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Advanced cell culture techniques such as 3D bio-printing and hydrogel-based cell embedding techniques harbor many new and exciting opportunities to study cells in environments that closely recapitulate in-vivo conditions. Researchers often study these environments using fluorescence microscopy to visualize the protein association with objects such as cells within the 3D environment, yet quantification of concentration profiles in the microenvironment has remained elusive. Here, we present a method to continuously measure the time-dependent concentration gradient of various biomarkers within a 3D cell culture assay using bead-based immunoassays to sequester and concentrate the fluorescence intensity of these tagged proteins. This assay allows for near real-time in situ biomarker detection and enables spatiotemporal quantification of biomarker concentration. Snapshots of concentration profiles can be taken, or time series analysis can be performed enabling time-varying biomarker production estimation. Example assays utilize an osteosacroma tumoroid as a case study for a quantitative single-plexed gel encapsulated assay, and a qualitative multi-plexed 3D bioprinted assay. In both cases, a time-varying cytokine concentration gradient is measured. An estimation for the production rate of the IL-8 cytokine per second per osteosarcoma cell results from fitting an analytical function for continuous point source diffusion to the measured concentration gradient and reveals that each cell produces approximately two IL-8 cytokines per second. Proper calibration and use of this assay is exhaustively explored for the case of diffusion-limited Langmuir kinetics of a spherical adsorber., Comment: 8 pages text with 3 figures, 3 pages appendix, 2 pages references, manuscript submitted to in vitro models

The National Aeronautics and Space Administration (NASA) Data Assimilation Office (DAO) at the Goddard Space Flight Center (GSFC) has developed the GEOS DAS, a data assimilation system that provides production support for NASA missions and will support NASA's Earth Observing System (EOS) in the coming years. The DAO's support of the EOS project along with the requirement of producing long-term reanalysis datasets with an unvarying system levy a large I/O burden on the future system. The DAO has been involved in prototyping parallel implementations of the GEOS DAS for a number of years and is now converting the production version from shared-memory parallelism to distributed-memory parallelism using the portable Message-Passing Interface (MPI). If the MPI-based GEOS DAS is to meet these production requirements, we must make I/O from the parallel system efficient. We have designed a scheme that allows efficient I/O processing while retaining portability, reducing the need for post-processing, and producing data formats that are required by our users, both internal and external. The first phase of the GEOS DAS Parallel I/O System (GPIOS) will expand upon the common method of gathering global data to a Single PE for output. Instead of using a PE also tasked with primary computation, a number of PEs will be dedicated to I/O and its related tasks. This allows the data transformations and formatting required prior to output to take place asynchronously with respect to the GEOS DAS assimilation cycle, improving performance and generating output data sets in a format convenient for our users. I/O PEs can be added as needed to handle larger data volumes or to meet user file specifications. We will show I/O performance results from a prototype MPI GCM integrated with GPIOS. Phase two of GPIOS development will examine ways of integrating new software technologies to further improve performance and build scalability into the system. The maturing of MPI-IO implementations and other supporting libraries such as parallel HDF should provide performance gains while retaining portability.

The National Aeronautics and Space Administration (NASA) Data Assimilation Office (DAO) at the Goddard Space Flight Center is moving its data assimilation system to massively parallel computing platforms. This parallel implementation of GEOS DAS will be used in the DAO's normal activities, which include reanalysis of data, and operational support for flight missions. Key components of GEOS DAS, including the gridpoint-based general circulation model and a data analysis system, are currently being parallelized. The parallelization of GEOS DAS is also one of the HPCC Grand Challenge Projects. The GEOS-DAS software employs several distinct grids. Some examples are: an observation grid- an unstructured grid of points at which observed or measured physical quantities from instruments or satellites are associated- a highly-structured latitude-longitude grid of points spanning the earth at given latitude-longitude coordinates at which prognostic quantities are determined, and a computational lat-lon grid in which the pole has been moved to a different location to avoid computational instabilities. Each of these grids has a different structure and number of constituent points. In spite of that, there are numerous interactions between the grids, e.g., values on one grid must be interpolated to another, or, in other cases, grids need to be redistributed on the underlying parallel platform. The DAO has designed a parallel integrated library for grid manipulations (PILGRIM) to support the needed grid interactions with maximum efficiency. It offers a flexible interface to generate new grids, define transformations between grids and apply them. Basic communication is currently MPI, however the interfaces defined here could conceivably be implemented with other message-passing libraries, e.g., Cray SHMEM, or with shared-memory constructs. The library is written in Fortran 90. First performance results indicate that even difficult problems, such as above-mentioned pole rotation- a sparse interpolation with little data locality between the physical lat-lon grid and a pole rotated computational grid- can be solved efficiently and at the GFlop/s rates needed to solve tomorrow's high resolution earth science models. In the subsequent presentation we will discuss the design and implementation of PILGRIM as well as a number of the problems it is required to solve. Some conclusions will be drawn about the potential performance of the overall earth science models on the supercomputer platforms foreseen for these problems.

The National Aeronautics and Space Administration (NASA) Data Assimilation Office (DAO) at the Goddard Space Flight Center (GSFC) has developed the GEOS DAS, a data assimilation system that provides production support for NASA missions and will support NASA's Earth Observing System (EOS) in the coming years. The GEOS DAS will be used to provide background fields of meteorological quantities to EOS satellite instrument teams for use in their data algorithms as well as providing assimilated data sets for climate studies on decadal time scales. The DAO has been involved in prototyping parallel implementations of the GEOS DAS for a number of years and is now embarking on an effort to convert the production version from shared-memory parallelism to distributed-memory parallelism using the portable Message-Passing Interface (MPI). The GEOS DAS consists of two main components, an atmospheric General Circulation Model (GCM) and a Physical-space Statistical Analysis System (PSAS). The GCM operates on data that are stored on a regular grid while PSAS works with observational data that are scattered irregularly throughout the atmosphere. As a result, the two components have different data decompositions. The GCM is decomposed horizontally as a checkerboard with all vertical levels of each box existing on the same processing element(PE). The dynamical core of the GCM can also operate on a rotated grid, which requires communication-intensive grid transformations during GCM integration. PSAS groups observations on PEs in a more irregular and dynamic fashion.

Research has been conducted on silicon nitride pin-on-disk sliding contacts at temperatures of up to 520 C, and four-ball rolling contacts with silicon nitride balls and 52100 steel or silicon nitride races at 590 C. These tests were conducted in a variety of gaseous environments in order to determine the effects of simulated engine exhaust gas on the carbonaceous gas decomposition lubrication scheme. In rolling tests with steel races and exhaust gas the wear track depth was roughly half that of tests run in nitrogen gas alone. The deposition of lubricous microcrystalline graphitic carbon on the rolling surfaces, generated from the carbon monoxide within the exhaust gas mixture, was verified by microfocused Raman spectroscopy. Ten-fold reductions in rolling wear could be achieved by the exhaust gas atmosphere in cases where water vapor was removed or not present. The exhaust gas mixture alone was not found to provide any lubricating effect on silicon nitride sliding contacts, where the rate of wear greatly exceeds the rate of carbon deposition. Directed admixture of acetylene (as low as 5% of the exhaust gas flow rates), has provided reductions in both wear volume and coefficient of friction by factors of 60X and 20X respectively for sliding contacts during the initial 80 m of sliding distance. Exhaust gas atmosphere with the acetylene admixture provided 65OX reductions in steady state wear rate compared to that measured for sliding contacts in dry N2. Such acetylene admixture also augments the ability of the exhaust gas atmosphere to lubricate high-temperature rolling contacts, with up to 25-fold reductions in wear track depth compared to those measured in the presence of N2 alone. In addition to providing some lubricating benefit itself, an important potential role of the exhaust gas from rich mixtures would be to shield bearings from 02. Such shielding enables surface deposition of lubricous pyrolytic carbon from the acetylene admixture, instead of combustion, rendering feasible the continuously replenished solid lubrication of high-temperature bearing surfaces.

A development status evaluation is presented for the aerodynamics of missile configurations with noncircular cross-sections and bank-to-turn maneuvering systems, giving attention to cases with elliptical and square cross-sections, as well as bodies with variable cross-sections. The assessment of bank-to-turn missile performance notes inherent stability/control problems. A summary and index are provided for aerodynamic data on monoplanar configurations, including those which incorporate airbreathing propulsion systems.

An evaluation is made of prospective missile applications for noncircular cross section bodies, and of recent developments in bank-to-turn missile configuration aerodynamics. The discussion encompasses cross-flow analysis techniques, as well as study results obtained for bodies with elliptical and square cross sections and with variable cross sections. Attention is given to both the performance advantages and the stability and control problems of bank-to-turn missile configurations; the aerodynamic data presented for monoplanar configurations extend to those incorporating airbreathing propulsion systems.

The shear stress fringe value of orthotropic birefringent model materials is presently used in an asymmetric four-point bend test that employs a notched specimen and is a modification of the Iosipescu test (1967). While in the case of a unidirectionally reinforced glass-polyester model material, shear stress fringe values obtained from 90- and 120-deg (sharp and radiused) notches are reasonably close to the values obtained for an off-axis tensile specimen, no conclusions can be drawn on the influence of the notch parameters, due to the peculiarities of the photoelastic response of the inhomogeneous orthotropic model material. The failure modes nevertheless indicate that a notch radius, and the 120-deg notch angle, reduce the stress concentration. Comparisons are made with finite element results.

Previously cited in issue 12, p. 1851, Accession no. A82-27106

This paper presents a survey of engineering computational methods and experimental programs used for estimating the aerodynamic characteristics of missile configurations. Emphasis is placed on those methods which are suitable for preliminary design of conventional and advanced concepts. An analysis of the technical approaches of the various methods is made in order to assess their suitability to estimate longitudinal and/or lateral-directional characteristics for different classes of missile configurations. Some comparisons between the predicted characteristics and experimental data are presented. These comparisons are made for a large variation in flow conditions and model attitude parameters. The paper also presents known experimental research programs developed for the specific purpose of validating analytical methods and extending the capability of data-base programs.

This paper presents an overview of the role NASA has had and continues to pursue in providing missile aerodynamic technology. In the past, NASA has provided considerable support to the missile industry and the military. The support has generally taken the form of theoretical aerodynamic analyses, experimental studies to provide solutions for specific problems, and the documentation of existing foreign missile systems and domestic missiles. In 1975, NASA shifted its missile-related efforts in aerodynamics from this largely service role to one of conducting more basic research. The areas of research include: innovative methods for roll control of cruciform missiles, airbreathing missiles with maneuver requirements, and an advanced generation of monoplanar missiles for efficient supersonic carriage and delivery.

Two different classes of missile aeroprediction programs have been recently developed. The first class of programs provides rapid engineering predictions and includes MISSILE1 and MISSILE2 applicable to missile configurations with axisymmetric bodies. The second class of programs consists of the DEMON series, including a simplified version NSWCDM, designed to calculate detailed loadings acting on supersonic missiles which may have non-circular body cross sections. Both classes account for high angles of attack and track vortices from canard or wing section to the tail section. Extensive comparisons with experimental data are presented including nonlinear effects of canard control.

An experimental investigation was conducted to provide a systematic set of lateral-directional stability data for a simplified wing-body model with a series of vertical-tail arrangements. The study was made at Mach numbers from 1.60 to 2.86 at nominal angles of attack from -8 to 12 deg and Reynolds number of 8.2 million per meter. Comparisons at zero angle of attack were made with three existing theoretical methods (MISLIFT - a second-order shock expansion and panel method; APAS - a slender body and first order panel method; and PAN AIR - a higher order panel method) and comparisons at angle of attack were made with PAN AIR. The results show that PAN AIR generally provides accurate estimates of these characteristics at moderate angles of attack for complete configurations with either single or twin vertical tails. APAS provides estimates for complete configurations at zero angle of attack. However, MISLIFT only provides estimates for the simplest body-vertical-tail configurations at zero angle of attack.

For abstract, see A81-37375.

This paper presents selected results of a panel loads study conducted as part of the final phase of an extensive investigation of an air-breathing missile concept employing wrap-around aerodynamic surfaces. Typical results for M = 2.36 are presented for the fin load results, plus a brief review of basic results of the previously reported tests. Vapor screen results are also discussed. The present results indicate that the fin load characteristics are nearly identical for planar and curved fins having the same projected planform and would permit the use of planar-surface predictions for supersonic speeds in the preliminary design stages of missiles employing wrap-around curved fins.

Some criteria for external carriage of missiles for fighter aircraft intended for aerial combat missions and for fighter-interceptor missions are considered. The mission requirements discussed include the short-range fighter-interceptor, the short-range interceptor, the medium-range interceptor, and the long-range interceptor. Missiles types considered to be compatible with the various point mission designs include the short-range missile, the medium-range missile, and the long-range missile. From the study, it appears that point mission design aircraft can be arranged in such a way that the required external-store arrangement will not impair the stability of the aircraft. An extensive reference list of NASA external store research is included.

The results of calculations necessary to extrapolate performance data on an XB-70-1 wind tunnel model to full scale at Mach numbers from 0.76 to 2.53 are presented. The extrapolation was part of a joint program to evaluate performance prediction techniques for large flexible supersonic airplanes similar to a supersonic transport. The extrapolation procedure included: interpolation of the wind tunnel data at the specific conditions of the flight test points; determination of the drag increments to be applied to the wind tunnel data, such as spillage drag, boundary layer trip drag, and skin friction increments; and estimates of the drag items not represented on the wind tunnel model, such as bypass doors, roughness, protuberances, and leakage drag. In addition, estimates of the effects of flexibility of the airplane were determined.

Three missile body shapes tested at Mach numbers of 1.50, 2.16, and 2.86 with angles of attack up to 30 degrees are described. The flow characteristics for each body shape are examined. The measured aerodynamic forces and moments are presented. The use of flow visualization techniques are described and the results such as vortex effects are discussed.

The results of pressure distribution tests conducted in a wind tunnel are presented without analysis. The data were obtained for trapezoidal aft tail control surfaces on a wingless missile model at Mach numbers of 1.60, 2.36, and 3.70 for angles of attack from -4 to 20 deg model roll angles from 0 to 90 deg and tail deflections of 0 and -15 deg. The test Reynolds number used was 6.6 million per meter.

Pressure coefficients were obtained in the Langley Unitary Plan wind tunnel for a wingless missile with a clipped delta tail. The angle of attack was varied from -4 deg to 20 deg, model roll angle was varied from 0 deg to 90 deg in 22.5 deg increments, and tail deflections were 0 deg to - 15 deg. The pressures were measured on two adjacent tail surfaces using 91 pressure orifices per tail surface. Results are presented in plotted and tabular form.

The results of pressure distribution tests are presented without analysis. The test Reynolds number used was 6.6. x 10 to the 6th power per meter.

An investigation was conducted of a monoplanar maneuverable missile concept having a nose forebody with a circular cross section and a centerbody and afterbody with elliptical cross sections. The tests involved several component changes and were conducted in the low Mach number test section of the Langley Unitary Plan wind tunnel at Mach numbers of 1.60, 2.16, and 2.86, at angles of attack ranging from -4 degs to 28 degs and at sideslip angles ranging from -4 degs to 8 degs. The most significant result was that at the highest Mach number (2.86), the configuration with the infrared nose produced nearly twice the axial force as the same configuration with the radar nose. The cranked wing had a destabilizing effect on the longitudinal stability and had no effect on the lateral-directional stability. The nose strakes had no effect longitudinally and were detrimental to the lateral-directional stability.

An experimental investigation has been made at supersonic Mach numbers to determine the feasibility of using a ram-air-spoiler roll control device on a typical canard control missile configuration. As a basis for roll control comparisons, conventional aileron controls on the tail fins were also tested. Results are presented which indicate that the addition of nacelles on the missile tail fins resulted in satisfactory roll control effectiveness and only small changes in basic missile stability. The ram-air-spoiler roll control effectiveness is relatively constant over the range of vehicle attitudes and Mach numbers investigated.

Effects of wing planform modifications on the longitudinal aerodynamic characteristics of a fixed span, maneuverable cruciform missile configuration were determined. A basic delta planform and two alternate trapezoidal planforms having progressively increasing tip chords were included. Data were obtained for angles of attack up to approximately -32 deg, model roll angles of 0 deg to 45 deg, and tail control deflections of 0 deg and -20 deg. The experimental drag due to lift was compared with linear values.

Delta wing point-design fighters with two pylon mounted missiles and aft tail controls (similar to several Soviet designs) have been investigated for a Mach number range from about 0.6 to 2.0. Whereas minimum drag penalties that are expected with the addition of external stores do occur, the effects at higher lifts, corresponding to maneuvering flight, are less severe and often favorable. The drag-due-to-lift factor is less with stores on although the lift curve slope is unaffected. The longitudinal stability level is reduced by the addition of stores while the pitch control effectiveness is unchanged. The directional stability was generally reduced at subsonic speeds and increased at supersonic speeds by the addition of stores but sufficiently high stability levels are obtainable that are compatible with the longitudinal maneuvering limits. Some examples of the potential maneuvering capability in terms of normal acceleration and turn radius are included.

An investigation was made to determine the aerodynamic characteristics of an airbreathing missile configuration having a forward located inlet. Control was provided by cruciform tail surfaces. Aerodynamic data were obtained for the body-tail configuration alone and with planar or cruciform wings. At Mach numbers from 2.86 to 4.63, the model had internal flow. At Mach numbers from 1.70 to 2.86, the internal duct was closed, and an inlet fairing was installed to simulate boost conditions.

Theoretical method for calculating aerodynamic loading on wing-body combinations

Experimental surface pressure distributions on spherically blunted cone at Mach numbers from 1.50 to 4.63

Computational method for surface pressures on blunt bodies of revolution at small angles of attack in supersonic flow

Theory and operation of nuclear magnetometer

Program was written to include integration of surface pressure in order to obtain axial-force, normal-force, and pitching-moment coefficients. Program was written in CDC FORTRAN for the CDC-6600 computer system.

The effect of nose shape and tail length on the static stability of a fin-stabilized projectile has been investigated in the Langley Unitary Plan with tunnel at angles of attack to about 12 deg for a Mach number range from 1.5 to 2.5. The tests were made at a constant Reynolds number of 6.56 x 1,000,000 per meter. The results of the investigation showed that nose shape had no effect on the static stability. Increasing the tail length resulted in a progressively stabilizing tendency. However, only the 1.5-caliber-tail-length configuration was stable over the test angle-of-attack range at Mach number 1.5. This configuration was marginally stable or unstable at the higher Mach numbers, and the shorter configurations were unstable at all Mach numbers for either part of or the entire test angle-of-attack range.

Method, developed and coded for digital computation, predicts aerodynamic loading on configurations for which linear theory assumptions are violated. Program is written in FORTRAN 4 for use on CDC-6000 series computers.