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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 469 records · Page 26
All sourcesNASA NTRSDOE OSTI
NASA NTRS1945

A Systematic Investigation of Pressure Distributions at High Speeds over Five Representative NACA Low-Drag and Conventional Airfoil Sections

Pressure distributions determined from high-speed wind-tunnel tests are presented for five NACA airfoil sections representative of both low-drag and conventional types. Section characteristics of lift, drag, and quarter-chord pitching moment are presented along with the measured pressure distributions for the NACA 65sub2-215 (a=0.5), 66sub2-215 (a=0.6), 0015, 23015, and 4415 airfoils for Mach numbers up to approximately 0.85. A critical study is made of the airfoil pressure distributions in an attempt to formulate a set of general criteria for defining the character of high speed flows over typical airfoil shapes. Comparisons are made of the relative characteristics of the low-drag and conventional airfoils investigated insofar as they would influence the high-speed performance and the high-speed stability and control characteristics of airplanes employing these wing sections.

Graham, Donald J↗
NASA NTRS1956

Lift, Drag, Static Stability, and Buffet Boundaries of a Model of the McDonnell F3H-1N Airplane at Mach Numbers from 0.40 to 1.27, TED No. NACA DE 351

The National Advisory Committee for Aeronautics has conducted a flight test of a model approximating the McDonnell F3H-lN airplane configuration to determine its pitch-up and buffet boundaries, as well as the usual longitudinal stability derivatives obtainable from the pulsed- tail technique. The test was conducted by the freely flying rocket- boosted model technique developed at the Langley Laboratory; results were obtained at Mach numbers from 0.40 to 1.27 at corresponding Reynolds numbers of 2.6 x 10(exp 6) and 9.0 x 10(exp 6). The phenomena of pitch-up, buffet, and maximum lift were encountered at Mach numbers between 0.42 and 0.85. The lift-curve slope and wing-root bending-moment slope increased with increasing angle of attack, whereas the static stability decreased with angle of attack at subsonic speeds and increased at transonic speeds. There was little change in trim at low lift at transonic speeds.

Crabill, Norman L.↗
NASA NTRS1972

Flight-determined derivatives and dynamic characteristics of the CV-990 airplane

Flight-determined longitudinal and lateral-directional stability and control derivatives are presented for the CV-990 airplane for various combinations of Mach number, altitude, and flap setting throughout the flight envelope up to a Mach number of 0.87. Also presented are the dynamic characteristics of the aircraft calculated from the flight-obtained derivatives and the measured phugoid characteristics. The derivative characteristics were obtained from flight records of longitudinal and lateral-directional transient oscillation maneuvers by using a modified Newton-Raphson digital derivative determination technique. Generally the derivatives exhibited consistent variation with lift coefficient in the low-speed data and with Mach number and altitude in the high-speed data. Many also varied with flap deflection, notably spoiler effectiveness and directional stability.

Gilyard, G. B.↗
NASA NTRS1973

Design of various fixed-geometry water-lubricated hydrodynamic journal bearings for maximum stability

This publication is the result of over 260 fractional-frequency-whirl stability tests on a variety of fixed-geometry journal bearings. It is intended principally as a guide in the selection and design of antiwhirl bearings that must operate at high speeds and low loads in low-viscosity fluids such as water or liquid metals. However, the various fixed-geometry configurations can be employed as well in applications where other lubricants, such as oil, are used and fractional-frequency whirl is a problem. The important parameters that effect stability are discussed for each bearing type, and design curves to facilitate the design of optimum-geometry bearings are included. A comparison of the stability of the different bearing configurations tested is also given.

Schuller, F. T.↗
NASA NTRS1977

Low-speed power effects on advanced fighter configurations with two-dimensional deflected thrust

Wind-tunnel studies at the Langley Research Center have shown that significant increases in maximum lift coefficient and stability and decreases in drag due to lift are obtained when two-dimensional vectored thrust is used in conjuction with a close-coupled canard. The configuration tested was somewhat above the theoretical minimum drag due to lift because of the sharp leading edge on the biconvex airfoil used on the wing and canard. An effort to design a new configuration which will approach the minimum drag due to lift while maintaining high-lift configurations was completed. The resulting model will incorporate a realistic planform, airfoil section, and twist for a transonic maneuvering configuration.

Paulson, J. W., Jr.↗
NASA NTRSDate not supplied

Flying Qualities Analysis and Piloted Simulation Testing of a Lift+Cruise Vehicle with Propulsion Failures in Hover and Low-Speed Conditions

The recent emergence of electric-Vertical Take-Off and Landing (eVTOL) vehicles for Urban Air Mobility (UAM) applications has resulted in a wide variety of configurations with unique stability and control characteristics. NASA is currently conducting research to develop conceptual design tools to accelerate public acceptance of these vehicles which includes requirements for safety during failure scenarios. This paper summarizes progress toward a toolbox for predicting flying qualities of eVTOL vehicles during critical propulsion failures that could impact the allowable design of the vehicle geometry or control system. Key topics include unique vulnerabilities of eVTOL/multirotor vehicles to propulsion failures, relevant flying qualities design metrics, and simulation modeling requirements for assessing flying qualities degradation due to failures. Results of a piloted simulation study conducted in the NASA Ames Vertical Motion Simulator (VMS) are presented. The VMS experiment was designed to assess and validate key handling qualities and safety design metrics for propulsion failures. These results show the correlation between control system design requirements and the degradation in handling qualities for various propulsion failures.

George Altamirano↗
NASA NTRS1985

Electromagnetic hot ion beam instabilities - Quasi-linear theory and simulation

This paper considers the quasi-linear theory of the right- and left-hand resonant electromagnetic instabilities driven by a hot ion beam streaming parallel to a magnetic field in a homogeneous Vlasov plasma. Using the single-mode approximation, the time evolutions of important parameters are obtained to show that for the range of parameters considered, reduction of the beam speed and formation of temperature anisotropies are the most significant factors in the quasi-linear stabilization process. Combining both instabilities in a quasi-linear study is found to produce a roughly equal mixture of both polarizations and relatively isotropic conditions for tenuous beam densities and low initial beam drift speeds. Computer simulations are used to compare with the quasi-linear results. The simulations justify the single-mode assumption, verify that quasi-linear changes are the means of saturation for the parameter range of concern, and check the nonlinear evolution of the system when both modes are present.

Rogers, B.↗
DOE OSTI2020

Fluid Dynamics Effects on Microstructure Prediction in Single-Laser Tracks for Additive Manufacturing of IN625

Single-track laser fusion were simulated using a heat-transfer-solidification-only (HTS) model and its extension with fluid dynamics (HTS_FD) model using a parallel open-source code, which included laminar fluid dynamics, flat-free surface of the molten alloy, heat transfer, phase-change, evaporation, and surface tension phenomena. The results illustrate that the fluid dynamics affects the solidification and ensuing microstructure. For the HTS_FD simulations, thermal gradient, G was found to exhibit a maximum at the extremity of the solidified pool ( i.e. , at the free surface), while for HTS simulations, G exhibited a maximum around the entire edge of the solidified pool. HTS_FD simulations predicted a wider range of cooling rates than the HTS simulations, exhibited an increased spread in the solidification speed, V variation within the melt-pool with respect to the HTS model results. Primary dendrite arm spacing (PDAS) were evaluated based on power law correlations and marginal stability theory models using the ( G , V ) from HTS and HTS_FD simulations to quantify the effect of the fluid dynamics on the microstructure. At low-laser powers and low-scan speeds, the PDAS obtained with the fluid dynamics model (HTS_FD) was larger by more than 30 pct with respect to the PDAS calculated with the simple HTS model. A new PDAS correlation, i.e. , \( \lambda_{1} \left[ {\mu {\text{m}}} \right] = 832\;G\left[ {\text{K/m}} \right]^{ - 0.5} V\left[ {\text{m/s}} \right]^{ - 0.25} \) , which uses the ( G , V ) results from the HTS_FD model was developed and validated against experimental results.

36 MATERIALS SCIENCE↗
NASA NTRS1990

Instability and transition; Proceedings of the Workshop, Hampton, VA, May 15-June 9, 1989. Vols. 1 & 2

The proceedings of the workshop on instability and transition, sponsored by the Institute for Computer Applications in Science and Engineering and the Langley Research Center, during May-June 1989 are presented. The proceedings contain panel discussions and the research papers resulting from the collaborative projects started during the workshop and completed shortly thereafter. Six panels are represented: theory, low-speed experiment, high-speed experiment, computation, receptivity, and roughness panels. Focus is concentrated on receptivity, advanced asymptotics, boundary-layer stability, free shear flows, compressible stability and transition, transition simulation, and transition modeling. The panel summaries and the position papers by the panelists as well as research papers of the experiment group and the receptivity group are discussed.

Hussaini, M. Y.↗
NASA NTRS1992

Current status of computational methods for transonic unsteady aerodynamics and aeroelastic applications

The status of computational methods for unsteady aerodynamics and aeroelasticity is reviewed. The key features of challenging aeroelastic applications is discussed in terms of the flowfield state - low angle high speed flows and high angle vortex dominated flows. The critical role played by viscous effects in determining aeroelastic stability for conditions of incipient flow separation is stressed. The need for a variety of flow modeling tools, from linear formulations to implementations of the Navier-Stokes equations, is emphasized. Estimates of computer run times for flutter calculations using several computational methods are given. Applications of these methods for unsteady aerodynamic and transonic flutter calculations for airfoils, wings, and configurations are summarized. Finally, recommendations are made concerning future research directions.

Edwards, John W.↗
NASA NTRS1956

Aerodynamic Characteristics and Flying Qualities of a Tailless Triangular-wing Airplane Configuration as Obtained from Flights of Rocket-propelled Models at Transonic and Supersonic Speeds

A flight investigation of rocket-powered models of a tailless triangular-wing airplane configuration was made through the transonic and low supersonic speed range at the Langley Pilotless Aircraft Research Station at Wallops Island, Va. An analysis of the aerodynamic coefficients, stability derivatives, and flying qualities based on the results obtained from the successful flight tests of three models is presented.

DAMPING DERIVATIVES - STABILITY↗
NASA NTRS1993

Computational fluid mechanics

Two papers are included in this progress report. In the first, the compressible Navier-Stokes equations have been used to compute leading edge receptivity of boundary layers over parabolic cylinders. Natural receptivity at the leading edge was simulated and Tollmien-Schlichting waves were observed to develop in response to an acoustic disturbance, applied through the farfield boundary conditions. To facilitate comparison with previous work, all computations were carried out at a free stream Mach number of 0.3. The spatial and temporal behavior of the flowfields are calculated through the use of finite volume algorithms and Runge-Kutta integration. The results are dominated by strong decay of the Tollmien-Schlichting wave due to the presence of the mean flow favorable pressure gradient. The effects of numerical dissipation, forcing frequency, and nose radius are studied. The Strouhal number is shown to have the greatest effect on the unsteady results. In the second paper, a transition model for low-speed flows, previously developed by Young et al., which incorporates first-mode (Tollmien-Schlichting) disturbance information from linear stability theory has been extended to high-speed flow by incorporating the effects of second mode disturbances. The transition model is incorporated into a Reynolds-averaged Navier-Stokes solver with a one-equation turbulence model. Results using a variable turbulent Prandtl number approach demonstrate that the current model accurately reproduces available experimental data for first and second-mode dominated transitional flows. The performance of the present model shows significant improvement over previous transition modeling attempts.

Hassan, H. A.↗
NASA NTRS2011

Effects of Cavities and Protuberances on Transition over Hypersonic Vehicles

Surface protuberances and cavities on a hypersonic vehicle are known to cause several aerodynamic or aerothermodynamic issues. Most important of all, premature transition due to these surface irregularities can lead to a significant rise in surface heating. To help understand laminar-turbulent transition induced by protuberances or cavities on a Crew Exploration Vehicle (CEV) surface, high-fidelity numerical simulations are carried out for both types of trips on a CEV wind tunnel model. Due to the large bluntness, these surface irregularities reside in an accelerating subsonic boundary layer. For the Mach 6 wind tunnel conditions with a roughness Reynolds number Re(sub kk) of 800, it was found that a protuberance with a height to boundary layer thickness ratio of 0.73 leads to strong wake instability and spontaneous vortex shedding, while a cavity with identical geometry only causes a rather weak flow unsteadiness. The same cavity with a larger Reynolds number also leads to similar spontaneous vortex shedding and wake instability. The wake development and the formation of hairpin vortices for both protuberance and cavity were found to be qualitatively similar to that observed for an isolated hemisphere submerged in a subsonic, low speed flat-plate boundary layer. However, the shed vortices and their accompanying instability waves were found to be slightly stabilized downstream by the accelerating boundary layer along the CEV surface. Despite this stabilizing influence, it was found that the wake instability spreads substantially in both wall-normal and azimuthal directions as the flow is evolving towards a transitional state. Similarities and differences between the wake instability behind a protuberance and a cavity are investigated. Computations for the Mach 6 boundary layer over a slender cylindrical roughness element with a height to the boundary layer thickness of about 1.1 also shows spontaneous vortex shedding and strong wake instability. Comparisons of detailed flow structures associated with protuberances at subsonic and supersonic edge Mach numbers indicate distinctively different instability mechanisms.

Chang, Chau-Lyan↗
NASA NTRS1976

Summary of flight tests to determine the spin and controllability characteristics of a remotely piloted, large-scale (3/8) fighter airplane model

An unpowered, large, dynamically scaled airplane model was test flown by remote pilot to investigate the stability and controllability of the configuration at high angles of attack. The configuration proved to be departure/spin resistant; however, spins were obtained by using techniques developed on a flight support simulator. Spin modes at high and medium high angles of attack were identified, and recovery techniques were investigated. A flight support simulation of the airplane model mechanized with low speed wind tunnel data over an angle of attack range of + or - 90 deg. and an angle of sideslip range of + or - 40 deg. provided insight into the effects of altitude, stability, aerodynamic damping, and the operation of the augmented flight control system on spins. Aerodynamic derivatives determined from flight maneuvers were used to correlate model controllability with two proposed departure/spin design criteria.

Holleman, E. C.↗
NASA NTRS1956

The Aerodynamic Characteristics in Pitch of a 1/15-Scale Model of the Grumman F11F-1 Airplane at Mach Numbers of 1.41, 1.61, and 2.01, TED No. NACA DE 390

Tests have been made in the Langley 4- by 4-foot supersonic pressure tunnel at Mach numbers of 1.41, 1.61, and 2.01 to determine the static longitudinal stability and control characteristics of various arrangements of the Grumman F11F-1 airplane. Tests were made of the complete model and various combinations of its component parts and, in addition, the effects of various body modifications, a revised vertical tail, and wing fences on the longitudinal characteristics were determined. The results indicate that for a horizontal-tail incidence of -10 deg the trim lift coefficient varied from 0.29 at a Mach number of 1.61 to 0.23 at a Mach number of 2.01 with a corresponding decrease in lift-drag trim from 3.72 to 3.15. Stick-position instability was indicated in the low-supersonic-speed range. A photographic-type nose modification resulted in slightly higher values of minimum drag coefficient but did not significantly affect the static stability or lift-curve slope. The minimum drag coefficient for the complete model with the production nose remained essentially constant at 0.047 throughout the Mach number range investigated.

Driver, Cornelius↗
NASA NTRS2012

Quiet High Speed Fan II (QHSF II): Final Report

This report details the aerodynamic, mechanical, structural design and fabrication of a Honey Engines Quiet High Speed Fan II (lower hub/tip ratio and higher specific flow than the Baseline I fan). This fan/nacelle system incorporates features such as advanced forward sweep and an advanced integrated fan/fan exit guide vane design that provides for the following characteristics: (1) Reduced noise at supersonic tip speeds, in comparison to current state-of-the-art fan technology; (2) Improved aeroelastic stability within the anticipated operating envelope; and (3) Aerodynamic performance consistent with current state-of-the-art fan technology. This fan was fabricated by Honeywell and tested in the NASA Glenn 9- by 15-Ft Low Speed Wind Tunnel for aerodynamic, aeromechanical, and acoustic performance.

Kontos, Karen↗
NASA NTRS1977

Effects of temperature transients at fan inlet of a turbofan engine

The effects of fan inlet temperature transients on the performance and stability of a turbofan engine were determined. The experiment was conducted at 90 and 74 percent of low-pressure-rotor military speed (9525 rpm) and with fan inlet temperature distortions having circumferential extents of 90 deg, 180 deg, 270 deg, and 360 deg. Temperature transients were controlled by varying the magnitude and rate of change of the inlet temperature rise. The engine response ranged from a momentary compressor pressure disturbance to low-pressure-compressor stall. The compressor distortion limits decreased with decreasing low-pressure-rotor speed and increased with increasing circumferential extent of distortion. Analysis of the data suggests strongly that the distortion limits of the compressor are a function of a critical magnitude of inlet temperature rise and are independent of the temperature rise rate.

Abdelwahab, M.↗