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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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Propeller slipstream effects as determined from wing pressure distribution of a large-scale six-propeller vtol model at static thrust
Thin film plastic bags use to insulate cryogenic propellant by boil-off of propellant nearest heat leak
Improve method for Apollo propellant system decontamination and propellant tank drying
Propellants and propellant mixtures effects on MPD arc jet performance in test for ions and neutrals mixture acceleration to high velocity
Ultrasmall nanomotors (<100 nm) are highly desirable nanomachines for their size-specific advantages over their larger counterparts in applications spanning nanomedicine, directed assembly, active sensing, and environmental remediation. While there are extensive studies on motors larger than 100 nm, the design and understanding of ultrasmall nanomotors have been scant due to the lack of high-resolution imaging of their propelled motions with orientation and shape details resolved. Here, we report the imaging of the propelled motions of catalytically powered ultrasmall nanomotors─hundreds of them─at the nanometer resolution using liquid-phase transmission electron microscopy. These nanomotors are Pt nanoparticles of asymmetric shapes (“tadpoles” and “boomerangs”), which are colloidally synthesized and observed to be fueled by the catalyzed decomposition of NaBH4 in solution. Statistical analysis of the orientation and position trajectories of fueled and unfueled motors, coupled with finite element simulation, reveals that the shape asymmetry alone is sufficient to induce local chemical concentration gradient and self-diffusiophoresis to act against random Brownian motion. Our work elucidates the colloidal design and fundamental forces involved in the motions of ultrasmall nanomotors, which hold promise as active nanomachines to perform tasks in confined environments such as drug delivery and chemical sensing.
Mathematical analysis of flame propagation at high pressures over surface of igniting solid propellants and propellant ingredients in oxygen/inert atmospheres
Vapor phase cleaning method for Apollo propellant system decontamination and propellant tank drying
The combustion characteristics of hypergolic propellants are described. A research project was conducted to determine if the reaction control system engine propellants on Apollo spacecraft undergo explosive reaction when subjected to conditions present in the engine at the time of ignition. Combustion characteristics pertinent to the hard-start phenomenon are considered. The thermal stability of frozen mixtures of hydrazine-based fuels with nitrogen tetroxide was analyzed. Results of the tests are presented in the form of tables and graphs.
The effect of front-to-rear propeller spacing on the interaction noise of a counterrotation propeller model was measured at cruise conditions. The data taken at an axial Mach number of 0.80 behaved as expected: interaction noise was reduced with increased spacing. The data taken at M=0.76 and M=0.72 did not behave as expected. At some of the test conditions the noise was unchanged; others even showed noise increases with increased spacing. A possible explanation, involving the amount of downstream blade area impacted by the tip vortex, is presented.
The object of this report, the last of a series of four on propeller design, is to describe a simple method for determining whether the strength of a propeller of a standard form is sufficient for safe operation. An approximate method of stress analysis is also given.
This report is the fifth of a series giving the results obtained from wind tunnel tests on the interference drag and propulsive efficiency of nacelle-propeller-wing combinations. This report gives results of tests of an NACA cowled air-cooled engine nacelle with tractor propeller located in 12 positions with reference to a Clark Y biplane cellule.
As the conditions of the ultimate wake are concerned both theoretically and practically, the magnitude of the slipstream contraction has been calculated. It will be noted that the contraction in a representative case is of the order of only 1 percent of the propeller diameter. In consequence, all calculations need involve only first-order effects. Curves and tables are given for the contraction coefficient of two-blade and four-blade propellers for various values of the advance ratio; the contraction coefficient is defined as the contraction in the diameter of the wake helix in terms of the wake diameter at infinity. The contour lines of the wake helix are also shown at four values of the advance ratio in comparison with the contour lines for an infinite number of blades.
This report presents the results of an investigation conducted in the Langley 300 mph 7- by 10-foot wind tunnel for the purpose of determining the aerodynamic characteristics of a model wing-propeller combination, and of the wing and propeller separately at angles of attack up to 90 degrees. The tests covered thrust coefficients corresponding to free-stream velocities from zero forward speed to the normal range of cruising speeds. The results indicate that increasing the thrust coefficient increases the angle of attack for maximum lift and greatly diminishes the usual reduction in lift above the angle of attack for maximum lift.
The goal of this project was to perform a preliminary design of a long term, reusable transportation system between earth and Mars which would be capable of providing both artificial gravity and shelter from solar flare radiation. The heart of this system was assumed to be a Cycler spacecraft propelled by an ion propulsion system. The crew transfer vehicle was designed to be propelled by a nuclear-thermal propulsion system. Several Mars transportation system architectures and their associated space vehicles were designed.
Experimental combustion efficiencies of eleven propellant combinations were determined as a function of chamber length. Efficiencies were measured in terms of characteristic exhaust velocities at three chamber lengths and in terms of gas velocities. The data were obtained in a nominal 200-pound-thrust rocket engine. Injector and engine configurations were kept essentially the same to allow comparison of the performance. The data, except for those on hydrazine and ammonia-fluorine, agreed with predicted results based on the assumption that vaporization of the propellants determines the rate of combustion. Decomposition in the liquid phase may be.responsible for the anomalous behavior of hydrazine. Over-all heat-transfer rates were also measured for each combination. These rates were close to the values predicted by standard heat-transfer calculations except for the combinations using ammonia.
An altitude-wind-tunnel investigation has been made to determine the performance of a Curtiss 732-1C2-0 four-blade propeller on a YP-47M airplane at high blade loadings and engine power. Propeller characteristics were obtained for a range of power coefficients from 0.30 to 1.00 at free-stream Mach numbers of 0.40 and .50.
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