Showing papers on "Propellant published in 1986"

Pressurized spray dispenser having valved mixing chamber

Abstract: A spray dispenser for a liquid to be sprayed has a container provided with an inlet valve adapted for connection to a source of pressurized driving gas for pressurizing a head space above the liquid in the container. A spray head including a spray nozzle is mounted to the container and connects through a dispensing valve to a dip tube extending to the vicinity of the container bottom, such that opening of the dispensing valve causes pressurized discharge of the liquid through the spray nozzle. A mixing chamber is provided between the dip tube and the dispensing valve, which mixing chamber communicates with the head space by means of one or more spring loaded valves which admit propellant gas from the head space into the mixing chamber so long as the head space pressure exceeds a minimum preset level sufficient to ensure adequate dispersion or misting of the spray discharge thereby warning the user of the need to replenish the propellant by a clear transition between a satisfactory discharge spray and an inadequate coalesced liquid discharge.

Multiple-propellant air vehicle and propulsion system

Abstract: A propulsion system and method having means for cooling the combustor liner and throat liner of a rocket casing by endothermic pyrolysis of a hydrocarbon fuel in the fuel passageway which is adjacent to and surrounding the combustor liner and the throat liner. Means is provided for high heat flux to the combustor liner and throat liner from combustion within the rocket casing whereby the temperature of the liners exceeds their thermal limits. Catalyst means are utilized for the endothermic pyrolysis. Hydrocarbon fuel is passed through the fuel passageway; heat is provided from the combustion of fuel in the combustion chamber to the fuel passageway by radiation through the combustor liner and the throat liner; and the hydrocarbon fuel is heated at a temperature sufficient to cause the endothermic pyrolysis of the hydrocarbon in the fuel passageway. By the propulsion system and method, heat is removed from the combustion chamber through the combustion liner and throat liner, and the temperature in the combustion chamber is reduced at the combustor liner and throat liner so that the thermal limit of the combustor liner and throat liner is not exceeded. The endothermic pyrolysis of the hydrocarbon fuel produces an improved fuel product which has a higher combustion rate, a higher combustion temperature and/or results in a fuel product having a lower molecular weight than the hydrocarbon fuel.

Coupling between acoustic velocity oscillations and solid-propellantcombustion

Abstract: : The fluid mechanics of the coupling between acoustic waves and the solid propellant combustion have been studied in a cold flow simulation of a solid propellant rocket motor. The axial variation of the oscillatory heat flux shows substantial deviations from the behavior expected from both the linear and nonlinear heuristic models of velocity coupling that are currently used in combustion stability analyses. Keywords: solid propellants, combustion, acoustics, instability.

Well treating system 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.

Ultrasonic Measurements of Solid Propellant Burning Rates in Nozzleless Rocket Motors

Abstract: Burning rates at several axial locations along the grain port of subscale nozzleless motors were measured by ultrasonic transducers that detect the instantaneous web thickness of the propellant. Two experimental devices were used for this: an axisymmetric nozzleless motor loaded with a metallized composite propellant and a twodimensional window nozzleless setup loaded with a nonmetallized composite propellant. The ultrasonic transducers give an estimate of the grain deflection during motor pressurization. In both setups, the nocrossflow burning rate measured in the head-end region agrees with standard strand burner data, and high erosive burning rates are found in the aft-end region. These erosive burning data have been represented in a format that includes the main experimental variables: port radius or channel width, no-crossflow burning rate, and mean crossflow velocity. The results reveal that the generally observed trends of threshold specific mass flow rate, no-crossflow burning rate sensitivity, and scale dependence are also demonstrated in nozzleless motors.

Internally-plasticized polyethers from substituted oxetanes

Abstract: Internally plasticized elastomeric binders for projectile propellants are cured polyethers containing THF, CYMMO plus BMEMO, BEEMO and/or OMMO mer units. The polyethers are fluid at room temperature and are miscible with nitrate ester plasticizers. Propellant binders prepared from these polyethers are curable, have low Tg 's, good mechanical properties, and Isp 's comparable to PEG-based binders.

Two Asymptotic Models for Solid Propellant Combustion

Abstract: —We derive two different asymptotic models which describe the nonsteady, nonplanar burning of certain types of homogeneous solid propellants. Motivated in part by recent work on ammonium perchlorate deflagration, we assume, in the first model, that a fraction of the pro-pellant is pyrolyzed directly to product gases at a solid/gas interface, while the remainder sublimes and burns in the gas phase. In the second model, there is a thin liquid layer between the solid and gas, with combustion occurring in both the liquid and gas phases. Our analysis exploits the largeness of activation energies to derive flame sheet models analogous to those derived for strictly gaseous and strictly condensed deflagrations. For the special case of steady, planar burning, we obtain expressions for the regression rate eigenvalue as a function of the various parameters in the problem. However, a linear stability analysis of this basic solution shows that, for sufficiently large values of a certain grouping of parameters...

Investigation of high pressure solid propellant combustion chemistryusing emission spectroscopy

Abstract: A survey of the chemiluminescent emission in the range from 280 and 800 nm from the flames of ammonium perchlorate (AP)and HMX-based solid propellants has been performed at pressures from atmospheric to 7 MPa (1000 psig). The AP propellant flame showed the emission of CH, CN, NH, and OH at atmospheric pressure (under nitrogen), as well as emission from several trace impurities such as Na, K, and Ca. As the pressure was increased, the banded molecular emission of all molecules except OH was rapidly obscured by a continuum that spanned the range 350-550 nm. In contrast, the HMX propellant showed CN, NH, and OH emission at pressures up to 7 MPa. CH emission was not detected in the HMX flame at any pressure; C2 emission was not detected in either propellant flame. Spatial intensity distributions of emitting species were obtained, showing OH and atomic emission spatially distributed throughout the propellant flame and CN and NH emission confined to the region near the surface.

Spraying and dosing apparatus

Abstract: In spraying liquid herbicides and insecticides, a predetermined dose is delivered in each operation of a dosing apparatus. The dosing apparatus comprises a piston-and-cylinder combination supported on a hangle. The piston is biased in one direction by a spring or by pressurized propellant, and movement of the piston in this one direction causes the liquid to flow into the cylinder through a check valve. Actuation of a trigger-operated valve releases pressurized propellant into the cylinder on the opposite side of the piston (or alternatively into an auxiliary cylinder) causing movement of the piston and discharge of the liquid through a second check valve and then through a spray nozzle. The dosing apparatus includes a portable backpack structure molded of a plastics material, and is especially suited for portable operation in remote locations. The backpack is formed with a reservoir for reagent and a compartment for removably holding a pressurized propellant container.

Combustion wave structures of ammonium perchlorate composite propellants

Abstract: The combustion wave structures of ammonium perchlorate composite propellants have been studied experimentally by means of thermal analysis and microthermocouple techniques. Two types of binders are used in order to examine the impact of binders on the burning rate characteristics. The binders used are hydroxy terminated polybutadiene (HTPB) and hydroxy terminated polyester (HTPE). The oxygen concentration of the HTPE is much higher than that of the HTPB. The results of thermal analysis show that HTPB decomposes exothermically and HTPE decomposes endothermically. The reaction time in the gas phase just above the burning surface is calculated based on the data obtained by the temperature distribution measurements in the combustion waves. The reaction rate of HTPE propellant is found to be higher than that of the HTPB propellant. Though the heat feedback from the gas phase to the burning surface of the HTPB propellant is smaller than that of HTPE propellant, the burning rate of HTPB propellant appears to be higher than that of HTPE propellant. The results obtained show that the chemical properties of binders play an important role not only in the reaction rate in the gas phase but also in the heat release at the burning surface.

Combined space vehicle fuel cell and space station structural building component

Abstract: A combined space vehicle fuel cell and modular space station structural building component which provides containment of propellants during launch and thereafter provides one of a plurality of modules which can be interconnected once orbit is achieved for constructing a space station or space platform. The combined space vehicle fuel cell and modular space station structural building component is configured so that a first pressure vessel for containment of one propellant is preferably concentrically positioned within a second pressure vessel for another propellant, which can result in a shorter fuel cell. Intervessel structure is included for interconnecting the pressure vessels but is lightweight due to the concentric configuration of the pressure vessels and is preferably concentrated in the aft end of the fuel cell for providing a rigid structure for receiving thrust through any and all thrust attaches to a spacecraft. The fuel cell configuration results in a lower center of gravity, which enables more vertically oriented thrust to be applied to the space vehicle. Docking structure is incorporated into the combined space vehicle fuel cell and modular space station structural building component for connection with other similar fuel cells after orbit is achieved for constructing a space station or space platform. Preferably, equipment which is not susceptible to damage through contact with the one propellant is pre-installed on the earth to alleviate the need to install such equipment after orbit is achieved, which results in a reduction in the payload requirements for space station construction. The fuel cell can also include a storage compartment for payload.

Performance characteristics of ring-cusp thrusters with xenon propellant

Abstract: The performance characteristics and operating envelope of several 30-cm ring-cusp ion thrusters with xenon propellant were investigated. Results indicate a strong performance dependence on the discharge chamber boundary magnetic fields and resultant distribution of electron currents. Significant improvements in discharge performance over J-series divergent-field thrusters were achieved for large throttling ranges, which translate into reduced cathode emission currents and reduced power dissipation which should be of significant benefit for operation at thruster power levels in excess of 10 kW. Mass spectrometry of the ion beam was documented for both the ring-cusp and J-series thrusters with xenon propellant for determination of overall thruster efficiency, and lifetime. Based on the lower centerline values of doubly charged ions in the ion beam and the lower operating discharge voltage, the screen grid erosion rate of the ring-cusp thruster is expected to be lower than the divergent-field J-series thruster by a factor of 2.

Rocket staging system

Abstract: An improved rocket staging system for missiles and the like wherein a carriage borne rocket engine assembly is sequentially employed within separate, generally aligned oxidizer stages which are generally coaxially disposed about the central rocket engine and its associated carriage. A central fuel tank is surrounded by several separate, cooperating, generally ring-shaped oxidizer tanks generally coaxially disposed about the rocket periphery. A plurality of oxidizer delivery lines run through each of the outer tanks and up to the top of the fuel tank, where a flexible hose brings oxidizer down to the engine carriage. As fuel is consumed, the rocket motor carriage slides upwardly inside the fuel tank in response to thrust. When the carriage is firmly seated inside the next higher oxidizer tank and all of the propellant has been removed from the lowest tank, the lowest tank is jettisoned to discard unnecessary mass. Thus when a stage is jettisoned, its oxidizer lines disconnect from those of the next higher stage and check valves in the lower endsd of the lines in the next stage prevent significant oxidizer spillage. Oxidizer intake ports such as solenoid valves mounted on the oxidizer delivery lines in each stage are kept open in the lowermost stage and closed in all other stages to allow oxidizer to be drawn only from the lowermost tank.

Apparatus and method for the acceleration of projectiles to hypervelocities

Abstract: A projectile is initially accelerated to a supersonic velocity and then injected into a launch tube filled with a gaseous propellant. The projectile outer surface and launch tube inner surface form a ramjet having a diffuser, a combustion chamber and a nozzle. A catalytic coated flame holder projecting from the projectile ignites the gaseous propellant in the combustion chamber thereby accelerating the projectile in a subsonic combustion mode zone. The projectile then enters an overdriven detonation wave launch tube zone wherein further projectile acceleration is achieved by a formed, controlled overdriven detonation wave capable of igniting the gaseous propellant in the combustion chamber. Ultrahigh velocity projectile accelerations are achieved in a launch tube layered detonation zone having an inner sleeve filled with hydrogen gas. An explosive, which is disposed in the annular zone between the inner sleeve and the launch tube, explodes responsive to an impinging shock wave emanating from the diffuser of the accelerating projectile thereby forcing the inner sleeve inward and imparting an acceleration to the projectile. For applications wherein solid or liquid high explosives are employed, the explosion thereof forces the inner sleeve inward, forming a throat behind the projectile. This throat chokes flow behind, thereby imparting an acceleration to the projectile.

Flow turning and admittance correction - An experimental comparison

Abstract: The proper Way of incorporating viscosity-related acoustic losses into solid propellant rocket motor linear stability analyses has been an open question for years. Mainly, two distinct theories are proposed in the literature and are referred to as the flow turning and admittance correction approaches. The two theories are briefly presented, then competitively used to derive composite solid propellant pressure-coupled response functions from experiments. The experimental setup is succinctly described and experimental results are presented. Comparisons between the theoretical approaches are made on the basis that the proper approach should yield equal values for the propellant response function, irrespective of particular grain geometry, provided that the same mean chamber pressure and frequency are achieved in the chamber. Results favor the admittance correction approach, as the flow turning approach seems to overestimate viscous losses.

Liner for a solid propellant rocket motor

Abstract: A curable liner for a solid propellent rocket motor includes a radiopaque material uniformly blended therein to provide a sufficient density difference between the propellant and insulator to enhance non-destructive X-ray evaluation of the liner propellant interface. Up to 10% powdered tungsten is added to a liner component mixture prior to incorporation in a rocket motor. Utilizing this radiopaque material in the liner mixture allows detection of previously undetectable voids, disbands or flaws at the liner propellant interface, thereby reducing the potential for failure of the rocket motor during operation.

The deflagration-to-detonation transition process for high-energy propellants - A review

Abstract: Introduction T HE deflagration-to-detonation transition (DDT) in solid energetic materials has been of interest to researchers for many decades since it has a variety of areas of applicability, ranging from industrial to military. A recent example of the former is the detonation that occurred during the production of propelling charges for hunting ammunition. In the military area, the applicability extends from gun systems and projectile impact hazards to rocket motors. Hence, the DDT process has been investigated for both voidless (cast) and porous systems. We shall use the rocket motor environment to assess the DDT hazards associated with high-energy propellants. In the area of solid propellant rocket motors, an oft-asked question is "Will a cast, well-manufactured rocket propellant grain undergo a transition to detonation from the burning mode?" The answer is no, to the best of our knowledge. Although there are few journal articles directly providing this assessment, our knowledge of the DDT mechanism in gases, liquids, and solids provides a rationale for reaching this conclusion. The rationale is based on the thesis that the deflagration process must ultimately produce a shock wave to drive the system to detonation.' That is, the shock-to-detonation transition (SDT) is the final stage in any DDT process. Consequently, in the solid-propellant rocket motor situation, the confinement provided by the motor case must be sufficient to allow the pressure from deflagration to build up to a sufficiently high shock pressure to initiate the cast propellant. As will be shown below, these shock amplitudes cannot be reached for cast propellant systems confined in rocket motor cases.

High molecular weight polycaprolactone prepolymers used in high-energy formulations

Abstract: Poly(caprolactone) polymers are provided having molecular weights of 4000 or higher, are used to form propellant binders. When the high molecular weight poly(caprolactone) polymers are used in propellant formulations and cured with isocyanates of adequate functionality, propellants are produced having improved stress and strain characteristics relative to presently formulated propellants having binders of cured, lower molecular weight poly(caprolactone) polymers.

Low Pressure Burning of Ammonium Perchlorate Composite Propellants

Abstract: The combustion wave structure of ammonium perchlorate (AP) composite propellants was observed by microphotographs and the heat transfer process from the gas phase to the con-densed phase was determined by microthermocouples. Since the thickness of the combustion wave increased with decreasing pressure, the experiments were conducted at low pressures below I atm in order to examine the structure as detailed as possible. It has been determined that the reaction zone in the gas phase consists of heterogeneous flamelets produced by the decomposed AP monopropellant flames and the decomposed gases of polymeric fuel binder. The thickness of the reaction zone decreases with decreasing the concentration of binder at a constant pressure. The heat feedback from the gas phase to the condensed phase and the heat release at the burning surface are very dependent on the type of binder used. The reaction rate in the gas phase is greater and the heat release at the burning surface is smaller for the binder (HTPE)...

High-energy compositions having castable thermoplastic binders

Abstract: A propellant formulation includes energetic particulate solids dispersed in a binder system of high molecular weight 1,2 syndiotactic butadiene and a plasticizer. The propellant is prepared by mixing above the melting temperature of the butadiene and without the use of solvents. The propellant is castable without curing.

Trident II first and second stage internal insulation

Abstract: A rocket motor insulator of varied thickness with an integrated flap to rce propellant stresses.

Priming method for rimfire cartridge

Abstract: A priming method for rimfire cartridges is disclosed in which a propellant solution is disposed adjacent a centrifugally located primer material in a rimfire cartridge. The solvent is then evaporated to leave a propellant film near the primer material in order to provide slower, more uniform ignition to thereby allow use of propellant powders which would otherwise be too fast or too sensitive.

Process and apparatus for producing plastic-bound propellant powders and explosives

Abstract: A process and apparatus for producing plastic-bound propellant powders and explosives in crystalline form, with the apparatus including an extruder comprising a casing with a feed opening, optionally a solvent supply opening and one or two extruder shafts with kneading and conveying segments. For processing the plastic binders, which polymerize photochemically or under X-rays, a casing section transparent for the rays is provided, with polymerization within the extruder being initiated by UV/VIS or X-radiation sources arranged around it and the radiation intensity and/or the wavelength of the radiation are controlled as a function of the pressure difference over a given path in a compression zone of the extruder, in such a manner that the propellant or explosive strand or strands leave the extruder in a dimensionally stable and cuttable manner.

Solid rocket motor propellants with reticulated structures embedded therein to provide variable burn rate characteristics

Abstract: A solid propellant grain which comprises an homogeneous mass of propellant material including an oxidant. A variable burn rate is provided by embedding a reticulated structure in a portion only of the propellant mass. The reticulated structure may be coated with a high thermal conductivity material to provide an increased burning rate. The coat of material is selected to preferably provide improved bonding to the propellant mass.