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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 325 records · Page 18
All sourcesNASA NTRSDOE OSTI
NASA NTRS1983

Flutter and forced response of mistuned rotors using standing wave analysis

A standing wave approach is applied to the analysis of the flutter and forced response of tuned and mistuned rotors. The traditional traveling wave cascade airforces are recast into standing wave arbitrary motion form using Pade approximants, and the resulting equations of motion are written in the matrix form. Applications for vibration modes, flutter, and forced response are discussed. It is noted that the standing wave methods may prove to be more versatile for dealing with certain applications, such as coupling flutter with forced response and dynamic shaft problems, transient impulses on the rotor, low-order engine excitation, bearing motion, and mistuning effects in rotors.

Bundas, D. J.↗
NASA NTRS2013

Effects of the Orion Launch Abort Vehicle Plumes on Aerodynamics and Controllability

Characterization of the launch abort system of the Multi-purpose Crew Vehicle (MPCV) for control design and accurate simulation has provided a significant challenge to aerodynamicists and design engineers. The design space of the launch abort vehicle (LAV) includes operational altitudes from ground level to approximately 300,000 feet, Mach numbers from 0-9, and peak dynamic pressure near 1300psf during transonic flight. Further complicating the characterization of the aerodynamics and the resultant vehicle controllability is the interaction of the vehicle flowfield with the plumes of the two solid propellant motors that provide attitude control and the main propulsive impulse for the LAV. These interactions are a function of flight parameters such as Mach number, altitude, dynamic pressure, vehicle attitude, as well as parameters relating to the operation of the motors themselves - either as a function of time for the AM, or as a result of the flight control system requests for control torque from the ACM. This paper discusses the computational aerodynamic modeling of the aerodynamic interaction caused by main abort motor and the attitude control motor of the MPCV LAV, showing the effects of these interactions on vehicle controllability.

Vicker, Darby↗
NASA NTRS2001

Combustion of Metals in Carbon Dioxide and Reduced-Gravity Environments

Ongoing exploration and future mission2001110444 s to Mars have given impetus to research on the use of natural resources of the planet. Since carbon dioxide (CO2) constitutes approximately 95% of the Mars atmosphere and since it reacts directly and vigorously with several metals, this investigation focuses on metal-CO2 reactions as a possible combination for rocket-propellant production and energy generation. Magnesium (Mg) has been initially selected as the metal fuel owing to its low ignition temperature and high specific impulse and burning rate in CO2. Our studies in this field started with low gravity (g) combustion tests of Mg in O2, CO2, and CO. Reduced gravity provided a clear picture of the burning phenomena by eliminating the intrusive buoyant flows in high-temperature metal reactions and by removing the destructive effect of gravity on the shape of molten metal samples. Suspended cylindrical metal samples of 2, 3, and 4-mm in diameter and length were radiatively ignited in low-g to generate free-floating samples exhibiting a spherically symmetric flame with increasing metal-oxide accumulation in an outer shell. For the Mg-CO2 combination, burning times twice as long as in normal-g and five times longer than in Mg-O2 flames were observed, revealing a diffusion-controlled reaction. The burning time is proportional to the square of the sample diameter. In tests conducted with pure CO, combustion was not possible without constant heating of the sample due to the formation of a thick carbon-containing coating around the Mg sample generated by surface reactions. The following work presents two new studies that attempt to explain some of the low-g experimental observations. First, a simplified one-dimensional, quasi-steady numerical model is developed to obtain temperature, species concentrations, and burning rates of the spherically symmetric diffusion flame around the Mg sample burning in O2 and CO2. Second, a Planar Laser Induced Fluorescence (PLIF) technique is implemented to provide spatially resolved measurements of magnesium oxide (MgO) in the reaction zone of Mg samples burning in O2 and CO2. These experiments reveal fundamental differences between the two combustion systems.

Branch, M. C.↗
NASA NTRS1972

Dual deflectable beam strip engine development.

This paper describes a dual beam thruster that has been designed, constructed, and tested. The system is suitable for two-axes attitude control and is comprised of two orthogonal strips, each capable of producing 0.30 mlb thrust and beam deflections of more than plus or minus 20 deg. The nominal specific impulse for the thruster is 5000 sec, and the thrust level from each strip can be varied from 0 to 100%. Neutralizer filaments that were developed and life tested over 2000 hours producing more than 40 mA of electron emission per watt of input power are also discussed. The system power required for clean ionizers is approximately 200 W.

Dulgeroff, C. R.↗
NASA NTRS1990

Gamma guidance of trajectories for coplanar, aeroassisted orbital transfer

The optimization and guidance of trajectories for coplaner, aeroassisted orbital transfer (AOT) from high Earth orbit (HEO) to low Earth orbit (LEO) are examined. In particular, HEO can be a geosynchronous Earth orbit (GEO). It is assumed that the initial and final orbits are circular, that the gravitational field is central and is governed by the inverse square law, and that at most three impulses are employed: one at HEO exit, one at atmospheric exit, and one at LEO entry. It is also assumed that, during the atmospheric pass, the trajectory is controlled via the lift coefficient. The presence of upper and lower bounds on the lift coefficient is considered. First, optimal trajectories are computed by minimizing the total velocity impulse (hence, the propellant consumption) required for AOT transfer. The sequential gradient-restoration algorithm (SGRA) is used for optimal control problems. The optimal trajectory is shown to include two branches: a relatively short descending flight branch (branch 1) and a long ascending flight branch (branch 2). Next, attention is focused on guidance trajectories capable of approximating the optimal trajectories in real time, while retaining the essential characteristics of simplicity, ease of implementation, and reliability. For the atmospheric pass, a feedback control scheme is employed and the lift coefficient is adjusted according to a two-stage gamma guidance law. Further improvements are possible via a modified gamma guidance which is more stable with respect to dispersion effects arising from navigation errors, variations of the atmospheric density, and uncertainties in the aerodynamic coefficients than gamma guidance trajectory. A byproduct of the studies on dispersion effects is the following design concept. For coplaner aeroassisted orbital transfer, the lift-range-to-weight ratio appears to play a more important role than the lift-to-drag ratio. This is because the lift-range-to-weight ratio controls mainly the minimum altitude (hence, the peak heating rate) of the guidance trajectory; on the other hand, the lift-to-drag ratio controls mainly the duration of the atmospheric pass of the guidance trajectory.

Miele, A.↗
NASA NTRS1992

Electromagnetic scattering in clouds

Techniques used to explain the nature of the optical effects of clouds on the light produced by lightning include a Monte Carlo simulation, an equivalent medium approach, and methods based on Boltzmann transport theory. A cuboidal cloud has been considered using transform methods and a diffusion approximation. Many simplifying assumptions have been used by authors to make this problem tractable. In this report, the cloud will have a spherical shape and its interior will consist of a uniform distribution of identical spherical water droplets. The source will be modeled as a Hertz dipole, electric or magnetic, inside or outside the cloud. An impulsive source is used. Superposition may be employed to obtain a sinusoid within an envelope which describes a lightning event. The problem is investigated by transforming to the frequency domain, obtaining Green's functions, and then using the Cagniard-DeHoop method to symbolically recover the time domain solution.

Solakiewicz, Richard↗
NASA NTRS2011

Variability of Thermosphere and Ionosphere Responses to Solar Flares

We investigated how the rise rate and decay rate of solar flares affect the thermosphere and ionosphere responses to them. Model simulations and data analysis were conducted for two flares of similar magnitude (X6.2 and X5.4) that had the same location on the solar limb, but the X6.2 flare had longer rise and decay times. Simulated total electron content (TEC) enhancements from the X6.2 and X5.4 flares were 6 total electron content units (TECU) and approximately 2 TECU, and the simulated neutral density enhancements were approximately 15% -20% and approximately 5%, respectively, in reasonable agreement with observations. Additional model simulations showed that for idealized flares with the same magnitude and location, the thermosphere and ionosphere responses changed significantly as a function of rise and decay rates. The Neupert Effect, which predicts that a faster flare rise rate leads to a larger EUV enhancement during the impulsive phase, caused a larger maximum ion production enhancement. In addition, model simulations showed that increased E x B plasma transport due to conductivity increases during the flares caused a significant equatorial anomaly feature in the electron density enhancement in the F region but a relatively weaker equatorial anomaly feature in TEC enhancement, owing to dominant contributions by photochemical production and loss processes. The latitude dependence of the thermosphere response correlated well with the solar zenith angle effect, whereas the latitude dependence of the ionosphere response was more complex, owing to plasma transport and the winter anomaly.

Qian, Liying↗
NASA NTRS1981

Impulsive radio discharges near Saturn

Nonthermal radio emissions from the Saturn system were first detected by the Voyager planetary radio astronomy (PRA) experiment on board Voyager 1 in January 1980. Since then emission between 100 kHz and 1 MHz from the planet, termed Saturn kilometric radiation (SKR), has been received almost continuously. A description is presented of eight characteristics which have been fairly well defined by the Voyager 1 encounter. These include a very flat broadband frequency spectrum, a period of approximately 10 h 10 min, a change in the envelope shape of episodes between pre and postencounter, an intensity population structure typical of plural populations, and an episodic structure of a width of approximately 180 deg. It was found that postencounter episodes continue for about three times as long as preencounter ones, and that postencounter bursts are left-circularly polarized at high frequencies. At least one episode shows the onset of high frequency events some time before that of lower frequency ones.

Evans, D. R.↗
NASA NTRS2017

Robust Exploration and Commercial Missions to the Moon Using NTR LANTR Propulsion and Lunar-Derived Propellants

The nuclear thermal rocket (NTR) has frequently been identified as a key space asset required for the human exploration of Mars. This proven technology can also provide the affordable access through cislunar space necessary for commercial development and sustained human presence on the Moon. In his post-Apollo Integrated Space Program Plan (1970-1990), Wernher von Braun, proposed a reusable nuclear thermal propulsion stage (NTPS) to deliver cargo and crew to the Moon to establish a lunar base before undertaking human missions to Mars. The NTR option was selected by von Braun because it was a demonstrated technology capable of generating both high thrust and high specific impulse (Isp 900 s) twice that of todays best chemical rockets. In NASAs Mars Design Reference Architecture (DRA) 5.0 study, the crewed Mars transfer vehicle used three 25 klbf Pewee engines the smallest and highest performing engine tested in the Rover program along with graphite composite fuel. Smaller, lunar transfer vehicles consisting of a NTPS using three approximately 16.5 klbf Small Nuclear Rocket Engines (SNREs), an in-line propellant tank, plus the payload can enable a variety of reusable lunar missions. These include cargo delivery and crewed lunar landing missions. Even weeklong tourism missions carrying passengers into lunar orbit for a day of sightseeing and picture taking are possible. The NTR can play an important role in the next phase of lunar exploration and development by providing an affordable in-space lunar transportation system (LTS) that can allow initial outposts to evolve into settlements supported by a variety of commercial activities such as in-situ propellant production used to supply strategically located propellant depots and transportation nodes. The utilization of iron-rich volcanic glass or lunar polar ice (LPI) deposits (each estimated at billions of metric tons) for propellant production can significantly reduce the launch mass requirements from Earth and can enable reusable, surface-based lunar landing vehicles (LLVs) using liquid oxygen/hydrogen (LOX/LH2) chemical rocket engines. Afterwards, LOX/LH2 propellant depots can be established in lunar equatorial and polar orbits to supply the LTS. At this point a modified version of the conventional NTR called the LOX-augmented NTR, or LANTR would be introduced into the LTS allowing bipropellant operation and leveraging the mission benefits of refueling with lunar-derived propellants for Earth return. The bipropellant LANTR engine utilizes the large divergent section of its nozzle as an afterburner into which oxygen is injected and supersonically combusted with nuclear preheated hydrogen emerging from the engines choked sonic throat essentially scramjet propulsion in reverse. By varying the oxygen-to-hydrogen mixture ratio, LANTR engines can operate over a range of thrust and Isp values while the reactor core power level remains relatively constant. Eventually, a LANTR-based LTS can enable a rapid commuter shuttle with one-way trip times to and from the Moon ranging from 36 to 24 hours. Even if only 1 of the extracted propellant from identified volcanic glass and polar ice deposits were available for use in lunar orbit, such a supply could support daily commuter flights to the Moon for many thousands of years! An evolutionary mission architecture is outlined and a variety of lunar missions and transfer vehicle designs are examined, along with the increasing demands on propellant production as mission complexity increases. A comparison of vehicle features and engine operating characteristics, for both NTR and LANTR engines, is also provided along with a brief discussion on the propellant production issues associated with using volcanic glass and LPI as source material.

Spacecraft design↗
NASA NTRS1987

Stochastic acceleration in the transrelativistic region and pion production in solar flares

The stochastic Fermi acceleration spectrum in the transrelativistic region obtained from a Monte-Carlo simulation for an energy-independent alpha(T) is much harder than the extension of the nonrelativistic analytic spectrum to this energy range for the same alpha(T). The latter, with alpha(T) = 0.043, was used to model the pion and nuclear line emissions for the impulsive phase of the 3 Jun. 1982 flare, as well as the 2.223 MeV emission from this flare. We find that the ratios of these three emissions for the Monte-Carlo spectrum with alpha(T) = 0.028 are essentially the same as those for the analytical spectrum with alpha(T) = 0.043. We also find that the acceleration time from approximately 30 MeV to approximately 1 GeV is less than or approximately = 10 s, consistent with the observations of the 3 Jun. 1982 flare.

Miller, James A.↗
NASA NTRS1993

Powdered aluminum and oxygen rocket propellants: Subscale combustion experiments

Aluminum combined with oxygen has been proposed as a potential lunar in situ propellant for ascent/descent and return missions for future lunar exploration. Engine concepts proposed to use this propellant have not previously been demonstrated, and the impact on performance from combustion and two-phase flow losses could only be estimated. Therefore, combustion tests were performed for aluminum and aluminum/magnesium alloy powders with oxygen in subscale heat-sink rocket engine hardware. The metal powder was pneumatically injected, with a small amount of nitrogen, through the center orifice of a single element O-F-O triplet injector. Gaseous oxygen impinged on the fuel stream. Hot-fire tests of aluminum/oxygen were performed over a mixture ratio range of 0.5 to 3.0, and at a chamber pressure of approximately 480 kPa (70 psia). The theoretical performance of the propellants was analyzed over a mixture ratio range of 0.5 to 5.0. In the theoretical predictions the ideal one-dimensional equilibrium rocket performance was reduced by loss mechanisms including finite rate kinetics, two-dimensional divergence losses, and boundary layer losses. Lower than predicted characteristic velocity and specific impulse performance efficiencies were achieved in the hot-fire tests, and this was attributed to poor mixing of the propellants and two-phase flow effects. Several tests with aluminum/9.8 percent magnesium alloy powder did not indicate any advantage over the pure aluminum fuel.

Meyer, Mike L.↗
NASA NTRS1985

An impulsive solar burst observed in H-alpha, microwaves, and hard X-rays

H-alpha, microwave, and hard X-ray observations of an unusually short duration impulsive spike burst are presented. The observations are analyzed, and it is found that the single spike is in fact composed of two separate acceleration episodes. The differences found in the time profiles for the two components stress the role of the decay rate and lead to a simple explanation for the often observed delay of the microwave peak. The approximate numbers of electrons responsible for the two types of emission are derived and compared.

Gary, D. E.↗
NASA NTRS2013

Multiwavelength Diagnostics of the Precursor and Main Phases of an M1.8 Flare on 2011 April 22

We study the temporal, spatial and spectral evolution of the M1.8 flare, which occurred in the active region 11195 (S17E31) on 2011 April 22, and explore the underlying physical processes during the precursor phase and their relation to the main phase. The study of the source morphology using the composite images in 131Å wavelength observed by the Solar Dynamics Observatory/Atmospheric Imaging Assembly and 6-14 kiloelectronvolts [from the Reuven Ramaty High Energy Solar Spectroscopic Imager (RHESSI)] revealed a multi-loop system that destabilized systematically during the precursor and main phases. In contrast, hard X-ray emission (20-50 kiloelectronvolts) was absent during the precursor phase, appearing only from the onset of the impulsive phase in the form of foot-points of emitting loops. This study also revealed the heated loop-top prior to the loop emission, although no accompanying foot-point sources were observed during the precursor phase. We estimate the flare plasma parameters, namely temperature (T), emission measure (EM), power-law index (gamma) and photon turn-over energy (to), and found them to be varying in the ranges 12.4-23.4 megakelvins, 0.0003-0.6 x 10 (sup 49) per cubic centimeter, 5-9 and 14-18 kiloelectronvolts, respectively, by forward fitting RHESSI spectral observations. The energy released in the precursor phase was thermal and constituted approximately 1 percent of the total energy released during the flare. The study of morphological evolution of the filament in conjunction with synthesized T and EM maps was carried out, which reveals (a) partial filament eruption prior to the onset of the precursor emission and (b) heated dense plasma over the polarity inversion line and in the vicinity of the slowly rising filament during the precursor phase. Based on the implications from multiwavelength observations, we propose a scheme to unify the energy release during the precursor and main phase emissions in which the precursor phase emission was originated via conduction front that resulted due to the partial filament eruption. Next, the heated leftover S-shaped filament underwent slow-rise and heating due to magnetic reconnection and finally erupted to produce emission during the impulsive and gradual phases.

gamma-rays↗
NASA NTRS2008

Pressure and Thrust Measurements of a High-Frequency Pulsed Detonation Tube

This paper describes measurements of a small-scale, high-frequency pulsed detonation tube. The device utilized a mixture of H2 fuel and air, which was injected into the device at frequencies of up to 1200 Hz. Pulsed detonations were demonstrated in an 8-inch long combustion volume, at about 600 Hz, for the quarter wave mode of resonance. The primary objective of this experiment was to measure the generated thrust. A mean value of thrust was measured up to 6.0 lb, corresponding to H2 flow based specific impulse of 2970 s. This value is comparable to measurements in H2-fueled pulsed detonation engines (PDEs). The injection and detonation frequency for this new experimental case was much higher than typical PDEs, where frequencies are usually less than 100 Hz. The compact size of the device and high frequency of detonation yields a thrust-per-unit-volume of approximately 2.0 pounds per cubic inch, and compares favorably with other experiments, which typically have thrust-per-unit-volume of order 0.01 pound per cubic inch. This much higher volumetric efficiency results in a potentially much more practical device than the typical PDE, for a wide range of potential applications, including high-speed boundary layer separation control, for example in hypersonic engine inlets, and propulsion for small aircraft and missiles.

Nguyen, N.↗
NASA NTRS1988

Performance of 10-kW class xenon ion thrusters

Presented are performance data for laboratory and engineering model 30 cm-diameter ion thrusters operated with xenon propellant over a range of input power levels from approximately 2 to 20 kW. Also presented are preliminary performance results obtained from laboratory model 50 cm-diameter cusp- and divergent-field ion thrusters operating with both 30 cm- amd 50 cm-diameter ion optics up to a 20 kW input power. These data include values of discharge chamber propellant and power efficiencies, as well as values of specific impulse, thruster efficiency, thrust and power. The operation of the 30 cm- and 50 cm-diameter ion optics are also discussed.

Patterson, Michael J.↗
NASA NTRS1988

Performance of 10-kW class xenon ion thrusters

Presented are performance data for laboratory and engineering model 30 cm-diameter ion thrusters operated with xenon propellant over a range of input power levels from approximately 2 to 20 kW. Also presented are preliminary performance results obtained from laboratory model 50 cm-diameter cusp- and divergent-field ion thrusters operating with both 30 cm- and 50 cm-diameter ion optics up to a 20 kW input power. These data include values of discharge chamber propellant and power efficiencies, as well as values of specific impulse, thruster efficiency, thrust and power. The operation of the 30 cm- and 50 cm-diameter ion optics are also discussed.

Patterson, Michael J.↗
NASA NTRS1989

Project Longshot

Project Longshot is an unmanned probe to our nearest star system, Alpha Centauri, 4.3 light years away. The Centauri system is a trinary system consisting of two central stars (A and B) orbiting a barycenter, and a third (Proxima Centauri) orbiting the two. The system is a declination of -67 degrees. The goal is to reach the Centauri system in 50 years. This time space was chosen because any shorter time would be impossible of the relativistic velocities involved, and any greater time would be impossible because of the difficulty of creating a spacecraft with such a long lifetime. Therefore, the following mission profile is proposed: (1) spacecraft is assembled in Earth orbit; (2) spacecraft escapes Earth and Sun in the ecliptic with a single impulse maneuver; (3) spacecraft changed declination to point toward Centauri system; (4) spacecraft accelerates to 0.1c; (5) spacecraft coasts at 0.1c for 41 years; (6) spacecraft decelerates upon reaching Centauri system; and (7) spacecraft orbits Centauri system, conducts investigations, and relays data to Earth. The total time to reach the Centauri system, taking into consideration acceleration and deceleration, will be approximately 50 years.

West, J. Curtis↗
NASA NTRS1990

Flow establishment in a generic scramjet combustor

The establishment of a quasi-steady flow in a generic scramjet combustor is studied for the case of a time varying inflow to the combustor. Such transient flow is characteristic of the reflected-shock tunnel and expansion-tube test facilities. Several numerical simulations of hypervelocity flow through a straight-duct combustor with either a side-wall-step fuel injector or a centrally-located strut injector are presented. Comparisons are made between impulsively started but otherwise constant flow conditions (typical of the expansion-tube or tailored operation of the reflected-shock tunnel) and the relaxing flow produced by the 'undertailored' operation of the reflected-shock tunnel. Generally the inviscid flow features, such as the shock pattern and pressure distribution, were unaffected by the time varying inlet conditions and approached steady state in approximately the times indicated by experimental correlations. However, viscous features, such as heat transfer and skin friction, were altered by the relaxing inlet flow conditions.

Jacobs, P. A.↗