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At least 649 records · Page 36
All sourcesNASA NTRSDOE OSTI
NASA NTRS1959

Experimental Investigation of Several Copper and Beryllium Hemispherical Models in Air at Stagnation Temperatures of 2,000 to 3,600 F

As part of an investigation by the National Aeronautics and Space Administration to determine the resistance to heating of various materials when used as a heat sink for hypersonic airframes, hemispherical nose-shape models of beryllium and copper have been tested in a Mach number 4 hot-air jet at stagnation temperatures of 2,000 F to 3,600 F and Reynolds numbers of 1.88 x 10(exp 6) to 2.93 x 10(exp 6). The experimental results of heating on the nose of the beryllium models agreed reasonably well with theoretical results, whereas heating on the nose of the copper models was almost twice that predicted by theory. Heating of the cylindrical wall behind the hemisphere agreed fairly well with that predicted by theory at lower temperatures. Beryllium produced a thin protective oxide when heated to its melting point with no tendency to ignite before melting. Copper produced a somewhat heavier layer of oxide upon heating, and ignited when heated to near its melting point. These tests indicate that beryllium is superior to copper as a heat-sink material because it absorbs more heat per unit weight, has greater resistance to oxidation in heated air, and does not ignite when heated in air up to its melting temperature.

Trout, Otto F., Jr.↗
NASA NTRS1959

Some Basic Aspects of Magnetohydrodynamic Boundary-Layer Flows

An appraisal is made of existing solutions of magnetohydrodynamic boundary-layer equations for stagnation flow and flat-plate flow, and some new solutions are given. Since an exact solution of the equations of magnetohydrodynamics requires complicated simultaneous treatment of the equations of fluid flow and of electromagnetism, certain simplifying assumptions are generally introduced. The full implications of these assumptions have not been brought out properly in several recent papers. It is shown in the present report that for the particular law of deformation which the magnetic lines are assumed to follow in these papers a magnet situated inside the missile nose would not be able to take up any drag forces; to do so it would have to be placed in the flow away from the nose. It is also shown that for the assumption that potential flow is maintained outside the boundary layer, the deformation of the magnetic lines is restricted to small values. The literature contains serious disagreements with regard to reductions in heat-transfer rates due to magnetic action at the nose of a missile, and these disagreements are shown to be mainly due to different interpretations of reentry conditions rather than more complicated effects. In the present paper the magnetohydrodynamic boundary-layer equation is also expressed in a simple form that is especially convenient for physical interpretation. This is done by adapting methods to magnetic forces which in the past have been used for forces due to gravitational or centrifugal action. The simplified approach is used to develop some new solutions of boundary-layer flow and to reinterpret certain solutions existing in the literature. An asymptotic boundary-layer solution representing a fixed velocity profile and shear is found. Special emphasis is put on estimating skin friction and heat-transfer rates.

Hess, Robert V.↗
NASA NTRS1959

Free-Flight Test of a Technique for Inflating an NASA 12-Foot-Diameter Sphere at High Altitudes

A free-flight test has been conducted to check a technique for inflating an NASA 12-foot-diameter inflatable sphere at high altitudes. Flight records indicated that the nose section was successfully separated from the booster rocket, that the sphere was ejected, and that the nose section was jettisoned from the fully inflated sphere. On the basis of preflight and flight records, it is believed that the sphere was fully inflated by the time of peak altitude (239,000 feet). Calculations showed that during descent, jettison of the nose section occurred above an altitude of 150,000 feet. The inflatable sphere was estimated to start to deform during descent at an altitude of about 120,000 feet.

Kehlet, Alan B.↗
NASA NTRS1961

Measurements of Aerodynamic Heat Transfer and Boundary-Layer Transition on a 15 deg. Cone in Free Flight at Supersonic Mach Numbers up to 5.2

Measurements of aerodynamic heat transfer have been made at several stations on the 15 deg total-angle conical nose of a rocket-propelled model in free flight at Mach numbers up to 5.2. Data are presented for a range of local Mach number just outside the boundary layer from 1.40 to 4.65 and a range of local Reynolds number from 3.8 x 10(exp 6) to 46.5 x 10(exp 6), based on length from the nose tip to a measurement station. Laminar, transitional, and turbulent heat-transfer coefficients were measured. The laminar data were in agreement with laminar theory for cones, and the turbulent data agreed well with turbulent theory for cones using Reynolds number based on length from the nose tip. At a nearly constant ratio of wall to local static temperature of 1.2 the Reynolds number of transition increased from 14 x 10(exp 6) to 30 x 10(exp 6) as Mach number increased from 1.4 to 2.9 and then decreased to 17 x 10(exp 6) as Mach number increased to 3.7. At Mach numbers near 3.5, transition Reynolds numbers appeared to be independent of skin temperature at skin temperatures very cold with respect to adiabatic wall temperature. The transition Reynolds number was 17.7 x 10(exp 6) at a condition of Mach number and ratio of wall to local static temperature near that for which three-dimensional disturbance theory has been evaluated and has predicted laminar boundary-layer stability to very high Reynolds numbers (approximately 10(exp 12)).

Rumsey, Charles B.↗
NASA NTRS1999

Boom Softening and Nacelle Integration on an Arrow-Wing High-Speed Civil Transport Concept

During the last cycle of concept design and wind-tunnel testing, the goal of the low-boom- shaped HSCT concepts (the B-935, the LB-16, and the LB- 1 8) was to meet mission requirements and generate shaped, ground-level pressure signatures with nose shock strengths of 1.0 psf or less. The wind-tunnel tests of these concepts produced results that were partially successful and encouraging although not fully up to expectations. In spite of this, however, these conceptual designs were overly optimistic and not acceptable because: the wing planforms had excessive area; the wing structural aspect ratio was too high; one concept had aft-fuselage rather than under-the-wing engines; and the gross takeoff weights were unrealistically low because of engines that were early, high-tech versions of later, revised, more-realistic engines. The need for reducing the ground-level overpressure shock strengths still existed; a need to be met within more restrictive guidelines of mission performance and gross takeoff weight limitations. Therefore, it was decided that the next conceptual design cycle would focus on decreased nose shock strengths, "boom softening," in the signatures of the Boeing and the McDonnell Douglas baseline concepts rather than low-boom concepts with shaped-signature designs. Overly-optimistic results were not the only problem with these low-sonic-boom concepts. Papers given at the 1994 Sonic-Boom Workshop had demonstrated that the problem of successful nacelle integration on HSCT concepts had only been partially solved. Wind-tunnel pressure signature data, from the HSCT-11B (a.k.a. the LB-18) wind-tunnel model, showed that the Langley HSCT design and analysis method had been successful in reducing the nacelle-volume disturbances in the flow field. This was due.to the engine nacelles mounted behind the wing trailing-edge on the aft fuselage so that no nacelle-wing interference-lift flow-field disturbances were generated. While acceptable from a sonic-boom research point of view, this concept was unacceptable from several practical and structural considerations. Preliminary wind-tunnel pressure signature data from the LB-16 wind-tunnel model, which had the engine nacelles mounted under the wings (the usual location), indicated that the application of the Langley nacelle-integration method had been only partially successful in the reduction of the nacelle-volume with nacelle-wing interference-lift pressure disturbances. So, "boom softening" had to also address the task of successful integration of the engine nacelles, with the engines in the required under-the-wing location. Unless this problem was solved, low-sonic-boom and low-drag modifications to the wing planform, the airfoil shape, and the fuselage longitudinal area distribution could be nullified if the nacelle disturbances added increments to the nose-shock strengths that were removed through component tailoring. In this paper, an arrow-wing boom-softened HSC7 concept which incorporated modifications to a baseline McDonnell Douglas concept is discussed. The analysis of the concept's characteristics will include estimates of weight, center of gravity, takeoff field length, mission range, and predictions of its ground-level sonic-boom pressure signature. Additional modifications which enhanced the softened-boom performance of this concept are also described as well as estimates of the performance penalties induced by these modifications.

Mack, Robert J.↗
NASA NTRS1996

Control of Flow Separation Using Adaptive Airfoils

A novel way of controlling flow separation is reported. The approach involves using an adaptive airfoil geometry that changes its leading edge shape to adjust to the instantaneous flow at high angles of attack such that the flow over it remains attached. In particular, a baseline NACA 0012 airfoil, whose leading edge curvature could be changed dynamically by 400% was tested under quasi-steady compressible flow conditions. A mechanical drive system was used to produce a rounded leading edge to reduce the strong local flow acceleration around its nose and thus reduce the strong adverse pressure gradient that follows such a rapid acceleration. Tests in steady flow showed that at M = 0.3, the flow separated at about 14 deg. angle of attack for the NACA 0012 profile but could be kept attached up to an angle of about 18 deg by changing the nose curvature. No significant hysteresis effects were observed; the flow could be made to reattach from its separated state at high angles by changing the leading edge curvature. Interestingly, the flow over a nearly semicircular nosed airfoil was separated even at low angles.

Chandrasekhara, M. S.↗
NASA NTRS1947

Results of Tests to Determine the Effect of a Conical Windshield on the Drag of a Bluff Body at Supersonic Speeds

Tests to evaluate the effect of a conical windshield on the drag of a bluff body at supersonic speeds were performed for the following configurations: a sharp nose fuselage with stabilizing fins,a blunt nose fuselage with a hemispherical shape, and a blunt nose fuselage with a conical point. Results of the drag coeeficient are described at Mach 1.0 and the greatest Mach number of 1.37.

Alexander, Sidney R.↗
NASA NTRS1961

Low-Subsonic-Speed Static Stability of Right-Triangular-Pyramid and Half-Cone Lifting Reentry Configurations

An investigation has been made to determine the low-subsonic-speed static stability characteristics of several right-triangular-pyramid and half-cone configurations. Also studied were the effects of various modifications, such as base extensions, nose shape, nose incidence, and ridge-line shape. The investigation showed that, in general, the models had satisfactory longitudinal and lateral stability. The basic pyramid model and the conical ridge-line model with or without a rounded nose had almost identical longitudinal and lateral stability characteristics and lift-drag ratios. The lift-drag ratios of the cylindrical ridge-line and half-cone models were considerably lower than those of the conical ridge-line model. The addition of a 20 degree boattail to the models increased the lift-drag ratios but decreased the directional stability, whereas a streamwise base extension was more effective in increasing the lift-drag ratios and increased the directional stability.

Ware, George M.↗
NASA NTRS2002

Effects of lesions of the caudal cerebellar vermis on cardiovascular regulation in awake cats

The vestibular system is known to participate in cardiovascular regulation during movement and postural alterations. The present study considered whether lesions of two regions of the posterior cerebellar vermis (the nodulus and uvula) that provide inputs to vestibular nucleus regions that affect control of blood pressure would alter cardiovascular responses during changes in posture. Blood pressure and heart rate were monitored in awake cats during nose-up tilts up to 60 degrees in amplitude before and following aspiration lesions of the nodulus or uvula; in most animals, cardiovascular responses were also recorded following the subsequent removal of vestibular inputs. Lesions of the nodulus or uvula did not affect baseline blood pressure or heart rate, although cardiovascular responses during nose-up tilts were altered. Increases in heart rate that typically occurred during 60 degrees nose-up tilt were attenuated in all three animals with lesions affecting both dorsal and ventral portions of the uvula; in contrast, the heart rate responses were augmented in the two animals with lesions mainly confined to the nodulus. Furthermore, following subsequent removal of vestibular inputs, uvulectomized animals, but not those with nodulus lesions, experienced more severe orthostatic hypotension than has previously been reported in cerebellum-intact animals with bilateral labyrinthectomies. These data suggest that the cerebellar nodulus and uvula modulate vestibulo-cardiovascular responses, although the two regions play different roles in cardiovascular regulation.

NASA Discipline Neuroscience↗
NASA NTRS1999

Effects of static orientation upon human optokinetic afternystagmus

"Normal" human subjects were placed in a series of 5 static orientations with respect to gravity and were asked to view an optokinetic display moving at a constant angular velocity. The axis of rotation coincided with the subject's rostro-caudal axis and produced horizontal optokinetic nystagmus and afternystagmus. Wall (1) previously reported that these optokinetic afternystagmus responses were not well characterized by parametric fits to slow component velocity. The response for nose-up, however, was larger than for nose-down. This suggested that the horizontal eye movements measured during optokinetic stimulation might include an induced linear VOR component as presented in the body of this paper. To investigate this hypothesis, another analysis of these data has been made using cumulative slow component eye position. Some subjects' responses had reversals in afternystagmus direction. These reversals were "filled in" by a zero slow component velocity. This method of analysis gives a much more consistent result across subjects and shows that, on average, responses from the nose-down horizontal (prone) orientation are greatly reduced (p < 0.05) compared to other horizontal and vertical orientations. Average responses are compared to responses predicted by a model previously used to predict successfully the responses to post-rotatory nystagmus after earth horizontal axis rotation. Ten of 11 subjects had larger responses in their supine than their prone orientation. Application of horizontal axis optokinetic afternystagmus for clinical otolith function testing, and implications for altered gravity experiments are discussed.

Non-NASA Center↗
NASA NTRS2005

Techniques for Transition and Surface Temperature Measurements on Projectiles at Hypersonic Velocities- A Status Report

A research effort to advance techniques for determining transition location and measuring surface temperatures on graphite-tipped projectiles in hypersonic flight in a ballistic range is described. Projectiles were launched at muzzle velocities of approx. 4.7 km/sec into air at pressures of 190-570 Torr. Most launches had maximum pitch and yaw angles of 2.5-5 degrees at pressures of 380 Torr and above and 3-6 degrees at pressures of 190-380 Torr. Arcjet-ablated and machined, bead-blasted projectiles were launched; special cleaning techniques had to be developed for the latter class of projectiles. Improved methods of using helium to remove the radiating gas cap around the projectiles at the locations where ICCD (intensified charge coupled device) camera images were taken are described. Two ICCD cameras with a wavelength sensitivity range of 480-870 nm have been used in this program for several years to obtain images. In the last year, a third camera, with a wavelength sensitivity range of 1.5-5 microns [in the infrared (IR)], has been added. ICCD and IR camera images of hemisphere nose and 70 degree sphere-cone nose projectiles at velocities of 4.0-4.7 km/sec are presented. The ICCD images clearly show a region of steep temperature rise indicative of transition from laminar to turbulent flow. Preliminary temperature data for the graphite projectile noses are presented.

Wilder, M. C.↗
NASA NTRS1937

Stability of Castering Wheels for Aircraft Landing Gears, Special Report

In many installations of castering rubber-tired wheels there is a tendency for the wheel to oscillate violently about the spindle axis. This phenomenon, popularly called 'shimmy,' has occurred in some airplane tail wheels and has been corrected in two ways: first by the application of friction in the spindles of the tail wheels; and, second, by locking the wheels while taxiing at high speeds. Shimmy is common with the large wheels used as nose wheels in tricycle landing gears and, since it is impossible to lock the wheels, friction in the nose-wheel spindle has been the sole means of correction. Because the nose wheel is larger than the conventional tail wheel and usually carries a greater load, the larger amounts of spindle friction necessary to prevent shimmy are objectionable. the present paper presents a theoretical and experimental study of the problem of the stability of castering wheels for airplane landing gears. On the basis of simplified assumptions induced from experimental observations, a theoretical study has been made of the shimmy of castering wheels. The theory is based on the discovery of a phenomenon called 'kinematic shimmy' and is compared quantitatively with the results of model experiments. Experimental checks, using a model having low-pressure tires, are reported and the applicability of the results to full scale is discussed. Theoretical methods of estimating the spindle viscous damping and spindle solid friction necessary to avoid shimmy - lateral freedom - is introduced.

Kantrowitz, Arthur↗
NASA NTRS2013

Analysis and Results from a Flush Airdata Sensing (FADS) System in Close Proximity to Firing Rocket Nozzles

This presentation presents information regarding the nose-cap flush airdata sensing (FADS) system on Orion's Pad Abort 1 (PA-1) vehicle. The purpose of the nose-cap FADS system was to test whether or not useful data could be obtained from a FADS system if it was placed in close proximity to firing rockets nozzles like the attitude control motor (ACM) nozzles on the PA-1 launch abort system (LAS). The nose-cap FADS systems use pressure measurements from a series of pressure ports which are arranged in a cruciform pattern and flush with the surface of the vehicle to estimate values of angle of attack, angle of side-slip, Mach number, impact pressure and free-stream static pressure.

Pad Abort 1↗
NASA NTRS2014

Self-Contained Compressed-Flow Generation Device for Use in Making Differential Measurements

A device used in making differential measurements of a flow includes a flow obstruction and a support arm. The flow obstruction's forward portion is a nose cone. The flow obstruction's aft portion is coupled to the nose cone. The support arm's first end is coupled to an exterior wall of a conduit, and its second end is coupled to the forward portion of the flow obstruction. The support arm positions the flow obstruction in the conduit such that a flow region is defined around its nose cone, and such that the support arm's first and second end are separated from one another with respect to a length dimension of the conduit. Measurement ports are provided in the support arm and flow obstruction. Manifolds extending through the flow obstruction and support arm couple the ports to points at the exterior wall of the conduit.

England, John Dwight↗
NASA NTRS1946

Tests of a 1/5-Scale Model of the Republic XP-84 Airplane (Army Project MX-578) in the Langley 300 MPH 7- by 10-Foot Tunnel

A 1/5-scale model of the Republic x-84 airplane (Army Project MX-578) was tested in the Langley 300 MPH 7- by 10-foot tunnel. The primary object of the tests was twofold: to determine, a practicable method of increasing the longitudinal 3tability in the landing configuration, and to investigate the effects on longitudinal and lateral Stability of various external stored (fuel tanks, bombs, and rockets). The effects of the fuselage dive brakes were also determined, and the critical Mach numbers of certain of the airplane components were estimated. The use of the revised horizontal tail (of larger aspect ratio and area than the original) seemed to be the most feasible expedient for materially increasing the longitudinal stability in the landing configuration. The neutral-point shifts produced by the various external stores were unstable, the largest shift being about 2.5 percent mean aerodynamic chord. No appreciable aerodynamic trim changes were caused by the external stores. From the standpoint of range, maximum s peed, and rate of climb, the advantages of mounting the fuel tanks at the wing tips rather than inboard beneath the wings were clearly demonstrated by the tests. The effective dihedral parameter was the only static lateral-stability derivative appreciably affected by the external stores. At high lift coefficients, the tip-mounted tanks caused a large increase in the effective dihedral parameter (about 40 increase at a lift coefficient of 1.0). This increase was held undesirable, because the tendency toward oscillatory instability that it would cause would be heightened by the increased moments of inertia resulting from the weight of the tanks when carrying fuel. The fuselage dive brakes, when deflected, caused a change in trim tending to nose the airplane up; the neutral point also moved rearward upon deflecting the dive brakes. The amount of elevator required to overcome the change in trim was well within the available range of deflection. It was estimated that a drive-brake deflection of 900 would.decrease the terminal Mach number in a vertical dive by about 0.1. The estimated critical Mach number of the V-front canopy was about 0.04 greater than that of the original canopy. Pressure-distribution tests disclosed severe pressure peaks inside the nose of the jet entrance duct. These peaks, which would lead to separation and consequently poor pressure recovery at, the engine, could be reduced by, using a smaller nose,radius and: a modified internal lip shape

Tucker, Warren A.↗
NASA NTRS2015

Receptivity of Hypersonic Boundary Layers to Acoustic and Vortical Disturbances (Invited)

Boundary-layer receptivity to two-dimensional acoustic and vortical disturbances for hypersonic flows over two-dimensional and axi-symmetric geometries were numerically investigated. The role of bluntness, wall cooling, and pressure gradients on the receptivity and stability were analyzed and compared with the sharp nose cases. It was found that for flows over sharp nose geometries in adiabatic wall conditions the instability waves are generated in the leading-edge region and that the boundary layer is much more receptive to slow acoustic waves as compared to the fast waves. The computations confirmed the stabilizing effect of nose bluntness and the role of the entropy layer in the delay of boundary layer transition. The receptivity coefficients in flows over blunt bodies are orders of magnitude smaller than that for the sharp cone cases. Wall cooling stabilizes the first mode strongly and destabilizes the second mode. However, the receptivity coefficients are also much smaller compared to the adiabatic case. The adverse pressure gradients increased the unstable second mode regions.

Balakumar, P.↗
NASA NTRS2018

Bow Shock in Merging Cluster A520: The Edge of the Radio Halo and the Electron-Proton Equilibration Timescale

We studied the prominent bow shock in the merging galaxy cluster A520 using a deep Chandra X-ray observation and archival VLA radio data. This shock is a useful diagnostic tool, owing to its clear geometry and relatively high Mach number. At the "nose" of the shock, we measure a Mach number of M=2.4(sup 0.4, sub -0.2). The shock becomes oblique away from the merger axis, with the Mach number falling to approx. 1.6 around 30 deg from the nose. The electron temperature immediately behind the shock nose is consistent with that from the Rankine-Hugoniot adiabat, and is higher (at a 95% confidence) than expected for adiabatic compression of electrons followed by Coulomb electron-proton equilibration, indicating the presence of equilibration mechanisms faster than Coulomb collisions. This is similar to an earlier finding for the Bullet cluster. We also combined four archival VLA data sets to obtain a better image of the cluster's giant radio halo at 1.4 GHz. An abrupt edge of the radio halo traces the shock front, and no emission is detected in the pre-shock region. If the radio edge were due only to adiabatic compression of relativistic electrons in pre-shock plasma, we would expect a pre-shock radio emission detectable in this radio data set; however, an interferometric artifact dominates the uncertainty, so we cannot rule this model out. Other interesting features of the radio halo include a peak at the remnant of the cool core, suggesting that the core used to have a radio minihalo, and a peak marking a possible region of high turbulence.

Wang, Qian H. S.↗
NASA NTRS2019

Predicted Ice Shape Formations on a Boundary Layer Ingesting Engine Inlet

Computational ice shapes were generated on the boundary layer ingesting engine nacelle of the D8 Double Bubble aircraft. The computations were generated using LEWICE3D, a well-known CFD icing post processor. A 50-bin global drop diameter discretization was used to capture the collection efficiency due to the direct impingement of water onto the engine nacelle. These discrete results were superposed in a weighted fashion to generate six drop size distributions that span the Appendix C and O regimes. Due to the presence of upstream geometries, i.e. the fuselage nose, the trajectories of the water drops are highly complex. Since the ice shapes are significantly correlated with the collection efficiency, the upstream fuselage nose has a significant impact on the ice accretion on the engine nacelle. These complex trajectories are caused by the ballistic nature of the particles and are thus exacerbated as particle size increases. Shadowzones are generated on the engine nacelle, and due to the curvature of the nose of the aircraft the shadowzone boundary moves from lower inboard to upper outboard as particle size increases. The largest particle impinging one the engine nacelle from the 50-bin discretization was the 47 um drop diameter. As a result, the MVD greater than 40 um Appendix O conditions were characterized by extremely low collection efficiency on the engine nacelle for these direct impingement simulations.

Porter, Christopher↗