Showing papers on "Propellant published in 1984"

Aluminum Agglomeration in Solid-Propellant Combustion

Abstract: -... Results of a series of tests on AP/aluminum/PBAN propellants, with particular attention to bimodal AP particle size distribution, show the effect of particle size of fine AP and the concentration of fine AP on aluminum agglomeration and ignition. Results are interpreted in terms of the distribution of aluminum in propellant microstructure and the proximity of AP-binder flamelets to precipitate ignition of the accumulating aluminum.

Spacecraft optimized arc rocket

Abstract: This invention relates to a thruster apparatus applicable to the environment of a space vehicle or satellite and operable for positioning such vehicle or satellite in the proper orbital location. The device utilizes a unique configuration of passageways to convey the propellant to a location adjacent an electrical arc forming device. The propellant, heated thereby, then travels out a nozzle section of the thruster to thereby produce thrust. If desired, an external heater may be provided to preheat the thruster to thereby contribute to greater efficiency in the use of propellant. Further, if desired, the thruster may include an accelerator extension.

Well treating method for stimulating recovery of fluids

Abstract: Subterranean oil and gas producing formations are fractured by providing one or more combustion gas generating units using rocket fuel type propellants disposed in a well casing at preselected depths. The well casing is filled with a compressible hydraulic fracturing fluid comprising a mixture of liquid, compressed gas, and propant material and precompressed to a pressure of about 1,000 psi or more greater than the fracture extension pressure at the depth of the zone to be fractured. At least one of the gas generating units is equipped with perforating shaped charges to form fluid exit perforations at the selected depth of the fracture zone. The gas generating units are simultaneously ignited to generate combustion gasses and perforate the well casing. The perforated zone is fractured by the rapid outflow of an initial charge of sand free combustion gas at the compression pressure followed by a charge of fracturing fluid laden with propant material and then a second charge of combustion gas. The column of precompressed fracturing fluid is discharged into the formation until the hydraulic extension pressure is reached and eventually the perforations sanded off.

In-line annular piston fixed bolt regenerative liquid propellant gun

Abstract: A regenerative liquid propellant gun structure in which the differential area piston is annular, having a peripheral cylindrical skirt extending away from the combustion chamber to define a propellant reservoir, and has an aperture permitting overrunning of a fixed bolt. The fixed bolt is shaped to define with the edge of the aperture a variable annual orifice for propellant injection as the piston moves. There is a second free piston overrunning the bolt and mating with the inside of the differential area piston to complete and provide for emptying of the reservoir. The structure also contains a spring to allow components to move responsive to increased combustion chamber pressure to provide for an initial movement of differential area piston relative to the bolt to start propellant injection and a fluid pressure means for movement of pistons after firings to facilitate reloading.

Status of Boron Combustion Research

Abstract: : This report describes the proceedings of an AFOSR-sponsored Specialists Meeting on Boron Combustion. The objectives of the meeting were to review current understanding and to recommend fundamental research needs concerning boron combustion. Combustion of both slurries and solid propellants containing boron was considered for airbreathing propulsion applications. Boron can provide more than twice the volumetric energy density of conventional hydrocarbon fuels for airbreathing propulsion systems--substantially improving the performance of volume-limited vehicles. Realizing this potential, however, has been difficult. Problems have been encountered in achieving adequate ignition, flame stability and combustion efficiency in practical-sized combustors. There is also evidence of energy trapping by combustion product vapors during energy conversion to generate propulsive forces, which reduces system performance. Both of these problems become more acute at low combustion chamber pressures which are associated with high-altitude operation. Several research problems were suggested in order to help resolve these difficulties. Originator supplied keywords include: Boron; Slurry; Solid propellant; Ignition; Flame stabilization; Flammability; Atomization; Secondary Breakup; Particle dispersion; Agglomerate; Percolation; Particle-Laden flows; Turbulent combustion.

Propellant tank resupply system

Abstract: A novel propellant resupply system for, and method of recharging the propellant tanks of, the attitude control system (ACS) of a spacecraft or the like are described, comprising first and second propellant conduits respectively interconnecting the outlets of the fuel and oxidizer pump inducers of the primary propulsion system with the fuel and oxidizer tanks for the ACS thruster engines, and controller valves within each propellant conduit for controllably diverting fuel and oxidizer under pressure from the fuel and oxidizer pumps to the thruster fuel and oxidizer tanks during operation of the primary propulsion system engines.

Computational modeling of jet induced mixing of cryogenic propellants in low-G

Abstract: The SOLA-ECLIPSE Code is being developed to enable computational prediction of jet induced mixing in cryogenic propellant tanks in a low-gravity environment. Velocity fields, predicted for scale model tanks, are presented which compare favorably with the available experimental data. A full scale liquid hydrogen tank for a typical Orbit Transfer Vehicle is analyzed with the conclusion that coupling an axial mixing jet with a thermodynamic vent system appears tobe a viable concept for the control of tank pressure. Previously announced in STAR as N84-25000

Modeling Slag Deposition in the Space Shuttle Solid Rocket Motor

Abstract: The burning of an aluminized propellant within a solid rocket motor produces liquid A1/A12O3 droplets on the propellant surface. Upon leaving the surface and experiencing collisions with other droplets large agglomerates may form. A significant percentage of these droplets and/or agglomerates may be deposited on the inner surfaces of the motor as a consequence of their inability to follow the gas streamlines through the nozzle. The study described herein develops an understanding of this deposition process through flow modeling and the subsequent accumulation and pooling of this slag material within the Space Shuttle solid rocket motor (SRM). From this , an analytical procedure was developed for predicting certain unknowns relating to the deposition and pooling processes. The results of an analysis using this procedure have been qualitatively verified by post-test observations of solidified slag within static horizontally fired motors. The results of the analysis were used to provide an explanation for the slag in the QM-2 and to predict the effects of flight acceleration on slag formation. The procedure is applicable to any rocket motor configuration, although IS was devloped for the SRM.

Aerosol producing device

Abstract: An apparatus and method are disclosed for applying an active ingredient as an aerosol. A transducer (22) for creating an aerosol includes a propellant inlet (96) and a chamber defined by at least one chamber wall (128) confluent with the propellant inlet. A bluff body (112) is provided for creating at least one vortex (152) in the chamber when the propellant enters the propellant inlet in a supersonic flow condition. An orifice (118, 146) spaced apart from the bluff body and separated therefrom by the chamber wall is provided for creating an exiting vortex (154).

Multiple purpose ammunition

Abstract: A multi-range ammunition unit for firing at a high angle and with preselected muzzle velocities. The ammunition unit includes at least two partial propellant charges, a first one of which is fixedly mounted to the bottom of the shell casing and the second one or ones of which are fixedly mounted to the bottom of the projectile of the ammunition unit. All partial propellant charges are coaxially arranged with respect to the shell casing. Mechanical and/or electrical ignition means are provided to first ignite the first propellant charge and then selectively sequentially ignite the second propellant charge or charges.

Constant pressure end burning gas generator

Abstract: End burning gas generators for use either as propulsion systems or means for generating large volume of gases for various purposes such as the generation of the fuel for a ducted rocket engine, the rapid inflation of air bags for personal protection or recovery of submersed items or the expulsion of projectiles from subsurface launch tubes are currently widely utilized. End burning gas generator grains are ignited at one end of a generally cylindrical charge mounted within a combustion chamber which is fixed with a suitable exhaust means. For many of the applications described above, it is desirable that the gas generator burn in a uniform manner such that a constant volume of gas is generated per unit of time so that the chamber pressure and the mass flow rate of gas remain constant. It has been observed, however, that rather than regressing at a uniform rate, the propellant grain burns in a manner which produces a convex cone, the angle of which increases with time thereby causing the burning surface to increase with time and to result in a progressive pressure trace within the combustion chamber and a corresponding continually increasing mass flow rate through the nozzle. According to this invention, it has been found that an end burning propellant grain can be caused to produce a substantially uniform pressure trace throughout its burn time if the interface between the grain and its container has a length greater than the axial length of the grain. The relationship between the length of the interface and the length of the grain is selected to compensate for the increased burning rate which is observed to occur at the interface and which produces the coning effect. The increased interfacial length can be easily obtained by the use of the corrugated member between the propellant grain and the case, the length across the corrugations providing the increase over the axial length of the grain. For typical gas generator propellant compositions, it has been found that if the interface is between 1.4 and 1.6 times the axial length of the grain that a substantially uniform pressure trace can be obtained. The corrugated interface also acts to improve the bond strength between the grain and the case.

Cased cartridge ammunition ignition booster

Abstract: A cased cartridge ammunition ignition booster includes a film of booster propellant that is case-bonded without insulation or liner to the inside wall and base of an ordinary brass, steel or plastic cartridge case and ignited by a conventional primer. Flame spread occurs from the case wall inward through the non case-bonded propellant bed, that is, the bed of the propelling charge.

Solid rocket motor with nozzle containing aromatic amide fibers

Abstract: A solid rocket motor is provided which comprises a rocket case and a centrally ported propellant grain comprising a main portion and a nozzle portion, wherein the main portion is a shaped and cured first propellant and wherein the nozzle portion comprises a shaped and cured second propellant composition having a lower burn rate than the first composition and having a plurality of aromatic amide fibers dispersed therethrough.

Industrial cartridge with separated deflagrating components

Abstract: An industrial cartridge has a case that contains a propellant charge of at least two successively arranged propellant charge powders deflagrating at differing speeds. The powders are separated from each other by a gas-permeable cover extending transversely to the case axis. The case of the cartridge contains a compressible seal at the opposite end to the base end of the case. An ignition-transmitting tube is axially arranged in the case. The tube includes, at the level of the second and/or last propellant charge powder as seen from the base, a cover which is thinner than the wall of the ignition-transmitting tube.

Process for forming a liner and cast propellant charge in a rocket motor casing

Abstract: Process for forming a liner and cast propellant charge in a rocket engine casing without separately precuring the liner, comprising the steps of providing a blocked curable liner composition which is the reaction product of a prepolymer, an isocyanate curing agent, and a blocking agent; applying the liner composition to an internal surface of a rocket casing; preheating the casing and liner assembly to simultaneously prepare the assembly for receiving a propellant, unblock the liner, and precure the liner to a tacky state; casting the propellant change into interfacial contact with the liner; and cocuring the liner and propellant compositions. The blocked, curable liner composition has a very long pot life which is terminated by heating the composition sufficiently to uncouple the blocking agent from the isocyanate curing agent and initiate a rapid cure. The liner composition does not require precuring, as it is unblocked by preheating the liner and rocket casing sufficiently to receive to cast propellant. The blocked liner composition has the following structure. ##STR1##

Telescoped ammunition with dual split cartridge case

Abstract: A telescoped ammunition round has a projectile positioned in the axial cavity of a propellant charge. A generally tubular casing surrounding the propellant charge has two longitudinal splits dividing the casing into two substantially identical halves which can separate during firing and sustain the pressure created by the propellant charge without deforming. An end cap assembly couples adjacent ends of the halves of the tubular casing to each other.

Thrust nozzle system

Abstract: The invention relates to a thrust nozzle system, especially for steering arojectile, having a nozzle arrangement (3) which is fed by a propellant source, for example a gas source. The nozzle system is arranged in a housing having at least one exhaust port (11) provided in the housing, and has a control (14) for steering a thrust jet (18) of the nozzle arrangement through the exhaust port. The invention provides a thrust nozzle system of simple construction which is especially suitable for a high miniaturization, and which permits a flexible thrust impulse forming. For this purpose the thrust nozzle system (2) has a rotating nozzle or a swinging nozzle body (3) which is rotatable relative to the housing about an axis, driven by the propellant, for example by the gas stream (P) from the gas source. The drive of the rotating nozzle body is preferably achieved by an acentric thrust nozzle (10) itself. Due to the low mass and hence low inertia of the nozzle body (3), it may be caused to rotate fast. A braking arrangement (14) is provided for the rotating nozzle body for steering the thrust jet (18) in a defined direction. Such a thrust nozzle system may serve for many uses, for example in conjunction with a secondary injection system or a hot gas motor.

Investigation of pyrotechnic MTV compositions for rocket motor igniters

Abstract: Introduction P compositions based on magnesiumTeflon (MT) formulations are characterized by many advantageous properties as rocket motor igniter materials": high-energy content, low hygroscopicity, high degree of safety in preparation, low temperature and pressure dependence of the burning rate, ease of igniter pellet or grain fabrication, favorable aging characteristics, stable burning at low pressures, and low production costs. Viton A copolymer is frequently added to the binary MT mix to increase homogeneity and facilitate product fabrication (by consolidation, extrusion, or injection molding), without a substantial effect on combustion thermochemistry. In addition to excellent resistance to environmental deterioration, magnesium-TeflonViton (MTV) compositions have been noted for good ignition source characteristics. The existence of very hot solid and liquid particles and condensible and reactive species in MTV combustion products enables fast heat transfer to a solid fuel/propellant surface or to combustible gas mixtures by almost all possible modes: conduction from impinging solid particles, forced convection from gaseous species, high thermal radiation, exothermic condensation and solidification, etc. "Soft" ignition of solidpropellant motors (defined by low motor pressurization rate and absence of an ignition peak), which is very desirable in certain propellant grain configurations, is readily obtainable with MT or MTV pellets or granules. For given specific motors and igniter mass flow rates, pyrogen igniters with MTV charges achieved shorter ignition delay times than propellant charges for similar flame temperatures.

Method for burning rate characterization of solid propellants

Abstract: A method for determining the burning rate of a propellant. The method involves the single firing of a subscale, propellant motor that has been modified so that the propellant motor has a tapered cylindrical port that produces a non-neutral pressure-time trace when burned. The pressure-time trace is initially progressive (pressure increases with time), and then regressive (pressure decreases with time). Unlike conventional motors, this motor operates over a range of burning rates; therefore, the burning rate behavior of the propellant can be characterized with a single motor firing. The burning rate of the propellant is extracted from the motor pressure-time history by a computer analysis package. The analysis package employs an optimization program which uses an internal ballistics model of the motor. The ballistics model is used to generate a theoretical pressure-time trace which can be compared with the digitized output signal from the actual motor. The optimization routine of the computer determines the propellant burning rate behavior by selecting the burning rate law which, when employed in the internal ballistics model of the motor, produces the best match between the computer generated and the actual motor pressure-time traces. Thus by using the tapered port motor and by reducing the data, the burning rate of a propellant can be characterized with a single motor firing.

Polyurethanes, the most versatile binder for solid composite propellants

Abstract: More than 600 years elapsed from the time black powder was first used as a rocket propellant until organic materials were effectively used in solid propellants. The use of nitrocellulose and organic binders, perchlorate salts, and aluminum has since resulted in the present mature propellant technology. Early development was influenced by the knowledgeobtainedfrom the basic polymerization chemistrywhich resulted in the rapidly growing plastic industry after World War II. The application ofthe polyurethane polymeric systems widely used as fibers and foams evolved to the modern development of propellants. The initial development occurred in the mid-1950s when it was recognized that the polyurethane cure is clean, straightforward, and does not generate unwanted byproducts. Additionally, the use of aluminum, which is compatible with this binder system, increased the ballistic performance of the propellant. Finally, the use of bonding agents improved the structural properties of the propellants so that most of the stringent mechanical properties required for tactical and strategic missile and space applications could be obtained. In this paper, the historical development ofthe four generations of propellants and the many building blocks for this propellant system are recorded. Applications of the polyurethane propellants in major weapon / systems are also summarized. Introduction

Laser Thermal Propulsion

Abstract: Laser thermal propulsion (LTP) is studied for the case in which laser power is absorbed by a small very high-temperature plasma (about 20,000 K) and transferred to the remainder of the pure hydrogen propellant by radiation and mixing. This concept could lead to the realization of a lightweight orbital transfer vehicle propulsion system having a specific impulse in the range 1000-2000 s. Approximately 12 percent of the input power may be radiated to the thruster walls, and 15 percent of the total propellant flow must be heated to 20,000 K to provide a bulk temperature of 5000 K prior to expansion. Three principal research issues identified are: (1) conditions for hydrogen plasma ignition, (2) control of the plasma position within the laser beam, plasma stability, and plasma absorption efficiency, and (3) characterization of the mixing of the plasma and buffer flows.

Combustion of RDX/AP composite propellants at low pressures

Abstract: The burning rate characteristics of RDX (cyclotrimethylene trinitramine)/AP (ammonium perchlorate) composite propellants were intermediate between the characteristics of RDX and AP composite propellants. Fine thermocouple traverses in the combustion zones revealed that the flame structure of the RDX/AP propellants consists of two types of flames: one type is the diffusion flame streams produced by the decomposed gases of the AP particles and the binder; the other type is the premixed flame produced by the decomposed gases of the RDX particles and/or the binder. Thus, the flame structure above the burning surface appeared to be very heterogeneous. The heat feedback from the gas phase to the burning surface was significantly reduced by the reduced reaction rate of the RDX premixed flame. Therefore the burning rate of the RDX/AP propellants was lower than that of the AP propellants. The burning rate of the RDX/AP propellants was also affected by the type of binder used. The effect of the binder on the physical and chemical processes of the RDX/AP propellant burning was the alteration of the burning surface structure and the alteration of the burning rate of the AP particles in the propellants.

Propellant charge casing

Abstract: The invention relates to a combustible propellant charge casing particularly useful for cartridge ammunition. In order to simplify the manufacture of ammunition and to improve its storage capacity, a propellant charge casing made from a synthetic foil is provided. The invention achieves particularly advantageous mechanical strength and stability in the propellant charge structure when the casing is formed from a shrinking foil.

Thermal protection for propellant grains

Abstract: A thin frangible cork barrier conforming to the contour of the burning surface of individual solid propellant grains in a multiple grain rocket motor prevents premature ignition of the protected grain by the heat of combustion of previously ignited adjacent grains and is expelled through the rocket motor nozzle upon ignition of the protected grain. In a second embodiment, the thermal protection afforded by the frangible cork barrier is augmented by bonding to a thermally-protective elastomer layer which is in turn bonded to the propellant grain burning surface. By virtue of the composition and structure of the embodiments, thermal protection is afforded. Upon ignition of a protected grain the thermal barrier breaks up so as to pass readily through the throat of a small motor nozzle.

Hydraulic waste propellant macerator and method of use

Abstract: A method and apparatus for hydraulically macerating and recovering soluble component from a solid material such as waste solid propellant that is semihard. The solid material is fed into a perforated enclosure where it is held until it is macerated such that it can pass through perforations in the enclosure. While held in the enclosure, the solid material is agitated and exposed to jets of liquid solvent under a pressure which is equal to at least about 1000 psig to thereby macerate the solid material and force it through the perforations and to dissolve soluble component such as ammonium perchlorate from solid propellant. In one embodiment, the propellant residue is conveyed to a residue discharge end of the apparatus by a sloping helical conveyor. The helical conveyor provides individual contact stages for counter-current extraction and washing of ammonium perchlorate from the residue with recycled solvent.

Propellant acquisition device

Abstract: A propellant acquisition device for a maneuvering engine in a satellite vehicle in which a propellant tank or container is provided with a propellant trap located internally of the tank adjacent a tank outlet to the maneuvering engine. The trap device is positioned to receive propellant and become refilled upon maneuvering engine use which causes propellant in the tank to migrate to the trap device and overcome the on-orbit drag on the propellant which causes it to migrate away from the trap device.