Showing papers on "Propellant published in 1982"

Review of Solid-Propellant Ignition Studies

Abstract: AN extensive review of the literature on solid-propell ant ignition was made to establish the state-of-the-art. Various ignition theories, experimental measurements, and ignition criteria were critically examined. The review was summarized in easy-to-read tabular form to facilitate comparison between various studies. The effects of important parameters on ignition processes were also discussed. Major technological gaps were identified and areas for future studies recommended. Contents The study of the ignition processes of solid propellants is important for many combustion and propulsion applications. An extensive review of research work performed in this area was conducted 14 years ago by Price et al. * Because many ignition studies have been conducted in the interim, a detailed survey of literature subsequent to the review paper of Price et al.1 is presented by the authors in Ref. 2. This synoptic of Ref. 2 (in which over 100 publications are cited) brings together the developments to date and the difficulties encountered under a unified view in order to establish the stateof-the-art in solid-propellant ignition. In general, ignition of a solid propellant is a complex phenomenon which involves many physicochemical processes, as depicted in Fig. 1. The ignition consists of the following sequence of events: 1) energy transfer to the propellant by an external stimulus which can be thermal, chemical, or mechanical; 2) heating and subsequent decomposition of the solid phase; 3) diffusion of vaporized gases into the surrounding atmosphere; and 4) subsurface, heterogeneous, and/or gas-phase reactions. When the net heat evolved from chemical reactions overcomes heat losses, sustained ignition is achieved. It is generally understood that ignition is incomplete if steady-state combustion does not follow the ignition event after the removal of external energy stimulus. The time period from the start of external stimulus to the instant of sustained ignition is called ignition delay^. Generally, it is controlled by three characteristic time intervals, viz., inert heating time, mixing (diffusion plus convection) time, and reaction time. Ignition delay, however, is not simply the algebraic sum of these three characteristic time intervals since there is no clear demarcation between the mixing process and the chemical reactions; these processes may have some overlapping periods. Ignition delay is one of the most important parameters in the study of ignition. However, it is very difficult to identify precisely the instant of sustained ignition.

Synthetic polymer-propellant compositions forming cold foamed structures having a temperature at least 30° C. below ambient temperature and containing open and/or closed cells

Abstract: Synthetic polymer-propellant compositions are provided that are capable of forming foamed structures having a temperature at least 30° C. below ambient temperature and containing open and/or closed cells, which may optionally contain a material which is deposited in the pores and/or walls of the structure as the structure is formed, and comprising: (a) a film-forming synthetic polymer is an amount within the range from about 2% to about 30% by weight of the composition; (b) at least one liquefied propellant boiling below -10° C; (c) the total propellant being in an amount within the range from about 50% to about 90% by weight of the composition; and having a heat of vaporization of at least 55 calories per gram; the propellant being capable of dissolving the synthetic polymer at least in the presence of a co-solvent that is soluble in the propellant and in solutions of the synthetic polymer in the propellant at ambient temperature; and (d) at least one nonsolvent that is soluble in the propellant solution but in which the synthetic polymer is insoluble in an amount within the range from about 1% to about 85% by weight of the composition; the composition forming upon volatilization of propellant at ambient temperature and pressure a foamed structure containing open and/or closed cells and having a temperature at least 30° C. below ambient temperature.

Modeling Calculations for HMX Composite Propellants

Abstract: HMX (cyclotetramethyline tetranitramine) composite propellants burn at rates significantly lower than HMX monopropellant. To model the behavior of these propellants, a new model was developed within the framework of the Beckstead-Derr-Price (BDP) modeling approach. A time-averaging approach has been developed assuming that the propellant burns through alternate layers of binder and oxidizer at significantly different rates. The model has been compared in detail with experimental results from 17 HMX/HTPB propellants. Both the data and the model show that there is only a small dependence of rate on particle size. The model predicts that the rates of HMX/HTPB propellants will converge with increasing solids loadings, and that above -85% solids there is very little change in rate for varying formulations. The interpretation of the data using the model indicates three predominant mechanisms leading to the peculiar characteristics of HMX propellants. First, the HMX binder diffusion flame is an energy poor flame that robs energy from the products that would otherwise result from the monopropellant flame. Second, there appears to be a significant ignition delay time associated with large particles that impedes the overall rate. Third, the binder rate appears to be very significant. The model indicates that changing the rate using conventional catalysis approaches would be very difficult, since the rate is more dependent on binder decomposition characteristics than on the oxidizer.

Three stage rocket vehicle with parallel staging

Abstract: A three stage rocket vehicle (11) having a large forward propellant tank (12) and a small aft propellant tank (13) axially aligned. Secured to the rear end of the aft propellant tank (13) is an engine mount structure (14) carrying rocket engines (15). Offset and secured to the propellant tanks (12,13) is a payload structure (18). The propellants from the large forward tank (12) are fed into the aft propellant tank (13) and the rocket engines (15) are fed propellants from the aft propellant tank (13). This arrangement enables the vehicle to parallel stage its use of engines and components.

Low flow constant rate pump

Abstract: A low flow, constant rate pump is provided which comprises a fixed volume container (10), the inner walls (12) of said container defining a receptacle for a liquid to be dispensed, said container (10) being provided with means (16) to form an unobstructed opening (18) through which said liquid may be caused to flow continuously at a low-flow, contant rate, a propellant chamber (19) within said container (10) formed of a material which is permeable to vapor from a propellant (20) of the liquid-vapor type, a supply of propellant (20) of the liquid-vapor type confined within said chamber, said propellant initially being under pressure sufficient to maintain a major portion thereof in liquid form at normal temperatures.

Propellant binders cure catalyst

Abstract: A multiple component catalyst system for curing of energetic urethane bins for solid fuel propellants and gas generators based upon curing of glycidyl-azide polymer and isocynate curative mixtures Void or bubble free propellant grains are obtained by employing a cure catalyst composed of a mixture of triphenyl bismuth and dibutyltin dilaurate, preferably in a respective ratio of about 10:1 by weight The void free propellant grains have burn characteristics acceptable for missile propulsion applications

Concepts for the Design of AN Antimatter Annihilation Rocket

Abstract: Matter-antimatter annihilation is considered for spacecraft propulsion. Annihilation produces considerably more energy per unit mass of propellant than any other known means of energy production. An antimatter annihilation rocket requires several systems and components that are unique to its nature. Among these are an antimatter storage system, a means to extract the antimatter from storage, a system to transport the antimatter to the rocket engine, and the engine wherein annihilation occurs and thrust is produced. Design concepts of these systems and components are presented and discussed.

Turbulent Flow Analysis of Erosive Burning of Cylindrical Composite Solid Propellants

Abstract: An axisymmetric turbulent boundary layer was analyzed in order to investigate erosive burning in composite solid-propellant rocket motors. It was found that there is a strong interaction between the core flow acceleration and turbulence level in the boundary layer. The increase of turbulence near the surface of the propellant plays an important role in the erosive burning mechanism. Reducing the port diameter makes a rocket motor more sensitive to erosive burning. When the port diameter is uniform, the erosive burning rate increases toward the aft end of the rocket motor. This trend is less pronounced or even reversed when the port diameter is divergent.

Combustion Processes of Propellants with Embedded Metal Wires

Abstract: Research was made with double-base propellants with embedded metal wires to understand how the burning rate changes along the silver wire and to obtain what factors control the burning rate of the propellants. The burning rate along the wire is influenced greatly by the kind of metal wire, wire size, and propellant compositions. With double-base propellants having varied burning characteristics, burning rate along the silver wire was measured and the effects of the luminous flame, dark zone, and the fizz zone were studied. By means of microthermocouples, the temperature profiles in the propellant near the silver wire were measured. Based on the test results, it was clear that the factors which control the burning rate along the silver wire are the temperature gradient in the fizz zone and the temperature in the dark zone. The luminous flame zone that stands above the burning surface affects neither the burning rate of the propellant nor the burning rate along the silver wire.

Modeling of nonlinear combustion instability in solid propellant rocket motors

Abstract: A comprehensive model of nonlinear longitudinal combustion instability in solid rocket motors has been developed. The two primary elements of this stability analysis are a finite difference solution of the two-phase flow in the combustion chamber and a coupled solution of the nonlinear transient propellant burning rate. Although quasi-one dimensional, the model has been generalized to treat realistic variable cross-section and partial length grains. An excellent finite difference shock capturing technique—a combination of the Lax-Wendroff, Hybrid and Artificial Compression schemes—gives the analysis the ability to treat the multiple shock-wave type of instabilities that are frequently observed in reduced smoke solid rocket motors. Ad hoc velocity coupling models were also incorporated into the analysis. Solutions are presented demonstrating that pressure oscillations in unstable solid rocket motors (with metallized as well as unmetallized propellants) reach the same limit cycle (amplitude and waveform) independent of the characteristics of the initiating disturbance. Results obtained with the velocity coupling models demonstrate the ability to analytically predict triggering, DC pressure shifts, modulated amplitude limit cycle, and strongly nonlinear waveforms; phenomena that have all been observed in actual solid rocket motor firings.

Dual pressure solid propellant control system

Abstract: A dual pressure solid propellant control system capable of operating at temperatures of approximately 3000° F. and at multiple pressures. A solid propellant gas generator is connected to a plurality of valve clusters by a manifold, the valves and manifold being of a high temperature material and each valve being independently operable. A pressure feedback loop maintains the system pressure at a commanded value by effectively increasing or decreasing the gas exit area by varying the pulse duration modulation of the valves.

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.

Propellant augmented pressurized gas dispensing device

Abstract: A propellant augmented, gas dispensing device is charged with a pressurized gas during assembly by inserting the gas in its solid form into the pressure chamber and thereafter sealing the pressure chamber. The gas is chosen so that it is in its gaseous state at normal ambient conditions. In another aspect, an improved pressure relief assembly includes a burst disc affixed to a flow passage between the pressurized gas chamber and the propellant chamber for normally preventing flow through the passage. In its preferred form, the burst disc carries an annular score line. A support member is provided on the propellant side of the burst disc and is placed in supporting contact with the burst disc adjacent the score line. The support member has a central opening communicating with the passageway to the propellant chamber. The score line is located outwardly from the central opening. The support structure allows the disc to rupture in the region of the central opening upon the occurrence of a predetermined high pressure in the gas chamber, while permitting the disc to rupture along the score line when the pressure difference between the propellant chamber and the gas chamber exceeds a predetermined low pressure on the propellant side of the disc.

Physicochemical processes of HMX propellant combustion

Abstract: The combustion processes of HMX composite propellants were studied in order to obtaina wide spectrum of burning rates. Four types of HMX composite propellants were formulated using four different types of inert binders: polyester (HTPS), polyether (HTPE), polyacetylene (HTPA), and polybutadiene (HTPB). Each has different physical and chemical properties and hydroxy terminated groups in its structure. Experimental results revealed that the burning rate of the propellants is strongly dependent on the type of binder used. When HTPE binder was replaced with HTPA, the burning rate approximately doubled. When a low-oxygen content binder (HTPB) was used, a large amount of carbonaceous material was formed on the burning surface, making the structures of the burning surface and the gas phase very heterogeneous. However, when a high-oxygen content binder (HTPE, HTPS, or HTPA) was used, these structures became homogeneous because of the formation of a molten layer of homogeneously mixed HMX/binder on the burning surface. It has been found that the combustion mode of HMX/HTPB propellants should be differentiated from that of the other three types of propellants. Thermocouple traverses in the combusition waves revealed that an increased burning rate when using HTPA binder, is due to the increased reaction rate in the first-stage reaction zone just above the burning surface. From the experimental results obtained in this study the physical and chemical processes of HMX composite propellant combustion can be deduced.

Study on Mixed Glycols‐based Semi‐flexible Unsaturated Polyester Resins for Inhibition of Rocket Propellants

Abstract: Seven semiflexible unsaturated polyester resins based on isophthalic acid, maleic anhydride and polyethylene glycol-200 and propylene glycol (in different proportions) have been synthesised and characterised for gel time, exotherm peak temperature, tensile strength, percentage elongation, bond strength, nitroglycerine absorption, water absorption, flame resistance and voltile losses. The effect of glycol variation on these properties has also been discussed. Based on the data for various characteristics, resin 4 has been selected for inhibition and static evaluation of propellant. The propellant sustainers containing 2 NDPA have been inhibited with resin 4 without the application of any barrier coats and statically fired at ambient, cold and hot temperatures. The results of the static firings prove the worthiness of the new direct bonding inhibition system.

Insensitive high energetic explosive formulations

Abstract: A cast cured propellant and explosive with a higher volume percentage of ymer resulting in improved mechanical and safety properties is made from glycidyl azide polymer, an energetic plasticizer and HMX or RDX. Aluminum powder can also be added.

Process and apparatus for the production of carbon black

Abstract: Carbon black is produced in a flow reaction by spraying a hydrocarbon containing liquid feedstock with the aid of a propellant gas into a stream of hot reaction gases produced by burning a fuel, the feedstock-propellant gas-jet enters the reaction zone with a spraying angle that is greater than the spreading angle of a free jet. There is also described an apparatus for carrying out the process which comprises a binary injector supporting an atomizing nozzle whose head has several channels which are adjusted from zero degrees to different angles to the longitudinal axis of the nozzle.

A concept for light-powered flight

Abstract: The concept for an advanced flight transport vehicle propelled by variable cycle laser propulsion engines is described. The vehicle is designed for efficient propulsion both within the atmosphere (by momentum exchange) and in space (as a rocket). Pulsed laser power is absorbed directly into gaseous reaction propellants by electrical gas breakdown and inverse Bremsstrahlung. Small scale experiments on each key engine component have been performed, and their performance characteristics are known. Innovation results from the synergism of the various engine parts. The resultant system can demonstrate substantial performance improvements over conventional chemical thrusters and flight vehicle configurations--in a future era of plentiful beamed space power.

Propellant-active carrier system for water-based paints

Abstract: A single phase propellant solvent solution comprising dimethyl ether, water, water-soluble polar organic solvents, preferably mixtures of aliphatic alcohol and water-soluble polar organic slow drying coalescing solvent, is used to provide non-flammable aerosol water-based paint compositions. Water-solubilized oil-modified film-forming polymers are dissolved in the propellant-solvent solution and are readily dispersed from an aerosol container onto a substrate surface to form a water-resistant and durable film thereon upon air drying.

Mathematical modeling of metal particle ignition in the high-temperature flow behind a shock

Abstract: The question of the mathematical description of the flow of a mixture of a gas and solid particles, with the ignition and combustion of these latter taken into account under conditions when the continuous phase has a sufficiently high temperature, occurs in different branches of science and engineering. As examples, we mention the processes of combustion and detonation of solid propellants with metallized admixtures, when particle burnup is possible behind the flame (detonation) front. The phenomena occurring during displacement of aerosols along channels of apparatus under emergency situations are also of interest from this viewpoint. In particular, the process of propagation of detonationlike combustion waves here attracts agreat deal of attention.

Combined rocket motor warhead

Abstract: A solid fuel rocket in which the rocket casing acts as a warhead. Solid propellant rocket motors require relatively heavy cases to contain the 1000-2000 psi combustion pressure needed for efficient performance. This casing can be used as a fragmentation warhead by forming longitudinal grooves in the elongated rocket casing, causing the casing to fracture along the grooves and allowing the pressure within the casing to disperse the fragments. The resulting strip-like fragments are particularly useful against "soft" equipment targets, such as anti-vehicle and anti-radar applications. Several different methods for rupturing the case along spaced, parallel longitudinal lines are disclosed. This system eliminates the need for a separate warhead including case and explosive at the cost of a slight increase in propellant case thickness and weight.

Valve construction particularly for ram jet rocket engines and to a ram jet rocket engine construction

Abstract: A ram jet rocket includes a precombustion chamber having a solid fuel propellant which burns off so that the gases generated are deficient in oxygen when they pass through a central aperture of a valve body into a main combustion chamber in which they are further burned. The valve body aperture is closed by a bore control member which is urged into a proper setting position by a push rod mounted behind the ball downstream of the aperture. The valve body defines an annular nozzle extending from the aperture downstream through one or more elbow passages into a corresponding number of axially extending passages into the main combustion chamber. The valve construction is characterized by at least one by-pass passage which extends from the upstream end of the valve body into the elbow passage which leads to the axial passage.

Slurry casting method for double base propellants

Abstract: A slurry casting process is described capable of producing crosslinked double base propellant having improved burning rates and high specific impulse. Such propellants can be formulated to be smokeless. Improved burning rates are achieved by incorporating into a slurry of double base composition, casting powder granules containing 20% to 75% by weight of small particle ammonium perchlorate. The casting powder granules substantially retain their indentity in the cured propellant matrix. The casting powder granules have a high burning rate and are uniformly distributed throughout the propellant. The granules are responsible for increasing the burning rate of the entire crosslinked double base propellant composition of the invention.