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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 109 records · Page 6
All sourcesNASA NTRSDOE OSTI
NASA NTRS1983

Turbulent flame propagation and combustion in spark ignition engines

Pressure measurements synchronized with high-speed motion-picture records of flame propagation have been made in a transparent-piston engine. The data show that the initial expansion speed of the flame front is close to that of a laminar flame. As the flame expands, its speed rapidly accelerates to a quasi-steady value comparable with that of the turbulent velocity fluctuations in the unburned gas. During the quasi-steady propagation phase, a significant fraction of the gas behind the visible front is unburned. Final burnout of the charge may be approximated by an exponential decay in time. The data have been analyzed in a model-independent way to obtain a set of empirical equations for calculating mass burning rates in spark-ignition engines. The burning equations contain three parameters: the laminar burning speed, a characteristic speed (uT), and a characteristic length (lT). The laminar burning speed is known from laboratory measurements. Tentative correlations relating uT and lT to engine geometry and operating variables have been derived from the engine data.

Beretta, G. P.↗
NASA NTRS1983

Latchup in CMOS devices from heavy ions

It is noted that complementary metal oxide semiconductor (CMOS) microcircuits are inherently latchup prone. The four-layer n-p-n-p structures formed from the parasitic pnp and npn transistors make up a silicon controlled rectifier. If properly biased, this rectifier may be triggered 'ON' by electrical transients, ionizing radiation, or a single heavy ion. This latchup phenomenon might lead to a loss of functionality or device burnout. Results are presented from tests on 19 different device types from six manufacturers which investigate their latchup sensitivity with argon and krypton beams. The parasitic npnp paths are identified in general, and a qualitative rationale is given for latchup susceptibility, along with a latchup cross section for each type of device. Also presented is the correlation between bit-flip sensitivity and latchup susceptibility.

Soliman, K.↗
NASA NTRS1984

T/R switch design for short-range measurements, part 6.1A

The positive intrinsic negative (PIN) diode switch which is designed to protect the receiver from burnout or damage on transmission and channel the echo signal to the receiver on reception is outlined. The receiver must be protected firmly. A schematic diagram of a transformer rectifier (TR-ATR) switch for the Urbana Radar is shown. The T/R switch consists of a half wavelength coaxial cavity with tuning condenser and PIN diodes. Two UM4300 PIN diodes were mounted between the inner and outer conductor. The dc biasing voltage required for the PIN diodes is supplied by a control circuit. On transmission, the PIN diodes are forward biased to about 0.5 amperes. On reception, about 10 volts reverse voltage is applied to the diodes, which produces an initial reverse current to speed the recovery time. The T/R switch characteristics are estimated and the result of testing at different peak transmitter powers from 410 kW to 1500 kW is shown.

Yu, B.↗
NASA NTRS1982

Scout trajectory error propagation computer program

Since 1969, flight experience has been used as the basis for predicting Scout orbital accuracy. The data used for calculating the accuracy consists of errors in the trajectory parameters (altitude, velocity, etc.) at stage burnout as observed on Scout flights. Approximately 50 sets of errors are used in Monte Carlo analysis to generate error statistics in the trajectory parameters. A covariance matrix is formed which may be propagated in time. The mechanization of this process resulted in computer program Scout Trajectory Error Propagation (STEP) and is described herein. Computer program STEP may be used in conjunction with the Statistical Orbital Analysis Routine to generate accuracy in the orbit parameters (apogee, perigee, inclination, etc.) based upon flight experience.

Myler, T. R.↗
NASA NTRS1986

Combustion theory for liquids with a free surface. 3: Special problems

Two special problems concerning the combustion of liquids with a free surface, i.e., flame quenching during the mixing of a burning liquid inside a container and liquid burnout from a porous layer, are analyzed using a quasi-one-dimensional model. The critical parameters corresponding to the quenching of a burning fluid with a free surface are determined. Determinations are also made of the limiting pressure gradients corresponding to the transition from the combustion mode where the liquid evaporates from the surface of a porous layer to the mode where the phase transition surface lies inside the porous layer.

Milkov, S. N.↗
NASA NTRS1987

A model for fluid flow during saturated boiling on a horizontal cylinder

A model has been developed to represent the vapor removal pattern in the vicinity of a cylinder during nucleate flow boiling across a horizontal cylinder. The model is based on a potential flow representation of the liquid and vapor regions and an estimate of the losses that should occur in the flow. Correlation of the losses shows a weak dependence on the Weber number and a slightly stronger dependence on the saturated liquid-to-vapor density ratio. The vapor jet thickness, which is crucial to the prediction of the burnout heat flux, and the shape of the vapor film are predicted. Both are verified by qualitative experimental observations.

Kheyrandish, K.↗
NASA NTRS1987

Space Shuttle guidance for multiple main engine failures during first stage

This paper presents contingency abort guidance schemes recently developed for multiple Space Shuttle main engine failures during the first two minutes of flight (first stage). The ascent and entry guidance schemes greatly improve the possibility of the crew and/or the Orbiter surviving a first stage contingency abort. Both guidance schemes were required to meet certain structural and controllability constraints. In addition, the systems were designed with the flexibility to allow for seasonal variations in the atmosphere and wind. The ascent scheme guides the vehicle to a desirable, lofted state at solid rocket booster burnout while reducing the structural loads on the vehicle. After Orbiter separation from the solid rockets and the external tank, the entry scheme guides the Orbiter through one of two possible entries. If the proper altitude/range/velocity conditions have been met, a return-to-launch-site 'Split-S' maneuver may be attempted. Otherwise, a down-range abort to an equilibrium glide and subsequent crew bailout is performed.

Sponaugle, Steven J.↗
NASA NTRS1987

Transition boiling heat transfer and the film transition regime

The Berenson (1960) flat-plate transition-boiling experiment has been recreated with a reduced thermal resistance in the heater, and an improved access to those portions of the transition boiling regime that have a steep negative slope. Tests have been made in Freon-113, acetone, benzene, and n-pentane boiling on horizontal flat copper heaters that have been mirror-polished, 'roughened', or teflon-coated. The resulting data reproduce and clarify certain features observed by Berenson: the modest surface finish dependence of boiling burnout, and the influence of surface chemistry on both the minimum heat flux and the mode of transition boiling, for example. A rational scheme of correlation yields a prediction of the heat flux in what Witte and Lienhard (1982) previously identified as the 'film-transition boiling' region. It is also shown how to calculate the heat flux at the boundary between the pure-film, and the film-transition, boiling regimes, as a function of the advancing contact angle.

Ramilison, J. M.↗
NASA NTRS1988

Programmable Pulser

User controls number of clock pulses to prevent burnout. New digital programmable pulser circuit in three formats; freely running, counted, and single pulse. Operates at frequencies up to 5 MHz, with no special consideration given to layout of components or to terminations. Pulser based on sequential circuit with four states and binary counter with appropriate decoding logic. Number of programmable pulses increased beyond 127 by addition of another counter and decoding logic. For very large pulse counts and/or very high frequencies, use synchronous counters to avoid errors caused by propagation delays. Invaluable tool for initial verification or diagnosis of digital or digitally controlled circuity.

Baumann, Eric↗
NASA NTRS1988

Hypersonic drone design: A multidisciplinary experience

Efforts were focused on design problems of an unmanned hypersonic vehicle. It is felt that a scaled hypersonic drone is necessary to bridge the gap between present theory on hypersonics and the future reality of the National Aerospace Plane (NASP) for two reasons: to fulfill a need for experimental data in the hypersonic regime, and to provide a testbed for the scramjet engine which is to be the primary mode of propulsion for the NASP. Three areas of great concern to NASP design were examined: propulsion, thermal management, and flight systems. Problem solving in these areas was directed towards design of the drone with the idea that the same design techniques could be applied to the NASP. A seventy degree swept double delta wing configuration, developed in the 70's at NASA Langley, was chosen as the aerodynamic and geometric model for the drone. This vehicle would be air-launched from a B-1 at Mach 0.8 and 48,000 feet, rocket boosted by two internal engines to Mach 10 and 100,000 feet, and allowed to cruise under power of the scramjet engine until burnout. It would then return to base for an unpowered landing. Preliminary energy calculations based upon the flight requirements give the drone a gross launch weight of 134,000 lb. and an overall length of 85 feet.

Source record↗
NASA NTRS1988

A new hydrodynamic prediction of the peak heat flux from horizontal cylinders in low speed upflow

Flow-boiling data have been obtained for horizontal cylinders in saturated acetone, isopropanol, and water, yielding heat flux vs. wall superheat boiling curves for the organic liquids. A region of low speed upflow is identified in which long cylindrical bubbles break off from the wake with regular frequency. The Strouhal number of bubble breakoff is a function only of the Froude number in any liquid, and the effective wake thickness in all liquids is a function of the density ratio and the Froude number. A low speed flow boiling burnout prediction procedure is presented which yields accurate results in widely dissimilar liquids.

Ungar, E. K.↗
NASA NTRS1990

Wrapped Wire Detects Rupture Of Pressure Vessel

Simple, inexpensive technique helps protect against damage caused by continuing operation of equipment after rupture or burnout of pressure vessel. Wire wrapped over area on outside of vessel where breakthrough most likely. If wall breaks or burns, so does wire. Current passing through wire ceases, triggering cutoff mechanism stopping flow in vessel to prevent further damage. Applied in other situations in which pipes or vessels fail due to overpressure, overheating, or corrosion.

Hunt, James B.↗
NASA NTRS1990

Fast round-trip Mars trajectories

This paper is concerned with the effect of limiting the overall duration or else the one-way flight time of a round trip to Mars, as reflected in the sum of impulsive velocity increments required of the spacecraft propulsion system. Ignition-to-burnout mass ratios for a hypothetical single stage spacecraft, obtained from the rocket equation by combining these delta-V sums with appropriate values of specific impulse, are used to evaluate the relative effectiveness of four high-thrust propulsion alternatives. If the flight crew goes to the surface of Mars and stays there for the duration of their stopover, it is much cheaper (in terms of delta-V) to minimize their zero-g exposure by limiting the interplanetary transit time of a conjunction-class mission (round trip time = 800-1000 days, Mars stopover = 450-700 days) than to impose the same limit on an opposition-class mission (round trip time less than 600 days, stopover = 40 days). Using solid-core nuclear thermal propulsion to fly a conjunction-class mission, for a moderate mass penalty the interplanetary transit time (each way) probably could be limited to something in the range of 4 to 6 months, depending on the launch year.

Wilson, Sam↗
NASA NTRS1991

Combustion of liquid fuel droplets in supercritical conditions

A comprehensive analysis of liquid-fuel droplet combustion in both sub- and super-critical environments has been conducted. The formulation is based on the complete conservation equations for both gas and liquid phases, and accommodates finite-rate chemical kinetics and a full treatment of liquid-vapor phase equilibrium at the droplet surface. The governing equations and the associated interface boundary conditions are solved numerically using a fully coupled, implicit scheme with the dual time-stepping integration technique. The model is capable of treating the entire droplet history, including the transition from the subcritical to the supercritical state. As a specific example, the combustion of n-pentane fuel droplets in air is studied for pressures of 5-140 atm. Results indicate that the ambient gas pressure exerts significant control of droplet gasification and burning processes through its influences on the fluid transport, gas/liquid interface thermodynamics, and chemical reactions. The droplet gasification rate increases progressively with pressure. However, the data for the overall burnout time exhibits a significant variation near the critical burning pressure, mainly as a result of reduced mass-diffusion rate and latent heat of vaporization with increased pressure. The influence of droplet size on the burning characteristics is also noted.

Shuen, J. S.↗
NASA NTRS1992

Ignition and combustion characteristics of metallized propellants

During this reporting period, theoretical work on the secondary atomization process was continued and the experimental apparatus was improved. A one-dimensional model of a rocket combustor, incorporating multiple droplet size classes, slurry combustion, secondary atomization, radiation heat transfer, and two-phase slip between slurry droplets and the gas flow was derived and a computer code was written to implement this model. The STANJAN chemical equilibrium solver was coupled with this code to yield gas temperature, density, and composition as functions of axial location. Preliminary results indicate that the model is performing correctly, given current model assumptions. Radiation heat transfer in the combustion chamber is treated as an optically-thick participating media problem requiring a solution of the radiative transfer equation. A cylindrical P sub 1 approximation was employed to yield an analytical expression for chamber-wall heat flux at each axial location. The code exercised to determine the effects of secondary atomization intensity, defined as the number of secondary drops produced per initial drop, on chamber burnout distance and final Al2O3 agglomerate diameter. These results indicate that only weak secondary atomization is required to significantly reduce these two parameters. Stronger atomization intensities were found to yield decreasing marginal benefits. The experimental apparatus was improved to reduce building vibration effects on the optical system alignment. This was accomplished by mounting the burner and the transmitting/receiving optics on a single frame supported by vibration-isolation legs. Calibration and shakedown tests indicate that vibration problems were eliminated and that the system is performing correctly.

Mueller, D. C.↗
NASA NTRS1992

The Superheat Phenomenon in the Combustion of Magnesium Particles

Magnesium is known to be a likely fuel for engines that could work in the CO2 atmospheres of Mars and Venus. The present paper reports temperature measurements of magnesium samples during combustion in CO2. The burning sample temperature increases with the decrease in the initial size. The temperature of the 1-mm samples is 300-400 K higher than the boiling point of magnesium. The stability of the superheated drop is explained by the presence of a porous shell on the surface. An attempt has been made to describe vaporization on the superheated drop by the Knudsen-Langmuir equation. During combustion at high-pressure fragment ejection of the flame is observed in high-speed motion pictures. This phenomenon is shown to be connected with the drop superheat. The repeated fracture of the outer shell formed in the flame ensures the complete burnout of metal particles at high pressure.

Shafirovich, E. IA.↗
NASA NTRS1992

Combustion of liquid-fuel droplets in supercritical conditions

A comprehensive analysis of liquid-fuel droplet combustion in both subcritical and supercritical environments has been conducted. The formulation is based on the complete conservation equations for both gas and liquid phases, and accommodates variable thermophysical properties, finite-rate chemical kinetics, and a full treatment of liquid-vapor phase equilibrium at the drop surface. The governing equations and associated interfacial boundary conditions are solved numerically using a fully coupled, implicit scheme with the dual time-stepping integration technique. The model is capable of treating the entire droplet history, including the transition from the subcritical to supercritical state. As a specific example, the combustion of n-pentane fuel droplets in air is studied for pressures in the range of 5-140 atm. Results indicate that the ambient gas pressure exerts significant control of droplet gasification and burning processes through its influence on fluid transport, gas-liquid interfacial thermodynamics, and chemical reactions. The droplet gasification rate increases progressively with pressure. However, the data for the overall burnout time exhibit a considerable change in the combustion mechanism at the critical pressure, mainly as a result of reduced mass diffusivity and latent heat of vaporization with increased pressure.

Shuen, J. S.↗
NASA NTRS1992

Lunar scout: A Project Artemis proposal

The results of a student project to design a lunar lander in the context of a specifically defined mission are presented. The Lunar Scout will be launched from Cape Canaveral, Florida onboard a Delta II launch vehicle. The Delta II will carry the lander and its payload to a 1367 km orbit. Once it reaches that altitude, a STAR 48A solid rocket motor will kick the spacecraft into a lunar trajectory. After burnout of the lunar insertion motor, it will be jettisoned from the spacecraft. The flight from the earth to the moon will take approximately 106.4 hours. During this time the battery, which was fully charged prior to launch, will provide all power to the spacecraft. Every hour, the spacecraft will use its sun sensors and star trackers to update its position, maintain some stabilization and relay it back to earth using the dipole antennas. At the start of its lunar trajectory, the spacecraft will fire one of its 1.5 N thrusters to spin in at a very small rate. The main reason for this is to prevent one side of the spacecraft from overheating in the sun. When the spacecraft nears the moon, it will orient itself for the main retro burn. At an altitude of 200 km, a 4400 N bipropellant liquid thruster will ignite to slow the spacecraft. During the burn, the radar altimeter will be turned on to guide the spacecraft. The main retro rocket will slow the lander to 10 m/s at an approximate altitude of 40 km above the moon. From there, the space craft will use four 4.5 N hydrazine vertical thrusters and 1.5 N horizontal thrusters to guide the spacecraft to a soft landing. Once on the ground, the lander will shutoff the radar and attitude control systems. After the debris from the impact has settled, the six solar panels will be deployed to begin recharging the batteries and to power up the payload. The feedhorn antenna will then rotate to fix itself on the earth. Once it moves, it will stay in that position for the spacecraft's lifetime. The payload will then be activated to begin the lunar mission.

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