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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 361 records · Page 20
All sourcesNASA NTRSDOE OSTI
NASA NTRS1974

Wind tunnel blockage and support interference effects on winged-body models at Mach numbers from 0.6 to 1.0

Three sting-mounted winged-body models with tunnel blockages of 0.1, 1.0, and 2.0 percent were tested in the Lewis Research Center's 8- by 6- Foot Supersonic Wind Tunnel. Fuselage pressures were obtained over a Mach number range of 0.6 to 1.0 at angles of attack from 0 deg to 4 deg. Two other types of model support were investigated, which included simulated wing-tip and fuselage support-strut mountings. The effects of tunnel porosity and sidewall geometry were also investigated. Model blockage effects were small up to M sub 0 = 0.95. At higher speeds the major blockage effect observed was a displacement of the local transonic terminal shocks on the model. The effects of the wing-tip type of model support were small up to M sub 0 = 0.95, but disturbances were observed on the fuselage at higher speeds. Changes in local tunnel porosity were effective in reducing the disturbances up to M sub 0 = 0.975, but a change in sidewall geometry was not.

Blaha, B. J.↗
NASA NTRS1953

A study of the use of leading-edge notches as a means for improving the low-speed pitching-moment characteristics of a thin 45-degree swept wing of aspect ratio 4

In an effort to develop a simple but effective means for controlling the stall progression of swept wings, a low-speed investigation has been undertaken primarily to determine the effectiveness of leading-edge notches. The investigation included consideration of notch size, location, and shape, as well as combined effects of notch—chord-extension and notch-fence arrangements. The wing used in the study had 45° quarter-chord sweep, aspect ratio 4, taper ratio 0.3, and NACA 65A006 airfoil sections streamwise. The results of the investigation indicated that although substantial improvements were obtained in the wing pitching-moment characteristics at moderate angles of attack, by use of outboard leading-edge notches, none of these notches produced desirable wing stability characteristics throughout the angle-of-attack range. The wing pitching-moment characteristics were relatively insensitive to notch size and to small changes in span-wise location. The use of a notch at the inboard end of a chord-extension increased the effectiveness of the chord-extension near stall and produced the same results as a chord-extension having an equivalent inboard overhang. Small localized notch—chord-extension combinations were found to produce the same pitching-moment characteristics as much larger chord-extension configurations. A leading-edge root cutout used in conjunction with a small outboard notch—chord-extension combination resulted in fairly desirable overall wing pitching-moment characteristics.

Joseph Weil↗
NASA NTRS1994

Influence of Incidence Angle on Sound Generation by Airfoils Interacting with High-Frequency Gusts

A theoretical model is developed for the sound generated when a convected vortical or entropic gust encounters an airfoil at nonzero angle of attack. The theory is based on a linearization of the Euler equations about the steady, subsonic flow past the airfoil. High-frequency gusts, whose wavelengths are short compared to the airfoil chord, but long compared to the displacement of the mean-flow stagnation point from the leading edge, are considered. The analysis utilizes singular perturbation techniques and involves four asymptotic regions. Local regions, which scale on the gust wavelength, are present at the airfoil leading and trailing edges. Behind the airfoil a "transition" region, which is similar to the transition zone between illuminated and shadow zones in optical problems, is present. In the outer region, far away from the airfoil edges and wake, the solution has a geometric-acoustics form. The primary sound generation is found to be concentrated in the local leading-edge region. The trailing edge plays a secondary role as a scatterer of the sound generated in the leading-edge region. Parametric calculations are presented which illustrate that moderate levels of airfoil steady loading can significantly affect the sound field produced by airfoil-gust interactions.

Matthew R. Myers↗
DOE OSTI2023

Spatio-Temporal Deep Graph Network for Event Detection, Localization, and Classification in Cyber-Physical Electric Distribution System

This work proposes a deep graph learning framework to identify, locate, and classify power, cyber, and cyber power events at the distribution system level. The proposed algorithm jointly exploits spatial, temporal, and node-level cyber and physical data features. The developed graph neural network, together with a deep autoencoder, utilizes physical measurements from distribution level phasor measurement units and cyber data from communication network logs. The spatial structure of the synchrophasor measurements and network is incorporated through a weighted adjacency matrix. The temporal structure is incorporated by defining a spatial operation in the gated recurrent unit. This spatio-temporal learning element resides inside a power event detection, localization, and classification module that provides the degree of confidence for an event label. To accurately pinpoint the location of an event to the nearest bus equipped with a measurement unit, a combination of squared error and proximity score is utilized. Also included is a cyber event detection module that employs heteroskedasticity to analyze the significance of various cyber features during different types of attacks. Finally, a dual-bit cyber-power decision table determines the nature of the event. The proposed method is validated on two distribution systems modeled in OPAL-RT/Hypersim with limited phasor measurement units for different possible physical and cyber events. Further analyses include comparison with other state-of-the-art methods and validation in the presence of measurement noise. As a result, our method outperforms existing approaches and achieves an average detection accuracy of 97.97%, F1-score of 96.88%, precision of 96.53%, and recall of 98.57%.

24 POWER TRANSMISSION AND DISTRIBUTION↗
NASA NTRS1950

Investigation of Aerodynamic and Icing Characteristics of Water-Inertia-Separation Inlets for Turbojet Engines

The results of an investigation of several internal water-inertia-separation inlets consisting of a main duct and an alternate duct designed to prevent automatically the entrance of large quantities of water into a turbojet engine in icing conditions are presented. Total-pressure losses and icing characteristics for a direct-ram inlet and the inertia-separation inlets are compared at similar aerodynamic and simulated icing conditions. Complete ice protection for inlet guide vanes could not be achieved with the inertia-separation inlets investigated. Approximately 8 percent of the volume of water entering the nacelles remained. In the air passing into the compressor inlet. Heavy alternate-duct-elbow ice formations caused by secondary inertia separation resulted in rapid total-pressure losses and decreases in mass flow. The duration in an icing condition for an inertia-separation- inlet, without local surface heating, was increased approximately four times above that for a direct-ram inlet with a compressor-inlet screen. For normal nonicing operation, the inertia-separation- inlet total-pressure losses were comparable to a direct-ram installation. The pressure losses and the circumferential uniformity of the mass flow in all the inlets were relatively independent of angle of attack. Use of an inertia-separation inlet would in most cases require a larger diameter nacelle than a direct-ram inlet in order to obtain an alternate duct sufficiently large to pass the required engine air flow at duct Mach numbers below 1.0 at the minimum area.

VonGlahn, Uwe↗
NASA NTRS1989

Surface pressure distributions on a delta wing undergoing large amplitude pitching oscillations

Wind tunnel experiments were performed on a 70 deg sweep delta wing to determine the effect of a sinusoidal pitching motion on the pressure field on the suction side of the wing. Twelve pressure taps were placed from 35 to 90 percent of the chord, at 60 percent of the local semi-span. Pressure coefficients were measured as a function of Reynolds number and pitch rate. The pressure coefficient was seen to vary at approximately the same frequency as the pitching frequency. The relative pressure variation at each chord location was comparable for each case. The average pressure distribution through each periodic motion was near the static distribution for the average angle of attack. Upon comparing the upstroke and downstroke pressures for a specific angle of attack, the downstroke pressures were slightly larger. Vortex breakdown was seen to have the most significant effect at the 40 to 45 percent chord location, where a decrease in pressure was apparent.

Thompson, Scott A.↗
NASA NTRS2001

Cranked Arrow Wing (F-16XL-1) Flight Flow Physics with CFD Predictions at Subsonic and Transonic Speeds

The computational fluid dynamics (CFD) modeling used has produced reasonably good global upper-surface pressure coefficient comparisons with measured flight data at both transonic and subsonic speeds at the angles of attack presented. Boundary layer comparisons showed the profiles to be reasonably well predicted inboard and under the primary vortex system. However, the secondary vortex profile was not well predicted either at the anticipated separation point or under the secondary vortex. Moreover, the flight data showed there to be a vortex/boundary-layer interaction that occurred in the vicinity of the secondary vortex. The spanwise distribution of local skin friction measured data was reasonably well predicted, especially away from the wing leading-edge. Lastly, predicted and measured flight-pressures, as well as flight-image data, for the F-16XL-1 airplane are now available via the World Wide Web.

Lamar, John E.↗
NASA NTRS2003

Cranked Arrow Wing (F-16XL-1) Flight Flow Physics with CFD Predictions at Subsonic and Transonic Speeds

The CFD modeling used has produced reasonably good global upper-surface pressure coefficient comparisons with measured flight data at both transonic and subsonic speeds at the angles of attack presented. Boundary layer comparisons showed the profiles to be reasonably well predicted inboard and under the primary vortex system. However, the secondary vortex profile was not well predicted either at the anticipated separation point or under the secondary vortex. Moreover, the flight data showed there to be a vortex/boundary-layer interaction that occurred in the vicinity of the secondary vortex. The spanwise distribution of local skin friction measured data was reasonably well predicted, especially away from the wing leading-edge. Lastly, predicted and measured flight-pressures, as well as flight-image data, for the F-16XL-1 airplane are now available via the World Wide Web.

Lamar, John E.↗
NASA NTRS1981

Force and moment, flow-visualization, and boundary-layer tests on a shuttle orbiter model at Mach 6

Force and moment, flow visualization, and boundary layer state tests were conducted using two 0.004 scale shuttle orbiter models. The force and moment tests were conducted for an angle of attack range from 20 to 40 deg and for Reynolds numbers based on reference length from 0.4 million to 3.6 million. Schlieren photographs were obtained for each angle of attack and Reynolds number. The boundary layer state tests, which were conducted using hot film sensors mounted in a separate model, were conducted over the same range of conditions as the force tests. Test results were combined to show that changes in the boundary layer on a typical hypersonic force test model affect measurement of the axial force coefficient and that the state of the local boundary layer is important for interpreting hypersonic aerodynamic test results.

Calloway, R. L.↗
NASA NTRS1982

Aerothermodynamic environment of a Titan aerocapture vehicle

The extent of convective and radiative heating for a Titan aerocapture vehicle is investigated. The flow in the shock layer is assumed to be axisymmetric, steady, viscous, and compressible. It is further assumed that the gas is in chemical and local thermodynamic equilibrium and tangent slab approximation is used for the radiative transport. The effect of the slip boundary conditions on the body surface and at the shock wave are included in the analysis of high-altitude entry conditions. The implicit finite difference techniques is used to solve the viscous shock-layer equations for a 45 degree sphere cone at zero angle of attack. Different compositions for the Titan atmosphere are assumed, and results are obtained for the entry conditions specified by the Jet Propulsion Laboratory.

Tiwari, S. N.↗
NASA NTRS1990

Numerical study of vortical flow over a sideslipping delta wing

The three-dimensional vortical viscous flow past a sideslipping delta wing at a high angle of attack is numerically investigated. A computational method for calculating low-speed viscous flowfields is developed. The time-iterative method uses an implicit upwind-relaxation finite-difference algorithm with a nonsingular eigensystem to solve the preconditioned, three-dimensional, incompressible Navier-Stokes equations in curvilinear coordinates. An algebraic turbulence model is implemented to account for the eddy viscosity. The technique of local time stepping is incorporated to accelerate the rate of convergence to a steady-state solution. Computed results are compared with experimental data.

Hsu, C.-H.↗
NASA NTRS2006

Design and Flight Evaluation of a New Force-Based Flow Angle Probe

A novel force-based flow angle probe was designed and flight tested on the NASA F-15B Research Testbed aircraft at NASA Dryden Flight Research Center. The prototype flow angle probe is a small, aerodynamic fin that has no moving parts. Forces on the prototype flow angle probe are measured with strain gages and correlated with the local flow angle. The flow angle probe may provide greater simplicity, greater robustness, and better access to flow measurements in confined areas relative to conventional moving vane-type flow angle probes. Flight test data were obtained at subsonic, transonic, and supersonic Mach numbers to a maximum of Mach 1.70. Flight conditions included takeoff, landing, straight and level flight, flight at higher aircraft angles of attack, and flight at elevated g-loadings. Flight test maneuvers included angle-of-attack and angle-of-sideslip sweeps. The flow angle probe-derived flow angles are compared with those obtained with a conventional moving vane probe. The flight tests validated the feasibility of a force-based flow angle measurement system.

Corda, Stephen↗
DOE OSTI2021

Self-Security for Grid-Interactive Smart Inverters Using Steady-State Reference Model

Smart inverters exchange information with other devices through a shared communication link, making the inverters more prone to receive harmful commands from external parties. This erroneous data can be received due to an anomaly in the system, such as a device fault, unintentional utility operator action, or a cyber-attack. In this paper, a device-level self-security strategy is implemented using reference models for a grid-interactive inverter to examine the incoming power setpoints, detect the anomalies, and protect the system accordingly. The PQ setpoints received from the utility supervisory controller are autonomously examined using the inverter’s normal and stable operating regions before engaging the setpoints to the inverter’s local controller. Grid parameters are estimated in real-time during the examination process. The efficacy of the self-security algorithm is tested using a three-phase 3-kVA SiC-MOSFET inverter and a 12-kW NHR 9410 regenerative grid emulator. The results verify that the proposed method can detect harmful PQ setpoints that can cause abnormal or unstable inverter operation.

Gursoy, Mehmetcan↗
DOE OSTI2022

Cyber-Physical Resiliency for Wind Power Generation

During the next three years, wind power generation is expected to add more generation capacity to the national electrical grid than any other energy sector. With new wind turbine designs rated power over 10 MW and wind farms reaching over 1 GW capacities, the consequences of cyber-attacks on wind power generation are becoming increasingly more critical. Moreover, the known vulnerabilities of wind turbine control systems and the potential damaging effects of intrusions, motivate an urgent protection improvement for the wind power generation sector. This program has developed a variety of new adaptive defense technologies that enable wind power generation systems to survive sophisticated cyberattacks by enhancing the control systems capabilities of detection, localization, and accommodation. The introduction of these technologies in the on-shore and the emerging off-shore market will result in a significantly more reliable and secure wind power infrastructure.

17 WIND ENERGY↗
DOE OSTI2022

Phase Transfer‐Mediated Degradation of Ether‐Based Localized High‐Concentration Electrolytes in Alkali Metal Batteries

Abstract Localized high‐concentration electrolytes (LHCEs) have attracted interest in alkali metal batteries due to the advantages of forming stable solid‐electrolyte interphases (SEIs) on anodes and good chemical/electrochemical stability. Herein, a new degradation mechanism is revealed for ether‐based LHCEs that questions their compatibility with alkali metal anodes (Li, Na, and K). Specifically, the ether solvent reacts with alkali metals to generate solvated electrons (e s − ) that attack hydrofluoroether co‐solvents to form a series of byproducts. The ether solvent essentially acts as a phase‐transfer reagent that continuously transfers electrons from solid‐phase metals into the solution phase, thus inhibiting the formation of stable SEI and leading to continuous alkali metal corrosion. Switching to an ester‐based solvating solvent or intercalation anodes such as graphite or molybdenum disulfide has been shown to avoid such a degradation mechanism due to the absence of e s − .

Chen, Xiaojuan↗
DOE OSTI2023

Influence of native oxide film on corrosion behavior of additively manufactured stainless steel 316L

The influence of the native oxide film on passive film properties and localized corrosion of additively manufactured SS 316L was studied in 1 wt% HCl by XPS characterization, electrochemical polarization curves, and post-test morphology analysis by SEM. Increased Cr oxidation kinetics was observed in the as-polished sample with the native oxide film resulting in formation of an overall more protective and compact film compared to the cathodically-activated sample. Electrochemical analysis showed that corrosive attack varied between dislocation cell boundaries to cell interiors depending on the initial surface state and polarization conditions. In conclusion, a corrosion mechanism is proposed to explain this variation.

36 MATERIALS SCIENCE↗
NASA NTRS1980

Aerodynamic-structural analysis of dual bladed helicopter systems

The aerodynamic and structural feasibility of the birotor blade concept is assessed. The inviscid flow field about the dual bladed rotor was investigated to determine the aerodynamic characteristics for various dual rotor blade placement combinations with respect to blade stagger, gap, and angle of attack between the two blades. The boundary layer separation on the rotors was studied and three dimensional induced drag calculations for the dual rotor system are presented. The thrust and power requirements of the rotor system were predicted. NASTRAN, employed as the primary modeling tool, was used to obtain a model for predicting in plane bending, out of plane bending, and the torsional behavior of the birotors. Local hub loads, blade loads, and the natural frequencies for the birotor configuration are discussed.

Selberg, B. P.↗
NASA NTRS1993

Transition effects on compressible dynamic stall of transiently pitching airfoils

Experimental results and analysis of the effects of boundary layer tripping on dynamic stall of a transiently pitching airfoil are presented. At low Mach numbers, the tripped airfoil exhibits qualitative similarity with the behavior of the untripped airfoil. However, the local supersonic flow at Mach numbers greater than 0.3 is significantly modified by the trip leading to vastly different shock/boundary layer interactions, dynamic stall onset and vortex formation angles. The formation of the laminar separation bubble is found to have a favorable influence in delaying dynamic stall on the untripped airfoil flow. In both Mach number regimes, the tripped flow actually stalls at slightly lower angles of attack. Further experimentation with three trips on an oscillating airfoil showed that the dynamic stall process is very sensitive to the state of the turbulence in the boundary layer. This sensitivity points to a need for finer turbulence modeling techniques for use in dynamic stall computations.

Wilder, M. C.↗