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Results for “PROPELLANT OXIDIZER”

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 NTRS1962

An Experimental Investigation of the Performance of the Nitrogen Tetroxide-Hydrazine System in the Oxidizer-Rich and Fuel-Rich Regions

The results of an experimental program directed toward determining some of the operational characteristics of the nitrogen tetroxide-hydrazine propellant system under oxidizer-rich and fuel-rich conditions are reported. Data are presented for the mixture-ratio ranges of 0 to 0.55 at a nominal chamber pressure of 300 psia and characteristic chamber length of 250 in., and at mixture ratios of 5.9 to 13.0 at nominal chamber pressures of 400 and 500 psia and characteristic chamber lengths of 100 and 3900 in. The relationship of the actual performance data obtained in each region with those predicted from thermochemical performance calculations is presented. From this comparison, it is concluded that in neither region are equilibrium conditions obtained and that utilization of performance data obtained from assumptions of equilibrium will lead to serious errors. In the oxidizer-rich region investigated, temperatures considerably below those predicted occur throughout the range of mixture ratios investigated because of the lack of exothermic dissociation of the nitrogen oxides. In the fuel-rich region, below a mixture ratio of 0.3, temperatures considerably higher than those predicted occur because of the lack of endothermic dissociation of ammonia.

Chilenski, J. J.↗
NASA NTRS2011

High energy, low temperature gelled bi-propellant formulation

The present invention is a bi-propellant system comprising a gelled liquid propane (GLP) fuel and a gelled MON-30 (70% N.sub.2O.sub.4+30% NO) oxidizer. The bi-propellant system is particularly well-suited for outer planet missions greater than 3 AU from the sun and also functions in earth and near earth environments. Additives such as powders of boron, carbon, lithium, and/or aluminum can be added to the fuel component to improve performance or enhance hypergolicity. The gelling agent can be silicon dioxide, clay, carbon, or organic or inorganic polymers. The bi-propellant system may be, but need not be, hypergolic.

Di Salvo, Roberto↗
NASA NTRS2000

Examples of Unsteady Combustion in Non-Metallized Propellants

For over thirty years the T-Burner has been the de facto standard for determining the unstable response of solid propellants. This paper reviews some of the different approaches that have been used, and briefly summarizes advantages and disadvantages of the different techniques. Data from a variety of sources have been reviewed and summarized showing the effect of various test conditions and propellant formulation variations. Much of the data taken in the 60's and early 70's were with so-called research propellants, with relatively low solids loading and usually monomodal AP as the oxidizer. Later data were obtained with propellants containing solids loading more typical and practical propellants. Several distinct differences were observed in the data for the different types of propellants. Data varying AP particle size distributions indicate that either additional fine AP or finer AP usually increases the propellant response. A significant reduction in the response can be achieved by reducing the amount of fine AP, while holding burning rate constant either by adding a catalyst or by varying the overall AP size distribution. A significant amount of data has been obtained for reduced smoke propellants containing stability additives. The data indicate that the additives provide significant particle damping, and most often, also reduce the propellant response. Data from double base propellants indicate a much broader response over a wide frequency range. This implies that double base propellants should be more unstable in tangential modes than composite propellants. This trend seems to be consistent with observed motor data. Measuring the response of metallized propellants has been a challenge due to the large damping due to metal oxide particles in the gas. Data from metallized propellants are not presented here, but will be summarized in a future paper.

Beckstead, M. W.↗
NASA NTRS1994

A Theoretical Evaluation of Secondary Atomization Effects on Engine Performance for Aluminum Gel Propellants

A one-dimensional model of a gel-fueled rocket combustion chamber has been developed. This model includes the processes of liquid hydrocarbon burnout, secondary atomization. aluminum ignition, and aluminum combustion. Also included is a model of radiative heat transfer from the solid combustion products to the chamber walls. Calculations indicate that only modest secondary atomization is required to significantly reduce propellant burnout distances, aluminum oxide residual size and radiation heat wall losses. Radiation losses equal to approximately 2-13 percent of the energy released during combustion were estimated. A two-dimensional, two-phase nozzle code was employed to estimate radiation and nozzle two-phase flow effects on overall engine performance. Radiation losses yielded a 1 percent decrease in engine I(sub sp). Results also indicate that secondary atomization may have less effect on two-phase losses than it does on propellant burnout distance and no effect if oxide particle coagulation and shear induced droplet breakup govern oxide particle size. Engine I(sub sp) was found to decrease from 337.4 to 293.7 seconds as gel aluminum mass loading was varied from 0-70 wt percent. Engine I(sub sp) efficiencies, accounting for radiation and two-phase flow effects, on the order of 0.946 were calculated for a 60 wt percent gel, assuming a fragmentation ratio of 5.

Mueller, D. C.↗
NASA NTRS1994

Ignition and combustion characteristics of metallized propellants, phase 2

Experimental and analytical investigations focusing on aluminum/hydrocarbon gel droplet secondary atomization and its effects on gel-fueled rocket engine performance are being conducted. A single laser sheet sizing/velocimetry diagnostic technique, which should eliminate sizing bias in the data collection process, has been designed and constructed to overcome limitations of the two-color forward-scatter technique used in previous work. Calibration of this system is in progress and the data acquisition/validation code is being written. Narrow-band measurements of radiant emission, discussed in previous reports, will be used to determine if aluminum ignition has occurred in a gel droplet. A one-dimensional model of a gel-fueled rocket combustion chamber, described in earlier reports, has been exercised in conjunction with a two-dimensional, two-phase nozzle code to predict the performance of an aluminum/hydrocarbon fueled engine. Estimated secondary atomization effects on propellant burnout distance, condensed particle radiation losses to the chamber walls, and nozzle two phase flow losses are also investigated. Calculations indicate that only modest secondary atomization is required to significantly reduce propellant burnout distances, aluminum oxide residual size, and radiation heat losses. Radiation losses equal to approximately 2-13 percent of the energy released during combustion were estimated, depending on secondary atomization intensity. A two-dimensional, two-phase nozzle code was employed to estimate radiation and nozzle two phase flow effects on overall engine performance. Radiation losses yielded a one percent decrease in engine Isp. Results also indicate that secondary atomization may have less effect on two-phase losses than it does on propellant burnout distance and no effect if oxide particle coagulation and shear induced droplet breakup govern oxide particle size. Engine Isp was found to decrease from 337.4 to 293.7 seconds as gel aluminum mass loading was varied from 0-70 wt percent. Engine Isp efficiencies, accounting for radiation and two phase flow effects, on the order of 0.946 were calculated for a 60 wt percent gel, assuming a fragmentation ratio of five.

Mueller, D. C.↗
NASA NTRS1993

Small rocket research and technology

Small chemical rockets are used on nearly all space missions. The small rocket program provides propulsion technology for civil and government space systems. Small rocket concepts are developed for systems which encompass reaction control for launch and orbit transfer systems, as well as on-board propulsion for large space systems and earth orbit and planetary spacecraft. Major roles for on-board propulsion include apogee kick, delta-V, de-orbit, drag makeup, final insertions, north-south stationkeeping, orbit change/trim, perigee kick, and reboost. The program encompasses efforts on earth-storable, space storable, and cryogenic propellants. The earth-storable propellants include nitrogen tetroxide (NTO) as an oxidizer with monomethylhydrazine (MMH) or anhydrous hydrazine (AH) as fuels. The space storable propellants include liquid oxygen (LOX) as an oxidizer with hydrazine or hydrocarbons such as liquid methane, ethane, and ethanol as fuels. Cryogenic propellants are LOX or gaseous oxygen (GOX) as oxidizers and liquid or gaseous hydrogen as fuels. Improved performance and lifetime for small chemical rockets are sought through the development of new predictive tools to understand the combustion and flow physics, the introduction of high temperature materials to eliminate fuel film cooling and its associated combustion inefficiency, and improved component designs to optimize performance. Improved predictive technology is sought through the comparison of both local and global predictions with experimental data. Results indicate that modeling of the injector and combustion process in small rockets needs improvement. High temperature materials require the development of fabrication processes, a durability data base in both laboratory and rocket environments, and basic engineering property data such as strength, creep, fatigue, and work hardening properties at both room and elevated temperature. Promising materials under development include iridium-coated rhenium and a ceramic composite of mixed hafnium carbide and tantalum carbide reinforced with graphite fibers.

Schneider, Steven↗
NASA NTRS2018

A 6U Hybrid Propulsion Module for CUBESATS

A 6U hybrid propulsion module is presented. It is designed to be used as a stage for a 6U CubeSat with a combined wet mass up to 25 kg. Hybrid rocket propulsion is utilized because it is responsive, high performance, capable of multiple burns and safe/non-toxic. Hybrid rockets employ a solid fuel and liquid or gaseous oxidizer. In this case, the fuel is PolyMethyl MethAcrylate and the oxidizer is gaseous oxygen. This propellant combination produces an Isp of greater than 300 seconds. The gaseous oxidizer simplifies the design and is used in dual mode with cold gas thrusters for attitude control. In general, hybrid propulsion is higher performance and higher thrust than conventional options considered for CubeSats (e.g. monopropellants), making it ideal for larger manuevers such as plane changes. The Attitude Control System has been sized to ensure that the CubeSat can be controlled under the thrust loads of the propulsion system. A hybrid rocket motor of this size is currently being tested at the Jet Propulsion Laboratory. Thirty-four successful tests have been completed at JPL with this propellant combination during the last 2.5 years. Current tests include long duration tests (> 95 seconds) and multiple starts (five). These tests and applicability to potential mission requirements will be discussed. Future test plans to move the motor into a vacuum facility to achieve a Technology Readiness Level of 5 by the end of September 2018 will be discussed, and an overview of the proposed TRL 5 to 6 maturity plan will also be presented.

Nakazono, B.↗
NASA NTRS1990

Performance of unconventional propellants

This research involves the theoretical calculations of rocket performance for exotic propellants at various operating conditions, such as chamber pressure, pressure ratios, and oxidizer-to-fuel ratios. Exotic propellants are materials that may not normally be used as propellants on earth due to their low performance characteristics or other factors. The majority of the work was done using the Gordon and McBride CET 86 Program in both a mainframe version and personal computer versions. In addition, the Lockheed/Air Force Solid Propellant Theoretical Performance Program for the IBM PS/2, which handles condensed product species better, was also used.

Rascon, Mario↗
NASA NTRS2011

Oxygen Extraction from Regolith Using Ionic Liquids

An important concern with long-duration manned space travel is the need to furnish enough materials to the vehicle, as well as the crew, for the duration of the mission. By extracting oxygen from the oxides present in regolith, propellant and life support could be supplied to the vehicle and the crew while in space, thereby limiting the amount of supplies needed prior to lift-off. Using a class of compounds known as ionic liquids, we have been able to lower the electrolysis operating temperature from 1600 C (molten oxide electrolysis) to less than 200 C, making this process much more feasible in terms of energy consumption and materials handling. To make this process ready for deployment into space, we have investigated what steps of the process would be affected by the low-gravity environment in space. In the lab, the solubilization of lunar regolith simulant in ionic liquid produces water vapor that is normally distilled out of solution and subsequently electrolyzed for oxygen production. This distillation is not possible in space, so we have tested a method known as pervaporation and have suggested a way this technique could be incorporated into a reactor design.

Barrios, Elizabeth A.↗
NASA NTRS1977

Analytical model of the combustion of multicomponent solid propellants

Multiple flame models derived for simple composite propellants are extended to describe the combustion of propellants containing multimodal particle sizes, mixed oxidizers and monopropellant binders. Models combining the component contributions to propellant surface structure, flame structure and energy distribution are based in part upon experimental observations and in part upon hypotheses constrained to provide reasonable agreement with measured burning rate characteristics. The methods employed consist of superposition, interaction and iteration. The computerized model is applied to explain the effects of multiple ingredients and to discuss burning rate tailoring problems of current interest.

Cohen, N. S.↗
NASA NTRS1989

Space Shuttle RCS Oxidizer Leak Repair for STS-26

Following propellant loading of the Space Shuttle's reaction control system (RCS) for mission STS 26, an oxidizer leak was detected in the left orbital maneuvering system (OMS) pod, where the RCS is located. Subsequent investigation determined that the leak was isolated at a mechanical Dynatube fitting near the RCS nitrogen tetroxide tank. An intense effort was initiated to design, fabricate, and qualify a sealing device to stop the oxidizer leak externally so that the Space Shuttle launch could proceed. It was discovered that sealing devices called clamshells were widely used throughout the petrochemical and power generation industries to stop leaks developed in large diameter pipes which carry steam or other hazardous fluids. These clamshells are available in different diameters and strengths and are placed around the pipe at the location of the leak. A sealing compound is then injected under high pressure into the clamshell to stop the leak. This technology was scaled down and applied to the problem of stopping the leak on the Orbiter, which was on a half-inch diameter line in a nearly inaccessible location. Many obstacles had to be overcome such as determining that the sealing material would be compatible with the nitrogen tetroxide and ensuring that the clamshell would actually fit around the Dynatube fitting without interfering with other lines which were in close proximity. The effort at the NASA Johnson Space Center included materials compatibility testing of several sealants, design of a clamshell to fit in the confined compartment, and manufacture and qualification of the flight hardware. A clamshell was successfully placed around the Dynatube fitting on the Orbiter and the oxidizer leak was terminated. Then it was decided to apply this technology further and design clamshells for other mechanical fittings onboard the Orbiter and develop sealing compounds which will be compatible with fuels such as monomethyl hydrazine (MMH). The potential exists for using this type of sealing device in numerous other applications throughout the aerospace industry.

Delventhal, R. A.↗
NASA NTRS1982

Porous plate analog burner study of composite solid propellant flame structure

A porous plate burner system for studying the structure and interaction of diffusion flames from adjacent unlike sources of oxidizer to simulate the behavior of solid propellant flames is presented. The passage of oxidizer gases through ports of determined diameters represents decomposing oxidizer crystals, and are adjustable for various oxidizer/fuel ratio experiments. Tests with air, ethane, oxygen, and oxygen-air mixtures are described, with parameters of flow rate, oxygen/fuel ratio, and pressure. Coarse ports are found to be oxidizer rich, while fine ports are stoichiometric. More heat feedback is found with fine ports, indicating that fine particle propellant systems burn faster unless fuel allocation is such that the reduced temperature overcomes the reduced diffusion length. Further observations on the interactions of flames are reported.

Strand, L. D.↗
NASA NTRS1992

A neural network-based estimator for the mixture ratio of the Space Shuttle Main Engine

In order to properly utilize the available fuel and oxidizer of a liquid propellant rocket engine, the mixture ratio is closed loop controlled during main stage (65 percent - 109 percent power) operation. However, because of the lack of flight-capable instrumentation for measuring mixture ratio, the value of mixture ratio in the control loop is estimated using available sensor measurements such as the combustion chamber pressure and the volumetric flow, and the temperature and pressure at the exit duct on the low pressure fuel pump. This estimation scheme has two limitations. First, the estimation formula is based on an empirical curve fitting which is accurate only within a narrow operating range. Second, the mixture ratio estimate relies on a few sensor measurements and loss of any of these measurements will make the estimate invalid. In this paper, we propose a neural network-based estimator for the mixture ratio of the Space Shuttle Main Engine. The estimator is an extension of a previously developed neural network based sensor failure detection and recovery algorithm (sensor validation). This neural network uses an auto associative structure which utilizes the redundant information of dissimilar sensors to detect inconsistent measurements. Two approaches have been identified for synthesizing mixture ratio from measurement data using a neural network. The first approach uses an auto associative neural network for sensor validation which is modified to include the mixture ratio as an additional output. The second uses a new network for the mixture ratio estimation in addition to the sensor validation network. Although mixture ratio is not directly measured in flight, it is generally available in simulation and in test bed firing data from facility measurements of fuel and oxidizer volumetric flows. The pros and cons of these two approaches will be discussed in terms of robustness to sensor failures and accuracy of the estimate during typical transients using simulation data.

Guo, T. H.↗
NASA NTRS1977

Experimental fatigue life investigation of cylindrical thrust chambers

Twenty-two cylindrical test sections of a cylindrical rocket thrust chamber were fabricated and 21 of them were cycled to failure to explore the failure mechanisms, determine the effects of wall temperature on cyclic life, and to rank the material life characteristics for comparison with results from isothermal tests of 12 alloys at 538 C. Cylinder liners were fabricated from OFHC copper, Amzirc, and NAR1loy-Z. Tests were conducted at a chamber pressure of 4.14 MW/sq m using hydrogen-oxygen propellants at an oxidant-fuel ratio of 6.0, which resulted in an average throat heat flux of 54 MW/sq m. The cylinders were cooled with liquid hydrogen at an average rate of 0.91 Kg/sec. All failures were characterized by a thinning of the cooling channel wall at the centerline and eventual failure by tensile rupture. Cyclic life rankings of the materials based on temperature do not agree with published rankings based on uniaxial, isothermal strain tests.

Quentmeyer, R. J.↗