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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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At least 109 records · Page 6

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NASA NTRS1986

An experimental evaluation of S-duct inlet-diffuser configurations for turboprop offset gearbox applications

A test program, utilizing a large scale model, was run in the NASA Lewis Research Center 10- by 10-ft wind tunnel to examine the influence on performance of design parameters of turboprop S-duct inlet/diffuser systems. The parametric test program investigated inlet lip thickness, inlet/diffuser cross-sectional geometry, throat design Mach number, and shaft fairing shape. The test program was run at angles of attack to 15 deg and tunnel Mach numbers to 0.35. Results of the program indicate that current design techniques can be used to design inlet/diffuser systems with acceptable total pressure recovery, but several of the design parameters, notably lip thickness (contraction ratio) and shaft fairing cross section, must be optimized to prevent excessive distortion at the compressor face.

Mcdill, Paul L.↗

NASA NTRS1996

Navier-Stokes Simulation of the Canard-Wing-Body Longitudinal Dynamic Stability Characteristics

Many modern aircraft are canard-configured for aircraft control and improved aerodynamic performance. Canards can often enhance aircraft cruise performance, maneuverability and agility. For close-coupled canard configurations, the aerodynamic interaction between the canard and wing significantly changes the flow characteristics of the wing. In unsteady flow, such changes in the flow structure and performance of wings can be quite pronounced. Accurate modeling of the unsteady aerodynamics is essential for potential CFD design and analysis of such configurations. A time-accurate numerical simulation is performed to study the unsteady aerodynamic interaction between a canard and wing with emphasis on the effects of the canard on the configuration's dynamic response characteristics. The thin-layer Reynolds-averaged Navier-Stokes Equations with various turbulence models are used in this study. Computations are made on a generic, analytically-defined, close-coupled canard-wing-body configuration which has been the subject of numerous previously published experimental studies during the 1970's to mid-80's. More recently, a series of steady-flow simulations has been performed and published by the author. In the current study, the configuration is given prescribed ramp and oscillatory motions in order to predict characteristics such as the damping-in-pitch and oscillatory longitudinal stability parameters. The current computations are made at high-subsonic and transonic Mach numbers, moderate angles-of- attack from -4 to 20 degrees, and at various pitch rates and reduced frequencies. Comparisons of pressures and integrated force quantities (e.g. lift, drag, pitching moment and selected dynamic parameters) are made with other published computational results and available experimental data. Results showing the unsteady effects of the canard on surface pressures, integrated forces, canard-wing vortex interaction and vortex breakdown will be presented.

Tu, Eugene L.↗

NASA NTRS2002

Technology Development for Deployable Aerodynamic Decelerators at Mars

Parachutes used for Mars landing missions are only certified for deployment at Mars behind blunt bodies flying at low angles of attack, Mach numbers up to 2.2, and dynamic pressures of up to 800 Pa. NASA is currently studying entry vehicle concepts for future robotic missions to Mars that would require parachutes to be deployed at higher Mach numbers and dynamic pressures. This paper demonstrates the need for expanding the parachute deployment envelope, and describes a three-phase technology development activity that has been initiated to address the need. The end result of the technology development program will be a aerodynamic decelerator system that can be deployed at Mach numbers of up to 3.1 and dynamic pressures of up to 1400 Pa.

Masciarelli, James P.↗

NASA NTRS1992

A computational examination of directional stability for smooth and chined forebodies at high-alpha

Computational Fluid Dynamics (CFD) has been used to study aircraft forebody flowfields at low-speed, angle-of-attack conditions with sideslip. The purpose is to define forebody geometries which provide good directional stability characteristics under these conditions. The flows over the experimentally investigated F-5A forebody and chine type configuration, previously computed by the authors, were recomputed with better grid topology and resolution. The results were obtained using a modified version of CFL3D (developed at NASA Langley) to solve either the Euler equations or the Reynolds equations employing the Baldwin-Lomax turbulence model with the Degani-Schiff modification to account for massive crossflow separation. Based on the results, it is concluded that current CFD methods can be used to investigate the aerodynamic characteristics of forebodies to achieve desirable high angle-of-attack characteristics. An analytically defined generic forebody model is described, and a parametric study of various forebody shapes was then conducted to determine which shapes promote a positive contribution to directional stability at high angle-of-attack. An unconventional approach for presenting the results is used to illustrate how the positive contribution arises. Based on the results of this initial parametric study, some guidelines for aerodynamic design to promote positive directional stability are presented.

Ravi, Ramakrishnan↗

NASA NTRS1921

The law relating to air currents

Discussed here are the aerodynamics of a subdivided wing section. The emphasis is upon the increase of lift with more acute angles of attack. Also discussed are wind tunnel tests of the relations among wind resistance, lift, angle of attack, and velocity.

Source record↗

NASA NTRS1978

An aerodynamic analysis of several hypersonic research airplane concepts from M = 0.2 to 6.0

Several conceptual hypersonic research airplanes, designed within the constraints of a B-52 launch aircraft, have been studied experimentally and analytically at Mach numbers from 0.2 to 6.0. Vehicles built to these criteria for Mach 6 cruise were shown to be feasible, if careful attention was paid to the low speed lift, drag, and high angle of attack stability to assure successful landings and transonic pitch angle maneuvers. The integrated scramjet engine drag was high at subsonic speeds and appears to be constant with Reynolds number. The variable geometry airfoil used previously to improve directional stability was shown to be equally adaptable to the improvement of longitudinal stability. The vortex lattice theory gave good subsonic predictions of lift, drag due to lift, and pitching moments. Wind tunnel tests must be relied on for the drag at zero lift, trim, static margins and lateral-directional stability. The Gentry Hypersonic Arbitrary Body Program gave good predictions of the trends of lift, drag, and pitching moments with angle of attack at Mach numbers above 3, but the level of the values were not consistently predicted. No currently available theory or program gave accurate predictions of directional stability or dihedral effects at hypersonic speeds.

Penland, J. A.↗

NASA NTRS1980

The prediction of pressure distributions on an arrow-wing configuration including the effect of camber, twist, and a wing fin

Wind tunnel tests of an arrow wing body configuration consisting of flat, twisted, and cambered twisted wings were conducted at Mach numbers from 0.40 to 2.50 to provide an experimental data base for comparison with theoretical methods. A variety of leading and trailing edge control surface deflections were included in these tests, and in addition, the cambered twisted wing was tested with an outboard vertical fin to determine its effect on wing and control surface loads. Theory experiment comparisons show that current state of the art linear and nonlinear attached flow methods were adequate at small angles of attack typical of cruise conditions. The incremental effects of outboard fin, wing twist, and wing camber are most accurately predicted by the advanced panel method PANAIR. Results of the advanced panel separated flow method, obtained with an early version of the program, show promise that accurate detailed pressure predictions may soon be possible for an aeroelasticity deformed wing at high angles of attack.

Bobbitt, P. J.↗

NASA NTRS1993

Plasma chamber testing of APSA coupons for the SAMPIE flight experiment

Among the solar cell technologies to be tested in space as part of the Solar Array Module Plasma Interactions Experiment (SAMPIE) will be the Advanced Photovoltaic Solar Array (APSA). Several prototype twelve cell coupons were built for NASA using different blanket materials and mounting techniques. The first conforms to the baseline design for APSA which calls for the cells to be mounted on a carbon loaded Kapton blanket to control charging in GEO. When deployed, this design has a flexible blanket supported around the edges. A second coupon was built with the cells mounted on Kapton-H, which was in turn cemented to a solid aluminum substrate. A final coupon was identical to the latter but used germanium coated Kapton to control atomic oxygen attack in LEO. Ground testing of these coupons in a plasma chamber showed considerable differences in plasma current collection. The Kapton-H coupon demonstrated current collection consistent with exposed interconnects and some degree of cell snapover. The other two coupons experienced anomalously large collection currents. This behavior is believed to a consequence of enhanced plasma sheaths supported by the weakly conducting carbon and germanium used in these coupons. The results reported here are the first experimental evidence that the use of such materials can result in power losses to high voltage space power systems.

Hillard, G. Barry↗

NASA NTRS1994

Neural Network for Positioning Space Station Solar Arrays

As a shuttle approaches the Space Station Freedom for a rendezvous, the shuttle's reaction control jet firings pose a risk of excessive plume impingement loads on Freedom solar arrays. The current solution to this problem, in which the arrays are locked in a feathered position prior to the approach, may be neither accurate nor robust, and is also expensive. An alternative solution is proposed here: the active control of Freedom's beta gimbals during the approach, positioning the arrays dynamically in such a way that they remain feathered relative to the shuttle jet most likely to cause an impingement load. An artificial neural network is proposed as a means of determining the gimbal angles that would drive plume angle of attack to zero. Such a network would be both accurate and robust, and could be less expensive to implement than the current solution. A network was trained via backpropagation, and results, which compare favorably to the current solution as well as to some other alternatives, are presented. Other training options are currently being evaluated.

Graham, Ronald E.↗

NASA NTRS1990

Expert systems for superalloy studies

There are many areas in science and engineering which require knowledge of an extremely complex foundation of experimental results in order to design methodologies for developing new materials or products. Superalloys are an area which fit well into this discussion in the sense that they are complex combinations of elements which exhibit certain characteristics. Obviously the use of superalloys in high performance, high temperature systems such as the Space Shuttle Main Engine is of interest to NASA. The superalloy manufacturing process is complex and the implementation of an expert system within the design process requires some thought as to how and where it should be implemented. A major motivation is to develop a methodology to assist metallurgists in the design of superalloy materials using current expert systems technology. Hydrogen embrittlement is disasterous to rocket engines and the heuristics can be very complex. Attacking this problem as one module in the overall design process represents a significant step forward. In order to describe the objectives of the first phase implementation, the expert system was designated Hydrogen Environment Embrittlement Expert System (HEEES).

Workman, Gary L.↗

NASA NTRS1998

Recent Dynamic Measurements and Considerations for Aerodynamic Modeling of Fighter Airplane Configurations

As airplane designs have trended toward the expansion of flight envelopes into the high angle of attack and high angular rate regimes, concerns regarding modeling the complex unsteady aerodynamics for simulation have arisen. Most current modeling methods still rely on traditional body axis damping coefficients that are measured using techniques which were intended for relatively benign flight conditions. This paper presents recent wind tunnel results obtained during large-amplitude pitch, roll and yaw testing of several fighter airplane configurations. A review of the similitude requirements for applying sub-scale test results to full-scale conditions is presented. Data is then shown to be a strong function of Strouhal number - both the traditional damping terms, but also the associated static stability terms. Additionally, large effects of sideslip are seen in the damping parameter that should be included in simulation math models. Finally, an example of the inclusion of frequency effects on the data in a simulation is shown.

Brandon, Jay M.↗

NASA NTRS2011

Slat Noise Simulations: Status and Challenges

Noise radiation from the leading edge slat of a high-lift system is known to be an important component of aircraft noise during approach. NASA's Langley Research Center is engaged in a coordinated series of investigations combining high-fidelity numerical simulations and detailed wind tunnel measurements of a generic, unswept, 3-element, high-lift configuration. The goal of this effort is to provide a validated predictive capability that would enable identification of the dominant noise source mechanisms and, ultimately, help develop physics inspired concepts for reducing the far-field acoustic intensity. This paper provides a brief overview of the current status of the computational effort and describes new findings pertaining to the effects of the angle of attack on the aeroacoustics of the slat cove region. Finally, the interplay of the simulation campaign with the concurrently evolving development of a benchmark dataset for an international workshop on airframe noise is outlined.

Choudhari, Meelan M.↗

NASA NTRS1976

Simulator study of the effectiveness of an automatic control system designed to improve the high-angle-of-attack characteristics of a fighter airplane

A piloted, fixed-base simulation was conducted to study the effectiveness of some automatic control system features designed to improve the stability and control characteristics of fighter airplanes at high angles of attack. These features include an angle-of-attack limiter, a normal-acceleration limiter, an aileron-rudder interconnect, and a stability-axis yaw damper. The study was based on a current lightweight fighter prototype. The aerodynamic data used in the simulation were measured on a 0.15-scale model at low Reynolds number and low subsonic Mach number. The simulation was conducted on the Langley differential maneuvering simulator, and the evaluation involved representative combat maneuvering. Results of the investigation show the fully augmented airplane to be quite stable and maneuverable throughout the operational angle-of-attack range. The angle-of-attack/normal-acceleration limiting feature of the pitch control system is found to be a necessity to avoid angle-of-attack excursions at high angles of attack. The aileron-rudder interconnect system is shown to be very effective in making the airplane departure resistant while the stability-axis yaw damper provided improved high-angle-of-attack roll performance with a minimum of sideslip excursions.

Gilbert, W. P.↗

NASA NTRS1983

Subsonic roll oscillation experiments on the Standard Dynamics Model

The experimental determination of the subsonic roll derivatives of the Standard Dynamics Model, which is representative of a current fighter aircraft configuration, is described. The direct, cross and cross-coupling derivatives are presented for angles of attack up to 41 deg and sideslip angles in the range from -5 deg to 5 deg, as functions of oscillation frequency. The derivatives exhibited significant nonlinear trends at high incidences and were found to be extremely sensitive to sideslip angle at angles of attack near 36 deg. The roll damping and dynamic cross derivatives were highly frequency dependent at angles of attack above 30 deg. The highest values measured for the dynamic cross and cross-coupling derivatives were comparable in magnitude with the maximum roll damping. The effects of oscillation amplitude and Mach number were also investigated, and the direct derivatives were correlated with data from another facility.

Beyers, M. E.↗

NASA NTRS1993

Large eddy simulation of a boundary layer with concave streamwise curvature

One of the most exciting recent developments in the field of large eddy simulation (LES) is the dynamic subgrid-scale model. The dynamic model concept is a general procedure for evaluating model constants by sampling a band of the smallest scales actually resolved in the simulation. To date, the procedure has been used primarily in conjunction with the Smagorinsky model. The dynamic procedure has the advantage that the value of the model constant need not be specified a priori, but rather is calculated as a function of space and time as the simulation progresses. This feature makes the dynamic model especially attractive for flows in complex geometries where it is difficult or impossible to calibrate model constants. The dynamic model was highly successful in benchmark tests involving homogeneous and channel flows. Having demonstrated the potential of the dynamic model in these simple flows, the overall direction of the LES effort at CTR shifted toward an evaluation of the model in more complex situations. The current test cases are basic engineering-type flows for which Reynolds averaged approaches were unable to model the turbulence to within engineering accuracy. Flows currently under investigation include a backward-facing step, wake behind a circular cylinder, airfoil at high angles of attack, separated flow in a diffuser, and boundary layer over a concave surface. Preliminary results from the backward-facing step and cylinder wake simulations are encouraging. Progress on the LES of a boundary layer on a concave surface is discussed. Although the geometry of a concave wall is not very complex, the boundary layer that develops on its surface is difficult to model due to the presence of streamwise Taylor-Gortler vortices. These vortices arise as a result of a centrifugal instability associated with the convex curvature.

Lund, Thomas S.↗

NASA NTRS1992

An experimental/computational study of sharp fin induced shock wave/turbulent boundary layer interactions at Mach 5 - Experimental results

A combined experimental/computational study has been performed of sharp fin induced shock wave/turbulent boundary layer interactions at Mach 5. The current paper focuses on the experiments and analysis of the results. The experimental data include mean surface heat transfer, mean surface pressure distributions and surface flow visualization for fin angles of attack of 6, 8, 10, 12, 14 and 16-degrees at Mach 5 under a moderately cooled wall condition. Comparisons between the results and correlations developed earlier show that Scuderi's correlation for the upstream influence angle (recast in a conical form) is superior to other such correlations in predicting the current results, that normal Mach number based correlations for peak pressure heat transfer are adequate and that the initial heat transfer peak can be predicted using pressure-interaction theory.

Rodi, Patrick E.↗

NASA NTRS1976

An investigation to determine the pressure distribution on the 0.0137 scale solid rocket booster forebody (MSFC model 467) at angles of attack at or near 90 deg and high Reynolds numbers in the MSFC High Reynolds Number Wind Tunnel (SA29F)

An aerodynamic investigation was conducted in the MSFC High Reynolds Number Wind Tunnel to determine the pressure distribution over the foresection of the current 146 inch diameter shuttle SRB. The test model consisted of a 0.0137 scale version of the SRB nose cone and a forward portion of the cylindrical body which was approximately 2.7 calibers in length. The pressure distributions are plotted as a function of longitudinal station ratioed to body diameter and circumferential location for each angle of attack and Mach number. A Reynolds number variation study was made for Mach numbers of 0.4 and 0.6 at an angle of attack of 270 deg and roll angle of 180 deg.

Ramsey, P. E.↗

NASA NTRS1987

Effect of Reynolds number variation on aerodynamics of a hydrogen-fueled transport concept at Mach 6

Two separate tests have been made on the same blended wing-body hydrogen-fueled transport model at a Mach number of about 6 and a range of Reynolds number (based on theoretical body length) of 1.577 to 55.36 X 10 to the 6th power. The results of these tests, made in a conventional hypersonic blowdown tunnel and a hypersonic shock tunnel, are presented through a range of angle of attack from -1 to 8 deg, with an extended study at a constant angle of attack of 3 deg. The model boundary layer flow appeared to be predominately turbulent except for the low Reynolds number shock tunnel tests. Model wall temperatures varied considerably; the blowdown tunnel varied from about 255 F to 340 F, whereas the shock tunnel had a constant 70 F model wall temperature. The experimental normal-force coefficients were essentially independent of Reynolds number. A current theoretical computer program was used to study the effect of Reynolds number. Theoretical predictions of normal-force coefficients were good, particularly at anticipated cruise angles of attack, that is 2 to 5 deg. Axial-force coefficients were generally underestimated for the turbulent skin friction conditions, and pitching-moment coefficients could not be predicted reliably.

Penland, Jim A.↗