Showing papers on "Propellant published in 1996"

Ion-thruster plume modeling for backflow contamination

Abstract: An axisymmetric model of the plume and backflow contamination from an ion-thruster plume is presented. Components included are primary beam ions, neutral propellant efflux, thermal propellant ions created mainly by charge-exchange collisions between primary beam ions and neutral propellant, nonpropellant efflux sputtered from thruster components, and neutralizing electrons. The plasma hybrid particle-in-cell technique is applied to both propellant charge-exchange ions and nonpropellant efflux ions produced within the beam and to their transport out of the beam to the regions surrounding a model spacecraft. Simulation results for plume properties such as ion density, beam potential, and ion flow angle are compared with experimental data. It is shown that the charge-exchange ions formed in the beam are accelerated outward by strong radial beam electric fields and form two distinct energy populations—one with a significant backstreaming velocity component.

Parietal vortex shedding as a cause of instability for long solid propellant motors - Numerial simulations and comparisons with firing tests

Abstract: This paper describes a new mechanism leading to vortex shedding instabilities in long (large L/D) solid propellant motors. This mechanism is termed "parietal vortex shedding" and has been discovered thanks to numerical simulations of the unsteady, 2D, compressible, Navier-Stokes equations. It seems to involve hydrodynamic instabilities of the mean flow velocity profiles, corresponding to injection induced internal flow (so called Culick or Taylor profiles), that couple with the acoustic frequencies of the chamber. Although this mechanism is found to be very powerful, it seems to need some background noise to feed it. The presented simulations can explain observed instabilities in configurations, without segmentation or without protruding inhibitor rings, of a simplified subscale setup of the Ariane 5 MPS P230 solid propellant motor. Detailed comparisons are proposed and the influence of the propellant combustion response and of the turbulence of the flow field are analyzed by means of recently developed models. INTRODUCTION-OBJECTIVES not precisely known if inhibitor rings are or not destroyed or completely slack; on the other hand non segmented motors have also given rise to a same type of instability as vortex shedding : this is the case of one of the configuration (LP3D) of the LP3 set up presented in a preceding paper. As it is known ' ' , classical linear acoustic balance computations do not give reliable stability predictions in complex internal geometries (such as in the P230 motor), and then an effort was carried out to perform full numerical simulations of the unsteady, compressible internal flow fields. On the other hand, motor internal flows are mostly non-observable, and without numerical simulations it is not possible to describe the path of the aerodynamic instability development. The object of the present paper is to show hydrodynamic instabilities (which drive pressure oscillations) in configurations without shear flows induced by inhibitor rings, and to explain how these instabilities occur by means of numerical simulations. LP3D AND LP3E TEST CASES This work is part of the overall research effort, supported by CNES, accompanying the development of the Ariane 5 P230 MPS solid propellant motor (program ASSM for Aerodynamics of Segmented Solid Motors) and makes use of experimental results obtained during the combustion stability assessment program carried out for BPD and CNES, by delegation of ESA. In this scope, it is the continuation of earlier works about numerical simulations in solid propellant rocket motors ". The Ariane 5 motor, as other large segmented motors (U.S. Space Shuttle and Titan SRM) has been reported to exhibit pressure and thrust oscillations. Until recently, it was believed that such instability was exclusively due to the segmented design : inhibitor rings induce shear layers and vortex shedding driven oscillations. Nevertheless it is * Research scientist, Energetics Dept. 1 Project manager, Energetics Dept., Member AIAA Copyright © 1996 by the American Institute of Aeronautics and Astronautics, Inc., All rights reserved. The test cases are based on two configurations of the LP3 motor. The LP3 motor, which has been presented in reference , is a simplified 1/15 subscale set-up of the Ariane 5 P230 motor used to test several inter-segment arrangements. The configurations of interest here, called LP3D and LP3E, have no prominent obstacles at the mid chamber point, see figure 1. Both configurations have a cylindrical chamber of length 1632 mm and inner diameter 203 mm. At ignition, the main propellant burning surface is cylindrical, inner diameter 90 mm, with a chamfered surface near the aft end. The supersonic outlet nozzle is submerged, with a throat diameter of 56.5 mm, then the total length of the motor is 1650 mm. LP3D and LP3E have a small forward segment of 225 mm length, cylindrical for LP3D (169.5

Aerosol drug formulations containing vegetable oils

Abstract: Pharmaceutical compositions for aerosol delivery comprising (a) a medicament, (b) a non-chlorofluorocarbon propellant, and (c) a vegetable oil or a pharmaceutically acceptable derivative thereof, as well as a method for preparing such compositions in which unwanted aggregation of the medicament is prevented without the use of surfactants, protective colloids or cosolvents.

HAN-based monopropellant assessment for spacecraft

Abstract: The growing cost of space missions, the need for increased mission performance, and concerns associated with environmental issues are changing rocket design and propellant selection criteria. Whereas a propellant's performance was once defined solely in terms of specific impulse and density, now environmental safety, operability, and cost are considered key drivers. Present emphasis on these considerations has heightened government and commercial launch sector interest in Hydroxylammonium Nitrate (HAN)-based liquid propellants as options to provide simple, safe, reliable, low cost, and high performance monopropellant systems.

Investigation of propellant inefficiencies in a pulsed plasma thruster

Abstract: : A Pulsed Plasma Thruster (PPT) benefits from the inherent engineering simplicity and reduced tankage fraction gained by storing the propellant as a solid. The solid is converted to the gaseous state during an electric surface discharge. Previous research has concluded that the bulk of the propellant expands gas-dynamically from the chamber at low directed velocity, with possibly as little as 10% ionized and efficiently accelerated to high velocity using electromagnetic forces. The two velocity components result in a low propellant utilization efficiency. Critical to improving the PPT efficiency is preventing the formation of the low-velocity propellant and/or developing a means of accelerating it electromagnetically. In the present work measurements are made of the solid propellant conversion to the gaseous state with the intent of better understanding the formation process. By better understanding the propellant conversion it is hoped that future PPTs can be designed with significantly increased propellant efficiencies.

Propellant composition for automotive safety applications

Abstract: Clean burning, non-self extinguishing propellant compositions for use in hybrid automotive air bag systems are disclosed. The propellant compositions are based on a mixture of a crystalline nitramine propellant, an energetic or non-energetic binder and one or a combination of an oxidizing propellant and an energetic plasticizer.

Condensed-phase kinetics of cyclotrimethylenetrinitramine by modeling the T-jump/infrared spectroscopy experiment

Abstract: During the combustion of solid propellants, explosives, or pyrotechnics, the condensed phase experiences heating rates that may exceed 20,000 K/s. At such high heating rates, the thermal decomposition behavior of the energetic material could be affected by its rate of decomposition. To simulate the high heating rate environment, the T-jump experiment was developed for use with Fourier-transform infrared spectroscopy. The T-jump experiment utilizes electrical resistance heating of a thin Pt filament on which a small amount of the energetic test sample is placed. This work describes a heat transfer model of the filament and sample, a model of the current's control circuit, and global decomposition and heat release mechanisms of Cyclotrimethylenetrinitramine (RDX), which is an energetic ingredient .used in propellants and explosives. Comparisons of model calculations with experimental data reveal an excellent agreement. Similarly, the predicted time to rapid heat release for the highly energetic RDX sample also shows a good agreement with experimental results. Thus the use of the developed model in conjunction with experiments should be a useful tool in studying the thermal decomposition behavior of energetic materials under combustion-like conditions.

Analysis of Hall-effect thrusters and ion engines for orbit transfer missions

Abstract: Analytical methods were combined with actual thruster data to create a model used to predict the performance of systems based on two types of electric propulsion thrusters, Hall-effect thrusters and ion engines, for several orbit transfer missions. Two missions were trip time constrained: a LEO-GEO transfer and a LEO constellation transfer. Hall thrusters were able to deliver greater payload due to their higher overall specific power. For the power limited orbit topping mission, the choice of thruster is dependent on the user’s need. Ion engines can deliver the greatest payload due to their higher specific impulse, but they do so at the cost of higher trip time. Study of reusable electric orbit transfer vehicle systems indicates that they can offer payload mass gains over chemical systems, but that these gains are less than those offered by other electric propulsion transfer scenarios due to the necessity of carrying propellant for return trips. Additionally, solar array degradation leads to increased trip time for subsequent reusable transfers. * Research Aerospace Engineer, Member AIAA ** Group Leader, USAF Electric Propulsion Lab, Member AIAA This paper is declared a work of the US Government and is not subject to copyright protection in the United States. INTRODUCTION: The US Air Force has recently completed several studies to investigate the potential advantages of advanced space propulsion for several orbit transfer scenarios. The first study investigated advanced propulsion concepts for expendable orbit transfer vehicles and concluded that the potential launch vehicle downsizing that resulted from the use of high specific impulse thrusters provided significant cost savings over base line chemical launch vehicle/upper stage systems. The second study looked at reusable advanced upper stages and preliminary indications are that while there remains the potential for launch vehicle downsizing, it is significantly reduced compared to expendable systems. This difference was largely due to the added propellant required to perform the round trip mission from low-earth orbit to geostationary orbit. Both studies pointed out advantages for advanced electric propulsion systems based on xenon propellant. The objective of this paper is to analyze the tradeoffs between Hall-effect thrusters and ion engines as a high power propulsion system for orbit transfer missions. Both the Hall-effect thruster and the gridded ion engine are classified as electrostatic thrusters and operate on heavy noble gases, primarily xenon. These electric propulsion devices are capable of specific impulses ranging from approximately 1500 to 4000 seconds, compared to chemical systems which typically operate in the range of 300 to 400 seconds. Electric propulsion is a type of rocket propulsion for space vehicles and satellites which utilizes electric and/or magnetic processes to accelerate a propellant at a much higher specific impulse than attainable using classical chemical propulsion. The concomitant reduction in required propellant mass results in increased payload mass capability. The method of analysis used in this study is based on the model developed by Messerole. It has been modified to reflect the most current information on thruster development levels and

Fourier transform Raman spectroscopy of some energetic materials and propellant formulations

Abstract: : Fourier transform Raman (FTR) spectroscopy employing near infrared (NIR) laser radiation at 9,394.5/cm is used to characterize neat energetic materials and several propellant formulations. Raman spectra are reported over the region from 100/cm to 3,000/cm, relative to the Rayleigh line. The technique is extended to the study of crystalline components of propellant formulations during heating. The utility of the technique in determining the principal crystalline ingredient in a propellant formulation is demonstrated. jg p3

Barrel assembly with axially stacked projectiles

Abstract: A barrel assembly (10) having a plurality of projectiles (11) stacked axially within the barrel (12) together with discrete selectively ignitable propellant charges (13) for propelling the projectiles (11) sequentially through the muzzle of the barrel (12) is provided with adjacent projectiles (11) separated from one another by locating means (13) independent of the projectiles. The locating means may be a solid propellant charge (13) located between adjacent projectiles or it may be a rigid casing (122) for the propellant. When subject to an in-barrel load a rear skirt portion of the active projectile (11) is expanded outwardly by the interaction between an inwardly reducing recess (14) formed in the rear end of a projectile and the nested complementary leading portion of the propellant charge (13) or propellant casing (122).

Flame stabilization in cryogenic propellant combustion

Abstract: Cryogenic combustion is of considerable technological interest in propulsion applications. Cryogenic propellants used in rocket engines provide the high performance needed for spacecraft launching and have operated safely for a number of years, but the processes that control combustion in such devices are still not well understood. Among the many important issues, flame stabilization constitutes one basic problem. This question is investigated in this article by imaging the flame originating from a single, coaxial injector fed with liquid oxygen and gaseous hydrogen. Results of experiments carried out on a facility for cryogenic propellant combustion research operated by ONERA are used to characterize the mechanisms that control the flame-holding process at atmospheric pressure, 5, and 10 bars. Data acquired correspond to elastic scattering by the spray, emission of OH radicals, and planar laser-induced fluorescence of these radicals. Fluorescence is obtained by pumping the X2II (v″=0)→A2Σ(v′=1) band of OH, and off-resonance light radiation is observed. This database provides the general structure of the flame in the injector near-field, and may be used to determine the position of the flame stabilization region. Simultaneous acquisition of laser-induced fluorescence and elastic scattering was used to locate the flame in respect to the liquid. It is shown that in all cases investigated the flame is initiated at a close distance from the injector exhaust plane.

Hot-spot ignition of condensed phase energetic materials

Abstract: A micromechanics approach to hot-spot formation and growth to detonation in condensed-phase energetic materials is presented. A numerical model based on fundamental conservation principles is developed to examine the dynamic and thermodynamic processes that occur in a generalized heterogeneous, energetic material subjected to weak shock loading. The work focuses on the thermal/mechanical processes that act to transfer compression work of the shock wave into localized high-temperature ignition sites. A special interest of this research is to determine the dominant physical processes occurring at different times during hot-spot formation. Processes such as viscoplastic heating, phase change, finite rate condensed-phase decomposition, gas-phase heating, and heat transfer between the void and the condensed-phase are included in the model. Results for cyclotrimethylene trinitramine (C3H6N6O6), a common ingredient in high-energy solid rocket propellants, show that viscoplastic heating is an effective mechanism for producing high-temperature regions in the energetic material adjacent to a shock-collapsed void. Furthermore, it is shown that under certain initial conditions (pore size, shock pressure, etc.), localized heating can lead to the release of chemical energy that exceeds the energy dissipated by heat losses, and that melting and the variation of condensed-phase viscosity and yield strength can greatly affect the dynamics of pore collapse.

Dual chamber nonazide gas generator

Abstract: A gas generator (10) utilizes two segregated propellant container/combustion chambers (62 and 64), each having a plurality of nonazide propellant grains (40 and 66) therein, and an igniter (24) for igniting only the propellant grains (40) located within the first combustion chamber (62). The nonazide propellant produces enough heat energy to subsequently ignite the segregated propellant grains (66) by forced convection and/or heat conduction.

Arcjet Thruster Development

Abstract: For several years an intensive program has been in progress at the University of Stuttgart to investigate and develop thermal arcjets for propellants including ammonia, nitrogen-hydrogen mixtures simulating hydrazine, and hydrogen. Since hydrogen yields the highest specific impulse /sp and best efficiencies TJ, special emphasis was placed on this propellant. Arcjet power levels between 0.7-150 kW have been studied, including water- and radiation-cooled laboratory models and flight hardware. Results yielded a maximal attainable 7sp as a function of the design and power level and showed that increasing power increased /sp. Radiation-cooled arcjets show better 17 and 7sp than water-cooled devices, but raise technical problems because of the high temperatures of the thrusters, which require the use of special refractory materials. Proper arcjet optimization was done with a thorough thermal analysis, including the propellant flow. A further improvement of these thrusters was reached by regenerative cooling and by optimizing the constrictor contour. The constrictor flow is modeled by a three-channel model, the results of which are compared with experimental data. A new two-dimensional computational fluid dynamics (CFD) approach for hydrogen arcjet thrusters is presented. In 1996 a 0.7-kW ammonia arcjet is scheduled for a flight on the P3-D AMSAT satellite.

Studies on Flame Spread with Sudden Expansions of Ports of Solid Propellant Rockets under Elevated Pressure.

Abstract: A detailed experimental study on flame spread over non-uniform ports of solid propellant rockets has been carried out. An idealised. 2-dimensional laboratory motor was used for the experimental study with the aid of cinephotography. Freshly prepared rectangular HTPB propellant with backward facing step was used as the specimenfor this study. It has been shown conclusively that under certain conditions of step location. step height and port height which govern the velocity of gases at the step by the partially ignited propellant surface. secondary ignition may occur far downstream of the step. This is very likely to be within the recirculatingflow region. The secondary ignition gives rise to two additional flame fronts one of which spreads backwardat a relatively lower velocity,presumably due to the low reverse velocities present around the separation zone. This phenomenon is likely to play an important role in the ignition transient of solid propellant rockets with non-uniform ports.

Low temperature autoigniting propellant composition

Abstract: A low temperature autoigniting composition for use in a mobile occupant restraint system comprising, a low temperature melting oxidizer and a fuel, wherein the low temperature autoigniting composition autoignites in the temperature range of about 130° C. to about 175° C. In a preferred embodiment, the composition comprises a low temperature melting oxidizer, a fuel, and a catalyst, wherein the composition autoignites in the temperature range of about 130° C. to about 150° C. Preferably, the oxidizer comprises about 20 to about 70 percent by weight of the composition, the fuel comprises about 10 to about 50 percent by weight of the composition, and the catalyst comprises about 2 to about 50 percent by weight of the composition. The autoignition propellants of the invention are designed to function at low temperatures and before heat damage to an airbag deployment mechanism can occur, for example, during a fire.

Liquid-fueled inflator with a porous containment device

Abstract: A liquid fuelled inflator (10) for a vehicle passive restraint airbag system is enhanced by storing the liquid fuel or propellant in a porous solid structure (22). The structure provides enhanced control over the combustion process by its influence over dispersal of the fuel or propellant into the combustion chamber (14). Porous structures which meter the dispersal of fuel or propellant into the combustion chamber, and frangible porous structures which propel the fuel throughout the combustion chamber are described.

Theoretical and Experimental Study of Composite Solid Propellant Combustion

Abstract: A burning rate law is proposed to relate the overall burning rate of composite solid propellant to the burning rates of individual components and the geometry of the burning surface. The complex three dimensional burning surface structure is simplified into unit cells in order to allow computational investigation of the mixing mechanism near the burning surface. The conservation equations of mass, momentum and mixture fraction have been solved numerically to obtain the velocity and mixture fraction field above the surface. The averaged mixture fraction on the burning surface of oxidiser and of the binder is correlated with both the surface geometry and Reynolds number. Experimental results for detailed surface structure and burning rate for specially formulated ammonium perchlorate (AP) composite propellant grains have been obtained at different pressure. The experimental results are explained using the averaged mixture fraction on the surface. It is concluded that the combustion of AP composite propellan...

Hybrid-LO2-LH2 low cost launch vehicle

Abstract: A low-cost rocket or thruster has a low-cost propellant injector, in which fluid fuel and oxidizer are injected into a combustion chamber. The walls of the combustion chamber are protected from the high temperatures of the combustion by a grain of solid propellant, the surface of which tends to melt andor vaporize in the presence of combustion temperatures, and thereby protects the walls of the chamber. The low-cost propellant injector may not mix the fluid fuel and oxidizer effectively, so that pockets of noncombusted gas may occur within the chamber. The ratio of fluid fuel and oxidizer is selected to be slightly oxidizer-rich, so that any pockets of unburned gas tend to be oxygen-rich. When the pockets come into contact with the solid fuel, the excess oxygen combusts with the gaseous solid fuel, and when the mixture is near stoichiometric, the fluid fuel combusts. Thus, an inexpensive bipropellant injector together with a solid fuel grain simultaneously allows use of solid fuel instead of insulation to protect the combustion chamber walls, and the combustion is stabilized by the presence of the solid fuel. Ideally, the amount of excess oxidizer should be sufficient to completely combust the solid fuel. The preferred propellants are LO2 and LH2, and the preferred solid fuel is HTPB.

Liquid propellant airbag inflator with auto injection combustion chamber

Abstract: An improved liquid propellant airbag inflator including a tubular sidewall internally defining a combustion chamber and a propellant reservoir separated by a piston. A propellant injection port connecting the combustion chamber to the propellant reservoir is defined by a propellant guide surface of the tubular sidewall receiving the piston. According to one embodiment of the invention, the propellant injection port is provided in the form of a plurality of longitudinal grooves. According to another embodiment of the invention, an inner diameter of the propellant guide surface is larger than an outside diameter of the piston so that a gap, acting as the propellant injection port, is created between the propellant guide surface and the piston. The present invention also provides an improved liquid propellant airbag inflator including a tubular sidewall internally defining a combustion chamber and a propellant reservoir separated by a piston, wherein the sidewall defines a series of gas exhaust ports communicating with the combustion chamber. A cylindrical exhaust sleeve extends from the piston into the combustion chamber and seals the gas exhaust ports until the inflator is initiated and the piston is forced into the propellant reservoir. The present invention additionally provides an improved liquid propellant airbag inflator including a tubular sidewall internally defining a coolant reservoir, a combustion chamber and a propellant reservoir, with a propellant piston separating the propellant reservoir from the combustion chamber, and a coolant piston, independent from the propellant piston, separating the coolant reservoir from the combustion chamber.

Transcritical Liquid Oxygen Droplet Vaporization; Effect on Rocket Combustion Instability

Abstract: The response of a liquid oxygen droplet to oscillatory ambient conditions consistent with liquid-rocket engines over a wide range of frequencies is computed. Two configurations are considered: 1) isolated droplets and 2) droplets in an array. The consequences of nonuniformities introduced in the gas phase by neighboring droplets on the droplet response are evaluated. The potential of gasification as the ratecontrolling mechanism was evaluated through the computation of a response factor derived from the Rayleigh criterion. Computations show that the peak frequency for the computed response factor is mainly correlated to the droplet lifetime and also depends on the type of flow that the droplet is experiencing (with or without reversal). Consequently, droplet secondary atomization, which causes a substantial droplet lifetime reduction, induces a significant (one order of magnitude) shift in the peak frequency. As a result, the frequency of the maximum response factor is too high to correspond to the acoustic frequencies of the typical modes for standard cryogenic rocket engine chambers. Since droplets are likely to undergo secondary atomization in the stripping regime for most of their lifetime in these engines, this phenomenon explains the observed better stability of such engines compared to storable propellant engines. It was also shown that the droplet gasification process, whether undergoing stripping or not, can drive combustion instabilities for the longitudinal mode, under certain simplifying assumptions. The effects of mean pressure and pressure fluctuations on the droplet response were also evaluated.

Variable nonazide gas generator having multiple propellant chambers

Abstract: A gas generator (10) utilizes at least three segregated propellant container/combustion chambers (22, 32 and 34), each having a plurality of nonazide propellant grains (20, 36 and 38) therein, and an igniter (16) for igniting only the propellant grains (20) located within the first combustion chamber (22). The nonazide propellant produces enough heat energy to subsequently ignite the segregated propellant grains (36 and 38) by forced convection and/or heat conduction. The output inflation profile can be tailored to optimally cover a range of 10 to 90 percentile vehicle occupants.

New, stable medicinal compositions for generating propellant-free aerosols

Abstract: The invention relates to ethanol-containing pharmaceutical preparations for the production of propellant-free aerosols.

Process for the filing of a propellant chamber of a gas driven pump

Abstract: A process for the filling of the pumping chamber of a gas pressure driven medicine pump. A pillow filled with the propellant is formed which consists of a material permeable to the propellant. The pillow filled with the propellant is inserted into the propellant chamber which is evacuated. Then, the propellant chamber is sealed gas tight. Subsequently, the propellant diffuses through the wall of the pillow into the propellant chamber. The process according to the invention offers the possibility of filling the pumping chamber in a simple manner without allowing the penetration of foreign gases.

Hybrid inflator and related propellants

Abstract: A hybrid inflator for an automotive inflatable safety system is disclosed. In one embodiment, a mixture of an inert gas (e.g., argon) and oxygen are contained within the inflator housing and a hybrid propellant (i.e., ballistic properties similar to double-base and long-term stability similar to LOVA) is included in the gas generator.