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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 451 records · Page 25
All sourcesNASA NTRSDOE OSTI
NASA NTRS1959

Flight Studies of Problems Pertinent to Low-Speed Operation of Jet Transports

Flight studies have been made of the low-speed operational regime of jet transports in order to assess potential operating problems. The study was performed utilizing a large multiengine jet airplane having geometric characteristics fairly representative of the jet transports; however, to insure general applicability of the results, the aerodynamic characteristics of the test airplane were varied to simulate a variety of jet-transport airplanes. The specific areas investigated include those of the take-off and landing, and the relation of these maneuvers to the 1 g stall speed and stalling characteristics. The take-off studies included evaluation of the factors affecting the take-off speed and attitude, including the effects of premature rotation and of over-rotation on ground run required. The approach and landing studies pertained to such factors as: desirable lateral-directional damping characteristics; lateral-control requirements; space-positioning limitations during approach under VFR or IFR conditions and requirements for glide-path controls; and evaluation of factors affecting the pilot's choice of landing speeds. Specific recommendations and some indication of desirable characteristics for the jet transports are advanced to alleviate possible operational difficulties or to improve operational performance in the low-speed range.

Fischel. Jack↗
NASA NTRS1959

Approach and Landing Investigation at Lift-Drag Ratios of 2 to 4 Utilizing a Straight-Wing Fighter Airplane

A series of landings was performed with a straight-wing airplane to evaluate the effect of low lift-drag ratios on approach and landing characteristics. Landings with a peak lift-drag ratio as low as 3 were performed by altering the airplane configuration (extending speed brakes, flaps, and gear and reducing throttle setting). As lift-drag ratio was reduced, it was necessary either to make the landing pattern tighter or to increase initial altitude, or both. At the lowest lift-drag ratio the pilots believed a 270 deg overhead pattern was advisable because of the greater ease afforded in visually positioning the airplane. The values of the pertinent flare parameters increased with the reduction of lift-drag ratio. These parameters included time required for final flare; speed change during final flare; and altitude, glide slope, indicated airspeed, and vertical velocity at initiation of final flare. The pilots believed that the tolerable limit was reached with this airplane in the present configuration, and that if, because of a further reduction in lift-drag ratio, more severe approaches than those experienced in this program were attempted, additional aids would be required to determine the flare-initiation point.

Matranga, Gene J.↗
NASA NTRS1959

Launch, Low-Speed, and Landing Characteristics Determined from the First Flight of the North American X-15 Research Airplane

The first flight of the North American X-15 research airplane was made on June 8, 1959. This was accomplished after completion of a series of captive flights with the X-15 attached to the B-52 carrier airplane to demonstrate the aerodynamic and systems compatibility of the X-15//B-52 combination and the X-15 subsystem operation. This flight was planned as a glide flight so that the pilot need not be concerned with the propulsion system. Discussions of the launch, low-speed maneuvering, and landing characteristics are presented, and the results are compared with predictions from preflight studies. The launch characteristics were generally satisfactory, and the X-15 vertical tail adequately cleared the B-52 wing cutout. The actual landing pattern and landing characteristics compared favorably with predictions, and the recommended landing technique of lowering the flaps and landing gear at a low altitude appears to be a satisfactory method of landing the X-15 airplane. There was a quantitative correlation between flight-measured and predicted lift-drag-ratio characteristics in the clean configuration and a qualitative correlation in the landing configuration. A longitudinal-controllability problem, which became severe in the landing configuration, was evident throughout the flight and, apparently, was aggravated by the sensitivity of the side-located control stick. In the low-to-moderate angle-of-attack range covered, the longitudinal and directional stability were indicated to be adequate.

Finch, Thomas W.↗
NASA NTRS1974

The Effect of Porosity on the Lift and Drag of Bird's Wings

Bird's wings are porous. The mass flux through the wings for various simulated wing loadings was measured and found to average about 2 x 10(exp -4) g/s.sq cm for one feather thickness at the average pigeon wing loading of 45 newtons/sq m. The calculated effect of the flux on lift was found to be less than 1/2 of 1%. Experiments at a typical pigeon flight Reynolds number of 10(exp 5) on flat plates with 3.5% and 14.5% open area spread over the trailing 58% of the chord showed that it is possible to both increase and decrease profile drag, relative to the drag on a non-porous flat plate, by < or = 5%. Stall on the porous plates occurred up to 3 deg earlier than stall on the non-porous plates. From lift calculations and drag measurements it is shown that the effects of porosity on the gliding characteristics of birds are negligible.

Lichter, Seth↗
NASA NTRS1965

Preliminary Study of Steep Instrument Approach of Three Conventional Aircraft

Preliminary studies of steep instrument approaches have been conducted with three conventional aircraft. Six-degree approach paths were found to be acceptable. A margin of 3 deg to 4 deg in maximum glide slope was available to allow for flight-path corrections. Control of the lateral-directional axis was the most difficult problem for pilots making either steep or conventional approaches; split-axis control reduced the pilot workload during this operation.

Hall, Albert W.↗
NASA NTRS1965

V/STOL Aircraft Operation in the Terminal Area

Some aspects of V/STOL operation in a terminal area have been presented. A V/STOL instrument approach with present-day displays and guidance systems requires about 5 minutes at low speeds, but this time could be cut to 1 1/2 to 2 minutes if the displays and guidance systems were improved. In order to keep approach-pattern time to a minimum, the conversion process should be simplified by interconnecting as many operations as possible with one control. Adequate vectoring for the lifting system, attitude stabilization to prevent excessive wandering, and automatic glide-path control for approaches over 6 deg. will also be required.

Reeder, John P.↗
NASA NTRS1995

A Light-Weight Inflatable Hypersonic Drag Device for Planetary Entry

The author has analyzed the use of a light-weight inflatable hypersonic drag device, called a ballute, (balloon + parachute) for flight in planetary atmospheres, for entry, aerocapture, and aerobraking. Studies to date include missions to Mars, Venus, Earth, Saturn, Titan, Neptune and Pluto. Data on a Pluto lander and a Mars orbiter will be presented to illustrate the concept. The main advantage of using a ballute is that aero deceleration and heating in atmospheric entry occurs at much smaller atmospheric density with a ballute than without it. For example, if a ballute has a diameter 10 times as large as the spacecraft, for unchanged total mass, entry speed and entry angle,the atmospheric density at peak convective heating is reduced by a factor of 100, reducing the peak heating by a factor of 10 for the spacecraft, and a factor of about 30 for the ballute. Consequently the entry payload (lander, orbiter, etc) is subject to much less heating, requires a much reduced thermal protection system (possibly only an MLI blanket), and the spacecraft design is therefore relatively unchanged from its vacuum counterpart. The heat flux on the ballute is small enough to be radiated at temperatures below 800 K or so. Also, the heating may be reduced further because the ballute enters at a more shallow angle, even allowing for the increased delivery angle error. Added advantages are a smaller mass ratio of entry system to total entry mass, and freedom from the low-density and transonic instability problems that conventional rigid entry bodies suffer, since the vehicle attitude is determined by the ballute, usually released at continuum conditions (hypersonic for an orbiter, and subsonic for a lander). Also, for a lander the range from entry to touchdown is less, offering a smaller footprint. The ballute derives an entry corridor for aerocapture by entering on a path that would lead to landing, and releasing the ballute adaptively, responding to measured deceleration, at a speed computed to achieve the desired orbiter exit conditions. For a lander an accurate landing point could be achieved by providing the lander with a small gliding capacity, using the large potential energy available from being subsonic at high altitude. Alternatively the ballute can be retained to act as a parachute or soft-landing device, or to float the payload as a buoyant aerobot. As expected, the ballute has smaller size for relatively small entry speeds, such as for Mars, or for the extensive atmosphere of a low-gravity planet such as Pluto. The author will discuss presently available ballute materials and a development program of aerodynamic tests and materials that would be required for ballutes to achieve their full potential.

McRonald, Angus D.↗
NASA NTRS2000

A Light-Weight Inflatable Hypersonic Drag Device for Planetary Entry

The author has analyzed the use of a light-weight inflatable hypersonic drag device, called a ballute, for flight in planetary atmospheres, for entry, aerocapture, and aerobraking. Studies to date include Mars, Venus, Earth, Saturn, Titan, Neptune and Pluto, and data on a Pluto lander and a Mars orbiter will be presented to illustrate the concept. The main advantage of using a ballute is that aero, deceleration and heating in atmospheric entry occurs at much smaller atmospheric density with a ballute than without it. For example, if a ballute has a diameter 10 times as large as the spacecraft, for unchanged total mass, entry speed and entry angle,the atmospheric density at peak convective heating is reduced by a factor of 100, reducing the heating by a factor of 10 for the spacecraft and a factor of 30 for the ballute. Consequently the entry payload (lander, orbiter, etc) is subject to much less heating, requires a much reduced thermal. protection system (possibly only an MLI blanket), and the spacecraft design is therefore relatively unchanged from its vacuum counterpart. The heat flux on the ballute is small enough to be radiated at temperatures below 800 K or so. Also, the heating may be reduced further because the ballute enters at a more shallow angle, even allowing for the increased delivery angle error. Added advantages are less mass ratio of entry system to total entry mass, and freedom from the low-density and transonic instability problems that conventional rigid entry bodies suffer, since the vehicle attitude is determined by the ballute, usually released at continuum conditions (hypersonic for an orbiter, and subsonic for a lander). Also, for a lander the range from entry to touchdown is less, offering a smaller footprint. The ballute derives an entry corridor for aerocapture by entering on a path that would lead to landing, and releasing the ballute adaptively, responding to measured deceleration, at a speed computed to achieve the desired orbiter exit conditions. For a lander an accurate landing point could be achieved by providing the lander with a small gliding capacity, using the large potential energy available from being subsonic at high altitude. Alternatively the ballute can be retained to act as a parachute or soft-landing device, or to float the payload as a buoyant aerobot. As expected, the ballute has smaller size for relatively small entry speeds, such as for Mars and Titan, or for the extensive atmosphere of a low-gravity planet such as Pluto. Details of a ballute to place a small Mars orbiter and a small Pluto lander will be given to illustrate the concept. The author will discuss presently available ballute materials and a development program of aerodynamic tests and materials that would be required for ballutes to achieve their full potential.

McRonald, Angus D.↗
NASA NTRS2000

X-38 Landing Gear Skid Test Report

NASA incorporates skid-equipped landing gear on its series of X-38 flight test vehicles. The X-38 test program is the proving ground for the Crew Return Vehicle (CRV) a gliding parafoil-equipped vehicle designed to land at relatively low speeds. The skid-equipped landing gear is designed to attenuate the vertical landing energy of the vehicle at touchdown using crushable materials within the struts themselves. The vehicle then slides out as the vehicle horizontal energy is dissipated through the skids. A series of tests was conducted at Edwards Airforce Base (EAFB) in an attempt to quantify the drag force produced while "dragging" various X-38 landing gear skids across lakebed regions of varying surface properties. These data were then used to calculate coefficients of friction for each condition. Coefficient of friction information is critical for landing analyses as well as for landing gear load and interface load analysis. The skid specimens included full- and sub-scale V201 (space test vehicle) nose and main gear designs, a V131/V 132 (atmospheric flight test vehicles) main gear skid (actual flight hardware), and a newly modified, full-scale V201 nose -ear skid with substantially increased edge curvature as compared to its original design. Results of the testing are discussed along with comments on the relative importance of various parameters that influence skid stability and other dynamic behavior.

Gafka, George K.↗
NASA NTRS1994

Model Based Vision for Aircraft Position Determination

This paper investigates the use of imaging sensors to estimate the position of an aircraft with respect to the runway during landing. Passive vision techniques to estimate aircraft position during landing rely on the known runway model, images acquired by onboard imaging sensor, orientation information provided by the inertial navigation system and the position estimate provided by devices such as the global positioning system. Point features in the runway model are compared with the onboard sensor images of the features and the difference between the two is used to correct the aircraft position and orientation. In this paper the sensitivity of point features is examined as a means of determining the accuracy of such position estimation techniques. Expressions are derived for the sensitivity of image point to errors in the position and orientation of the sensor. Using these, the sensitivity of the image to aircraft position and orientation errors along a typical landing glide path is studied. A least squares technique based on this sensitivity analysis is described for the correction of position and orientation estimates. The final version of the paper will include results from the application of this analysis to real image sequences collected in flight.

Sridhar, Banavar↗
NASA NTRS1994

Design of a Flush Airdata System (FADS) for the Hypersonic Air Launched Option (HALO) Vehicle

This paper presents a design study for a pressure based Flush airdata system (FADS) on the Hypersonic Air Launched Option (HALO) Vehicle. The analysis will demonstrate the feasibility of using a pressure based airdata system for the HALO and provide measurement uncertainty estimates along a candidate trajectory. The HALO is a conceived as a man-rated vehicle to be air launched from an SR-71 platform and is proposed as a testbed for an airbreathing hydrogen scramjet. A feasibility study has been performed and indicates that the proposed trajectory is possible with minimal modifications to the existing SR71 vehicle. The mission consists of launching the HALO off the top of an SR-71 at Mach 3 and 80,000 ft. A rocket motor is then used to accelerate the vehicle to the test condition. After the scramjet test is completed the vehicle will glide to a lakebed runway landing. This option provides reusability of the vehicle and scramjet engine. The HALO design will also allow for various scramjet engine and flowpath designs to be flight tested. For the HALO flights, measurements of freestream airdata are considered to be a mission critical to perform gain scheduling and trajectory optimization. One approach taken to obtaining airdata involves measurement of certain parameters such as external atmospheric winds, temperature, etc to estimate the airdata quantities. This study takes an alternate approach. Here the feasibility of obtaining airdata using a pressure-based flush airdata system (FADS) methods is assessed. The analysis, although it is performed using the HALO configuration and trajectory, is generally applicable to other hypersonic vehicles. The method to be presented offers the distinct advantage of inferring total pressure, Mach number, and flow incidence angles, without stagnating the freestream flow. This approach allows for airdata measurements to be made using blunt surfaces and significantly diminishes the heating load at the sensor. In the FADS concept a matrix of flush ports is placed in the vicinity of the aircraft nose, and the airdata are inferred indirectly from the measured pressures.

Whitmore, Stephen A.↗
NASA NTRS2001

Rheology of Diabase: Implications for Tectonics on Venus and Mars

Two important goals of our experimental investigation of the rheological behavior of diabase rocks were: (1) to determine flow laws describing their creep behavior over wide ranges of temperature, stress and strain rate and (2) to develop an understanding of the physical mechanisms by which these rocks flow under laboratory conditions. With this basis, a primary objective then was to construct constitutive equations that can be used to extrapolate from laboratory to planetary conditions. We specifically studied the rheological properties of both natural rock samples and synthetic aggregates. The former provided constraints for geologic systems, while the latter defined the relative contributions of the constituent mineral phases and avoided the influence of glass/melt found in natural samples. In addition, partially molten samples of crustal rock composition were deformed in shear to large strains (greater than 200%) important in crustal environments. The results of this research yielded essential rheological properties essential for models of crustal deformation on terrestrial planets, specifically Venus and Mars, as well as on the geodynamical evolution of these planets. Over the past three years, we also completed our investigation of the creep behavior of water ice with applications to the glaciers, ice sheets and icy satellites. Constitutive equations were determined that describe flow over a wide ranged of stress, strain rate, grain size and temperature. In the case of ice, three creep regimes were delineate. Extrapolation demonstrates that dislocation glide and grain boundary sliding processes dominate flow in ice I under planetary conditions and that diffusion creep is not an important deformation mechanism either in the laboratory or on icy satellites. These results have already been incorporated by other investigators into models describing, for example, the thickness and stability of the ice shell on Europa and to unravel long-standing discrepancies between field observations on glaciers and laboratory results.

Kohlstedt, David L.↗
NASA NTRS2002

Reduction of Base Drag on Launch Vehicles

Current reentry vehicle designs exhibit large amounts of base drag due to large base areas. These large base areas can arise from the integration of the propulsion system (X-33) or control surface placement (X-38). Large base drag limits the vehicle's cross-range capability and causes a large glide-slope angle. Fortunately, there appears to be a possible means of lowering the base drag on these vehicles. Based on early work on the subsonic aerodynamics of lifting bodies, it appears that the addition of small amounts of viscous fore-body drag can produce a significant reduction in base drag. Recent work suggests that this phenomenon also occurs in the transonic and supersonic flight regimes. This report summarizes a study designed to demonstrate the reduction of base drag through the addition of fore-body viscous drag. The present study has focused on the measurement of viscous fore-body drag and the demonstration of the relationship between fore-body viscous drag and base drag at Reynolds Numbers up to 2.5 x 10(exp 6). The results of the present work do not conclusively demonstrate that viscous fore-body drag reduces base drag. The apparent contradictory results of the present study are attributed to the different geometry used in the present study. However, the results suggest that the increased boundary layer thickness at separation caused by larger fore-body viscous drag somehow affects the vortex structure in the wake thereby reducing the base drag. More research is required to confirm this postulated mechanism.

Naughton, Jonathan W.↗
NASA NTRS2002

APEX 3D Propeller Test Preliminary Design

A low Reynolds number, high subsonic mach number flight regime is fairly uncommon in aeronautics. Most flight vehicles do not fly under these aerodynamic conditions. However, recently there have been a number of proposed aircraft applications (such as high altitude observation platforms and Mars aircraft) that require flight within this regime. One of the main obstacles to flight under these conditions is the ability to reliably generate sufficient thrust for the aircraft. For a conventional propulsion system, the operation and design of the propeller is the key aspect to its operation. Due to the difficulty in experimentally modeling the flight conditions in ground-based facilities, it has been proposed to conduct propeller experiments from a high altitude gliding platform (APEX). A preliminary design of a propeller experiment under the low Reynolds number, high mach number flight conditions has been devised. The details of the design are described as well as the potential data that will be collected.

Colozza, Anthony J.↗
NASA NTRS2002

Launch Vehicle Ascent Stage Separation Wind Tunnel Test

The Aerodynamic Research Facility (ARF) LGBB (Liquid Glide-back Booster) Stage Separation Test is part of the Multi-Center Second Generation In-House Tool Development Task. The ARF LGBB Stage Separation Test has been completed at MSFC (Marshall Space Flight Center). It includes the following: PSP (Project Study Plan) Feasibility Test; Isolated Force/Moment Data; Bimese Configuration Force/Moment Data; Schlieren Video. The LGBB Bimese Reference Configuration Analyses and Test Results In-Work to Develop Tools and Database. Preliminary results showed qualitative agreement with CFD (computational fluid dynamics) aerodynamic predictions. The preliminary results exhibit the complex nature of the stage separation aerothermal problem.

Bordelon, Wayne↗
NASA NTRS2002

On-Board Generation of Three-Dimensional Constrained Entry Trajectories

A methodology for very fast design of 3DOF entry trajectories subject to all common inequality and equality constraints is developed. The approach make novel use of the well known quasi-equilibrium glide phenomenon in lifting entry as a center piece for conveniently enforcing the inequality constraints which are otherwise difficulty to handle. The algorithm is able to generate a complete feasible 3DOF entry trajectory, given the entry conditions, values of constraint parameters, and final conditions in about 2 seconds on a PC. Numerical simulations with the X-33 vehicle model for various entry missions to land at Kennedy Space Center will be presented.

Shen, Zuojun↗
NASA NTRS1976

Hang Gliders for Sport

Hang gliding is growing rapidly. Free Flight produces 1,000 gliders a month and other companies are entering the field. Wing is simple to control, pulling back on control bar allows you to pick up speed and at the same time lowers your altitude. Pushing forward slows your speed and levels you off. Birdmen can choose from prone, upright or swing seat harnesses in either kits or ready-to-fly gliders.

Source record↗
NASA NTRS2003

Elevated Temperature Creep Deformation in Solid Solution <001> NiAL-3.6Ti Single Crystals

The 1100 to 1500 K slow plastic strain rate compressive properties of <001> oriented NiAl-3.6Ti single crystals have been measured, and the results suggests that two deformation processes exist. While the intermediate temperature/faster strain rate mechanism is uncertain, plastic flow at elevated temperature/slower strain rates in NiAl-3.6Ti appears to be controlled by solute drag as described by the Cottrell-Jaswon solute drag model for gliding b = a(sub 0)<101> dislocations. While the calculated activation energy of deformation is much higher (approximately 480 kJ/mol) than the activation energy for diffusion (approximately 290 kJ/mol) used in the Cottrell-Jaswon creep model, a forced temperature compensated - power law fit using the activation energy for diffusion was able to adequately (greater than 90%) predict the observed creep properties. Thus we conclude that the rejection of a diffusion controlled mechanism can not be simply based on a large numerical difference between the activation energies for deformation and diffusion.

Whittenberger, J. Daniel↗