Showing papers on "Burn rate (chemistry) published in 1998"

Effect of Propellant Temperature on Efficiency in the 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 and accelerated by an electric discharge across the propellant face. Previous research has concluded that as little as 10% of the consumed propellant is converted to plasma and efficiently accelerated. The remaining propellant is consumed in the form of late-time vaporization and particulate emission, creating minimal thrust. Critical to improving the PPT performance is improving the propellant utilization. The present work demonstrates one possible method of increasing the PPT propellant efficiency. By measuring the PPT thrust, propellant consumption, and propellant temperature while varying the power level, duration of the experimental run, and total propellant mass, a correlation is established between decreased propropellant temperature and increased propellant efficiency. The method is demonstrated by performance measurements at 60 W and S W, which show a 25% increase in thrust efficiency, while the propellant temperature decreases from 135 to 42 deg C. Larger increases in the efficiency may be realized on-orbit where operating temperatures are commonly subzero. The dependence of propellant consumption on temperature also creates systematic errors in laboratory measurements with short experimental runs, and orbit analyses where the PPT performance measured at one power level is linearly scaled to the power available on the spacecraft.

Pyrotechnic gas generant composition including high oxygen balance fuel

Abstract: A pyrotechnic gas generant composition including a high oxygen balance compound or fuel which is the resulting reaction product of aminoguanidine nitrate and nitric acid. Specifically, the resulting reaction product is a yellow precipitate that can be used alone, with or without oxidizers or other additives, for very rapid self-deflagration or in combination with oxidizers and additives. In each instance, the gas generant composition provides both high gas output and low production of solid combustion products. Further, the precipitant is relatively non-hygroscopic and has a high burn rate. Specifically, the gas generating composition is useful as a gas generator for an air bag of an occupant restraint system for an automobile, gun propellants, inflation and expulsion devices, flotation devices, pyrotechnics, fire suppression devices and smokeless, reduced smoke and smokey rocket propellants.

Advanced designs for high pressure, high performance solid propellant rocket motors

Abstract: A solid rocket propellant formulation with a burn rate slope of less than about 0.15 ips/psi (2.54 cms/69.102Pa) over a substantial portion of a pressure range of about 1,000 psi (69.102Pa) to about 7,000 psi (69.102Pa) and a temperature sensitivity of less than about 0.15 %/°F (0.15 %/0.56 °C) is provided. A high performance solid propellant rocket motor including the solid rocket propellant formulation is also provided. The rocket motor is encased in a high strength low weight motor casing which is further equipped with a nozzle throat constructed of material that has an erosion rate not more than about 2 to about 3 mils (2.54.10-3cm) per second during motor operation. The solid rocket propellant formulation can be cast in a grain pattern such that an all-boost thrust profile is achieved.

Study of Energetic Nitramine Extruded Double-Base Propellants

Abstract: This paper gives the results of an experimental study on nitramine extruded double-base (EDB)fonnulationscontaining up to 25 percent RDXin low and high calorimetric value double-base(DB)propellants. The effect of ballistic modifiers on the burn rate and pressure exponent ( 11) of promisingfonnulations has also been investigated. The data generated on various parameters reveal that ( i)nitraniine EDB propellants exhibit relatively superior thennal stability,' (ii) tensile strength andpercentage elongation are drastically altered if RDX concentration exceeds 15 per cent, (iii) 11 islowered significantly in the presence of ballistic modifiers, (iv) characteristic velocity (C*) values arehigher to that for the control tonnulation, and ( v) temperature sensitivity of burn rate is on the lowerside (0,20 -0.25 % / °C as against 0.40 % / °C) in the presence of ballistic modifiers.

Energetic oxetane propellants

Abstract: An energetic solid rocket motor propellant having one or more plateau regions of low operating pressure exponent is disclosed. The propellant is formulated from ingredients including an energetic polyoxetane, an effective amount of a plasticizer, an inorganic oxidizer in at least two discrete particle size ranges, and a refractory oxide burn rate modifier.

High pressure, high performance solid rocket hydroxy-terminated polybutadiene propellant formulations

Abstract: A solid rocket propellant formulation with a burn rate slope of less than about 0.15 ips/psi (2.54cm/69.102Pas) over a substantial portion of a pressure range of about 1,000 psi (69.102Pa) to about 7,000 psi (69.102pa) and a temperature sensitivity of less than about 0.15 %/°F is provided. The solid propellant formulation contains from 35 % to 55 % by weight ammonium perchlorate particles having an average size of 200 νm, from 25 % to 40 % by weight ammonium perchlorate particles having an average size in the range of from 2 νm to 50 νm, from 7 % to 10 % by weight hydroxy-terminated polybutadiene binder and at least one member selected from the group consisting of a co-oxidizer, a ballistic additive, and a polyisocyanate curative. A high performance solid propellant rocket motor including the solid rocket propellant formulation is also provided. The rocket motor is encased in a high strength low weight motor casing which is further equipped with a nozzle throat constructed of material that has an erosion rate of about 2 to about 3 mils per second during motor operation. The solid rocket propellant formulation can be cast in a grain pattern such that an all-boost thrust profile is achieved.

Application of Ultrasonic Technique for Measurement of Instantaneous Burn Rate of Solid Propellants .

Abstract: The ultrasonic pulse-echo technique has been applied for the measurement of instantaneous burnrate of aluminised composite solid propellants. The tests have been carried out on end-burning 30 mmthick propellant specimens at nearly constant pressure of about 1.9 MPa. Necessary software forpost-test data processing and instantaneous burn rate computations have been developed. The burnrates measured by the ultrasonic technique have been compared with those obtained from ballisticevaluation motor tests on propellant from the same mix. An accuracy of about +- 1 per cent ininstantaneous burn rate measurements and reproducibility of results have been demonstrated byapplying ultrasonic technique.

Burn Rate Modelling of Solid Rocket Propellants (Short Communication)

Abstract: A generalised model of burning of a solid rocket propellant based on kinetics of propellant hasbeen developed. A complete set of variables has been formed after examining the existing models.Buckingham theorem provides the functional form of the model, such that the existing models are thesubcases of this generalised model. This proposed model has been validated by an experimental data.

Burn Rate Modelling of Solid Rocket Propellants

Abstract: A generalised model of burning of a solid rocket propellant based on kinetics of propellant has been developed. A complete set of variables has been formed after examining the existing models. Buckingham theorem provides the functional form of the model, such that the existing models are the subcases of this generalised model. This proposed model has been validated by an experimental data.

The Influence of Propellant Loading Density on Computed Burn Rate in a Mini-Closed Bomb

Abstract: : An investigation was conducted to determine what loading density should be used to calculate propellant thermochemical properties used in closed-chamber data analysis to minimize the differences in computed burn rates observed as the propellant loading density in the closed chamber varies. A comparison between the traditional loading density of 0.2 g/sq cm and the actual propellant loading density was made. The traditional r = bP(exp n) burn rate law was used as the basis for the comparison.

Energy-driven model for HE initiation and burn

Abstract: A simple 2-state reactive flow HE burn model is described in which an approximate thermal energy is used in place of temperature to drive an Arrhenius-like rate expression. The product volume fraction and the exchange energy are determined by Newton-Raphson iteration under the twin requirements that reactant and product end up in mechanical (P+Q) equilibrium and that energy be rigorously conserved in the zone. The burn fraction is then adjusted by iterating the burn rate calculation. The rate expression is analytically integrable provided the rate coefficients can be taken as constant over a hydro cycle; we assume this to be true. Ignition is represented in two ways: by a void-collapse hot-spot model in porous zones and, in zones that are sufficiently energetic, by a direct-conversion reactant burn model. Neither the reactant nor the product EOS is part of the model prescription. This separates the rate law from the EOS parametrization and frees the user to choose any available EOSs to represent the reactant and product states. In particular, it is possible to model the reactant material with strength, which can be an important capability in threshold situations.