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Reports are presented at the meeting at the University of Arizona on the study of predictable and reliable solid rocket motors. The following subject areas were covered: present state and trends in the research of solid propellants; the University of Arizona program in solid propellants, particularly in mixing (experimental and analytical results are presented).
Data were calculated for equivalence ratios of 1 to 4, chamber pressures of 300 and 600 pounds per square inch absolute, and pressure ratios of 1 to 1500. Parameters included are specific impulse, combustion and exit temperatures, molecular weight, characteristic velocity, coefficient of thrust, ratio of nozzle-exit area to throat area, specific heat at constant pressure, isentropic exponent, viscosity, and thermal conductivity. A correlation is given which permits determination of performance for a wide range of chamber pressures. A method for obtaining specific impulse of JP-4 fuel with OF2 and O3-F2 mixtures is given.
Researchers and small companies have been investigating and developing new solid rocket motor propellant ingredients with renewed vigor. Lithium-aluminum alloy is an ingredient of particular interest, but most of the advertisement about its benefits are focused on defense tactical rocket applications. This paper develops a framework to evaluate this and other propellants against standard aluminized composite propellant for mass-limited missions like many NASA in-space applications. It shows a way to estimate density’s effect with an exponent less than 1 depending on mission and technology parameters, and an algorithm for directly calculating delta-velocity or payload performance improvement.
Microwave attenuation magnitude associated with exhaust plasma of nonaluminized and aluminized composite propellant
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.
Composite propellant manufacturing by pneumatic mix process
Composite solid propellant flame structure in vicinity of burning surface based on spectral radiation shadowgraph technique
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.
It is pointed out that in-process tests to verify quality and detect discrepant propellant which could compromise motor performance are essential elements of the solid composite propellant manufacturing process. The successful performance of the 260SL-1 and 260SL-2 motors aptly verified the controls used for manufacturing the propellant. The present investigation is concerned with the selected control parameters, and their relationships to composition and final propellant properties. Control performance is evaluated by comparison with processing data experienced in the manufacture of the propellant for the 260SL-1 motor. It is found that the in-process quality verification controls utilized in the propellant manufacturing process for the 260-in. diameter motor contributed significantly to the confidence of successful and predictable motor performance.
Composite solid propellant flame structure in vicinity of burning surface based on spectral radiation shadowgraph technique
Fast burning rates in thin film solid composite propellants composed of McCormick-Selph 510, 164 monopropellant, oxidizer and polyvinyl chloride binder
Alumina size distributions were obtained for both a coarse and a fine oxidizer composite propellant burning in a nitrogen atmosphere over the pressure range from atmospheric to 500 pounds per square inch. The amount of additive agglomeration was found to be significantly higher for the coarse oxidizer propellant and decreased with increasing pressure to the 0.3 power over the range from atmospheric pressure to 250 pounds per square inch. High-speed photographs of the burning propellant surface revealed that the additives moved on the surface with the average particle velocity decreasing with pressure to approximately the 0.3 power over the pressure range from atmospheric to 50 pounds per square inch. The empirical relation between particle velocity and pressure was used to modify an agglomeration criterion presented previously. Evaluation of the critical aluminum diameter required for agglomeration indicated that both propellant types used in this study should experience some agglomeration over the pressure range studied. The aluminum size required for agglomeration was found to increase with increasing pressure. The experimental findings of the critical aluminum diameter required for agglomeration were in reasonable agreement with the calculated data. The volume mean diameter of the alumina was found to decrease with increasing pressure.
Burning rate and decomposition mechanisms of ammonium perchlorate composite propellant
Explosive proclivities of composite propellants related to dynamic mechanical properties
The feasibility of reducing troublesome nozzle blockage (by condensation deposits) in laboratory-scale solid rockets by addition of a silicone oil as a propellant ingredient was explored experimentally. An aluminized composite propellant and its counterpart with 1% silicone oil replacing part of the binder were fired in a 63.5 mm diameter, end-burning, all-metal burner. Pressure-time histories were recorded for all of the tests by a Taber gauge mounted at the downstream end of the chamber; temperature-time data at the nozzle throat were obtained in some of the runs by thermocouples having junctions positioned at the wall but insulated from the metal. Deposition of condensables on the nozzle walls causing a progressive increase in the chamber pressure with time was noted. The fraction of firings exhibiting practically no condensation was 59% with silicone and 32% without. On the average, temperature readings at the nozzle throat were higher with the silicone propellants. Although various phenomena may contribute to these findings, the results are not understood completely.
A transient ballistics and combustion model is derived to represent the closed vessel experiment that is widely used to characterize the high pressure burning rates of solid propellants. The model is applied to explain why burning rates deduced from the closed vessel are in basic agreement with those measured from an equilibrium strand burner in the case of homogeneous propellants, but differ significantly in the case of nitramine composite propellants. Thermal profile time lag effects become small at high pressure because the burning rates become high. However, the development of the burning surface structure of those nitramine propellants which exhibit shifts in pressure exponent causes the mass burning rate to lag and then exceed the equilibrium value. It is necessary to consider this mechanism in applications dealing with high pressures and pressurization rates.
Studies of several General Aviation aircraft indicated that the application of advanced technologies to General Aviation propellers can reduce fuel consumption in future aircraft by a significant amount. Propeller blade weight reductions achieved through the use of composites, propeller efficiency and noise improvements achieved through the use of advanced concepts and improved propeller analytical design methods result in aircraft with lower operating cost, acquisition cost and gross weight.
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.