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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 343 records · Page 19
All sourcesNASA NTRSDOE OSTI
DOE OSTI2023

Plasma Pressure Distribution of Ions and Electrons in the Inner Magnetosphere During CIR Driven Storms Observed During Arase Era

Here, using Arase observations of the inner magnetosphere during 26 CIR-driven geomagnetic storms with minimum Sym-H between –33 and –86 nT, we investigated ring current pressure development of ions (H + , He + , O + ) and electron during prestorm, main, early recovery and late recovery phases as a function of L-shell and magnetic local time. It is found that during the main and early recovery phase of the storms the ion pressure is asymmetric in the inner magnetosphere, leading to a strong partial ring current. The ion pressure becomes symmetric during the late recovery phase. H + ions with energies of ~20–50 keV and ~50–100 keV contribute more to the ring current pressure during the main phase and early/late recovery phase, respectively. O + ions with energies of ~10–20 keV contribute significantly during main and early recovery phase. These are consistent with previous studies. The electron pressure was found to be asymmetric during the main, early recovery and late recovery phase. The electron pressure peaks from midnight to the dawn sector. Electrons with energy of <50 keV contribute to the ring current pressure during the main and early recovery phase of the storms. Overall, the electron contribution to the total ring current is found to be ~11% during the main and early recovery phases. However, the electron contribution is found to be significant (~22%) in the 03–09 MLT sector during the main and early recovery phase. The results indicate an important role of electrons in the ring current build up.

79 ASTRONOMY AND ASTROPHYSICS↗
NASA NTRS1923

Pressure Distribution over the Rudder and Fin of an Airplane in Flight

This investigation was carried out by the National Advisory Committee for Aeronautics at the Langley Memorial Aeronautical Laboratory for the purpose of determining the loads which occur on the vertical tail surfaces in flight. The method consisted in measuring the pressures simultaneously at 28 points on the rudder and fin by means of a recording multiple manometer. The results show that the maximum load encountered in stunting was 7 pounds per square foot on the rudder and fin, and it is probable that this might rise to 10 pounds per square foot in a violent barrel roll; but in steady flight the average loads do not exceed 0.6 pound per square foot. The maximum load on the rudder and fin may occur at the same instant as the maximum load on the horizontal tail surfaces and the maximum normal acceleration. The torsional moment about the axis of the fuselage due to the rudder and fin may rise as high as 250 foot-pound. The results obtained from this investigation have a direct application to the design of the rudder, fin, and fuselage.

Norton, F H↗
NASA NTRS1939

Effect of transition in cross-sectional shape on the development of the velocity and pressure distribution of turbulent flow in pipes

With regard to the change in shape of the cross section while the area remains constant, no investigation results are as yet available. Such an investigation will be the subject of the present paper. For this purpose it is necessary to consider the velocity and pressure relations over each entire cross section so that we are confronted with a three-dimensional problem.

Mayer, Edwin↗
NASA NTRS1921

The Pressure Distribution over the Horizontal Tail Surfaces of an Airplane II

This investigation was undertaken to determine whether the results obtained upon model tail surfaces can be used to accurately predict loads upon the full-sized tail; and also to find the distribution of load when large elevator angles are used, as the loads from such angles can not be obtained readily in free flight. The method consisted in using a metal horizontal tail surface inside of which small air passages, connecting with a series of holes in the surface, led the pressure off from the tail in rubber tubes. In this way the pressure at each of these holes was measured by a manometer at several angles of attack and several to the loading under similar conditions in the full-sized airplane and the manner of distribution is quite similar in the two cases when there is no slip stream.

Norton, F H↗
NASA NTRS1931

The Pressure Distribution over the Wings and Tail Surfaces of a PW-9 Pursuit Airplane in Flight

This report presents the results of an investigation to determine (1) the magnitude and distribution of aerodynamic loads over the wings and tail surfaces of a pursuit-type airplane in the maneuvers likely to impose critical loads on the various subassemblies of the airplane structure. (2) To study the phenomenon of center of pressure movement and normal force coefficient variation in accelerated flight, and (3) to measure the normal accelerations at the center of gravity, wing-tip, and tail, in order to determine the nature of the inertia forces acting simultaneously with the critical aerodynamic loads. The results obtained throw light on a number of important questions involving structural design. Some of the more interesting results are discussed in some detail, but in general the report is for the purpose of making this collection of airplane-load data obtained in flight available to those interested in airplane structures.

Rhode, Richard↗
NASA NTRS1930

Flight Tests on U.S.S. Los Angeles. Part I : Full Scale Pressure Distribution Investigation

The primary purpose of this investigation was to obtain simultaneous data on the loads and stress experience in flight by the U. S. S. Los Angeles which could be used in rigid airship structure design. A secondary object of the investigation was to determine the turning and drag characteristics of the airship. The aerodynamic loading was obtained by measuring the pressure at 95 locations on the tail surfaces, 54 on the hull, and 5 on the passenger car. These measurements were made during a series of maneuvers consisting of turns and reversals in smooth air and during a cruise in rough air which was just short of squall proportions. The results of the pressure measurements on the hull indicate that the forces on the forebody of an airship are relatively small. The tail surface measurements show conclusively that the forces caused by gusts are much greater than those caused by horizontal maneuvers. In this investigation the tail surface loadings caused by gusts closely approached the designed loads of the tail structure. The turning and drag characteristics will be reported in separate reports.

De France, S J↗
NASA NTRS1973

Effect of blockage ratio on drag and pressure distributions for bodies of revolution at transonic speeds

Experimental data were obtained in two wind tunnels for 13 models over a Mach number range from 0.70 to 1.02. Effects of increasing test-section blockage ratio in the transonic region near a Mach number of 1.0 included change in the shape of the drag curves, premature drag creep, delayed drag divergence, and a positive increment of pressures on the model afterbodies. Effects of wall interference were apparent in the data even for a change in blockage ratio from a very low 0.000343 to an even lower 0.000170. Therefore, models having values of blockage ratio of 0.0003 - an order of magnitude below the previously considered safe value of 0.0050 - had significant errors in the drag-coefficient values obtained at speeds near a Mach number of 1.0. Furthermore, the flow relief afforded by slots or perforations in test-section walls - designed according to previously accepted criteria for interference-free subsonic flow - does not appear to be sufficient to avoid significant interference of the walls with the model flow field for Mach numbers very close to 1.0.

Couch, L. M.↗
NASA NTRS1983

Analysis of surface pressure distributions on two elliptic missile configurations

The state-of-the-art methods for predicting missile aerodynamic characteristics do not accurately predict the loads of missile configurations with bodies of elliptic cross section. An investigation of this problem found significant nonlinear flow disturbance on the windward surface of a 3:1 elliptic body at Mach 2.50 in addition to the nonlinear vortical flows which develop on the leeside. A nonlinear full-potential flow method (NCOREL) was found to provide extremely accurate pressure estimates for attached-flow conditions and the vortex prediction method contained in the state-of-the-art method (NOSEVTX) was shown to accurately calculate body vortices and leeside pressures.

Allen, J. M.↗
NASA NTRS1975

An investigation in the NASA MSFC 14-inch trisonic wind tunnel to determine the pressure distribution over the components of a 0.004 scale version of the Rockwell MCR 0074 baseline shuttle ascent configuration (IA32F), volume 1

Data were obtained for Mach numbers from 0.6 to 3.48, angles of attack from -10 to 10 degrees, and angles of sideslip from -10 to 10 degrees at zero angle of attack. Also, -4 and 4 degrees sideslip were run for an angle of attack of -5 and 5 degrees. Aerodynamic configurations of the solid rocket motors, external tank, and orbiter are shown. Graphs of plotted pressure data (pressure coefficients) for the external tank and solid rocket motors are given. A description of the test facility is included.

Ramsey, P. E.↗
NASA NTRS1947

Preliminary Results of Altitude-Wind-Tunnel Investigation of X24C-4B Turbojet Engine. I - Pressure and Temperature Distributions

Pressures and temperatures throughout the X24C-4B turbojet engine are presented in both tabular and graphical forms to show the effect of altitude, flight Mach number, and engine speed on the internal operation of the engine. These data were obtained in the NACA Cleveland altitude wind tunnel at simulated altitudes from 5000 to 45,000 feet, simulated flight Mach numbers from 0.25 to 1.08, and engine speeds from 4000 to 12,500 rpm. Location and detail drawings of the instrumentation installed at seven survey stations in the engine are shown. Application of generalization factors to pressures and temperatures at each measuring station for the range of altitudes investigated showed that the data did not generalize above an altitude of 25,000 feet. Total-pressure distribution at the compressor outlet varied only with change in engine speed. At altitudes above 35,000 feet and engine speeds above 11,000 rpm, the peak temperature at the turbine-outlet annulus moved inward toward the root of the blade, which is undesirable from blade-stress considerations. The temperature levels at the turbine outlet and the exhaust-nozzle outlet were lowered as the Mach number was increased. The static-pressure measurements obtained at each stator stage of the compressor showed a pressure drop through the inlet guide vanes and the first-stage rotor at high engine speeds. The average values measured by the manufacturer's instrumentation werein close agreement with the average values obtained with NACA instrumentation.

Prince, William R.↗
NASA NTRS1984

Computational-experimental pressure distributions on a transonic, low-aspect-ratio wing

A generic, transonic, supercritical, low aspect-ratio wing was tested at design incidence of 5 deg at Mach numbers from 0.25 to 0.96. Oil-flow studies at the design Mach number of 0.85 showed local-flow separation, which, in retrospect, might have been avoided. At Mach 0.82 with unseparated flow, the surface-flow angles were less than 10 deg. Predictions with the FLO22 transonic potential code are good. Lift interference is strong without tunnel-wall suction. Evidence from this study shows that wings that are optimized for mild shock waves and pressure-recovery gradients generally have small, three-dimensional flow at conditions for unseparated flow.

Keener, E. R.↗
NASA NTRS1987

Experimental evaluation of blockage ratio and plenum evacuation system flow effects on pressure distribution for bodies of revolution in 0.1 scale model test section of NASA Lewis Research Center's proposed altitude wind tunnel

An experimental investigation was conducted in the slotted test section of the 0.1-scale model of the proposed Altitude Wind Tunnel to evaluate wall interference effects at tunnel Mach numbers from 0.70 to 0.95 on bodies of revolution with blockage rates of 0.43, 3, 6, and 12 percent. The amount of flow that had to be removed from the plenum chamber (which surrounded the slotted test section) by the plenum evacuation system (PES) to eliminate wall interference effects was determined. The effectiveness of tunnel reentry flaps in removing flow from the plenum chamber was examined. The 0.43-percent blockage model was the only one free of wall interference effects with no PES flow. Surface pressures on the forward part of the other models were greater than interference-free results and were not influenced by PES flow. Interference-free results were achieved on the aft part of the 3- and 6-percent blockage models with the proper amount of PES flow. The required PES flow was substantially reduced by opening the reentry flaps.

Burley, Richard R.↗
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 NTRS1961

Normal-Force and Hinge-Moment Characteristics at Transonic Speeds of Flap-Type Ailerons at Three Spanwise Locations on a 4-Percent-Thick Sweptback-Wing-Body Model and Pressure-Distribution Measurements on an Inboard Aileron

An investigation has been conducted at the Langley 16-foot transonic tunnel to determine the loading characteristics of flap-type ailerons located at inboard, midspan, and outboard positions on a 45 deg. sweptback-wing-body combination. Aileron normal-force and hinge-moment data have been obtained at Mach numbers from 0.80 t o 1.03, at angles of attack up to about 27 deg., and at aileron deflections between approximately -15 deg. and 15 deg. Results of the investigation indicate that the loading over the ailerons was established by the wing-flow characteristics, and the loading shapes were irregular in the transonic speed range. The spanwise location of the aileron had little effect on the values of the slope of the curves of hinge-moment coefficient against aileron deflection, but the inboard aileron had the greatest value of the slope of the curves of hinge-moment coefficient against angle of attack and the outboard aileron had the least. Hinge-moment and aileron normal-force data taken with strain-gage instrumentation are compared with data obtained with pressure measurements.

Runckel, Jack F.↗