Showing papers on "Propellant published in 1997"

Vortex Shedding in a Large Solid Rocket Motor Without Inhibitors at the Segment Interfaces

Abstract: The internal acoustic response of the Titan IV Solid Rocket Motor Upgrade (SRMU) is analyzed using pressure oscillation time histories measured during four static e ring tests. Pressure oscillations for other large solid rocket motors are caused by vortex shedding at annular inhibitors and acoustic feedback resulting from impingement of the vortices on other inhibitors or on the solid rocket motor nozzle. The SRMU does not have inhibitors, but it is shown that acoustic feedback also dee nes its pressure oscillations. Vortices are shed around the cavity between the center and aft segments and impinge on the nozzle entrance. The frequencies of the pressure oscillations vary about that of the motor fundamental acoustic mode and generally agree with a simple empirical relationship that has been used to model acoustic feedback. The effect of the SRMU pressure oscillations on dynamics and control of the Titan IV system is minor. I. Introduction A NEW solid rocket motor design called the Solid Rocket Motor Upgrade (SRMU) will be integrated with the Titan IV system beginning in late 1996. The SRMU has a lower inert weight and higher propellant weight than the current Titan IV solid rocket motor, and therefore, augments system performance. The SRMU has three segments (forward, center, and aft) and is the only large solid rocket motor in production that does not have annular inhibitors at the segment interfaces to prevent combustion and to support the propellant grain during burning. A. SRMU Static Firing Tests Five static e ring tests were conducted for qualie cation of the SRMU. The test articles included one preliminary qualie cation motor (called PQM19) and four qualie cation motors (called QM1, QM2, QM3, and QM4 ). The tests were conducted vertically, nozzle down at the Phillips Laboratory 1-125 Firing Complex, Edwards Air Force Base. 1 The test cone guration and the attachments of the solid rocket motor to the test stand simulated e ight. The physical differences between the e ve test motors were minor and are believed to have had a negligible effect on combustion stability. The static e ring tests were conducted at propellant bulk temperatures in the range 36.5‐ 937F (2.5‐ 33.97C) to assess the effect on motor performance. The static e ring test instrumentation included Condec gauges (CCC-406) of absolute pressure on the forward closure (measurement PCF1 ) and on the aft dome (measurement PCA1), and a Kistler gauge (202M122) of oscillatory pressure (measurement PD4 ) on the forward closure. 2 Measurements PCF1 and PCA1 were recorded by a Neff digital acquisition system, while measurement PD4 was recorded by a Metraplex FM system. The Neff system applied a 100-Hz low-pass e lter and recorded measurements PCF1 and PCA1 at 250 samples/s.

An Eigenvalue Method for Computing the Burning Rates of RDX Propellants

Abstract: A mathematical model for a three-tiered system consisting of solid, liquid and gas is derived for studying the combustion of RDX propellants. The resulting nonlinear two-point boundary value problem is solved by Newton's method with adaptive gridding techniques. In this study the burning rate is computed as an eigenvalue, which can remove the uncertainty associated with employing evaporation and condensation rate laws in its evaluation. Results are presented for laser-assisted and self-deflagration of RDX monopropellants and are compared with experimental results. The burning rates are computed over a wide range of ambient pressures and compare well with experimental results from one to ninety atmospheres. The burning rate is found to be proportional to the pressure raised to the 0.76 power. Sensitivity of the burning rate to initial propellant temperture is calculated and found to be extremely low, in agreement with past theoretical predictions and experimental data. Results for laser-assisted combustion...

An experimental investigation of the spray issued from a pMDI using laser diagnostic techniques.

Abstract: This research was concerned with the experimental investigation of the spray issued from a pressurised metered-dose inhaler (pMDI) using laser diagnostic techniques and has been motivated by the urgent need to find suitable replacements to the environmentally destructive CFC propellants currently used in the device. The experimental work was conducted using phase-Doppler particle analysis (PDPA), a single particle light scattering technique that provides the simultaneous measurement of drop size, velocity, and concentration, yielding the most detailed temporal and spatial analysis of the pMDI spray to date. Three formulations were studied to compare the performance of an "ozone-friendly" hydrofluoroalkane propellant against that of a traditional CFC propellant mixture and a commercially available CFC formulation containing drug and surfactant. The PDPA analysis was complemented by a visual investigation of the near-orifice flow field using copper laserstrobe microcinematography to obtain informat...

Constitutive Equations for Solid Propellants

Abstract: Mechanical behavior of the Space Shuttle redesigned solid rocket motor (RSRM) propellant is studied from a phenomenological point of view. Motivated by the study of the experimental data three initially isotropic constitutive models have been developed. All models represent the effect of strain rate, superimposed hydrostatic pressure, and cyclic loading on the stress and dilatation response of the material. A particular emphasis is given to the prediction of volume dilatation. The model resulting in the best representation of the available data is calibrated using only a few tests. The predictions of the model are compared with experiments for several loading conditions not used in the calibration.

Vehicle occupant restraint apparatus

Abstract: To achieve programmed inflation of an airbag in a vehicle occupant restraint apparatus, movement of a piston (14), that separates a combustion chamber (20) from a liquid propellant reservoir (24), is utilized to vary the opening size of an injection port through which liquid propellant (26) is regeneratively pumped from the reservoir (24) into the combustion chamber (20) for combustion, such that the rate at which the bag inflation gas is generated during an inflation period may be controlled.

New aqueous medicament preparations for the production of propellent gas-free aerosols

Abstract: The invention relates to medicament preparations in the form of aqueous solutions for the production of propellent gas-free aerosols.

Pyrotechnic compositions and uses therefore

Abstract: Pyrotechnic compositions having combustion reaction products that include a high percentage of carbon dioxide at high temperatures. These pyrotechnic compositions include a fuel component having combustion product with a relatively high percentage of carbon to hydrogen content, such as an aromatic polycarboxylic anhydride fuel component. These pyrotechnic compositions may be employed to emit infrared radiation from a decoy flare or other device. The compositions may also be employed as a gas generating propellant for a projectile, such as a flare or rocket, or for other applications such as an automobile airbag. These pyrotechnic compositions emit infrared emissions with minimized short wavelength components, and produce substantially non-toxic combustion gases having relatively small amounts of water vapor.

Thermally stable nonazide automotive airbag propellants

Abstract: Thermally stable gas generant compositions incorporate a combination of nitroguanidine, one or more nonazide high-nitrogen fuels, and phase-stabilized ammonium nitrate or a similar nonmetallic oxidizer that, upon combustion, result in a greater yield of gaseous products per mass unit of gas generant, a reduced yield of solid combustion products, and acceptable burn rates, thermal stability, and ballistic properties. These compositions are especially suitable for inflating air bags in passenger-restraint devices.

Kinetic study of the reaction between hydroxyl-terminated polybutadiene and isophorone diisocyanate in bulk by quantitative FTIR spectroscopy

Abstract: A kinetic study of the reaction between a hydroxyl-terminated polybuta-diene (HTPB) and isophorone diisocyanate (IPDI) was carried out in the bulk stateby using quantitative Fourier transform infrared (FTIR) spectroscopy. The reaction isshown to obey a second-order rate law, being first order in both the HTPB and IPDIconcentrations. The activation parameters obtained from the evaluation of kinetic dataare D H ‡ †41.1 {0.4 kJ mol, D S ‡ †0198 {2JK 01 mol 01 and E a †43.8 {0.4 kJ mol 01 ,which are quite different from the solution values. However, they are in agreement withthe results obtained on propellants by torsional braid measurements. The large nega-tive value of the activation entropy is indicative of an associative mechanism, whichis in accord with the second-order rate law for the polyurethane formation. q 1997 John Wiley & Sons, Inc. J Appl Polym Sci 66: 1979–1983, 1997 Key words: HTPB; IPDI; diisocyanate; kinetics; polyurethane; FTIR spectroscopy INTRODUCTION other binders, such as poly(vinyl chloride). Fur-thermore, the low viscosity of HTPB facilitatesComposite propellants based on hydroxyl-termi- high solid loading in both fuel and oxidizer of thenated polybutadiene (HTPB) have become the propellant. At the final stage of a composite pro-workhorse propellants in present day solid rocket pellant preparation, a curing agent, usually a di-motors. Crystalline ammonium perchlorate (AP) isocyanate, is added to the heterogeneous mixtureand fine aluminum powder are used as oxidizer containing the prepolymer, fuel, oxidizer, andandmetallicfuel, respectively,intheHTPB-based other ingredients. The prepolymer and diisocya-composite rocket propellants. The solid loading, nate react with each other to form the polyure-which contributes to the rheological properties thane network [eq. (1)],which holds the fuel andand therefore affects the processability and me- oxidizerparticulates ina compositematrix havingchanical properties of the propellants, is largely mechanical and ballistic properties suitable fordetermined by the extent of polyurethane forma- rocket applications.tion. In this respect, the binder used in the propel-lant plays an important role. The polybutadienechain, in general, gives higher energy value andO'H 1 OCN O'C'N (1)better mechanical properties compared to the OHThe cure reaction of HTPB with diisocyanate hasbeen found to follow the second-order rate law.

Method of intrapulmonary administration of a narcotic drug

Abstract: An aerosol formulation of an aerosol propellant and a base form of a narcotic drug selected from the group consisting of fentanyl, sufentanil and remfentanyl is provided. Such a formulation allows for the drug to be dissolved within the propellant and used within a device which does not require the use of a lubricant. Formulations are also disclosed which include lubricants, wherein the lubricant and propellant are both either polar or both non-polar. Thus, the lubricant component does not act as a solvent or cosolvent, but rather acts as a lubricant for the valve used for dispersing the formulation to a patient. Typical non-polar propellants include chlorofluorocarbons, which are typically used in connection with non-polar lubricants such as saturated vegetable oils, e.g. fractionated coconut oils. Typical polar propellants include hydrofluoroalkanes, which are typically used in connection with polar lubricants such as polyethylene glycols.

Non-lethal cartridge with spin-stabilized projectile

Abstract: The invention is a non lethal weapon cartridge comprising a projectile and means for propelling the projectile through a weapon barrel. A munition of this type can be employed by soldiers during operations-other-than-war, such as riot control during humanitarian missions, or by law enforcement personnel when a lethal response is not warranted. The projectile comprises a full-bore projectile body fitted with a compliant nose. The projectile is designed to be spin-stabilized such that it will fly, and impact, nose first, while describing a ballistic trajectory. The projectile is intended to be launched from a rifled weapon tube. The rifling imparts the spin necessary to achieve dynamic stability. The propulsion system utilizes a modern smokeless propellant in combination with a high-low technique to produce consistent interior ballistics.

Challenges in Propellants and Combustion: 100 Years after Nobel.

Abstract: : The Final Proceedings for the Fourth International Symposium on Special Topics in Chemical Propulsion 4-ISICP, 27 May 1996 - 31 May 1996. The Topics covered include: chemical kinetics of propellant combustion, environmental considerations in combustion of solid and liquid propellants, commercial application in the combustion of energetic materials, effective utilization of propellants, combustion diagnostics, and recycling.

Nonlinear Stability Testing of Full-Scale Tactical Motors

Abstract: The U.S. Naval Air Warfare Center has participated in a program to develop an improved understanding of linear and nonlinear combustion instability in solid propellant rocket motors. One goal of this program was to develop a systematic database of motor stability data. This paper describes the nonlinear aspects of the motor e rings and analysis. The motors used had diameters of 127 mm and were 1.7 m long. The majority were loaded with an 88% solids reduced-smoke ammonium perchlorate propellant with a nominal burning rate of 6.1 mm/s at 6.9 MPa. Motor pressures ranged from 3.45 to 10.34 MPa and various grain geometries were tested. In addition, motors have been e red that contain 1% 8- mm aluminum oxide, 90- mm aluminum oxide, and 3- mm zirconium carbide as stability additives in place of 1% ammonium perchlorate. This paper discusses experimental methods of motor pulsing and various theoretical approaches to predict pulse amplitudes in motors. The paper also examines nonlinear acoustic motor response to various amplitudes of acoustic pulsing and the resultant characteristics of sustained nonlinear oscillations. Finally, some theoretical interpretations are presented from the experimental motor data. The results show that there is a direct relationship between the dc pressure shift and the magnitude of the acoustic oscillations.

Advanced Materials for Ultrahigh Temperature Structural Applications Above 2000 deg C

Abstract: : The primary incentive for developing ultrahigh temperature materials for liquid bi-propellant rocket engines lies in the minimization and/or elimination of fuel-film and regenerative cooling of combustion chambers. Cooling is currently required because the most commonly used material for rocket combustion chambers is niobium alloy coated with disilicide with upper limit of operation up to 1450 deg C which is only approximately 50 % of the propellant combustion temperature. Therefore, by developing an ultrahigh temperature material with temperature capabilities in the range of 2200 - 3000 deg C, the fuel-film and regenerative cooling can be significantly reduced and/or eliminated resulting in cleaner burning of rocket engine. Thus fuel utilization can be vastly improved, more payload can be sent to space, higher specific impulse (Isp) can be achieved and finally the cost of the rocket engine could be reduced.

Analysis of Flame Patterns in Cryogenic Propellant Combustion

Abstract: Design and optimization of high performance rocket engines may be improved by detailed studies of the basic combustion mechanisms. Much detailed information exists on elementary processes such as atomization, multiple jet interactions, vaporization of single droplets, structure of spray flames, ignition of nonpremixed systems etc. It is however important to approach the real conditions existing in rocket motors and to this purpose several facilities for cryogenic propellant combustion research have been designed and constructed. One experimental set-up designated as “Mascotte” is operated by ONERA and used for fundamental research as well as technical studies. This article describes results of experiments conducted in this facility by our laboratory. Two series of tests carried out during the last two years have provided a large set of images of combustion in a liquid oxygen/gaseous hydrogen coaxial injection geometry operating at atmospheric pressure and at 5 and 10 bars. The data correspond to laser ela...

Atmospheric fueled ion engine

Abstract: An environmentally compatible propulsion system for low maintenance and long term durations at high altitudes is provided which is capable of utilizing high altitude ambient gas as fuel and producing ozone as a by-product of propulsion. The ion engine propulsion system ionizes a portion of an ambient atmospheric fuel to create a negative ionic plasma for bombarding and accelerating the remaining portion of the ambient atmospheric gas in a focused and directed path to an ion thruster anode. The novel ion engines provided create a negative ionic plasma between a cathode ion thruster and a ring-shaped anode in a housing composed of an electrical insulative material in which the cathode ion thruster is charged to -18 to -110 kilovolts (kv) to utilize ambient atmospheric gas as the propellant.

Propellant mixtures and aerosols for micronising medicaments with compressed gas

Abstract: Propellant mixtures and medicament aerosols which contain them are described for micronising medicaments for pulmonary use. The propellant mixture is in a subcritical state and contains at least one component from a first class of propellant gasses and at least one component from a second class of propellant gasses. The first class includes propellant gasses with an evaporation enthalpy of 200 kJ/kg or less at 25 °C and a vapour pressure of 20 bars or more at 25 °C, and the second class includes propellant gasses with an evaporation enthalpy of 300 kJ/kg or more at 25 °C and a vapour pressure of 10 bars or less at 25 °C. By using this propellant mixture in a medicament aerosol, micronised medicaments are obtained in which approximately 80 % by weight of the generated particles have less than 8 νm diameter.

Substantially smoke-free and particulate free inflator for inflatable safety restraint system

Abstract: Provided is an inflator for providing propellant gases to inflate a bag of an automotive safety system for restraining an occupant of a vehicle in which the inflator is a pure pyrotechnic inflator that generates propellant gases that are substantially free of metal-containing particulate and condensable materials without requiring storage of a pressurized gas. The propellant gases are generated from a solid gas-generating propellant composition including two components. A first component is fuel-rich and has a fast burn rate. The second component is a oxidizer that, although having a slow burn rate, operates in combination with the first component at an acceptably fast burn rate to oxidize carbon monoxide and/or hydrogen produced during combustion of the first component to carbon dioxide and/or water, respectively.

Liquid propellant airbag inflator with dual telescoping pistons

Abstract: An airbag inflator includes a pair of telescoping pistons slidingly mounted within a housing. Pressurization of a combustion chamber by a pyrotechnic initiator acts on one piston to pressurize a hydraulic fluid chamber, which, in turn, acts on the other piston to pressurize a liquid propellant reservoir. Liquid propellant can then be regeneratively pumped from the reservoir into the combustion chamber for combustion to generate an airbag inflation gas. A hydraulic fluid damping chamber is utilized to exert a controllable retarding force on the one piston, such as to control the rate of airbag inflation gas generation.

Automated propellant blending

Abstract: An automated propellant blending apparatus and method uses closely metered addition of countersolvent to a binder solution with propellant particles dispersed therein to precisely control binder precipitation and particle aggregation. A profile of binder precipitation versus countersolvent-solvent ratio is established empirically and used in a computer algorithm to establish countersolvent addition parameters near the cloud point for controlling the transition of properties of the binder during agglomeration and finishing of the propellant composition particles. The system is remotely operated by computer for safety, reliability and improved product properties, and also increases product output.

Method and apparatus for dispensing liquid with gas

Abstract: A method and apparatus for dispensing a liquid, such as a fire suppressinggent, with a gas wherein both the liquid and a combustible propellant for generating the gas are stored in separate sealed compartments at atmospheric pressure. The liquid is stored in a chamber between an annular piston and a central pedestal containing a gas generating canister. A portion of the gas generated drives the piston to expel the liquid into a mixing chamber in the pedestal, and another gas portion is fed into the mixing chamber so as to mix with and propel the liquid through a nozzle. The liquid may be atomized or vaporized depending on its composition. Mixing of the liquid with the gas may be enhanced by tangentially injecting the liquid into the mixing chamber.

Solid solution vehicle airbag clean gas generator propellant

Abstract: An ignitable solid gas generating composition comprises a polyalkylammonium binder, usually polyvinylamine nitrate or polyethyleneimmonium nitrate, an oxidizer mixture comprising ammonium nitrate and a first additive which produces an eutectic melt which is liquid at a temperature well below the melting point of the ammonium nitrate as well as that of the first additive, and an additional quantity of the ammonium nitrate and a second additive. Further, combustion modifier additives may be added to the composition.

3-d grain design and ballistic analysis using the spp97 code

Abstract: The Solid Performance Program (SPP) has become the standard reference computer program throughout the United States for predicting the delivered performance of solid propellant rocket motors The code embodies a methodology for calculating the delivered performance of solid propellant rocket motors prescribed by JANNAF The nozzle performance methodology starts with the ideal performance and addresses each of the following performance loss mechanisms: finite rate chemical kinetics, nozzle throat erosion, nozzle submergence, nozzle flow divergence, two phase flow, combustion efficiency, and the nozzle wall boundary layer The Grain Design and Ballistics (GDB) module calculates the ideal pressure-thrust history, and subsequently modifies these values based on the nozzle performance efficiencies The GDB module combines existing 3-D grain design and ballistics methods with modern graphics and ease of use to provide a tool which allows even novice users to obtain successful results

Spacecraft attitude and velocity control thruster system

Abstract: A rocket propulsion system for spacecraft achieves greater economy, reliability and efficiency rocket by incorporating monopropellant RCS thrusters (1a-1f) for attitude control and bipropellant SCAT thrusters (5a-5c) for velocity control. Both sets of thrusters are designed to use the same liquid fuel, supplied by a pressurized non-pressure regulated tank, and operate in the blow down mode. In the propulsion system such station keeping and attitude control thrusters may function in conjunction with a large thrust apogee kick engine, which may also be of the SCAT thruster construction, that uses the same propellent fuel. Hydrazine and Binitrogen tetroxide are preferred as the fuel and oxidizer, respectively. The new system offers a simple conversion of existing monopropellant systems to a high performance bipropellant dual mode system without the extreme complexity and cost attendant to a binitrogen tetroxide--hydrazine bipropellant system.

Performance Tests of a Liquid Hydrogen Propellant Densification Ground System for the X33/RLV

Abstract: A concept for improving the performance of propulsion systems in expendable and single-stage-to-orbit (SSTO) launch vehicles much like the X33/RLV has been identified. The approach is to utilize densified cryogenic liquid hydrogen (LH2) and liquid oxygen (LOX) propellants to fuel the propulsion stage. The primary benefit for using this relatively high specific impulse densified propellant mixture is the subsequent reduction of the launch vehicle gross lift-off weight. Production of densified propellants however requires specialized equipment to actively subcool both the liquid oxygen and liquid hydrogen to temperatures below their normal boiling point. A propellant densification unit based on an external thermodynamic vent principle which operates at subatmospheric pressure and supercold temperatures provides a means for the LH2 and LOX densification process to occur. To demonstrate the production concept for the densification of the liquid hydrogen propellant, a system comprised of a multistage gaseous hydrogen compressor, LH2 recirculation pumps and a cryogenic LH2 heat exchanger was designed, built and tested at the NASA Lewis Research Center (LeRC). This paper presents the design configuration of the LH2 propellant densification production hardware, analytical details and results of performance testing conducted with the hydrogen densifier Ground Support Equipment (GSE).

Gas primed powder actuated tool

Abstract: There is provided a powder actuated tool effective to drive a metallic fastener into a workpiece, for example, a nail gun. The pressure to drive the fastener is generated by combustion of an flammable propellant mix that is ignited by a primer flash. The primer flash having been generated by ignition of a combustible gas by discharge of an electric arc in the tool barrel. The combustible gas is lead-free such that the operator and the environment are not exposed to dangerous lead residue during use. The use of a solid propellant generates higher pressure and more effective driving of the metallic fastener than is achieved with powder actuated tools driven solely by combustible gases.

Propulsion system containing a mixed-phase propellant and a method for propelling an object with the same

Abstract: A propulsion system with at least one storage chamber containing at least one solid propellant and at least one fluid propellant is disclosed. The fluid propellant is retained under pressurized conditions, such that depressurization of the storage chamber substantially homogeneously disperses the at least one solid propellant in the at least one fluid propellant. A mixed-phase propellant can thereby be fed to a combustion chamber. The pressurized conditions under which the at least one fluid propellant is retained can include supercritical or critical conditions, saturation conditions, and conditions sufficient to provide a compressed gas.