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At least 577 records · Page 32

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NASA NTRS1948

Ditching Tests of a 1/24-Scale Model of the Lockheed XR60-1 Airplane, TED No. NACA 235

The ditching characteristics of the Lockheed XR60-1 airplane were determined by tests of a 1/24-scale dynamic model in calm water at the Langley tank no. 2 monorail. Various landing attitudes, flap settings, speeds, and conditions of damager were investigated. The ditching behavior was evaluated from recordings of decelerations, length of runs, and motions of the model. Scale-strength bottoms and simulated crumpled bottoms were used to reproduce probable damage to the fuselage. It was concluded that the airplane should be ditched at a landing attitude of about 5 deg with flaps full down. At this attitude, the maximum longitudinal deceleration should not exceed 2g and the landing run will be bout three fuselage lengths. Damage to the fuselage will not be excessive and will be greatest near the point of initial contact with the water.

Fisher, Lloyd J.↗

NASA NTRS1948

Ditching Tests of Two Models of the Army B-36 Airplane

The ditching characteristics of the Army B-36 airplane were determined by testing 1/20- and 1/30-scale dynamic models in calm water in Langley tank no. 2 and at the outdoor catapult. The scope of the tests consisted of ditching the models at various conditions of simulated damage, landing attitudes, and speeds, with various flap settings using several degrees of restraint of the flap hinges. The ditching behavior was evaluated from recordings of deceleration, length of run, and motions of the models. The results showed that the airplane should be ditched at an attitude of about 9 deg with flaps full down. The probable ditching behavior will be a smooth run with a maximum longitudinal deceleration of 3g to 4g and a landing run of 4 to 5 fuselage lengths. Structural failure of the underside of the fuselage will not seriously affect the behavior of the airplane.

Fisher, Lloyd J.↗

NASA NTRS1950

Ditching Investigation of a 1/24-Scale Model of the Boeing B-47 Airplane

An investigation of a 1/24-scale dynamically similar model of the Boeing B-47 airplane was made to determine the ditching characteristics and proper ditching technique for the airplane. Various conditions of damage, landing attitude, flap setting, and speed were investigated. The behavior of the model was determined from visual observations, motion-picture records, and time-history deceleration records. The results of the investigation are presented in table form, photographs, and curves. The airplane should be ditched at the lowest speed and highest attitude consistent with adequate control; the flaps should be full down. The airplane will probably make a deep but fairly smooth run. The fuselage bottom will be damaged and partially filled with water; consequently, crew members should be assigned ditching stations near an exit in the upper or forward part of the fuselage. The nacelles may be expected to be torn away from the wing. In calm water the maximum decelerations will be about 3g and the landing run will be about 6 fuselage lengths.

Fisher, Lloyd J.↗

NASA NTRS1951

Ditching Investigation of a 1/24-Scale Model of the Douglas C-124 Airplane

An investigation of a 1/24- scale dynamically similar model of the Douglas C-124 airplane was made to determine the ditching characteristics and proper technique for ditching the airplane. Various conditions of damage, landing attitude, flap setting, and speed were investigated. The behavior of the model was determined from visual observations, motion- picture records, and time-history deceleration records. The results of the investigation are presented in table form, photographs, and curves. It was concluded on the basis of results from model tests with scale-strength bottoms (equivalent to 1150 pounds per square foot, full scale) that the airplane should be ditched at a medium nose-high landing attitude (near 7deg) with flaps full down. The airplane will probably make a smooth run with considerable damage resulting to the fuselage bottom just forward of the wing, but it is not likely that the water inflow will be overwhelming to personnel provided they are not in the belly compartment. Longitudinal decelerations in calm water will be about 2 1/2g and the landing run will be about four fuselage lengths.

Fisher, Lloyd J.↗

NASA NTRS1950

Ditching Investigation of a 1/14-Scale Model of the Northrop C-125 Airplane

An investigation of a 1/14-scale dynamically similar model of the Northrop C-125 airplane was made to determine the ditching characteristics and proper ditching technique for the airplane. Various conditions of damage, landing attitude, flap setting, and speed were investigated. The behavior of the model was determined from visual observations, motion-picture records, and time-history deceleration records. The results of the investigation are presented in table form, photographs, and curves. It was concluded that the airplane should be ditched at a nose-high landing attitude (near 8 deg) with flaps full down. The fixed landing gear will cause the airplane to dive. The fuselage will be damaged and will probably flood rapidly. Longitudinal decelerations will be about 4g and the length of landing run will be about two fuselage lengths. If the main landing gear could be jettisonable or retractable the ditching characteristics would be greatly improved.

Fisher, Lloyd J.↗

NASA NTRS1955

Ditching Investigation of a 1/20-Scale Model of the Boeing Stratocruiser Airplane (C-97)

An investigation of a 1/20-scale dynamically similar model of the Boeing Stratocruiser airplane (C-97) was made to determine the ditching characteristics and proper technique for ditching the airplane. Scale-strength bottoms were used to determine probable damage to the fuselage and the effect of damage on behavior. The behavior of the model was determined from visual observations, motion-picture records, and time-history deceleration records. Data are presented in a table, photographs, and curves. It was concluded that the airplane should be ditched at a medium nose-high landing attitude (near 6 deg) with landing flaps full down. The airplane will probably make a smooth run of medium depth with light spray and may even trim up slightly in the water. The fuselage will probably be damaged and the lower compartment filled with water. In calm water, the maximum longitudinal deceleration will be about 4g and the landing run will be about four fuselage lengths.

Fisher, Lloyd J.↗

NASA NTRS2003

Effect of High Impact or Non-impact Loading Activity on Bone Bending Stiffness and Mineral Density

Material properties of conical bone, including mineral density (BMD) and its geometry is closely related to its load-carrying capacity. These two primary components determine the strength of conical bone. High impact loading involving acceleration and deceleration movements used in gymnastics induce higher BMD of the affected bone compared to the non-impact acceleration and deceleration movements used in swimming. Study of these two groups of athletes on bone bending stiffness has not been reported. The purpose of this study was to compare differences in bone bending stiffness and BMD between competitive female synchronized swimmers and female gymnasts. Thirteen world class female synchronized swimmers (SYN) and 8 female gymnasts (GYM), mean age 21 +/- 2.9 yr. were recruited for this study. We used a mechanical response tissue analyzer (Gaitscan, NJ) to calculate EI, where E is Young's modulus of elasticity and I is the cross-sectional moment of inertia. EI was obtained from tissue response to a vibration probe placed directly on the skin of the mid-region of tibia and ulna. BMD of the heel and wrist were measured with a probe densitometer (PIXI, Lunor, WI). The SYN were taller than (p < 0.05) the GYM but weighed the same as the GYM. EI obtained from tibia and ulna of the SYN (291 +/- 159 and 41 +/- 19.4, respectively) were not significantly different from thc GYM (285 +/- 140 and 44 +/- 18.3, respectively). BMD of the heel and wrist in GYM were higher than in SYN (p < 0.001). High impact weight-bearing activities promote similar bone strength but greater BMD response than non-impact activities performed in a buoyant environment.

Liang, Michael T. C.↗

NASA NTRS2005

The Growth, Polarization and Motion of the Radio Afterglow from the Giant Flare from SGR 1806-20

The extraordinary giant flare (GF) of 2004 December 27 from the soft gamma repeater (SGR) 1806-20 was followed by a bright radio afterglow. We present an analysis of VLA observations of this radio afterglow from SGR 1806-20, consisting of previously reported 8.5 GHz data covering days 7 to 20 after the GF, plus new observations at 8.5 and 22 GHz from day 24 to 81. We find a deceleration in the expansion, from approximately 4.5 mas/day to less than 2.5 mas/day. The time of deceleration is roughly coincident with the rebrightening in the radio light curve, as expected to result when the ejecta from the GF sweeps up enough of the external medium, and transitions from a coasting phase to the Sedov-Taylor regime. The radio afterglow is elongated and maintains a 2:l axis ratio with an average position angle of -40 degrees (north through east), oriented perpendicular to the average intrinsic linear polarization angle. We also report on the discovery of motion in the flux centroid of the afterglow, at an average velocity of 0.26 plus or minus 0.03 c (assuming a distance of 15 kpc) at a position angle of -40 degrees. This motion, in combination with the growth and polarization measurements, suggests an initially asymmetric outflow, mainly from one side of the magnetar.

Taylor, G. B.↗

NASA NTRS2005

Aerocapture Performance Analysis of A Venus Exploration Mission

A performance analysis of a Discovery Class Venus Exploration Mission in which aerocapture is used to capture a spacecraft into a 300km polar orbit for a two year science mission has been conducted to quantify its performance. A preliminary performance assessment determined that a high heritage 70 sphere-cone rigid aeroshell with a 0.25 lift to drag ratio has adequate control authority to provide an entry flight path angle corridor large enough for the mission s aerocapture maneuver. A 114 kilograms per square meter ballistic coefficient reference vehicle was developed from the science requirements and the preliminary assessment s heating indicators and deceleration loads. Performance analyses were conducted for the reference vehicle and for sensitivity studies on vehicle ballistic coefficient and maximum bank rate. The performance analyses used a high fidelity flight simulation within a Monte Carlo executive to define the aerocapture heating environment and deceleration loads and to determine mission success statistics. The simulation utilized the Program to Optimize Simulated Trajectories (POST) that was modified to include Venus specific atmospheric and planet models, aerodynamic characteristics, and interplanetary trajectory models. In addition to Venus specific models, an autonomous guidance system, HYPAS, and a pseudo flight controller were incorporated in the simulation. The Monte Carlo analyses incorporated a reference set of approach trajectory delivery errors, aerodynamic uncertainties, and atmospheric density variations. The reference performance analysis determined the reference vehicle achieves 100% successful capture and has a 99.87% probability of attaining the science orbit with a 90 meters per second delta V budget for post aerocapture orbital adjustments. A ballistic coefficient trade study conducted with reference uncertainties determined that the 0.25 L/D vehicle can achieve 100% successful capture with a ballistic coefficient of 228 kilograms per square meter and that the increased ballistic coefficient increases post aerocapture V budget to 134 meters per second for a 99.87% probability of attaining the science orbit. A trade study on vehicle bank rate determined that the 0.25 L/D vehicle can achieve 100% successful capture when the maximum bank rate is decreased from 30 deg/s to 20 deg/s. The decreased bank rate increases post aerocapture delta V budget to 102 meters per second for a 99.87% probability of attaining the science orbit.

Starr, Brett R.↗

NASA NTRS1973

Free-Flight Test Results of Scale Models Simulating Viking Parachute/Lander Staging

This report presents the results of Viking Aerothermodynamics Test D4-34.0. Motion picture coverage of a number of Scale model drop tests provides the data from which time-position characteristics as well as canopy shape and model system attitudes are measured. These data are processed to obtain the instantaneous drag during staging of a model simulating the Viking decelerator system during parachute staging at Mars. Through scaling laws derived prior to test (Appendix A and B) these results are used to predict such performance of the Viking decelerator parachute during staging at Mars. The tests were performed at the NASA/Kennedy Space Center (KSC) Vertical Assembly Building (VAB). Model assemblies were dropped 300 feet to a platform in High Bay No. 3. The data consist of an edited master film (negative) which is on permanent file in the NASA/LRC Library. Principal results of this investigation indicate that for Viking parachute staging at Mars: 1. Parachute staging separation distance is always positive and continuously increasing generally along the descent path. 2. At staging, the parachute drag coefficient is at least 55% of its prestage equilibrium value. One quarter minute later, it has recovered to its pre-stage value.

Polutchko, Robert J.↗

NASA NTRS2006

Integrated Thermal Protection Systems and Heat Resistant Structures

In the early stages of NASA's Exploration Initiative, Snecma Propulsion Solide was funded under the Exploration Systems Research & Technology program to develop a CMC heatshield, a deployable decelerator, and an ablative heat shield for reentry vehicles. Due to changes within NASA's Exploration Initiative, this task was cancelled in early FY06. This paper will give an overview of the work that was accomplished prior to cancellation. The Snecma team consisted of MT Aerospace, Germany, and Materials Research & Design (MR&D), NASA Langley, NASA Dryden, and NASA Ames in the United States. An Apollo-type capsule was chosen as the reference vehicle for the work. NASA Langley generated the trajectory and aerothermal loads. Snecma and MT Aerospace began the design of a ceramic aft heatshield (CAS) utilizing C/SiC panels as the capsule heatshield. MR&D led the design of a C/SiC deployable decelerator, NASA Ames led the characterization of several ablators, NASA Dryden led the development of a heath management system and the high temperature structures testing, and NASA Langley led the insulation characterization. Though the task was pre-maturely cancelled, a significant quantity of work was accomplished.

Pichon, Thierry↗

NASA NTRS2005

Simulations of Dynamics and Transport during the September 2002 Antarctic Major Warming

A mechanistic model simulation initialized on 14 September 2002, forced by 100-hPa geopotential heights from Met Office analyses, reproduced the dynamical features of the 2002 Antarctic major warming. The vortex split on approx.25 September; recovery after the warming, westward and equatorward tilting vortices, and strong baroclinic zones in temperature associated with a dipole pattern of upward and downward vertical velocities were all captured in the simulation. Model results and analyses show a pattern of strong upward wave propagation throughout the warming, with zonal wind deceleration throughout the stratosphere at high latitudes before the vortex split, continuing in the middle and upper stratosphere and spreading to lower latitudes after the split. Three-dimensional Eliassen-Palm fluxes show the largest upward and poleward wave propagation in the 0(deg)-90(deg)E sector prior to the vortex split (coincident with the location of strongest cyclogenesis at the model's lower boundary), with an additional region of strong upward propagation developing near 180(deg)-270(deg)E. These characteristics are similar to those of Arctic wave-2 major warmings, except that during this warming, the vortex did not split below approx.600 K. The effects of poleward transport and mixing dominate modeled trace gas evolution through most of the mid- to high-latitude stratosphere, with a core region in the lower-stratospheric vortex where enhanced descent dominates and the vortex remains isolated. Strongly tilted vortices led to low-latitude air overlying vortex air, resulting in highly unusual trace gas profiles. Simulations driven with several meteorological datasets reproduced the major warming, but in others, stronger latitudinal gradients at high latitudes at the model boundary resulted in simulations without a complete vortex split in the midstratosphere. Numerous tests indicate very high sensitivity to the boundary fields, especially the wave-2 amplitude. Major warmings occurred for initial fields with stronger winds and larger vortices, but not smaller vortices, consistent with the initiation of wind-deceleration by upward-propagating waves near the poleward edge of the region where wave 2 can propagate above the jet core. Thus, given the observed 100-hPa boundary forcing, stratospheric preconditioning is not needed to reproduce a major warming similar to that observed. The anomalously strong forcing in the lower stratosphere can be viewed as the primary direct cause of the major warming.

winds↗

NASA NTRS1968

Flight Test of a 40-Foot Nominal-Diameter Disk-Gap-Band Parachute Deployed at a Mach Number of 1.91 and a Dynamic Pressure of 11.6 Pounds per Square Foot

A 40-foot (12.2 meter) nominal-diameter disk-gap-band parachute was flight tested as part of the NASA Supersonic Planetary Entry Decelerator Program (SPED-I). The test parachute was ejected by a deployment mortar from an instrumented payload at an altitude of 140,000 feet (42.5 kilometers). The payload was at a Mach number of 1.91 and the dynamic pressure was 11.6 pounds per square foot (555 newtons per square meter) at the time the parachute deployment mortar was fired. The parachute reached suspension line stretch in 0.43 second with a resultant snatch force loading of 1990 pounds (8850 newtons). The maximum parachute opening load of 6500 pounds (28,910 newtons) came 0.61 second later at a total elapsed time from mortar firing of 1.04 seconds. The first full inflation occurred at 1.12 seconds and stable inflation was achieved at approximately 1.60 seconds. The parachute had an average axial-force coefficient of 0.53 during the deceleration period. During the steady-state descent portion of the flight test, the average effective drag coefficient was also 0.53 and pitch-yaw oscillations of the canopy averaged less than 10 degrees in the altitude region above 100,000 feet (30.5 meters).

EXPERIMENTS↗

NASA NTRS1971

High Altitude Flight Test of a Reefed 12.2 Meter Diameter Disk-Gap-Band Parachute with Deployment at Mach Number of 2.58

A reefed 12.2-meter nominal-diameter (40-ft) disk-gap-band parachute was flight tested as part of the NASA Supersonic High Altitude Parachute Experiment (SHAPE) program. A three-stage rocket was used to drive the instrumented payload to an altitude of 43.6 km (143,000 ft), a Mach number of 2.58, and a dynamic pressure of 972 N/m(exp 2) (20.3 lb/ft(exp 2)) where the parachute was deployed by means of a mortar. The parachute deployed satisfactorily and reached a partially inflated condition characterized by irregular variations in parachute projected area. A full, stable reefed inflation was achieved when the system had decelerated to a Mach number of about 1.5. The steady, reefed projected area was 49 percent of the steady, unreefed area and the average drag coefficient was 0.30. Disreefing occurred at a Mach number of 0.99 and a dynamic pressure of 81 N/m(exp 2) (1.7 lb/ft(exp 2)). The parachute maintained a steady inflated shape for the remainder of the deceleration portion of the flight and throughout descent. During descent, the average effective drag coefficient was 0.57. There was little, if any, coning motion, and the amplitude of planar oscillations was generally less than 10 degrees. The film also shows a wind tunnel test of a 1.7-meter-diameter parachute inflating at Mach number 2.0.

EXPERIMENTS↗

NASA NTRS2007

Analysis of the Interactions of Planetary Waves with the Mean Flow of the Stratosphere

During the winter period, large scale waves (planetary waves) are observed to propagate from the troposphere into the stratosphere. Such wave events have been recognized since the 1 950s. The very largest wave events result in major stratospheric warmings. These large scale wave events have typical durations of a few days to 2 weeks. The wave events deposit easterly momentum in the stratosphere, decelerating the polar night jet and warming the polar region. In this presentation we show the typical characteristics of these events via a compositing analysis. We will show the typical periods and scales of motion and the associated decelerations and warmings. We will illustrate some of the differences between major and minor warming wave events. We will further illustrate the feedback by the mean flow on subsequent wave events.

Newman, Paul A.↗

NASA NTRS2008

Simulated Reentry Heating by Torching

The two first order reentry heating parameters are peak heating flux (W/cm2) and peak heat load (kJ/cm2). Peak heating flux (and deceleration, gs) is higher for a ballistic reentry and peak heat load is higher for a lifting reentry. Manned vehicle reentries are generally lifting reentries at nominal 1-5 gs so that personnel will not be crushed by high deceleration force. A few off-nominal manned reentries have experienced 8 or more gs with corresponding high heating flux (but below nominal heat load). The Shuttle Orbiter reentries provide about an order of magnitude difference in peak heating flux at mid-bottom (TPS tiles, approximately 6 W/cm2 or 5 BTU/ft2- sec) and leading edge (RCC, approximately 60 W/cm2 or 50 BTU/ft2- sec). Orion lunar return and Mars sample lander are of the same order of magnitude as orbiter leading edge peak heat loads. Flight temperature measurements are available for some orbiter TPS tile and RCC locations. Return-to-Flight on-orbit tile-repair-candidate-material-heating performance was evaluated by matching propane torch heating of candidate-materials temperatures at several depths to orbiter TPS tile flight-temperatures. Char and ash characteristics, heat expansion, and temperature histories at several depths of the cure-in-place ablator were some of the TPS repair material performance characteristics measured. The final char surface was above the initial surface for the primary candidate (silicone based) material, in contrast to a receded surface for the Apollo-type ablative heat shield material. Candidate TPS materials for Orion CEV (LEO and lunar return), and for Mars sample lander are now being evaluated. Torching of a candidate ablator material, PICA, was performed to match the ablation experienced by the STARDUST PICA heat shield. Torching showed that the carbon fiberform skeleton in a sample of PICA was inhomogeneous in that sample, and allowed measurements (of the clumps and voids) of the inhomogeneity. Additional reentry heating-performance characterizations of high temperature insulation materials were performed.

Harvey, Gale A.↗

NASA NTRS2007

The Spatial Variation of Polar Rain Electrons and its Cause

It is generally accepted that field aligned electrons in the solar wind can follow field lines connected to Earth and precipitate in the polar ionosphere where they are known as polar rain. Few-hundred eV, field-aligned electrons of the solar wind "strahl" carry the interplanetary heat flux moving out from the sun and these electrons precipitate in either the northern or southern hemisphere depending on the magnetic field direction. These electrons produce enhanced polar rain in one hemisphere or the other although weaker polar rain is usually produced in the opposite hemisphere by whatever electrons are moving in the opposite direction. Although much evidence exists for this simple free entry mechanism, it has also long been known that there are spatial variations in the energies and intensities of the precipitating electrons. The present work compares electron distribution functions measured by the ACE spacecraft in the solar wind with those measured by the DMSP spacecraft at 800 km altitude over the polar cap. It is found that shifting the DMSP distribution functions in energy by amounts ranging from 10's to a few hundred eV produces quite good agreement with simultaneous ACE measurements. Over most of the polar cap this DMSP energy shift must be positive to achieve this agreement, suggesting the electrons have been decelerated by a field aligned potential as they move from the solar wind to low altitudes. The largest shifts occur on the nightside and on the dawn or dusk side, with the latter depending on the plasma convection pattern which is controlled by the orientation of the IMF. Nearer the cusp the shift is smaller or even negative. Since more massive tailward flowing magnetosheath ions are unable io follow the field lines into the magnetotail like the electrons, a field aligned potential is expected to develop to exclude low energy electrons and prevent an excessive charge imbalance. Such a potential would also produce the deceleration of those electrons that reach low altitudes. This improved understanding of polar rain should increase the utility of polar rain measurements as a diagnostic of the magnetosphere magnetic field configuration.

Fairfield, D. H.↗

NASA NTRS1947

An Anaylsis of Control Requirements and Control Parameters for Direct-Coupled Turbojet Engines

Requirements of an automatic engine control, as affected by engine characteristics, have been analyzed for a direct-coupled turbojet engine. Control parameters for various conditions of engine operation are discussed. A hypothetical engine control is presented to illustrate the use of these parameters. An adjustable speed governor was found to offer a desirable method of over-all engine control. The selection of a minimum value of fuel flow was found to offer a means of preventing unstable burner operation during steady-state operation. Until satisfactory high-temperature-measuring devices are developed, air-fuel ratio is considered to be a satisfactory acceleration-control parameter for the attainment of the maximum acceleration rates consistent with safe turbine temperatures. No danger of unstable burner operation exists during acceleration if a temperature-limiting acceleration control is assumed to be effective. Deceleration was found to be accompanied by the possibility of burner blow-out even if a minimum fuel-flow control that prevents burner blow-out during steady-state operation is assumed to be effective. Burner blow-out during deceleration may be eliminated by varying the value of minimum fuel flow as a function of compressor-discharge pressure, but in no case should the fuel flow be allowed to fall below the value required for steady-state burner operation.

Novik, David↗