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Search indexed NASA NTRS and DOE OSTI research on propulsion, heat transfer, battery materials and energy systems. Follow report and document links to the original sources.

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All sourcesNASA NTRSDOE OSTI
NASA NTRS1992

A domain-specific design architecture for composite material design and aircraft part redesign

Advanced composites have been targeted as a 'leapfrog' technology that would provide a unique global competitive position for U.S. industry. Composites are unique in the requirements for an integrated approach to designing, manufacturing, and marketing of products developed utilizing the new materials of construction. Numerous studies extending across the entire economic spectrum of the United States from aerospace to military to durable goods have identified composites as a 'key' technology. In general there have been two approaches to composite construction: build models of a given composite materials, then determine characteristics of the material via numerical simulation and empirical testing; and experience-directed construction of fabrication plans for building composites with given properties. The first route sets a goal to capture basic understanding of a device (the composite) by use of a rigorous mathematical model; the second attempts to capture the expertise about the process of fabricating a composite (to date) at a surface level typically expressed in a rule based system. From an AI perspective, these two research lines are attacking distinctly different problems, and both tracks have current limitations. The mathematical modeling approach has yielded a wealth of data but a large number of simplifying assumptions are needed to make numerical simulation tractable. Likewise, although surface level expertise about how to build a particular composite may yield important results, recent trends in the KBS area are towards augmenting surface level problem solving with deeper level knowledge. Many of the relative advantages of composites, e.g., the strength:weight ratio, is most prominent when the entire component is designed as a unitary piece. The bottleneck in undertaking such unitary design lies in the difficulty of the re-design task. Designing the fabrication protocols for a complex-shaped, thick section composite are currently very difficult. It is in fact this difficulty that our research will address.

Punch, W. F., III↗
NASA NTRS2004

Design for a Unitary Graphite Composite Instrument Boom

This paper describes development of a Unitary graphite composite instrument boom that incorporates carpenter-tape like hinges for stowage. While light and stiff, graphite composite is not ordinarily thought of as a flexible material. This design has taken advantage of the stiffness of the composite in tubular geometry, yet leveraged its thin- section behavior to place flexibility at the required locations. Key is the proprietary layup, which results in a tough yet flexible hinge capable of rotating over 90 degrees in each direction. When the boom deploys, there is enough torque to overcome parasitic resistance from harness, etc. It will snap to the fully extended, rigid shape. The design has addressed materials issues such as out-of-plane bending, edge cracking, and interlaminar ply separation.

Alexander, Wes↗
NASA NTRS2004

Non-unitary probabilistic quantum computing

We present a method for designing quantum circuits that perform non-unitary quantum computations on n-qubit states probabilistically, and give analytic expressions for the success probability and fidelity.

quantum gate↗
NASA NTRS1992

Study of optical techniques for the Ames unitary wind tunnels. Part 1: Schlieren

Alignment procedures and conceptual designs for the rapid alignment of the Ames Unitary Wind Tunnel schlieren systems were devised. The schlieren systems can be aligned by translating the light source, the mirrors, and the knife edge equal distances. One design for rapid alignment consists of a manual pin locking scheme. The other is a motorized electronic position scheme. A study of two optical concepts which can be used with the schlieren system was made. These are the 'point diffraction interferometers' and the 'focus schlieren'. Effects of vibrations were studied.

Lee, George↗
NASA NTRS1979

Advanced vehicle separation apparatus

A method of obtaining test data from two independent models or bodies in a conventional wind tunnel is described. The system makes efficient use of wind tunnel test time with computer control performing complex coordinate transformations necessary for model positioning. The apparatus is designed to be used in any of the three Unitary Wind Tunnels at NASA-Ames Research Center. Mechanical design details and a brief description of the control system for the separation apparatus are presented.

Ospring, M. J.↗
NASA NTRS2017

Recent Advancements in the Infrared Flow Visualization System for the NASA Ames Unitary Plan Wind Tunnels

The following details recent efforts undertaken at the NASA Ames Unitary Plan wind tunnels to design and deploy an advanced, production-level infrared (IR) flow visualization data system. Highly sensitive IR cameras, coupled with in-line image processing, have enabled the visualization of wind tunnel model surface flow features as they develop in real-time. Boundary layer transition, shock impingement, junction flow, vortex dynamics, and buffet are routinely observed in both transonic and supersonic flow regimes all without the need of dedicated ramps in test section total temperature. Successful measurements have been performed on wing-body sting mounted test articles, semi-span floor mounted aircraft models, and sting mounted launch vehicle configurations. The unique requirements of imaging in production wind tunnel testing has led to advancements in the deployment of advanced IR cameras in a harsh test environment, robust data acquisition storage and workflow, real-time image processing algorithms, and evaluation of optimal surface treatments. The addition of a multi-camera IR flow visualization data system to the Ames UPWT has demonstrated itself to be a valuable analyses tool in the study of new and old aircraft/launch vehicle aerodynamics and has provided new insight for the evaluation of computational techniques.

Infrared Thermography↗
NASA NTRS2015

Institutional Schlieren: A Production-Level Wind Tunnel Test Measurement

The following details recent efforts undertaken at the NASA Ames Unitary Plan Wind Tunnel to design and deploy an advanced, institutional, production-level data system for the classical Schlieren-shadowgraph technique. Motivation for the selection of individual system components is discussed along with a software methodology that combines image acquisition and processing into a production-level wind tunnel test measurement. In general terms, a production-level measurement refers to any data system that is seamlessly integrated into the primary wind tunnel data system, and whose data products are available real-time (e.g. force and moment, pressure, temperature data). The advantage of integrating a measurement in such a manner is an immediate increase in data product efficiency, productivity, reliability, and quality. Coupled with these benefits and leveraging recent advancements in high-speed imaging and image processing, automated, synchronized, time-resolved Schlieren-shadowgraph imaging for dynamic flow phenomena is now a reality. This makes possible the synthesis of dynamic off-body imaging with unsteady on-body measurements to produce a uniquely descriptive data product invaluable to the modern researcher.

Shadowgraph↗
NASA NTRS2011

Supersonic Retro-Propulsion Experimental Design for Computational Fluid Dynamics Model Validation

The development of supersonic retro-propulsion, an enabling technology for heavy payload exploration missions to Mars, is the primary focus for the present paper. A new experimental model, intended to provide computational fluid dynamics model validation data, was recently designed for the Langley Research Center Unitary Plan Wind Tunnel Test Section 2. Pre-test computations were instrumental for sizing and refining the model, over the Mach number range of 2.4 to 4.6, such that tunnel blockage and internal flow separation issues would be minimized. A 5-in diameter 70-deg sphere-cone forebody, which accommodates up to four 4:1 area ratio nozzles, followed by a 10-in long cylindrical aftbody was developed for this study based on the computational results. The model was designed to allow for a large number of surface pressure measurements on the forebody and aftbody. Supplemental data included high-speed Schlieren video and internal pressures and temperatures. The run matrix was developed to allow for the quantification of various sources of experimental uncertainty, such as random errors due to run-to-run variations and bias errors due to flow field or model misalignments. Some preliminary results and observations from the test are presented, although detailed analyses of the data and uncertainties are still on going.

Berry, Scott A.↗
NASA NTRS1999

Cone-Probe Rake Design and Calibration for Supersonic Wind Tunnel Models

A series of experimental investigations were conducted at the NASA Langley Unitary Plan Wind Tunnel (UPWT) to calibrate cone-probe rakes designed to measure the flow field on 1-2% scale, high-speed wind tunnel models from Mach 2.15 to 2.4. The rakes were developed from a previous design that exhibited unfavorable measurement characteristics caused by a high probe spatial density and flow blockage from the rake body. Calibration parameters included Mach number, total pressure recovery, and flow angularity. Reference conditions were determined from a localized UPWT test section flow survey using a 10deg supersonic wedge probe. Test section Mach number and total pressure were determined using a novel iterative technique that accounted for boundary layer effects on the wedge surface. Cone-probe measurements were correlated to the surveyed flow conditions using analytical functions and recursive algorithms that resolved Mach number, pressure recovery, and flow angle to within +/-0.01, +/-1% and +/-0.1deg , respectively, for angles of attack and sideslip between +/-8deg. Uncertainty estimates indicated the overall cone-probe calibration accuracy was strongly influenced by the propagation of measurement error into the calculated results.

Won, Mark J.↗
NASA NTRS1999

Experimental Investigation of a Point Design Optimized Arrow Wing HSCT Configuration

The M2.4-7A Arrow Wing HSCT configuration was optimized for straight and level cruise at a Mach number of 2.4 and a lift coefficient of 0.10. A quasi-Newton optimization scheme maximized the lift-to-drag ratio (by minimizing drag-to-lift) using Euler solutions from FL067 to estimate the lift and drag forces. A 1.675% wind-tunnel model of the Opt5 HSCT configuration was built to validate the design methodology. Experimental data gathered at the NASA Langley Unitary Plan Wind Tunnel (UPWT) section #2 facility verified CFL3D Euler and Navier-Stokes predictions of the Opt5 performance at the design point. In turn, CFL3D confirmed the improvement in the lift-to-drag ratio obtained during the optimization, thus validating the design procedure. A data base at off-design conditions was obtained during three wind-tunnel tests. The entry into NASA Langley UPWT section #2 obtained data at a free stream Mach number, M(sub infinity), of 2.55 as well as the design Mach number, M(sub infinity)=2.4. Data from a Mach number range of 1.8 to 2.4 was taken at UPWT section #1. Transonic and low supersonic Mach numbers, M(sub infinity)=0.6 to 1.2, was gathered at the NASA Langley 16 ft. Transonic Wind Tunnel (TWT). In addition to good agreement between CFD and experimental data, highlights from the wind-tunnel tests include a trip dot study suggesting a linear relationship between trip dot drag and Mach number, an aeroelastic study that measured the outboard wing deflection and twist, and a flap scheduling study that identifies the possibility of only one leading-edge and trailing-edge flap setting for transonic cruise and another for low supersonic acceleration.

Narducci, Robert P.↗
NASA NTRS1958

Aerodynamic characteristics at mach numbers from 2.5 to 3.5 of a canard bomber configuration designed for supersonic cruise flight

Resilts have been obtained from an investigation in the Langley Unitary Plan wind tunnel at Mach numbers from 2.5 to 3.5 of a canard-type configuration designed for supersonic cruise flight. Tests extended over an angle-of-attack range from about -4 deg to 11 deg and an angle-of-sideslip range from -4 deg to 6 deg. For the present tests, the results indicate that forebody deflection was an efficient means of providing a sizable positive pitching-moment shift with little or no increase in drag. The test configuration had a trimmed lift-drag ratio of approximately 6.0 at Mach numbers near 3.0 and at a Reynolds number of 2.52 X 10(exp 6). The configuration was both longitudinally and directionally stable. The lift-drag ratios are believed to be somewhat low in as much as the models used for the present tests had large-grain size transition strips fixed to the various surfaces and these strips added wave drag. Also, the model boundary-layer diverter is oversized with respect to a full-scale configuration and therefore contributes additional drag.

Carmel, M. M.↗
NASA NTRS1958

Aerodynamic Characteristics at Mach Numbers from 2.5 to 3.5 of a Canard Bomber Configuration Designed for Supersonic Cruise Flight

Results have been obtained from an investigation in the Langley Unitary Plan wind tunnel at Mach numbers from 2.5 to 3.5 of a canard-type configuration designed for supersonic cruise flight. Tests extended over an angle-of-attack range from about -4 deg to 11 deg and an angle-of-sideslip range from -4 deg to 6 deg. For the present tests, the results indicate that forebody deflection was an efficient means of providing a sizable positive pitching-moment shift with little or no increase in drag. The test configuration had a trimmed lift-drag ratio of approximately 6.0 at Mach numbers near 3.0 and at a Reynolds number of 2.52 x 10(exp 6). The configuration was both longitudinally and directionally stable. The lift-drag ratios are believed to be somewhat low inasmuch as the models used for the present tests had large-grain-size transition strips fixed to the various surfaces and these strips added wave drag. Also, the model boundary- layer diverter is oversized with respect to a full - scale configuration and therefore contributes additional drag.

Carmel, Melvin R↗
NASA NTRS2012

Supersonic Retropropulsion Experimental Results from the NASA Ames 9- x 7-Foot Supersonic Wind Tunnel

Supersonic retropropulsion was experimentally examined in the Ames Research Center 9x7-Foot Supersonic Wind Tunnel at Mach 1.8 and 2.4. The experimental model, previously designed for and tested in the Langley Research Center Unitary Plan Wind Tunnel at Mach 2.4, 3.5 and 4.6, was a 5-in diameter 70-deg sphere-cone forebody with a 9.55-in long cylindrical aftbody. The forebody was designed to accommodate up to four 4:1 area ratio nozzles, one on the model centerline and the other three on the half radius spaced 120-deg apart. Surface pressure and flow visualization were the primary measurements, including high-speed data to investigate the dynamics of the interactions between the bow and nozzle shocks. Three blowing configurations were tested with thrust coefficients up to 10 and angles of attack up to 20-deg. Preliminary results and observations from the test are provided

Berry, Scott A.↗
NASA NTRS1990

A computational study of incipient leading-edge separation on a 65-deg delta wing at M = 1.60

A computational study on a 65-deg delta wing at a freestream Mach number of 1.60 has been conducted by obtaining conical Reynolds-averaged Navier-Stokes solutions on a parametric series of geometries which varied in leading-edge radius and/or circular-arc camber. The computational results showed that increasing leading-edge radius or camber can delay the onset of leading-edge separation on the leeside of a delta wing at a specific angle of attack. Reynolds number was varied from 1 x 10 to the 6th to 5 x 10 to the 6th for a turbulent boundary-layer and was shown to have a minor effect on the effectiveness of leading-edge radius and/or camber in delaying the onset of leading-edge separation. Both laminar and turbulent boundary-layer models were investigated at a Reynolds number of 1 x 10 to the 6th, and the predicted flow pattern was found to change from attached flow for the turbulent boundary-layer model to separated flow for the laminar boundary-layer model. Based upon these results, three wind-tunnel models have been designed to be tested in the Langley Unitary Plan Wind Tunnel.

Mcmillin, S. Naomi↗
NASA NTRS1999

Supersonic Aftbody Closure Wind-Tunnel Testing, Data Analysis, and Computational Results

This paper reports on the model, test, and results from the Langley Supersonic Aftbody Closure wind tunnel test. This project is an experimental evaluation of the 1.5% Technology Concept Aircraft (TCA) aftbody closure model (Model 23) in the Langley Unitary Plan Wind Tunnel. The baseline TCA design is the result of a multidisciplinary, multipoint optimization process and was developed using linear design and analysis methods, supplemented with Euler and Navier-Stokes numerical methods. After a thorough design review, it was decided to use an upswept blade attached to the forebody as the mounting system. Structural concerns dictated that a wingtip support system would not be feasible. Only the aftbody part of the model is metric. The metric break was chosen to be at the fuselage station where prior aft-sting supported models had been truncated. Model 23 is thus a modified version of Model 20. The wing strongback, flap parts, and nacelles from Model 20 were used, whereas new aftbodies, a common forebody, and some new tails were fabricated. In summary, significant differences in longitudinal and direction stability and control characteristics between the ABF and ABB aftbody geometries were measured. Correcting the experimental data obtained for the TCA configuration with the flared aftbody to the representative of the baseline TCA closed aftbody will result in a significant reduction in longitudinal stability, a moderate reduction in stabilizer effectiveness and directional stability, and a moderate to significant reduction in rudder effectiveness. These reductions in the stability and control effectiveness levels of the baseline TCA closed aftbody are attributed to the reduction in carry-over area.

Allen, Jerry↗
NASA NTRS1974

Effects of leading edge sweep angle and design lift coefficient on performance of a modified arrow wing at a design Mach number of 2.6

Wing models were tested in the high-speed section of the Langley Unitary Plan wind tunnel to study the effects of the leading-edge sweep angle and the design lift coefficient on aerodynamic performance and efficiency. The models had leading-edge sweep angles of 69.44 deg, 72.65 deg, and 75.96 deg which correspond to values of the design Mach-number-sweep-angle parameter (beta cotangent A) sub DES of 0.6, 0.75, and 0.9, respectively. For each sweep angle, camber surfaces having design lift coefficients of 0,0.08, and 0.12 at a design Mach number of 2.6 were generated. The wind-tunnel tests were conducted at Mach numbers of 2.3, 2.6, and 2.96 with a stagnation temperature of 338.7 K (150 F) and a Reynolds number per meter of 9.843 times 10 to the 6th power. The results of the tests showed that only a moderate sweeping of the wing leading edge aft of the Mach line along with a small-to-moderate amount of camber and twist was needed to significantly improve the zero-lift (flat camber surface) wing performance and efficiency.

Mack, R. J.↗
NASA NTRS1999

A Computational/Experimental Study of Two Optimized Supersonic Transport Designs and the Reference H Baseline

Two supersonic transport configurations designed by use of non-linear aerodynamic optimization methods are compared with a linearly designed baseline configuration. One optimized configuration, designated Ames 7-04, was designed at NASA Ames Research Center using an Euler flow solver, and the other, designated Boeing W27, was designed at Boeing using a full-potential method. The two optimized configurations and the baseline were tested in the NASA Langley Unitary Plan Supersonic Wind Tunnel to evaluate the non-linear design optimization methodologies. In addition, the experimental results are compared with computational predictions for each of the three configurations from the Enter flow solver, AIRPLANE. The computational and experimental results both indicate moderate to substantial performance gains for the optimized configurations over the baseline configuration. The computed performance changes with and without diverters and nacelles were in excellent agreement with experiment for all three models. Comparisons of the computational and experimental cruise drag increments for the optimized configurations relative to the baseline show excellent agreement for the model designed by the Euler method, but poorer comparisons were found for the configuration designed by the full-potential code.

Cliff, Susan E.↗
NASA NTRS2011

ARES I Aerodynamic Testing at the NASA Langley Unitary Plan Wind Tunnel

Small-scale force and moment and pressure models based on the outer mold lines of the Ares I design analysis cycle crew launch vehicle were tested in the NASA Langley Research Center Unitary Plan Wind Tunnel from May 2006 to September 2009. The test objectives were to establish supersonic ascent aerodynamic databases and to obtain force and moment, surface pressure, and longitudinal line-load distributions for comparison to computational predictions. Test data were obtained at low through high supersonic Mach numbers for ranges of the Reynolds number, angle of attack, and roll angle. This paper focuses on (1) the sensitivity of the supersonic aerodynamic characteristics to selected protuberances, outer mold line changes, and wind tunnel boundary layer transition techniques, (2) comparisons of experimental data to computational predictions, and (3) data reproducibility. The experimental data obtained in the Unitary Plan Wind Tunnel captured the effects of evolutionary changes to the Ares I crew launch vehicle, exhibited good agreement with predictions, and displayed satisfactory within-test and tunnel-to-tunnel data reproducibility.

Erickson, Gary E.↗