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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 163 records · Page 9
All sourcesNASA NTRSDOE OSTI
NASA NTRS2005

On the Application of Contour Bumps for Transonic Drag Reduction(Invited)

The effect of discrete contour bumps on reducing the transonic drag at off-design conditions on an airfoil have been examined. The research focused on fully-turbulent flow conditions, at a realistic flight chord Reynolds number of 30 million. State-of-the-art computational fluid dynamics methods were used to design a new baseline airfoil, and a family of fixed contour bumps. The new configurations were experimentally evaluated in the 0.3-m Transonic Cryogenic Tunnel at the NASA Langley Research center, which utilizes an adaptive wall test section to minimize wall interference. The computational study showed that transonic drag reduction, on the order of 12% - 15%, was possible using a surface contour bump to spread a normal shock wave. The computational study also indicated that the divergence drag Mach number was increased for the contour bump applications. Preliminary analysis of the experimental data showed a similar contour bump effect, but this data needed to be further analyzed for residual wall interference corrections.

Milholen, William E., II↗
NASA NTRS1987

Predictions of F-111 TACT aircraft buffet response and correlations of fluctuating pressures measured on aluminum and steel models and the aircraft

Results of buffet research that was conducted as part of the joint USAF/NASA F-111 TACT Research Program are presented. The correlation of wind tunnel and flight measurements of buffet excitation showed that there generally was good agreement between measurements of pressure fluctuations on the models and aircraft in regions of separated flow. At shock-wave boundaries of the separated flow, correlations of pressure fluctuations were not so good, due to Reynolds number and static elastic effects. The buffet prediction method, which applies a forcing function that is obtained by real-time integration of pressure time histories with the natural modes, is described. The generalized forces, including the effects of wing and tail, correlations of predicted and measured damping, and correlations of predicted and measured buffet response are presented. All presented data are for a Mach number of 0.8 with wing-sweep angles of 26 and 35 deg for a range of angles-of-attack that include buffet onset to high intensity buffeting. Generally, the buffet predictions were considered to be quite good particularly in light of past buffet-prediction experience.

Coe, Charles F.↗
NASA NTRS1990

Lyman alpha radiation in external galaxies

The Ly alpha line of atomic hydrogen is often a luminous component of the radiation emitted by distant galaxies. Except for those galaxies which have a substantial central source of non-stellar ionizing radiation, most of the Ly alpha radiation emitted by galaxies is generated within regions of the interstellar medium which are photoionized by starlight. Conversely, much of the energy radiated by photoionized regions is carried by the Ly alpha line. Only hot, massive stars are capable of ionizing hydrogen in the interstellar medium which surrounds them, and because such stars are necessarily short-lived, Ly alpha emission traces regions of active star formation. Researchers argue that the strength of the Ly alpha emission observed from external galaxies may be used to estimate quantitatively the dust content of the emitting region, while the Ly alpha line profile is sensitive to the presence of shock waves. Interstellar dust particles and shock waves are intimately associated with the process of star formation in two senses. First, both dust particles and shock waves owe their existence to stellar activity; second, they may both serve as agents which facilitate the formation of stars, shocks by triggering gravitational instabilities in the interstellar gas that they compress, and dust by shielding star-forming molecular clouds from the ionizing and dissociative effects of external UV radiation. By using Ly alpha observations as a probe of the dust content in diffuse gas at high redshift, we might hope to learn about the earliest epochs of star formation.

Neufeld, David A.↗
NASA NTRS1998

Inlet Performance Analysis Code Developed

The design characteristics of an inlet very much depend on whether the inlet is to be flown at subsonic, supersonic, or hypersonic speed. Whichever the case, the primary function of an inlet is to deliver free-stream air to the engine face at the highest stagnation pressure possible and with the lowest possible variation in both stagnation pressure and temperature. At high speeds, this is achieved by a system of oblique and/or normal shock waves, and possibly some isentropic compression. For both subsonic and supersonic flight, current design practice indicates that the inlet should deliver the air to the engine face at approximately Mach 0.45. As a result, even for flight in the high subsonic regime, the inlet must retard (or diffuse) the air substantially. Second, the design of an inlet is influenced largely by the compromise between high performance and low weight. This compromise involves tradeoffs between the mission requirements, flight trajectory, airframe aerodynamics, engine performance, and weight-all of which, in turn, influence each other. Therefore, to study the effects of some of these influential factors, the Propulsion System Analysis Office of the NASA Lewis Research Center developed the Inlet Performance Analysis Code (IPAC). This code uses oblique shock and Prandtl-Meyer expansion theory to predict inlet performance. It can be used to predict performance for a given inlet geometric design such as pitot, axisymmetric, and two-dimensional. IPAC also can be used to design preliminary inlet systems and to make subsequent performance analyses. It computes the total pressure, the recovery, the airflow, and the drag coefficients. The pressure recovery includes losses associated with normal and oblique shocks, internal and external friction, the sharp lip, and diffuser components. Flow rate includes captured, engine, spillage, bleed, and bypass flows. The aerodynamic drag calculation includes drags associated with spillage, cowl lip suction, wave, bleed, and bypass.

Jules, Kenol↗
NASA NTRS1993

Wall pressure fluctuations beneath swept shock wave/boundary layer interactions

An experimental research program providing basic knowledge and establishing a database on the fluctuating pressure loads produced on aerodynamic surfaces beneath 3D shock wave/boundary layer interactions is presented. A turbulent boundary layer on a flat plate is subjected to interactions with swept planar shock waves generated by sharp fins at angle of attack. Measurements are made for the first time in the aft areas of these interactions, showing fluctuating pressure levels as high as 160 dB.

Garg, S.↗
NASA NTRS1985

Collisionless shocks in the heliosphere: Reviews of current research

The present conference on heliospheric collisionless shocks considers such macrostructure-, microstructure-, and particle acceleration-related topics as interplanetary shock phenomena near and within 1 AU, as well as beyond, planetary bow shocks, shock formation and evolution in the solar atmosphere, MHD and gasdynamic theories for planetary bow waves, and subcritical collisionless shock waves. Also discussed are ion reflection, gyration, and dissipation at supercritical shocks, the numerical simulation of quasi-perpendicular collisionless shocks, electron distributions near collisionless shocks, the microtheory of collisionless shock current layers, plasma waves and instabilities, the electron foreshock, upstream suprathermal ions, and both diffusive and shock drift acceleration.

Tsurutani, Bruce T.↗
NASA NTRS2001

The CE/SE Method: a CFD Framework for the Challenges of the New Millennium

The space-time conservation element and solution element (CE/SE) method, which was originated and is continuously being developed at NASA Glenn Research Center, is a high-resolution, genuinely multidimensional and unstructured-mesh compatible numerical method for solving conservation laws. Since its inception in 1991, the CE/SE method has been used to obtain highly accurate numerical solutions for 1D, 2D and 3D flow problems involving shocks, contact discontinuities, acoustic waves, vortices, shock/acoustic waves/vortices interactions, shock/boundary layers interactions and chemical reactions. Without the aid of preconditioning or other special techniques, it has been applied to both steady and unsteady flows with speeds ranging from Mach number = 0.00288 to 10. In addition, the method has unique features that allow for (i) the use of very simple non-reflecting boundary conditions, and (ii) a unified wall boundary treatment for viscous and inviscid flows. The CE/SE method was developed with the conviction that, with a solid foundation in physics, a robust, coherent and accurate numerical framework can be built without involving overly complex mathematics. As a result, the method was constructed using a set of design principles that facilitate simplicity, robustness and accuracy. The most important among them are: (i) enforcing both local and global flux conservation in space and time, with flux evaluation at an interface being an integral part of the solution procedure and requiring no interpolation or extrapolation; (ii) unifying space and time and treating them as a single entity; and (iii) requiring that a numerical scheme be built from a nondissipative core scheme such that the numerical dissipation can be effectively controlled and, as a result, will not overwhelm the physical dissipation. Part I of the workshop will be devoted to a discussion of these principles along with a description of how the ID, 2D and 3D CE/SE schemes are constructed. In Part II, various applications of the CE/SE method, particularly those involving chemical reactions and acoustics, will be presented. The workshop will be concluded with a sketch of the future research directions.

Chang, Sin-Chung↗
NASA NTRS1995

Preliminary airborne measurements for the SR-71 sonic boom propagation experiment

SR-71 sonic boom signatures were measured to validate sonic boom propagation prediction codes. An SR-71 aircraft generated sonic booms from Mach 1.25 to Mach 1.6, at altitudes of 31,000 to 48,000 ft, and at various gross weights. An F-16XL aircraft measured the SR-71 near-field shock waves from close to the aircraft to more than 8,000 ft below, gathering 105 signatures. A YO-3A aircraft measured the SR-71 sonic booms from 21,000 to 38,000 feet below, recording 17 passes. The sonic booms at ground level and atmospheric data were recorded for each flight. Data analysis is underway. Preliminary results show that shock wave patterns and coalescence vary with SR-71 gross weight, Mach number, and altitude. For example, noncoalesced shock wave signatures were measured by the YO-3A at 21,000 ft below the SR-71 aircraft while at a low gross weight, Mach 1.25, and 31,000-ft altitude. This paper describes the design and execution of the flight research experiment. Instrumentation and flight maneuvers of the SR-71, F-16XL, and YO-3A aircraft and sample sonic boom signatures are included.

Haering, Edward A., Jr.↗
NASA NTRS1976

Aeroacoustics: Acoustic wave propagation; Aircraft noise prediction; Aeroacoustic instrumentation

The papers in this volume deal with recent research into acoustic-wave propagation through the atmosphere and progress in aeroacoustic instrumentation, facilities, and test techniques. Topics include the propagation of aircraft noise over long distances in the lower atmosphere, measured effects of turbulence on the rise time of a weak shock, sound scattering from atmospheric turbulence, saturation effects associated with sound propagation in a turbulent medium, and a computer model of the lightning-thunder process. Other papers discuss the development of a computer system for aircraft noise prediction; aircraft flyover noise measurements; and theories and methods for the prediction of ground effects on aircraft noise propagation, for the prediction of airframe aerodynamic noise, for turbine noise prediction, and for combustion noise prediction. Attention is also given to the use of Hartmann generators as sources of high-intensity sound in a large absorption flow-duct facility, an outdoor jet noise facility, factors in the design and performance of free-jet acoustic wind tunnels, and the use of a laser shadowgraph for jet noise diagnosis.

Schwartz, I. R.↗
NASA NTRS1982

Current status of IMS plasma wave research

The present investigation is concerned with a review of the status of magnetospheric plasma wave science as a result of the International Magnetospheric Study (IMS). The presence of an international effort has supported the development and completion of the numerous magnetospheric science spacecraft launched during the IMS, including GEOS, ISEE, and EXOS B. Ground-based VLF observations are considered along with coordinated ground-based and satellite observations. During the IMS, plasma wave research using satellite data has covered a wide range of subjects. Attention is given to magnetospheric electrostatic emissions, magnetospheric electromagnetic plasma waves, continuum radiation, auroral kilometric radiation, auroral zone plasma waves, plasma waves in the magnetosheath and near the mangetopause, and plasma waves at the bow shock.

Anderson, R. R.↗
NASA NTRS1986

Evaluation of diffuse-illumination holographic cinematography in a flutter cascade

Since 1979, the Lewis Research Center has examined holographic cinematography for three-dimensional flow visualization. The Nd:YAG lasers used were Q-switched, double-pulsed, and frequency-doubled, operating at 20 pulses per second. The primary subjects for flow visualization were the shock waves produced in two flutter cascades. Flow visualization was by diffuse-illumination, double-exposure, and holographic interferometry. The performances of the lasers, holography, and diffuse-illumination interferometry are evaluated in single-window wind tunnels. The fringe-contrast factor is used to evaluate the results. The effects of turbulence on shock-wave visualization in a transonic flow are discussed. The depth of field for visualization of a turbulent structure is demonstrated to be a measure of the relative density and scale of that structure. Other items discussed are the holographic emulsion, tests of coherence and polarization, effects of windows and diffusers, hologram bleaching, laser configurations, influence and handling of specular reflections, modes of fringe localization, noise sources, and coherence requirements as a function of the pulse energy. Holography and diffuse illumination interferometry are also reviewed.

Decker, A. J.↗
NASA NTRS2012

Recent Progress in Entry Radiation Measurements in the NASA Ames Electric ARC Shock Tube Facility

The Electric Arc Shock Tube (EAST) at NASA Ames Research Center is NASA's only working shock tube capable of obtaining conditions representative of entry in a multitude of planetary atmospheres. The facility is capable of mapping spectroscopic signatures of a wide range of planetary entries from the Vacuum Ultraviolet through Mid-Wave Infrared (120-5500 nm). This paper summarizes the tests performed in EAST for Earth, Mars and Venus entries since 2008, then focuses on a specific test case for CO2/N2 mixtures. In particular, the paper will focus on providing information for the proper interpretation of the EAST data.

Cruden, Brett A.↗
NASA NTRS2022

The Hyper-inflation Stage in the Coronal Mass Ejection Formation: A Missing Link That Connects Flares, Coronal Mass Ejections, and Shocks in the Low Corona

We analyze the formation and three-dimensional (3D) evolution of two coronal mass ejections (CMEs) and their associated waves in the low corona via a detailed multi-viewpoint analysis of extreme-ultraviolet observations. We analyze the kinematics in the radial and lateral directions and identify three stages in the early evolution of the CME: (1) a hyper-inflation stage, when the CME laterally expands at speeds of ∼1000 km/s, followed by (2) a shorter and slower expansion stage of a few minutes and ending with (3) a self-similar phase that carries the CME into the middle corona. The first two stages coincide with the impulsive phase of the accompanying flare, the formation and separation of an EUV wave from the CME, and the start of the metric type II radio burst. Our 3D analysis suggests that the hyper-inflation phase may be a crucial stage in the CME formation with wide-ranging implications for solar eruption research. It likely represents the formation stage of the magnetic structure that is eventually ejected into the corona, as the white-light CME. It appears to be driven by the injection of poloidal flux into the ejecting magnetic structure, which leads to the lateral (primarily) growth of the magnetic flux rope. The rapid growth results in the creation of EUV waves and eventually shocks at the CME flanks that are detected as metric type II radio bursts. In other words, the hyper-inflation stage in the early CME evolution may be the "missing" link between CMEs, flares, and coronal shocks.

Solar coronal waves↗
NASA NTRS1986

Three dimensional unsteady aerodynamics and aeroelastic response of advanced turboprops

A method for the prediction of steady and unsteady aerodynamic loads and aeroelastic response of advanced turboprops is presented. The aerodynamic analysis uses three dimensional unsteady linearized compressible flow theory to compute the blade pressure distribution. The aeroelastic analysis is based on a normal mode representation of the structure. The method is applicable to both conventional and advanced turbo-prop configurations, provided that blade stall and transonic shock waves are not important factors. Aerodynamic results are presented which validate the model in various limits by comparisons to alternative theories and experimental data. Finally, results of a stability analysis of an advanced turboprop are given, with comparisons to measurements made at NASA Lewis Research Center.

Williams, M. H.↗
NASA NTRS1992

Time dependent heat transfer rates in high Reynolds number hypersonic flowfields

Time dependent heat transfer rates have been calculated from time dependent temperature measurements in the vicinity of shock-wave boundary-layer interactions due to conical compression ramps on an axisymmetric body. The basic model is a cylindrical body with a 10 degree conical nose. Four conical ramps, 20, 25, 30, and 35 degrees serve as shock wave generators. Flowfield surveys have been made in the vicinity of the conical ramp vertex, the separation point, and the reattachment point. A significant effort was made to characterize the natural frequencies and relative powers of the resulting fluctuations in heat transfer rates. This research effort, sponsored jointly by NASA and the Air Force, was conducted in the Air Force Flight Dynamics Directorate High Reynolds Facility. The nominal freestream Mach number was 6, and the freestream Reynolds numbers ranged from 2.2 million/ft to 30.0 million/ft. Experimental results quantify temperature response and the resulting heat transfer rates as a function of ramp angle and Reynolds number. The temperature response within the flowfield appears to be steady-state for all compression ramp angles and all Reynolds numbers, and hence, the heat transfer rates appear to be steady-state.

Flanagan, Michael J.↗
NASA NTRS2000

Screech Noise Generation From Supersonic Underexpanded Jets Investigated

Many supersonic military aircraft and some of the modern civilian aircraft (such as the Boeing 777) produce shock-associated noise. This noise is generated from the jet engine plume when the engine nozzle is operated beyond the subsonic operation limit to gain additional thrust. At these underexpanded conditions, a series of shock waves appear in the plume. The turbulent vortices present in the jet interact with the shock waves and produce the additional shock-associated noise. Screech belongs to this noise category, where sound is generated in single or multiple pure tones. The high dynamic load associated with screech can damage the tailplane. One purpose of this study at the NASA Glenn Research Center at Lewis Field was to provide an accurate data base for validating various computational fluid dynamics (CFD) codes. These codes will be used to predict the frequency and amplitude of screech tones. A second purpose was to advance the fundamental physical understanding of how shock-turbulence interactions generate sound. Previously, experiments on shock-turbulence interaction were impossible to perform because no suitable technique was available. As one part of this program, an optical Rayleigh-scattering measurement technique was devised to overcome this difficulty.

Panda, Jayanta↗
NASA NTRS2015

Flexible Micropost Arrays for Shear Stress Measurement

Increased fuel costs, heightened environmental protection requirements, and noise abatement continue to place drag reduction at the forefront of aerospace research priorities. Unfortunately, shortfalls still exist in the fundamental understanding of boundary-layer airflow over aerodynamic surfaces, especially regarding drag arising from skin friction. For example, there is insufficient availability of instrumentation to adequately characterize complex flows with strong pressure gradients, heat transfer, wall mass flux, three-dimensionality, separation, shock waves, and transient phenomena. One example is the acoustic liner efficacy on aircraft engine nacelle walls. Active measurement of shear stress in boundary layer airflow would enable a better understanding of how aircraft structure and flight dynamics affect skin friction. Current shear stress measurement techniques suffer from reliability, complexity, and airflow disruption, thereby compromising resultant shear stress data. The state-of-the-art for shear stress sensing uses indirect or direct measurement techniques. Indirect measurements (e.g., hot-wire, heat flux gages, oil interferometry, laser Doppler anemometry, small scale pressure drag surfaces, i.e., fences) require intricate knowledge of the studied flow, restrictive instrument arrangements, large surface areas, flow disruption, or seeding material; with smaller, higher bandwidth probes under development. Direct measurements involve strain displacement of a sensor element and require no prior knowledge of the flow. Unfortunately, conventional "floating" recessed components for direct measurements are mm to cm in size. Whispering gallery mode devices and Fiber Bragg Gratings are examples of recent additions to this type of sensor with much smaller (m) sensor components. Direct detection techniques are often single point measurements and difficult to calibrate and implement in wind tunnel experiments. In addition, the wiring, packaging, and installation of delicate micro-electromechanical devices impede the use of most direct shear sensors. Similarly, the cavity required for sensing element displacement is sensitive to particulate obstruction. This work was focused on developing a shear stress sensor for use in subsonic wind tunnel test facilities applicable to an array of test configurations. The non-displacement shear sensors described here have minimal packaging requirements resulting in minimal or no disturbance of boundary layer flow. Compared to previous concepts, device installation could be simple with reduced cost and down-time. The novelty lies in the creation of low profile (nanoscale to 100 μm) micropost arrays that stay within the viscous sub-layer of the airflow. Aerodynamic forces, which are related to the surface shear stress, cause post deflection and optical property changes. Ultimately, a reliable, accurate shear stress sensor that does not disrupt the airflow has the potential to provide high value data for flow physics researchers, aerodynamicists, and aircraft manufacturers leading to greater flight efficiency arising from more in-depth knowledge on how aircraft design impacts near surface properties.

Wohl, Christopher J.↗