack growth rate data from laboratory specimens to obtain the effective stress-intensity factor against crack growth rate relations used in FASTRAN II. FASTRAN II is written in standard FORTRAN 77. It has been successfully compiled and implemented on Sun4 series computers running SunOS and on IBM PC compatibles running MS-DOS using the Lahey F77L FORTRAN compiler. Sample input and output data are included with the FASTRAN II package. The UNIX version requires 660K of RAM for execution. The standard distribution medium for the UNIX version (LAR-14865) is a .25 inch streaming magnetic tape cartridge in UNIX tar format. It is also available on a 3.5 inch diskette in UNIX tar format. The standard distribution medium for the MS-DOS version (LAR-14944) is a 5.25 inch 360K MS-DOS format diskette. The contents of the diskette are compressed using the PKWARE archiving tools. The utility to unarchive the files, PKUNZIP.EXE, is included. The program was developed in 1984 and revised in 1992. Sun4 and SunOS are trademarks of Sun Microsystems, Inc. IBM PC is a trademark of International Business Machines Corp. MS-DOS is a trademark of Microsoft, Inc. F77L is a trademark of the Lahey Computer Systems, Inc. UNIX is a registered trademark of AT&T Bell Laboratories. PKWARE and PKUNZIP are trademarks of PKWare, Inc.

Description: The Panel Analysis and Sizing Code (PASCO) was developed for the buckling and vibration analysis and sizing of prismatic structures having an arbitrary cross section. PASCO is primarily intended for analyzing and sizing stiffened panels made of laminated orthotropic materials and is of particular value in analyzing and sizing filamentary composite structures. When used in the analysis mode, PASCO calculates laminate stiffnesses, lamina stress and strains (including the effects of temperature and panel bending), buckling loads, vibration frequencies, and overall panel stiffness. When used in the sizing mode, PASCO adjusts sizing variables to provide a low-mass panel design that carries a set of specified loadings without exceeding buckling or material strength allowables and that meets other design requirements such as upper and lower bounds on sizing variables, upper and lower bounds on overall bending, extensional and shear stiffnesses, and lower bounds on vibration frequencies. Although emphasis in PASCO is placed on flat panels having several identical bays, the only restriction on configuration modeling is that the structure is assumed to be prismatic. In addition, it is assumed that loads and temperatures do not vary along the length of a panel. Because of their wide application in aerospace structures, stiffened panels are readily handled by PASCO. The panel cross section may be composed of an arbitrary assemblage of thin, flat, rectangular plate elements that are connected together along their longitudinal edges. Each plate element consists of a balanced symmetric laminate of any number of layers of orthotropic material. Any group of element widths, layer thicknesses, and layer orientation angles may be selected as sizing variables. Substructuring is available to increase the efficiency of the analysis and to simplify the modeling of complicated structures. The Macintosh version of PASCO includes an interactive, graphic preprocessor called MacPASCO. The main objective of MacPASCO is to make the use of PASCO faster, simpler, and less error-prone. By using a graphical user interface (GUI), MacPASCO simplifies the specification of panel geometry and reduces user input errors, thus making the modeling and analysis of panel designs more efficient. The user draws the initial structural geometry on the computer screen, then uses a combination of graphic and text inputs to: refine the structural geometry, specify information required for analysis such as panel load conditions, and define design variables and constraints for minimum-mass optimization. Composite panel design is an ideal application because the graphical user interface can: serve as a visual aid, eliminate the tedious aspects of text-based input, and eliminate many sources of input errors. The current version of MacPASCO does not implement all the modeling features of PASCO, but has been found to be sufficient for many users. Many difficulties common to text-based inputs are avoided because MacPASCO uses a GUI. First, the graphic displays eliminate syntax errors, like misplaced commas and incorrect command names, because there is no text-based syntax. Second, graphic displays allow the user to see the geometry as it is created and immediately detect and correct any errors. Third, MacPASCO's drawing tools have been designed to avoid modeling errors. Fourth, the graphic displays make revisions to existing structural designs much easier and less error-prone by eliminating the need for the user to conceptualize the text input as geometry. The user can work directly with the geometry displayed on the screen. Finally, MacPASCO automatically generates the correct PASCO input from the geometry displayed on the screen. This input file can be used with any machine version of PASCO to actually perform the analysis and sizing and to output results. The DEC VAX version of PASCO is written in FORTRAN IV for batch execution and has been implemented on a DEC VAX series computer. The Macintosh version of PASCO was developed for Macintosh II series computers with at least 2Mb of RAM running MPW Pascal 3.0 and Language Systems FORTRAN 2.0 under the MPW programming environment. It includes MPW compatible makefiles for compiling the source code. The Macintosh version uses input files compatible with versions of PASCO running on different platforms. MacPASCO is written in Macintosh Programmers Workbench 3.0, MPW Pascal 3.0, and MacAPP 2.0. The Pascal source code is included on the distribution diskette. MacAPP is a development library which is not included. MacPASCO requires a Mac Plus, SE/30, or MacII, IIx, IIcx, IIci, or IIfx running System 6.0 or greater. MacPASCO is System 7.0 compatible. A minimum of 2Mb of RAM is required for execution. The Macintosh version of PASCO is distributed on four 3.5 inch 800K Macintosh format diskettes. The DEC VAX version is distributed on a 9-track 1600 BPI magnetic tape. The PASCO program was developed in 1981, adapted to the DEC VAX in 1983 and to the Macintosh in 1991. MacPASCO was released in 1992.

Description: 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.

Description: 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.

Description: 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.

Description: 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.

Description: PLAN2D is a FORTRAN computer program for the plastic analysis of planar rigid frame structures. Given a structure and loading pattern as input, PLAN2D calculates the ultimate load that the structure can sustain before collapse. Element moments and plastic hinge rotations are calculated for the ultimate load. The location of hinges required for a collapse mechanism to form are also determined. The program proceeds in an iterative series of linear elastic analyses. After each iteration the resulting elastic moments in each member are compared to the reserve plastic moment capacity of that member. The member or members that have moments closest to their reserve capacity will determine the minimum load factor and the site where the next hinge is to be inserted. Next, hinges are inserted and the structural stiffness matrix is reformulated. This cycle is repeated until the structure becomes unstable. At this point the ultimate collapse load is calculated by accumulating the minimum load factor from each previous iteration and multiplying them by the original input loads. PLAN2D is based on the program STAN, originally written by Dr. E.L. Wilson at U.C. Berkeley. PLAN2D has several limitations: 1) Although PLAN2D will detect unloading of hinges it does not contain the capability to remove hinges; 2) PLAN2D does not allow the user to input different positive and negative moment capacities and 3) PLAN2D does not consider the interaction between axial and plastic moment capacity. Axial yielding and buckling is ignored as is the reduction in moment capacity due to axial load. PLAN2D is written in FORTRAN and is machine independent. It has been tested on an IBM PC and a DEC MicroVAX. The program was developed in 1988.

Description: Predictions of fatigue crack growth behavior can be made with the Fatigue Crack Growth Structural Analysis (FASTRAN II) computer program. As cyclic loads are applied to a selected crack configuration with an initial crack size, FASTRAN II predicts crack growth as a function of cyclic load history until either a desired crack size is reached or failure occurs. FASTRAN II is based on plasticity-induced crack-closure behavior of cracks in metallic materials and accounts for load-interaction effects, such as retardation and acceleration, under variable-amplitude loading. The closure model is based on the Dugdale model with modifications to allow plastically deformed material to be left along the crack surfaces as the crack grows. Plane stress and plane strain conditions, as well as conditions between these two, can be simulated in FASTRAN II by using a constraint factor on tensile yielding at the crack front to approximately account for three-dimensional stress states. FASTRAN II contains seventeen predefined crack configurations (standard laboratory fatigue crack growth rate specimens and many common crack configurations found in structures); and the user can define one additional crack configuration. The baseline crack growth rate properties (effective stress-intensity factor against crack growth rate) may be given in either equation or tabular form. For three-dimensional crack configurations, such as surface cracks or corner cracks at holes or notches, the fatigue crack growth rate properties may be different in the crack depth and crack length directions. Final failure of the cracked structure can be modelled with fracture toughness properties using either linear-elastic fracture mechanics (brittle materials), a two-parameter fracture criterion (brittle to ductile materials), or plastic collapse (extremely ductile materials). The crack configurations in FASTRAN II can be subjected to either constant-amplitude, variable-amplitude or spectrum loading. The applied loads may be either tensile or compressive. Several standardized aircraft flight-load histories, such as TWIST, Mini-TWIST, FALSTAFF, Inverted FALSTAFF, Felix and Gaussian, are included as options. FASTRAN II also includes two other methods that will help the user input spectrum load histories. The two methods are: (1) a list of stress points, and (2) a flight-by-flight history of stress points. Examples are provided in the user manual. Developed as a research program, FASTRAN II has successfully predicted crack growth in many metallic materials under various aircraft spectrum loading. A computer program DKEFF which is a part of the FASTRAN II package was also developed to analyze crack growth rate data from laboratory specimens to obtain the effective stress-intensity factor against crack growth rate relations used in FASTRAN II. FASTRAN II is written in standard FORTRAN 77. It has been successfully compiled and implemented on Sun4 series computers running SunOS and on IBM PC compatibles running MS-DOS using the Lahey F77L FORTRAN compiler. Sample input and output data are included with the FASTRAN II package. The UNIX version requires 660K of RAM for execution. The standard distribution medium for the UNIX version (LAR-14865) is a .25 inch streaming magnetic tape cartridge in UNIX tar format. It is also available on a 3.5 inch diskette in UNIX tar format. The standard distribution medium for the MS-DOS version (LAR-14944) is a 5.25 inch 360K MS-DOS format diskette. The contents of the diskette are compressed using the PKWARE archiving tools. The utility to unarchive the files, PKUNZIP.EXE, is included. The program was developed in 1984 and revised in 1992. Sun4 and SunOS are trademarks of Sun Microsystems, Inc. IBM PC is a trademark of International Business Machines Corp. MS-DOS is a trademark of Microsoft, Inc. F77L is a trademark of the Lahey Computer Systems, Inc. UNIX is a registered trademark of AT&T Bell Laboratories. PKWARE and PKUNZIP are trademarks of PKWare, Inc.

Description: 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.

Description: 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.

Description: The SNIP program is a FORTRAN computer code that generates NASTRAN structural model thermal loads when given SINDA (or similar thermal model) temperature results. SNIP correlates thermal nodes to structural elements to interface SINDA finite difference thermal models with NASTRAN finite element structural models. Node-to-element correlation includes determining which SINDA nodes should be related to each NASTRAN element and calculating a weighing factor for temperatures associated with each element-related thermal node. SNIP provides structural model thermal loads that accurately reflect thermal model results while reducing the time required to interface thermal and structural models as compared to other methods. SNIP uses thermal model geometry to search the three-dimensional space around each structural element for the nearest thermal nodes. Thermal model geometry is the combination of standard thermal model temperature results from SINDA and structural model geometry from NASTRAN. Thermal and structural models must both be defined in the same, single Cartesian coordinate system. The thermal nodes located nearest each element are used to determine element temperature for thermal distortion and stress analysis. The program shapes the three-dimensional search region while the user controls the size. With these region specifications, the numerical coding of thermal nodes, and the structural element numbers; the code can provide for the separation of substructures during correlation. The input to SNIP contains a file of thermal model temperature results and a physical location of each thermal node in three-dimensional space, combined in a SNIP-unique format. The input also contains a standard NASTRAN input deck for a model made up of plate, shell, beam, and bar elements. SNIP supports the CTRIA, CQUAD, CBAR, and CBEAM elements of NASTRAN. The user adjusts the input parameters in the source code which control the node-to-element correlation. The program outputs NASTRAN element temperature load cards for each element and NASTRAN case control cards for each temperature load set. SNIP also outputs a list of elements that contains the numbers of the SINDA nodes related to each NASTRAN element and the weight that is given to each node in temperature calculations. SNIP is written in ANSI standard FORTRAN 77. The PC version requires a PC FORTRAN compiler and has compiled successfully using Lahey FORTRAN v. 3.0. A core memory of 300k is recommended. The program was developed in 1987.

Description: 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.

Description: 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.

Description: 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.

Description: 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.

Description: 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.

Description: 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.

Description: 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.

Description: 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.

Description: 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.

Description: 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.

Description: 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.

Description: 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.

Description: 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.

Description: 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.

Description: 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.

Description: 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.

Description: NPLOT is an interactive computer graphics program for plotting undeformed and deformed NASTRAN finite element models (FEMs). Although there are many commercial codes already available for plotting FEMs, these have limited use due to their cost, speed, and lack of features to view BAR elements. NPLOT was specifically developed to overcome these limitations. On a vector type graphics device the two best ways to show depth are by hidden line plotting or haloed line plotting. A hidden line algorithm generates views of models with all hidden lines removed, and a haloed line algorithm displays views with aft lines broken in order to show depth while keeping the entire model visible. A haloed line algorithm is especially useful for plotting models composed of many line elements and few surface elements. The most important feature of NPLOT is its ability to create both hidden line and haloed line views accurately and much more quickly than with any other existing hidden or haloed line algorithms. NPLOT is also capable of plotting a normal wire frame view to display all lines of a model. NPLOT is able to aid in viewing all elements, but it has special features not generally available for plotting BAR elements. These features include plotting of TRUE LENGTH and NORMALIZED offset vectors and orientation vectors. Standard display operations such as rotation and perspective are possible, but different view planes such as X-Y, Y-Z, and X-Z may also be selected. Another display option is the Z-axis cut which allows a portion of the fore part of the model to be cut away to reveal details of the inside of the model. A zoom function is available to terminals with a locator (graphics cursor, joystick, etc.). The user interface of NPLOT is designed to make the program quick and easy to use. A combination of menus and commands with help menus for detailed information about each command allows experienced users greater speed and efficiency. Once a plot is on the screen the interface becomes command driven, enabling the user to manipulate the display or execute a command without having to return to the menu. NPLOT is also able to plot deformed shapes allowing it to perform post-processing. The program can read displacements, either static displacements or eigenvectors, from a MSC/NASTRAN F06 file or a UAI/NASTRAN PRT file. The displacements are written into a unformatted scratch file where they are available for rapid access when the user wishes to display a deformed shape. All subcases or mode shapes can be read in at once. Then it is easy to enable the deformed shape, to change subcases or mode shapes and to change the scale factor for subsequent plots. NPLOT is written in VAX FORTRAN for DEC VAX series computers running VMS. As distributed, the NPLOT source code makes calls to the DI3000 graphics package from Precision Visuals; however, a set of interface routines is provided to translate the DI3000 calls into Tektronix PLOT10/TCS graphics library calls so that NPLOT can use the standard Tektronix 4010 which many PC terminal emulation software programs support. NPLOT is available in VAX BACKUP format on a 9-track 1600 BPI DEC VAX BACKUP format magnetic tape (standard media) or a TK50 tape cartridge. This program was developed in 1991. DEC, VAX, VMS, and TK50 are trademarks of Digital Equipment Corporation. Tektronix, PLOT10, and TCS are trademarks of Tektronix, Inc. DI3000 is a registered trademark of Precision Visuals, Inc. NASTRAN is a registered trademark of the National Aeronautics and Space Administration. MSC/ is a trademark of MacNeal-Schwendler Corporation. UAI is a trademark of Universal Analytics, Inc.

Description: 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.

Description: LATDYN is a computer code for modeling the Large Angle Transient DYNamics of flexible articulating structures and mechanisms involving joints about which members rotate through large angles. LATDYN extends and brings together some of the aspects of Finite Element Structural Analysis, Multi-Body Dynamics, and Control System Analysis; three disciplines that have been historically separate. It combines significant portions of their distinct capabilities into one single analysis tool. The finite element formulation for flexible bodies in LATDYN extends the conventional finite element formulation by using a convected coordinate system for constructing the equation of motion. LATDYN's formulation allows for large displacements and rotations of finite elements subject to the restriction that deformations within each are small. Also, the finite element approach implemented in LATDYN provides a convergent path for checking solutions simply by increasing mesh density. For rigid bodies and joints LATDYN borrows extensively from methodology used in multi-body dynamics where rigid bodies may be defined and connected together through joints (hinges, ball, universal, sliders, etc.). Joints may be modeled either by constraints or by adding joint degrees of freedom. To eliminate error brought about by the separation of structural analysis and control analysis, LATDYN provides symbolic capabilities for modeling control systems which are integrated with the structural dynamic analysis itself. Its command language contains syntactical structures which perform symbolic operations which are also interfaced directly with the finite element structural model, bypassing the modal approximation. Thus, when the dynamic equations representing the structural model are integrated, the equations representing the control system are integrated along with them as a coupled system. This procedure also has the side benefit of enabling a dramatic simplification of the user interface for modeling control systems. Three FORTRAN computer programs, the LATDYN Program, the Preprocessor, and the Postprocessor, make up the collective LATDYN System. The Preprocessor translates user commands into a form which can be used while the LATDYN program provides the computational core. The Postprocessor allows the user to interactively plot and manage a database of LATDYN transient analysis results. It also includes special facilities for modeling control systems and for programming changes to the model which take place during analysis sequence. The documentation includes a Demonstration Problem Manual for the evaluation and verification of results and a Postprocessor guide. Because the program should be viewed as a byproduct of research on technology development, LATDYN's scope is limited. It does not have a wide library of finite elements, and 3-D Graphics are not available. Nevertheless, it does have a measure of "user friendliness". The LATDYN program was developed over a period of several years and was implemented on a CDC NOS/VE & Convex Unix computer. It is written in FORTRAN 77 and has a virtual memory requirement of 1.46 MB. The program was validated on a DEC MICROVAX operating under VMS 5.2.

Description: The increasing number of applications of fiber-reinforced composites in industry demands a detailed understanding of their material properties and behavior. A three-dimensional finite-element computer program called PAFAC (Plastic and Failure Analysis of Composites) has been developed for the elastic-plastic analysis of fiber-reinforced composite materials and structures. The evaluation of stresses and deformations at edges, cut-outs, and joints is essential in understanding the strength and failure for metal-matrix composites since the onset of plastic yielding starts very early in the loading process as compared to the composite's ultimate strength. Such comprehensive analysis can only be achieved by a finite-element program like PAFAC. PAFAC is particularly suited for the analysis of laminated metal-matrix composites. It can model the elastic-plastic behavior of the matrix phase while the fibers remain elastic. Since the PAFAC program uses a three-dimensional element, the program can also model the individual layers of the laminate to account for thickness effects. In PAFAC, the composite is modeled as a continuum reinforced by cylindrical fibers of vanishingly small diameter which occupy a finite volume fraction of the composite. In this way, the essential axial constraint of the phases is retained. Furthermore, the local stress and strain fields are uniform. The PAFAC finite-element solution is obtained using the displacement method. Solution of the nonlinear equilibrium equations is obtained with a Newton-Raphson iteration technique. The elastic-plastic behavior of composites consisting of aligned, continuous elastic filaments and an elastic-plastic matrix is described in terms of the constituent properties, their volume fractions, and mutual constraints between phases indicated by the geometry of the microstructure. The program uses an iterative procedure to determine the overall response of the laminate, then from the overall response determines the stress state in each phase of the composite material. Failure of the fibers or matrix within an element can also be modeled by PAFAC. PAFAC is written in FORTRAN IV for batch execution and has been implemented on a CDC CYBER 170 series computer with a segmented memory requirement of approximately 66K (octal) of 60 bit words. PAFAC was developed in 1982.

Description: 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.

Description: The Panel Analysis and Sizing Code (PASCO) was developed for the buckling and vibration analysis and sizing of prismatic structures having an arbitrary cross section. PASCO is primarily intended for analyzing and sizing stiffened panels made of laminated orthotropic materials and is of particular value in analyzing and sizing filamentary composite structures. When used in the analysis mode, PASCO calculates laminate stiffnesses, lamina stress and strains (including the effects of temperature and panel bending), buckling loads, vibration frequencies, and overall panel stiffness. When used in the sizing mode, PASCO adjusts sizing variables to provide a low-mass panel design that carries a set of specified loadings without exceeding buckling or material strength allowables and that meets other design requirements such as upper and lower bounds on sizing variables, upper and lower bounds on overall bending, extensional and shear stiffnesses, and lower bounds on vibration frequencies. Although emphasis in PASCO is placed on flat panels having several identical bays, the only restriction on configuration modeling is that the structure is assumed to be prismatic. In addition, it is assumed that loads and temperatures do not vary along the length of a panel. Because of their wide application in aerospace structures, stiffened panels are readily handled by PASCO. The panel cross section may be composed of an arbitrary assemblage of thin, flat, rectangular plate elements that are connected together along their longitudinal edges. Each plate element consists of a balanced symmetric laminate of any number of layers of orthotropic material. Any group of element widths, layer thicknesses, and layer orientation angles may be selected as sizing variables. Substructuring is available to increase the efficiency of the analysis and to simplify the modeling of complicated structures. The Macintosh version of PASCO includes an interactive, graphic preprocessor called MacPASCO. The main objective of MacPASCO is to make the use of PASCO faster, simpler, and less error-prone. By using a graphical user interface (GUI), MacPASCO simplifies the specification of panel geometry and reduces user input errors, thus making the modeling and analysis of panel designs more efficient. The user draws the initial structural geometry on the computer screen, then uses a combination of graphic and text inputs to: refine the structural geometry, specify information required for analysis such as panel load conditions, and define design variables and constraints for minimum-mass optimization. Composite panel design is an ideal application because the graphical user interface can: serve as a visual aid, eliminate the tedious aspects of text-based input, and eliminate many sources of input errors. The current version of MacPASCO does not implement all the modeling features of PASCO, but has been found to be sufficient for many users. Many difficulties common to text-based inputs are avoided because MacPASCO uses a GUI. First, the graphic displays eliminate syntax errors, like misplaced commas and incorrect command names, because there is no text-based syntax. Second, graphic displays allow the user to see the geometry as it is created and immediately detect and correct any errors. Third, MacPASCO's drawing tools have been designed to avoid modeling errors. Fourth, the graphic displays make revisions to existing structural designs much easier and less error-prone by eliminating the need for the user to conceptualize the text input as geometry. The user can work directly with the geometry displayed on the screen. Finally, MacPASCO automatically generates the correct PASCO input from the geometry displayed on the screen. This input file can be used with any machine version of PASCO to actually perform the analysis and sizing and to output results. The DEC VAX version of PASCO is written in FORTRAN IV for batch execution and has been implemented on a DEC VAX series computer. The Macintosh version of PASCO was developed for Macintosh II series computers with at least 2Mb of RAM running MPW Pascal 3.0 and Language Systems FORTRAN 2.0 under the MPW programming environment. It includes MPW compatible makefiles for compiling the source code. The Macintosh version uses input files compatible with versions of PASCO running on different platforms. MacPASCO is written in Macintosh Programmers Workbench 3.0, MPW Pascal 3.0, and MacAPP 2.0. The Pascal source code is included on the distribution diskette. MacAPP is a development library which is not included. MacPASCO requires a Mac Plus, SE/30, or MacII, IIx, IIcx, IIci, or IIfx running System 6.0 or greater. MacPASCO is System 7.0 compatible. A minimum of 2Mb of RAM is required for execution. The Macintosh version of PASCO is distributed on four 3.5 inch 800K Macintosh format diskettes. The DEC VAX version is distributed on a 9-track 1600 BPI magnetic tape. The PASCO program was developed in 1981, adapted to the DEC VAX in 1983 and to the Macintosh in 1991. MacPASCO was released in 1992.

Description: 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.

Description: 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.

Description: 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.

Description: The Program for Solving the General-Frequency Unsteady Two-Dimensional Transonic Small-Disturbance Equation, XTRAN2L, is used to calculate time-accurate, finite-difference solutions of the nonlinear, small-disturbance potential equation for two- dimensional transonic flow about airfoils. The code can treat forced harmonic, pulse, or aeroelastic transient type motions. XTRAN2L uses a transonic small-disturbance equation that incorporates a time accurate finite-difference scheme. Airfoil flow tangency boundary conditions are defined to include airfoil contour, chord deformation, nondimensional plunge displacement, pitch, and trailing edge control surface deflection. Forced harmonic motion can be based on: 1) coefficients of harmonics based on information from each quarter period of the last cycle of harmonic motion; or 2) Fourier analyses of the last cycle of motion. Pulse motion (an alternate to forced harmonic motion) in which the airfoil is given a small prescribed pulse in a given mode of motion, and the aerodynamic transients are calculated. An aeroelastic transient capability is available within XTRAN2L, wherein the structural equations of motion are coupled with the aerodynamic solution procedure for simultaneous time-integration. The wake is represented as a slit downstream of the airfoil trailing edge. XTRAN2L includes nonreflecting farfield boundary conditions. XTRAN2L was developed on a CDC CYBER mainframe running under NOS 2.4. It is written in FORTRAN 5 and uses overlays to minimize storage requirements. The program requires 120K of memory in overlayed form. XTRAN2L was developed in 1987.

Description: BUCKO is a computer program developed to predict the buckling load of a rectangular compression-loaded orthotropic plate with a centrally located cutout. The plate is assumed to be a balanced, symmetric laminate of uniform thickness. The cutout shape can be elliptical, circular, rectangular, or square. The BUCKO package includes sample data that demonstrates the essence of the program and its ease of usage. BUCKO uses an approximate one-dimensional formulation of the classical two-dimensional buckling problem following the Kantorovich method. The boundary conditions are considered to be simply supported unloaded edges and either clamped or simply supported loaded edges. The plate is loaded in uniaxial compression by either uniformly displacing or uniformly stressing two opposite edges of the plate. The BUCKO analysis consists of two parts: calculation of the inplane stress distribution prior to buckling, and calculation of the plate axial load and displacement at buckling. User input includes plate planform and cutout geometry, plate membrane and bending stiffnesses, finite difference parameters, boundary condition data, and loading data. Results generated by BUCKO are the prebuckling strain energy, inplane stress resultants, buckling mode shape, critical end shortening, and average axial and transverse strains at buckling. BUCKO is written in FORTRAN V for batch execution and has been implemented on a CDC CYBER 170 series computer operating under NOS with a central memory requirement of approximately 343K of 60 bit words. This program was developed in 1984 and was last updated in 1990.

Description: The NASCRIN program was developed for analyzing two-dimensional flow fields in supersonic combustion ramjet (scramjet) inlets. NASCRIN solves the two-dimensional Euler or Navier-Stokes equations in conservative form by an unsplit, explicit, two-step finite-difference method. A more recent explicit-implicit, two-step scheme has also been incorporated in the code for viscous flow analysis. An algebraic, two-layer eddy-viscosity model is used for the turbulent flow calculations. NASCRIN can analyze both inviscid and viscous flows with no struts, one strut, or multiple struts embedded in the flow field. NASCRIN can be used in a quasi-three-dimensional sense for some scramjet inlets under certain simplifying assumptions. Although developed for supersonic internal flow, NASCRIN may be adapted to a variety of other flow problems. In particular, it should be readily adaptable to subsonic inflow with supersonic outflow, supersonic inflow with subsonic outflow, or fully subsonic flow. The NASCRIN program is available for batch execution on the CDC CYBER 203. The vectorized FORTRAN version was developed in 1983. NASCRIN has a central memory requirement of approximately 300K words for a grid size of about 3,000 points.

Description: The problem of axisymmetric transonic flow is of interest not only because of the practical application to missile and launch vehicle aerodynamics, but also because of its relation to fully three-dimensional flow in terms of the area rule. The RAXBOD computer program was developed for the analysis of steady, inviscid, irrotational, transonic flow over axisymmetric bodies in free air. RAXBOD uses a finite-difference relaxation method to numerically solve the exact formulation of the disturbance velocity potential with exact surface boundary conditions. Agreement with available experimental results has been good in cases where viscous effects and wind-tunnel wall interference are not important. The governing second-order partial differential equation describing the flow potential is replaced by a system of finite difference equations, including Jameson's "rotated" difference scheme at supersonic points. A stretching is applied to both the normal and tangential coordinates such that the infinite physical space is mapped onto a finite computational space. The boundary condition at infinity can be applied directly and there is no need for an asymptotic far-field solution. The system of finite difference equations is solved by a column relaxation method. In order to obtain both rapid convergence and any desired resolution, the relaxation is performed iteratively on successively refined grids. Input to RAXBOD consists of a description of the body geometry, the free stream conditions, and the desired resolution control parameters. Output from RAXBOD includes computed geometric parameters in the normal and tangential directions, iteration history information, drag coefficients, flow field data in the computational plane, and coordinates of the sonic line. This program is written in FORTRAN IV for batch execution and has been implemented on a CDC 6600 computer with an overlayed central memory requirement of approximately 40K (octal) of 60 bit words. Optional plotted output can be generated for the Calcomp plotting system. The RAXBOD program was developed in 1976.

Description: Two separate and distinct theories are incorporated in this computer program to estimate the lift-induced pressures existent on a wing-body combination. These are (1) the second-order shock-expansion theory, which is used to obtain the lifting pressures on the body alone at small angles of attack, and (2) the linear-theory integral equations, which is used to evaluate the lifting pressures induced by the wing. These equations relate the local surface slope at a point on the lifting surface to the pressure differential at the point and the influence of the pressures upstream of the point. The numerical solution of these equations is effected by treating the wing-planform as a composite of elemental rectangles and applying summation techniques to satisfy the necessary integral relations. Most of the input required by this program is involved with the description of the missile planform geometry. The output consists of the computed value of the lifting pressure slope (the differential pressure coefficient per degree angle of attack) for each of the elements in the planform array. A force and moment summary is presented for the configuration under consideration.

Description: A modified strip analysis has been developed for rapidly predicting flutter of finite-span, swept or unswept wings at subsonic to hypersonic speeds. The method employs distributions of aerodynamic parameters which may be evaluated from any suitable linear or nonlinear steady-flow theory or from measured steady-flow load distributions for the underformed wing. The method has been shown to give good flutter results for a broad range of wings at Mach number from 0 to as high as 15.3. The principles of the modified strip analysis may be summarized as follows: Variable section lift-curve slope and aerodynamic center are substituted respectively, for the two-dimensional incompressible-flow values of 2 pi and quarter chord which were employed by Barmby, Cunningham, and Garrick. Spanwise distributions of these steady-flow section aerodynamic parameters, which are pertinent to the desired planform and Mach number, are used. Appropriate values of Mach number-dependent circulation functions are obtained from two-dimensional unsteady compressible-flow theory. Use of the modified strip analysis avoids the necessity of reevaluating a number of loading parameters for each value of reduced frequency, since only the modified circulation functions, and of course the reduced frequency itself, vary with frequency. It is therefore practical to include in the digital computing program a very brief logical subroutine, which automatically selects reduced-frequency values that converge on a flutter solution. The problem of guessing suitable reduced-frequency values is thus eliminated, so that a large number of flutter points can be completely determined in a single brief run on the computing machine. If necessary, it is also practical to perform the calculations manually. Flutter characteristics have been calculated by the modified strip analysis and compared with results of other calculations and with experiments for Mach numbers up to 15.3 and for wings with sweep angles from 0 degrees to 52.5 degrees, aspect ratios from 2.0 to 7.4, taper ratios from 0.2 to 1.0, and center-of-gravity positions between 34% chord and 59% chord. These ranges probably cover the great majority of wings that are of practical interest with the exception of very low-aspect-ratio surfaces such as delta wings and missile fins. This program has been implemented on the IBM 7094.

Description: The SUBAERF2 program was developed to provide for the aerodynamic analysis and design of low speed wing flap systems. SUBAERF2 is based on a linearized theory lifting surface solution. It is particularly well suited to configurations which, because of high speed flight requirements, must employ thin wings with highly swept leading edges. The program is applicable to wings with either sharp or rounded leading edges. This program is a new and improved version of LAR-13116 and LAR-12987, which it replaces. The low speed aerodynamic analysis method used in SUBAERF2 provides estimates of wing performance which include the effects of attainable leading-edge thrust and vortex lift. This basic aerodynamic analysis method has been improved to provide for the convenient, efficient and accurate treatment of simple leading-edge and trailing-edge flap systems. The user inputs flap geometry directly. Solutions can be found for various combinations of leading and trailing edge flap deflections. The program provides for the simultaneous analysis of up to 25 pairs of leading-edge and trailing-edge flap deflection schedules. A revised attainable thrust algorithm improves accuracy at the low Mach numbers sometimes encountered in wind tunnel testing. Also added is a means of estimating the distribution of leading edge separation vortex forces. The revised program has been particularly useful in the subsonic analysis of vehicles designed for supersonic cruise. The SUBAERF2 program is written in FORTRAN V for batch execution and has been implemented on a CDC 175 computer operating under NOS 2.4 with a central memory requirement of approximately 115K (octal) of 60 bit words. This program was originally developed in 1983 and later revised in 1988.

Description: WAVDRAG calculates the supersonic zero-lift wave drag of complex aircraft configurations. The numerical model of an aircraft is used throughout the design process from concept to manufacturing. WAVDRAG incorporates extended geometric input capabilities to permit use of a more accurate mathematical model. With WAVDRAG, the engineer can define aircraft components as fusiform or nonfusiform in terms of non-intersecting contours in any direction or more traditional parallel contours. In addition, laterally asymmetric configurations can be simulated. The calculations in WAVDRAG are based on Whitcomb's area-rule computation of equivalent-bodies, with modifications for supersonic speed. Instead of using a single equivalent-body, WAVDRAG calculates a series of equivalent-bodies, one for each roll angle. The total aircraft configuration wave drag is the integrated average of the equivalent-body wave drags through the full roll range of 360 degrees. WAVDRAG currently accepts up to 30 user-defined components containing a maximum of 50 contours as geometric input. Each contour contains a maximum of 50 points. The Mach number, angle-of-attack, and coordinates of angle-of-attack rotation are also input. The program warns of any fusiform-body line segments having a slope larger than the Mach angle. WAVDRAG calculates total drag and the wave-drag coefficient of the specified aircraft configuration. WAVDRAG is written in FORTRAN 77 for batch execution and has been implemented on a CDC CYBER 170 series computer with a central memory requirement of approximately 63K (octal) of 60 bit words. This program was developed in 1983.

Description: An approximate inverse solution is presented for the nonequilibrium flow in the inviscid shock layer about a vehicle in hypersonic flight. The method is based upon a thin-shock-layer approximation and has the advantage of being applicable to both subsonic and supersonic regions of the shock layer. The relative simplicity of the method makes it ideally suited for programming on a digital computer with a significant reduction in storage capacity and computing time required by other more exact methods. Comparison of nonequilibrium solutions for an air mixture obtained by the present method is made with solutions obtained by two other methods. Additional cases are presented for entry of spherical nose cones into representative Venusian and Matrian atmospheres. A digital computer program written in FORTRAN language is presented that permits an arbitrary gas mixture to be employed in the solution. The effects of vibration, dissociation, recombination, electronic excitation, and ionization are included in the program.

Description: Small areas of high heat transfer and pressure can occur on a vehicle surface due to the influence of an impinging shock on the local flow. A method was needed to determine peak pressure and heating of these areas. This package is a system of computer programs designed to calculate two-dimensional shock interference patterns for six types of interference flows. Results also include properties of the inviscid flow field and the inviscid-viscous interaction at the surface along with peak pressure and peak heating at the impingement point. The six types of interference flow patterns considered are: 1) Type I interference patterns, occurring when two weak shocks of opposite families, BS (bow shock) and IS (impingment shock), intersect when the flow upstream of the impingement point is supersonic, or in the case of a blunt body, takes place well below the sonic point. 2) Type II interference pattern occurs when two shocks of opposite families (bow shock and impinging shock) intersect. Both shocks are weak as in type I, but are of such strength that in order to turn the flow, a Mach reflection must exist in the center of the flow field with an embedded subsonic region occurring between the intersection points (A & B) and the accompanying shear layers. Type II interference occurs on a blunt body when the impinging shock intersects the bow shock near the sonic point. 3) Type III shock interference pattern occurs when a weak impinging shock intersects a strong detached bow shock. On a blunt body the shock intersection occurs near or above the lower sonic point. 4) Type IV interference can occur when the impinging shock intersects a strong bow shock ahead of a subsonic flow region. On a blunt body this shock intersection is located between the lower sonic point and just above the body axis. The impinging shock causes a displacement of the bow shock and the formation of a supersonic jet that is embedded in the subsonic region. A jet bow shock is produced when the jet impinges on the surface, creating a small region with high stagnation heating. 5) Type V interference involves the interaction of two weak shocks of the same family. The interaction produces a shear layer, a supersonic jet, and a transmitted impinging shock. On a blunt body the shock interaction occurs near the upper sonic point. 6) Type VI interference involves the intersection of two weak shocks of the same family, which leads to an entirely supersonic flow field. This type of interference is important because it provides a means for predicting the onset of type V. Peak-heating correlations for laminar and turbulent shock-boundary-layer interactions are included in the programs for types I, II, V, and VI interference patterns. Heating correlations for laminar and turbulent reattaching shear layers obtained from separation studies are included in the program for type III interference. This program is written in FORTRAN IV for batch execution and has been implemented on a CDC 6000 Series computer. This program was developed in 1973.

Description: A computer program has been written to obtain the wave and friction drag of configurations with bodies of revolution and fins. These inviscid flow fields are superimposed and the wave drag of the configuration is obtained by integration of the surface pressures. The friction drag is obtained from the viscous flow field of the body and a flat-plate friction analysis of the fins. The numerical solution of these flow fields, superposition, and integration to obtain total drag have been programmed for high-speed digital computation. A large portion of the input required by the program is involved with the description of the configuration geometry and the specific surface positions where pressures are to be evaluated. In addition to drag forces, an output is available whereby the pressure distributions on the body and fins can be obtained.

Description: This paper is on the control of nonlinear-nonstationary vibration of a frame-stringer structure resulting from high levels of excitaation from a nearby supersonic jet exhaust. The structure exhibits periodic, chaotic, or random behaviors when forced by high-intensity sound from a supersonic jet exhaust with shock loading superimposed on the broadband response. The time history of the pressure, showing the rotation and flapping of the shock structure in the jet column due to large-scale instabilities, indicates that the response is not only nonlinear but also nonstationary. The acoustic pressure radiated by the structure also contains shocks and the formation of harmonics with distance. Control of the structural response is achieved by actively forcing the structure with an actuator at the shock oscillation frequency whose amplitude is locked into a self-control cycle. Results show that the peak power level is reduced by a factor of 63, or 18 dB. As a result, new broadband components emerge with at least four harmonics. At accelerating and decelerating supersonic speeds, the exhaust from the jet induces higher transient loading on the nearby flexible structure due to the occurence of multiple shocks from the jet.

Description: We report the discovery of a series of infrared absorption bands between 3600 and 3100/cm (2.8-3.2 micrometers) in the spectrum of Io. Individual narrow bands are detected at 3553, 3514.5, 3438, 3423, 3411.5, and 3401/cm (2.815, 2.845, 2.909, 2.921, 2.931, and 2.940 micrometers, respectively). The positions and relative strengths of these bands, and the difference of their absolute strengths between the leading and trailing faces of Io, indicate that they are due to SO2. The band at 3438/cm (2.909 micrometers) could potentially have a contribution from an additional molecular species. The existence of these bands in the spectrum of Io indicates that a substantial fraction of the SO2 on Io must reside in transparent ices having relatively large crystal sizes. The decrease in the continuum observed at the high frequency ends of the spectra is probably due to the low frequency side of the recently detected, strong 3590/cm (2.79 micrometer) feature. This band is likely due to the combination of a moderately strong SO2 band and an additional absorption from another molecular species, perhaps H2O isolated in SO2 at low concentrations. A broad (FWHM approximately = 40-60/cm), weak band is seen near 3160/cm (3.16 micrometers) and is consistent with the presence of small quantities of H2O isolated in SO2-rich ices. There is no evidence in the spectra for the presence of H2O vapor on Io. Thus, the spectra presented here neither provide unequivocal evidence for the presence of H2O on Io nor preclude it at the low concentrations suggested by past studies.

Description: A thermoviscoplastic finite element method employing the Bodner-Partom constitutve model is used to investigate the response of simplified thermal-structural models to intense local heating. The computational method formulates the problem in rate and advances the solution in time by numerical integration. The thermoviscoplastic response of simplified structures with prescribed temperatures is investigated. With rapid rises of temperature, the nickel alloy structures display initially higher yield stresses due to strain rate effects. As temperatures approach elevated values, yield stress and stiffness degrade rapidly and pronounced plastic deformation occurs.

Description: A three-bay, space, cantilever truss is probabilistically evaluated to describe progressive buckling and truss collapse in view of the numerous uncertainties associated with the structural, material, and load variables that describe the truss. Initially, the truss is deterministically analyzed for member forces, and members in which the axial force exceeds the Euler buckling load are identified. These members are then discretized with several intermediate nodes, and a probabilistic buckling analysis is performed on the truss to obtain its probabilistic buckling loads and the respective mode shapes. Furthermore, sensitivities associated with the uncertainties in the primitive variables are investigated, margin of safety values for the truss are determined, and truss end node displacements are noted. These steps are repeated by sequentially removing buckled members until onset of truss collapse is reached. Results show that this procedure yields an optimum truss configuration for a given loading and for a specified reliability.

Description: The magnetic field configuration-states of the magnetotails of the planets Uranus and Neptune are compared. Earth's case is also briefly treated, as well as some related aspects of the other three magnetic planets. In Uranus' case, due to the large tilt (59 deg) of the planet's magnetic dipole with respect to its spin axis and the unusual obliquity of that axis, the angle of attack (alpha) of the solar wind with respect to dipole alignment goes through all possible angles, 0 deg to 180 deg, yielding a very broad spectrum of configuration-states of its tail. Cases are discussed where the planetary magnetic dipole is either aligned with the Sun-planet-line ('pole-on' state) or perpendicular to it and some intermediate states, for both Uranus and Neptune. Only Uranus experiences the pole-on state, which next occurs in November 1999 (+/- 2 months); last year (1993.2) it had the first 'perpendicular' state since Voyager encounter which resembles Earth's case. Neptune never has a pole-on configuration, but it gets as close as alpha = 14 deg from it; the next occurrence is early in 2003. At Voyager encounter Neptune's magnetotail apparently rapidly migrated through a broad spectrum of field structures with near extreme states resembling an Earth-like case on the one hand and a cylindrically symmetric one on the other. Magnetopause 'openness' should dramatically change in terms of the rapidly changing angle of attack throughout a planetary day for these two planets, and this has important implications for their magnetotails. Any future manetospheric mission plans for Uranus or Neptune should take in to consideration the allowed range of values for alpha for the epoch of interest; this is especially of concern for Uranus which has a pole-on state, and all possible alphas, around the middle of 2014, 20 years from now.

Description: We studied five new Antartic achondrites, MacAlpine Hills (MAC) 88177, Yamato (Y)74357, Y75274, Y791491 and Elephant Moraine (EET)84302 by mineralogical techniques to gain a better understanding of the mineral assemblages of a group of meteorites with an affinity to Lodran (stony-iron meteorite) and their formation processes. This group is being called lodranites. These meteorites contain major coarse-grained orthopyroxene (Opx) and olivine as in Lodran and variable amounts of FeNi metal and troilite etc. MAC88177 has more augite and less FeNi than Lodran; Y74357 has more olivine and contains minor augite; Y791491 contains in addition plagioclase. EET84302 has an Acapulco-like chondritic mineral assembladge and is enriched in FeNi metal and plagioclase, but one part is enriched in Opx and chromite. The EET84302 and MAC88177 Opx crystals have dusty cores as in Acapulco. EET84302 and Y75274 are more Mg-rich than other members of the lodranite group, and Y74357 is intermediate. Since these meteorites all have coarse-grained textures, similar major mineral assemblages, variable amounts of augite, plagioclase, FeNi metal, chromite and olivine, we suggest that they are related and are linked to a parent body with modified chondritic compositions. The variability of the abundances of these minerals are in line with a proposed model of the surface mineral assemblages of the S asteroids. The mineral assemblages can best be explained by differing degrees of loss or movements of lower temperature partial melts and recrystallization, and reduction. A portion of EET84302 rich in metal and plagioclase may represent a type of component removed from the lodranite group meteorites. Y791058 and Caddo County, which were studied for comparison, are plagioclase-rich silicate inclusions in IAB iron meteorites and may have been derived by similar process but in a different body.

Description: We have reduced high-titanium lunar mare soil and iron-rich lunar volcanic glass with hydrogen at temperatures of 900-1100 C. Ilmenite is the most reactive phase in the soil, exhibiting rapid and complete reduction at all temperatures. Ferrous iron in the glass is extensively reduced concurrent with partial crystallization. In both samples pyroxene and olivine undergo partial reduction along with chemical and mineralogical modifications. High-temperature reduction provides insight into the optical and chemical effects of lunar soil maturation, and places constraints on models of that process. Mare soil and volcanic glass are attractive feedstocks for lunar oxygen production, with achievable yields of 2-5 wt%.

Description: Recently Brown et al. (1991) showed that Triton's internal heat source could amount to 5-20% of the absorbed insolation on Triton, thus significantly affecting volatile transport and atmospheric pressure. Subsequently, Kirk and Brown (1991a) used simple analytical models of the effect of internal heat on the distribution of volatiles on Triton's surface, confirming the speculation of Brown et al. that Triton's internal heat flow could strongly couple to the surface volatile distribution. To further explore this idea, we present numerical models of the permanent distribution of nitrogen ice on Triton that include the effects of sunlight, the two-dimensional distribution of internal heat flow, the coupling of internal heat flow to the surface distribution of nitrogen ice, and the finite viscosity of nitrogen ice. From these models we conclude that: (1) The strong vertical thermal gradient induced in Triton's polar caps by internal heat-flow facilitates viscous spreading to lower latitudes, thus opposing the poleward transport of volatiles by sunlight, and, for plausible viscosities and nitrogen inventories, producing permanent caps of considerable latitudinal extent; (2) It is probable that there is a strong coupling between the surface distribution of nitrogen ice on Triton and internal heat flow; (3) Asymmetries in the spatial distribution of Triton's heat flow, possibly driven by large-scale, volcanic activity or convection in Triton's interior, can result in permanent polar caps of unequal latitudinal extent, including the case of only one permanent polar cap; (4) Melting at the base of a permanent polar cap on Triton caused by internal heat flow can significantly enhance viscous spreading, and, as an alternative to the solid-state greenhouse mechanism proposed by Brown et al. (1990), could provide the necessary energy, fluids, and/or gases to drive Triton's geyser-like plumes; (5) The atmospheric collapse predicted to occur on Triton in the next 20 years (Spencer, 1990) may be plausibly avoided because of the large latitudinal extent expected for permanent polar caps on Triton.

Description: In early 1994, the US will once again place a spacecraft in lunar orbit. Popularly called Clementine, this spacecraft will spend about two months mapping the Moon and then will travel on to encounter the near-Earth asteroid, 1620 Geographos. A brief description of the historical development of the mission; the lunar survey; Clementine's payload, equipment, and capabilties; and the encounter with Geographos is presented.

Description: The transitional flight characteristics of a geometrically simplified Short Take-Off Vertical Landing (STOVL) aircraft configuration have been measured in the NASA Ames 7- by 10-Foot Wind Tunnel. The experiment is the first in a sequence of tests designed to provide detailed data for evaluating the capability of computational fluid dynamics methods to predict the important flow parameters for powered lift. The model consists of a 60 deg cropped delta wing platform, blended fuselage and two circular in-line jets that exit perpendicularly from the flat lower surface. The measured flows have a maximum freestream Mach number of 0.2. Model angle of attack is varied between -10 and +20 deg. The flow is ambient temperature in both jet exits and the nozzle pressure ratios are varied between 1 and 3. The data presented includes forces and moments, pressures measured at 281 surface pressure ports and the pressures of the jets. Measurements of the flow are also made in the tunnel test section upstream and downstream of the model and at the jet exits to guide boundary condition selection for the planned computations. Flow visualization and total pressure measurements in the jet plumes provide a description of the three-dimensional jet efflux flowfield.

Description: This test was originally conducted to determine the effects of several empennage and afterbody parameters on the aft-end aerodynamic characteristics of a twin-engine fighter-type configuration. Model variables were as follows: horizontal tail axial location and incidence, vertical tail axial location and configuration (twin-vs single-tail arrangements), tail booms, and nozzle power setting. Jet propulsion was simulated by exhausting high-pressure, cold-flow air from the nozzles. Following a successful test conducted on a single engine nacelle model to validate a CFD code, this model was chosen to be instrumented with pressure taps on the afterbody and nozzles and used as a follow-on test, providing a more complex geometry for the CFD code validation. A more limited test matrix was run to collect the pressure data, employing only the twin-tail configuration and varying only the horizontal and vertical tail locations. Mach number was varied from 0.6 to 1.2. Nozzle pressure ratio was varied from jet-off to 8. Angle-of-attack varied from 0 to 8 deg.

Description: This test was initiated to provide validation data on low aspect ratio wings at transonic speeds. The test was conducted so that the data obtained would be useful in the validation of codes, and all boundary condition data required would be measured as part of the test. During the conduct of the test, the measured quantities were checked for repeatability, and when the data would not repeat, the cause was tracked down and either eliminated or included in the measurement uncertainty. The accuracy of the data was in the end limited by wall imperfections of the wind tunnel in which the test was run.

Description: The data presented in this contribution were obtained in the NASA Langley 16-Foot Transonic Tunnel. Multiple test entries were completed and the results have been completely reported in five NASA reports. The objective of the initial investigation was to determine the effect of empennage (tail) interference on the drag characteristics of an axisymmetric model with a single engine fighter aft-end with convergent divergent nozzles. Two nozzle power settings, dry and maximum afterburning, were investigated. Several empennage arrangements and afterbody modifications were investigated during the initial investigation. Subsequent investigations were used to determine the effects of other model variables including tail incidence, tail span, and nozzle shape. For the final investigation, extensive surface pressure instrumentation was added to the model in order to develop and understanding of the flow interactions associated with afterbody/empennage integration and also to provide data for code validation. Extensive computational analysis has been conducted on the staggered empennage configuration at a Mach number of 0.6 utilizing a three-dimensional Navier Stokes code. Most of the investigations were conducted at Mach numbers from 0.60 to 1.20 and at ratios of jet total pressure to free stream static pressure (nozzle pressure ratio) from 0.1 (jet off) to 8.0. Some angle of attack variation was obtained at jet off conditions.

Description: The purpose of this investigation is to provide a comprehensive data base for the validation of numerical simulations. The initial results of the study (single angle of attack) were presented in ref. 1, where the effects of various parameters and the adequacies of selected turbulence models were discussed. The objective of the present paper is to provide a tabulation of the experimental data. The data were obtained in the two-dimensional, transonic flowfield surrounding a supercritical airfoil. A variety of flows were studied in which the boundary layer at the trailing edge of the model was either attached or separated. Unsteady flows were avoided by controlling the Mach number and angle of attack. Surface pressures were measured on both the model and wind tunnel walls, and the flowfield surrounding the model was documented using a laser Doppler velocimeter (LDV). Although wall interference could not be completely eliminated, its effect was minimized by employing the following techniques. Sidewall boundary layers were reduced by aspiration, and upper and lower walls were contoured to accommodate the flow around the model and the boundary-layer growth on the tunnel walls. A data base with minimal interference from a tunnel with solid walls provides an ideal basis for evaluating the development of codes for the transonic speed range because the codes can include the wall boundary conditions more precisely than interference corrections can be made to the data sets.

Description: The computation of unsteady shock waves, which contribute significantly to noise generation in supersonic jet flows, is investigated. This paper focuses on the difficulties of computing slowly moving shock waves. Numerical error is found to manifest itself principally as a spurious entropy wave. Calculations presented are performed using a third order essentially nonoscillatory scheme. The effect of stencil biasing parameters and of two versions of numerical flux formulas on the magnitude of spurious entropy are investigated. The level of numerical error introduced in the calculation in quantified as a function of shock pressure ratio, shock speed, Courant number, and mesh density. The spurious entropy relative to the entropy jump across a static shock decreases with increasing shock strength and shock velocity relative to the grid, but is insensitive to Courant number. The structure of the spurious entropy wave is affected by the choice of flux formulas and algorithm biasing parameters. The effect of the spurious numerical waves on the calculation of sound amplification by a shock wave is investigated. For this class of problem, the acoustic pressure waves are relatively unaffected by the spurious numerical phenomena.

Description: Jovis Tholus, Ulysses Patera, and Biblis Patera, three small volcanoes in the Tharsis area of Mars, provide important insight into the evolution of volcanism on Mars. All three are interpreted to be shield volcanoes, indicating that shield volcansim was present from the outset in Tharsis. Jovis Tholus is the least complex with simple repeated outpouring of lavas and caldera-forming events. Ulysses Patera is dominated by a giant caldera within which is a line of cinder cones or domes suggesting terminal stages of volcanism in which the magma had either significant volatiles or increased viscosity. Biblis Patera is characterized by nested calderas which have expanded by block faulting of the flank; it also exhibits lava flows erupted onto the flanks from events along concentric fractures. These shields are different from the younger Tharsis Montes shields only in terms of the volume of erupted material. The limited shield volume suggests that the magma source which fed the shields was rapidly depleted. The relatively large size ofthe calderas probably results from relatively large, shallow magma bodies rather than significant burial of the flanks by younger lavas. Eruption rates consistent with typical terrestrial basaltic eruptiuon rates suggest that these volcanoes were probably built over time spans of 10(exp 4) to 10(exp 5) years. Stratigraphic ages range from Early to Upper Hesperian; absolute ages range from 1.9 to 3.4 Ga.