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At least 37 records · Page 2
All sourcesNASA NTRSDOE OSTI
NASA NTRS1999

Experimental and Numerical Study of Ammonium Perchlorate Counterflow Diffusion Flames

Many solid rocket propellants are based on a composite mixture of ammonium perchlorate (AP) oxidizer and polymeric binder fuels. In these propellants, complex three-dimensional diffusion flame structures between the AP and binder decomposition products, dependent upon the length scales of the heterogeneous mixture, drive the combustion via heat transfer back to the surface. Changing the AP crystal size changes the burn rate of such propellants. Large AP crystals are governed by the cooler AP self-deflagration flame and burn slowly, while small AP crystals are governed more by the hot diffusion flame with the binder and burn faster. This allows control of composite propellant ballistic properties via particle size variation. Previous measurements on these diffusion flames in the planar two-dimensional sandwich configuration yielded insight into controlling flame structure, but there are several drawbacks that make comparison with modeling difficult. First, the flames are two-dimensional and this makes modeling much more complex computationally than with one-dimensional problems, such as RDX self- and laser-supported deflagration. In addition, little is known about the nature, concentration, and evolution rates of the gaseous chemical species produced by the various binders as they decompose. This makes comparison with models quite difficult. Alternatively, counterflow flames provide an excellent geometric configuration within which AP/binder diffusion flames can be studied both experimentally and computationally.

Smooke, M. D.↗
NASA NTRS1973

Some experiments related to L-star instability in rocket motors

The role of solid phase heterogeneity on the low-pressure L-star instability of nonmetallized AP/PBAN propellants is explored. Four particle size distributions are employed in propellants that are otherwise identical. Over one hundred test firings were conducted in the 21/2 in. diameter L-star burner. Pressure time histories in the chamber and color movies of two firings constitute the raw data. An economical firing program was used which enables the interesting range of L-star values to be covered during a single firing (at a set mean pressure), through the variations in the depleting propellant volume. Time-independent combustion, Helmholtz mode, chuff mode, and the pressure-burst phenomena are revealed as the principal signatures. Of these, the Helmholtz mode is found to be the most ordered form of instability.

Kumar, R. N.↗
NASA NTRS1977

Combustion response modeling for composite solid propellants

A computerized mathematical model of the combustion response function of composite solid propellants was developed with particular attention to the contributions of the solid phase heterogeneity. The one-dimensional model treats the solid phase as alternating layers of ammonium perchlorate and binder, with an exothermic melt layer at the surface. Solution of the Fourier heat equation in the solid provides temperature and heat flux distributions with space and time. The problem is solved by conserving the heat flux at the surface from that produced by a suitable model of the gas phase. An approximation of the BDP flame model is utilized to represent the gas phase. By the use of several reasonable assumptions, it is found that a significant portion of the problem can be solved in closed form. A method is presented by which the model can be applied to tetramodal particle size distributions. A computerized steady-state version of the model was completed, which served to validate the various approximations and lay a foundation for the combustion response modeling. The combustion response modeling was completed in a form which does not require an iterative solution, and some preliminary results were acquired.

Source record↗
NASA NTRS1974

A study of the effects of solid phase reactions on the thermal degradation and ballistic properties of solid propellants

The thermal stability of perchlorate composite propellants was studied at 135 and 170 C. The experimental efforts were concentrated on determining the importance of heterogeneous oxidizer-fuel reactions in the thermal degradation process. The experimental approach used to elucidate the mechanisms by which the oxidizer fuel composites thermally degrade was divided into two parts: (1) keeping the fuel constant and varying the nature of the oxidizers, and (2) holding the oxidizer constant and varying the fuel components. The fuel component primarily utilized in the first phase was polyethylene. Oxidizers included KClO4, KClO3, NH4ClO4 and NH4ClO4 doped with materials such as chlorate, phosphate and arsenate. In the second phase the oxidizer used was primarily NH4ClO4 while the fuels included saturated and unsaturated polybutadiene prepolymers and a series of bonding agents. Techniques employed in the current study include thermogravimetric measurements, differential thermal analysis, infrared, mass spectrometry, electron microscopy, and appropriate wet chemical analysis.

Schmidt, W. G.↗
NASA NTRS2003

Monolithic Hydrogen Peroxide Catalyst Bed Development

With recent increased industry and government interest in rocket grade hydrogen peroxide as a viable propellant, significant effort has been expended to improve on earlier developments. This effort has been predominately centered in improving heterogeneous. typically catalyst beds; and homogeneous catalysts, which are typically solutions of catalytic substances. Heterogeneous catalyst beds have traditionally consisted of compressed wire screens plated with a catalytic substance, usually silver, and were used m many RCS applications (X-1, Mercury, and Centaur for example). Aerojet has devised a heterogeneous catalyst design that is monolithic (single piece), extremely compact, and has pressure drops equal to or less than traditional screen beds. The design consists of a bonded stack of very thin, photoetched metal plates, silver coated. This design leads to a high surface area per unit volume and precise flow area, resulting in high, stable, and repeatable performance. Very high throughputs have been demonstrated with 90% hydrogen peroxide. (0.60 lbm/s/sq in at 1775-175 psia) with no flooding of the catalyst bed. Bed life of over 900 seconds has also been demonstrated at throughputs of 0.60 lbm/s/sq in across varying chamber pressures. The monolithic design also exhibits good starting performance, short break-in periods, and will easily scale to various sizes.

Ponzo, J. B.↗
NASA NTRS1966

A study of heterogeneous detonations.

Heterogeneous detonations, discussing polydisperse and monodisperse spray detonations and liquid fuel film shock-induced combustion

COMBUSTION INSTABILITY↗
NASA NTRS1986

Mapping the Structure of Heterogeneous Materials

Image-processing microdensitometer/Fourier analyzer yields statistics of subcomponent distribution. Nondestructive method for studying structure heterogeneous materials uses energy-dispersive X-ray analysis in scanning electron microscope. Scanning microdensitometer/Fourier analyzer (SMFA) is applied to SEM images to obtain statistics about sample structure. Method originally developed for studying effect on combustion of fine structure of composite solid propellants.

Strand, L. D.↗
NASA NTRS1966

A study of heterogeneous detonations.

Heterogeneous detonations, discussing polydisperse and monodisperse spray detonations and liquid fuel film shock-induced combustion

COMBUSTION INSTABILITY↗
NASA NTRS1991

Heat transfer analysis of fuel assemblies in a heterogeneous gas core nuclear rocket

Heat transfer problems of a heterogeneous gaseous core nuclear rocket were studied. The reactor core consists of 1.5-m long hexagonal fuel assemblies filled with pressurized uranium tetrafluoride (UF4) gas. The fuel gas temperature ranges from 3500 to 7000 K at a nominal operating condition of 40 atm. Each fuel assembly has seven coolant tubes, through which hydrogen propellant flows. The propellant temperature is not constrained by the fuel temperature but by the maximum temperature of the graphite coolant tube. For a core achieving a fission power density of 1000 MW/cu m, the propellant core exit temperature can be as high as 3200 K. The physical size of a 1250 MW gaseous core nuclear rocket is comparable with that of a NERVA-type solid core nuclear rocket. The engine can deliver a specific impulse of 1020 seconds and a thrust of 330 kN.

Watanabe, Yoichi↗
NASA NTRS1972

The effect of solid phase reactions on the ballistic properties of propellants

The combustion of NHClO4 composite propellants was studied between 15 and 3000 psi. The emphasis was on determining the mechanisms by which the fuel components influence the burning rate of the composites. All combustion experiments were performed with pressed powder strands. The fuels affected the combustion mechanism of the composite through their effect on the oxidizer decomposition mechanism, the composite surface temperature, and by reacting heterogeneously with the oxidizer on the surface of the composite. The results indicate that the fuel component has a significant effect on the surface temperature in both catalyzed and uncatalyzed composites. The data suggest that the contribution, to the overall combustion reaction, of heterogeneous reactions increases with increasing pressure. There is also evidence that even in steady state combustion the ignition of individual oxidizer particles is an important parameter in determining the combustion characteristics of the composite.

Schmidt, W. G.↗
NASA NTRS1974

The hard start phenomena in hypergolic engines. Volume 4: The chemistry of hydrazine fuels and nitrogen tetroxide propellant systems

The various chemical reactions that occur and that could possibly occur in the RCS engines utilizing hydrazine-type fuel/nitrogen tetroxide propellant systems, prior to ignition (preignition), during combustion, and after combustion (postcombustion), and endeavors to relate the hard-start phenomenon to some of these reactions are discussed. The discussion is based on studies utilizing a variety of experimental techniques and apparatus as well as current theories of chemical reactions and reaction kinetics. The chemical reactions were studied in low pressure gas flow reactors, low temperature homogeneous- and heterogeneous-phase reactors, simulated two-dimensional (2-D) engines, and scaled and full size engines.

Miron, Y.↗
NASA NTRS1995

Atmospheric environmental implications of propulsion systems

Three independent studies have been conducted for assessing the impact of rocket launches on the earth's environment. These studies have addressed issues of acid rain in the troposphere, ozone depletion in the stratosphere, toxicity of chemical rocket exhaust products, and the potential impact on global warming from carbon dioxide emissions from rocket launches. Local, regional, and global impact assessments were examined and compared with both natural sources and anthropogenic sources of known atmospheric pollutants with the following conclusions: (1) Neither solid nor liquid rocket launches have a significant impact on the earth's global environment, and there is no real significant difference between the two. (2) Regional and local atmospheric impacts are more significant than global impacts, but quickly return to normal background conditions within a few hours after launch. And (3) vastly increased space launch activities equivalent to 50 U.S. Space Shuttles or 50 Russian Energia launches per year would not significantly impact these conclusions. However, these assessments, for the most part, are based upon homogeneous gas phase chemistry analysis; heterogeneous chemistry from exhaust particulates, such as aluminum oxide, ice contrails, soot, etc., and the influence of plume temperature and afterburning of fuel-rich exhaust products, need to be further addressed. It was the consensus of these studies that computer modeling of interactive plume chemistry with the atmosphere needs to be improved and computer models need to be verified with experimental data. Rocket exhaust plume chemistry can be modified with propellant reformulation and changes in operating conditions, but, based upon the current state of knowledge, it does not appear that significant environmental improvements from propellant formulation changes can be made or are warranted. Flight safety, reliability, and cost improvements are paramount for any new rocket system, and these important aspects cannot be compromised. A detailed environmental cost-benefit-risk analysis must be conducted before any new chemistry or changes in rocket operating conditions should be seriously considered for any future space or defense applications. This paper presents a summary of the results of environmental assessments contained in these independent studies.

Mcdonald, Allan J.↗
NASA NTRS2001

Combustion of Metals in Reduced-Gravity and Extraterrestrial Environments

As a result of the ongoing exploration of Mars and the several unmanned and possibly manned missions planned for the near future, increased attention has been given to the use of the natural resources of the planet for rocket propellant production and energy generation. Since the atmosphere of Mars consists of approximately 95% carbon dioxide (CO2), this gas is the resource of choice to be employed for these purposes. Since many metals burn vigorously with CO2, these may be used as an energy source or as propellants for a research vehicle on the surface of Mars. Shafirovich and Goldshleger conducted experiments with spherical particles up to 2.5 mm in diameter and found that the burning process was controlled by diffusion and that the particles exhibited pulsating combustion due to superheating of the Mg vapor trapped inside a protective oxide shell. They also proposed a reaction mechanism based on the gas-phase reaction, Mg + CO2 yields MgO + CO and the heterogeneous reaction Mg + CO yields MgO + C occurring on the sample surface. In all the above studies with large Mg particles, the burning process is invariably influenced by strong convective currents that accelerate the combustion reaction and shorten the burning times. Although these currents are nearly absent in the burning of small particles, the high emissivity of the flames, rapid reaction, and small length scales make the gathering of any useful information on burning rates and flame structure very difficult. The goal of this investigation is to provide a detailed study of flame structure by taking advantage of large, free-floating spherical metal samples and their corresponding long burning times available in a weightless environment. The use of reduced gravity is essential to eliminate the intrusive buoyant flows that plague high temperature metal reactions, to remove the destructive effect of gravity on the shape of molten metal samples, and to study the combustion behavior of metals in the presence of solid oxides undisturbed by natural convection. This work presents the most complete modeling of metal particle burning to date for Mg with CO2 and O2.

Branch, M. C.↗
NASA NTRSDate not supplied

Revisiting the Origin of Macromolecular Carbon (MMC) in Lunar Basalts 15556 & 10044

Volatile elements influence the geo-chemical evolution of planetary bodies and they are in magmas at every stage, from melting within planetary interiors to eruption at the surface. Analyses of lunar mare basalts supported the hypothesis that lunar mag-mas were depleted in volatiles (H-C-F-Cl-S), relative to their terrestrial analogs [1]. Nevertheless, several early studies of samples returned during the Apollo program proposed that the mare basalt eruptions, including the “fire fountain” eruptions, were propelled by the oxidation of magmatic graphite to CO (and/or CO2) gas [2, 3]. Seminal studies during the 1970’s measured the bulk concentration and isotopic compositions of C from Apollo 11 samples, and identified several carbonaceous compounds, including: (a) gaseous (CO, CO2, and traces of CH4), (b) metallic carbide, and (c) potentially elemental carbon [4-6]. These studies reported a relatively broad range of C contents (~100-400 μg/g) and isotopic values (δC13 = -30 to +20), and suggested that these heterogeneities can be explained by contribution from multiple factors, including: (a) indigenous carbon, (b) solar wind implantation, (c) bombardment and/or meteorite impact, and (d) terrestrial contamination [5,6]. However, unequivocal observations of magmatic graphite in lunar basalts have never been made. Macromolecular carbon (MMC)—graphitic carbon varying from nearly amorphous to highly crystalline varieties—was identified as inclusions hosted by igneous pyroxenes from Martian meteorites and were attributed to being indigenous to Mars [8]. The authors carefully considered the textural and mineralogical relationship of the MMC phases, and concluded that the subset of MMC located within and/or adjacent to cracks, or at a disrupted surface (e.g., cut) were most consistent with terrestrial contamination. The near absence of con-firmed instances of lunar magmatic MMC within the literature [9], combined with the wide range of isotopic values and bulk carbon contents measured in lunar bas-alts begs the question as to whether previously measured carbon is of an indigenous origin, or the result of terrestrial contamination. Using Raman spectroscopy, we have observed MMC in lunar basalts subjected to different forms of anthropogenic modification related to sample preparation including polished sections, sawn surfaces, and fractured surfaces adjacent to sawn sur-faces. We have observed MMC of unknown origin in all of these settings. Here we report the preliminary textural and spectroscopic characteristics of MMC hosted within the groundmass of Apollo 15 (15556) and Apollo 11 basalts (10044) as part of our ongoing investigation of the origin of these carbonaceous materials.

lunar↗
NASA NTRS1993

Ignition and combustion of lunar propellants

The ignition and combustion of Al, Mg, and Al/Mg alloy particles in 99 percent O2/1 percent N2 mixtures is investigated at high temperatures and pressures for rocket engine applications. The 20 micron particles contain 0, 5, 10, 20, 40, 60, 80, and 100 weight percent Mg alloyed with Al, and are ignited in oxygen using the reflected shock in a shock tube near the endwall. Using this technique, the ignition delay and combustion times of the particles are measured at temperatures up to 3250 K as a function of Mg content for oxygen pressures of 8.5, 17, and 34 atm. An ignition model is developed which employs a simple lumped capacitance energy equation and temperature and pressure dependent particle and gas properties. Good agreement is achieved between the measured and predicted trends in the ignition delay times. For the particles investigated, the contribution of heterogeneous reaction to the heating of the particle is found to be significant at lower temperatures, but may be neglected as gas temperatures above 3000 K. As little as 10 percent Mg reduces the ignition delay time substantially at all pressures tested. The particle ignition delay times decrease with increasing Mg content, and this reduction becomes less pronounced as oxidizer temperature and pressure are increased.

Burton, Rodney L.↗
NASA NTRS1999

Combustion of Metals in Reduced-Gravity and Extra Terrestrial Environments

The combustion of metals is a field with important practical applications in rocket propellants, high-temperature flames, and material synthesis. Also, the safe operation of metal containers in high-pressure oxygen systems and with cryogenic fuels and oxidizers remains an important concern in industry. The increasing use of metallic components in spacecraft and space structures has also raised concerns about their flammability properties and fire suppression mechanisms. In addition, recent efforts to embark on unmanned and manned planetary exploration, such as on Mars, have also renewed the interest in metal/carbon-dioxide combustion as an effective in situ resource utilization technology. In spite of these practical applications, the understanding of the combustion properties of metals remains far behind that of the most commonly used fuels such as hydrocarbons. The lack of understanding is due to the many problems unique to metal- oxidizer reactions such as: low-temperature surface oxidation prior to ignition, heterogeneous reactions, very high combustion temperatures, product condensation, high emissivity of products, and multi-phase interactions. Very few analytical models (all neglecting the influence of gravity) have been developed to predict the burning characteristics and the flame structure details. Several experimental studies attempting to validate these models have used small metal particles to recreate gravity-free conditions. The high emissivity of the flames, rapid reaction, and intermittent explosions experienced by these particles have made the gathering of any useful information on burning rates and flame structure very difficult. The use of a reduced gravity environment is needed to clarify some of the complex interactions among the phenomena described above. First, the elimination of the intrusive buoyant flows that plague all combustion phenomena is of paramount importance in metal reactions due to the much higher temperatures reached during combustion. Second, a low-gravity environment is absolutely essential to remove the destructive effect of gravity on the shape of a molten metal droplet in order to study a spherically symmetric condition with large bulk samples. The larger size of the spherical metal droplet and the longer burning times available in reduced gravity extend the spatial and temporal dimensions to permit careful probing of the flame structure and dynamics. Third, the influence of the radiative heat transfer from the solid oxides can be studied more carefully by generating a stagnant spherical shell of condensed products undisturbed by buoyancy.

Branch, M.C.↗
NASA NTRS1996

Parallelization of Rocket Engine System Software (Press)

The main goal is to assess parallelization requirements for the Rocket Engine Numeric Simulator (RENS) project which, aside from gathering information on liquid-propelled rocket engines and setting forth requirements, involve a large FORTRAN based package at NASA Lewis Research Center and TDK software developed by SUBR/UWF. The ultimate aim is to develop, test, integrate, and suitably deploy a family of software packages on various aspects and facets of rocket engines using liquid-propellants. At present, all project efforts by the funding agency, NASA Lewis Research Center, and the HBCU participants are disseminated over the internet using world wide web home pages. Considering obviously expensive methods of actual field trails, the benefits of software simulators are potentially enormous. When realized, these benefits will be analogous to those provided by numerous CAD/CAM packages and flight-training simulators. According to the overall task assignments, Hampton University's role is to collect all available software, place them in a common format, assess and evaluate, define interfaces, and provide integration. Most importantly, the HU's mission is to see to it that the real-time performance is assured. This involves source code translations, porting, and distribution. The porting will be done in two phases: First, place all software on Cray XMP platform using FORTRAN. After testing and evaluation on the Cray X-MP, the code will be translated to C + + and ported to the parallel nCUBE platform. At present, we are evaluating another option of distributed processing over local area networks using Sun NFS, Ethernet, TCP/IP. Considering the heterogeneous nature of the present software (e.g., first started as an expert system using LISP machines) which now involve FORTRAN code, the effort is expected to be quite challenging.

Cezzar, Ruknet↗