Development of a three-dimensional nonlinear viscoelastic constitutive model of solid propellant

A three dimensional nonlinear viscoelastic constitutive model for the solid propellant is developed. In their earlier work, the authors have developed an isotropic constitutive model and verified it for one dimensional case. In the present work, the validity of the model is extended to three-dimensional cases. Large deformation, dewetting and cyclic loading effects are treated as the main sources of nonlinear behavior of the solid propellant. Viscoelastic dewetting criteria is used and the softening of the solid propellant due to dewetting is treated by the modulus decrease. The nonlinearities during cyclic loading are accounted for by the functions of the octahedral shear strain measure. The constitutive equation is implemented into a finite element code for the analysis of propellant grains. A commercial finite element package ‘ABAQUS’ is used for the analysis and the model is introduced into the code through a user subroutine. The model is evaluated with different loading conditions and the predicted values are in good agreement with the measured ones. The resulting model applied to analyze a solid propellant grain for the thermal cycling load.

Thermal degradation of a composite solid propellant examined by DSC - Kinetic study

The thermal decomposition of ammonium perchlorate (AP)/hydroxyl-terminated-polybutadiene (HTPB), the AP/HTPB solid propellant, was studied at different heating rates in dynamic nitrogen atmosphere. The exothermic reaction kinetics was studied by differential scanning calorimetry (DSC) in non-isothermal conditions. The Arrhenius Parameters were estimated according to the Ozawa method. The calculated activation energy was 134.5 W mol(-1), the pre-exponential factor, A, was 2.04.10(10) min(-1) and the reaction order for the global composite decomposition was estimated in 0.7 by the kinetic Shimadzu software based on the Ozawa method. The Kissinger method for obtaining the activation energy value was also used for comparison. These results are discussed here.

TG studies of a composite solid rocket propellant based on HTPB-binder

Thermal decomposition kinetics of solid rocket propellants based on hydroxyl-terminated polybutadiene-HTPB binder was studied by applying the Arrhenius and Flynn-Wall-Ozawa's methods. The thermal decomposition data of the propellant samples were analyzed by thermogravimetric analysis (TG/DTG) at different heating rates in the temperature range of 300-1200 K. TG curves showed that the thermal degradation occurred in three main stages regardless of the plasticizer (DOA) raw material, the partial HTPB/IPDI binder and the total ammonium perchlorate decompositions. The kinetic parameters E-a (activation energy) and A (pre-exponential factor) and the compensation parameter (S-p) were determined. The apparent activation energies obtained from different methods showed a very good agreement.

EVALUATION OF MAGNESIUM AS A HALL THRUSTER PROPELLANT

In this study, the use of magnesium as a Hall thruster propellant was evaluated. A xenon Hall thruster was modified such that magnesium propellant could be loaded into the anode and use waste heat from the thruster discharge to drive the propellant vaporization. A control scheme was developed, which allowed for precise control of the mass flow rate while still using plasma heating as the main mechanism for evaporation. The thruster anode, which also served as the propellant reservoir, was designed such that the open area was too low for sufficient vapor flow at normal operating temperatures (i.e. plasma heating alone). The remaining heat needed to achieve enough vapor flow to sustain thruster discharge came from a counter-wound resistive heater located behind the anode. The control system has the ability to arrest thermal runaway in a direct evaporation feed system and stabilize the discharge current during voltage-limited operation. A proportional-integral-derivative control algorithm was implemented to enable automated operation of the mass flow control system using the discharge current as the measured variable and the anode heater current as the controlled parameter. Steady-state operation at constant voltage with discharge current excursions less than 0.35 A was demonstrated for 70 min. Using this long-duration method, stable operation was achieved with heater powers as low as 6% of the total discharge power. Using the thermal mass flow control system the thruster operated stably enough and long enough that performance measurements could be obtained and compared to the performance of the thruster using xenon propellant. It was found that when operated with magnesium, the thruster has thrust ranging from 34 mN at 200 V to 39 mN at 300 V with 1.7 mg/s of propellant. It was found to have 27 mN of thrust at 300 V using 1.0 mg/s of propellant. The thrust-to-power ratio ranged from 24 mN/kW at 200 V to 18 mN/kW at 300 volts. The specific impulse was 2000 s at 200 V and upwards of 2700 s at 300 V. The anode efficiency was found to be ~23% using magnesium, which is substantially lower than the 40% anode efficiency of xenon at approximately equivalent molar flow rates. Measurements in the plasma plume of the thruster—operated using magnesium and xenon propellants—were obtained using a Faraday probe to measure off-axis current distribution, a retarding potential analyzer to measure ion energy, and a double Langmuir probe to measure plasma density, electron temperature, and plasma potential. Additionally, the off axis current distributions and ion energy distributions were compared to measurements made in krypton and bismuth plasmas obtained in previous studies of the same thruster. Comparisons showed that magnesium had the largest beam divergence of the four propellants while the others had similar divergence. The comparisons also showed that magnesium and krypton both had very low voltage utilization compared to xenon and bismuth. It is likely that the differences in plume structure are due to the atomic differences between the propellants; the ionization mean free path goes down with increasing atomic mass. Magnesium and krypton have long ionization mean free paths and therefore require physically larger thruster dimensions for efficient thruster operation and would benefit from magnetic shielding.

BETs: Propellant less de orbiting of space debris by bare electrodynamic tethers

As a fundamental contribution to limiting the increase of debris in the Space environment, a three-year project started on 1 November 2010 financed by the European Commission under the FP-7 Space Programme. It aims at developing a universal system to be carried on board future satellites launched into low Earth orbit (LEO), to allow de-orbiting at end of life. The operational system involves a conductive tape-tether left bare of insulation to establish anodic contact with the ambient plasma as a giant Langmuir probe. The project will size the three disparate dimensions of a tape for a selected de-orbit mission and determine scaling laws to allow system design for a general mission. It will implement control laws to restrain tether dynamics in/off the orbital plane; and will carry out plasma chamber measurements and numerical simulations of tether-plasma interaction. The project also involves the design and manufacturing of subsystems: electron-ejecting plasma contactors, an electric control and power module, interface elements, tether and deployment mechanisms, tether tape/end-mass as well as current collection plus free-fall, and hypervelocity impact tests.

Development of a solid propellant inflation technique for the subcompact car passenger restraint system. Phase I final report.

National Highway Traffic Safety Administration, Washington, D.C.

Solid propellant processing factors in rocket motor design.

At head of title: NASA space vehicle design criteria (chemical propulsion).

Solid propellant grain design and internal ballistics.

Cover title.

Solid propellant selection and characterization.

At head of title: NASA space vehicle design criteria (chemical propulsion)

Titan II storable propellant handbook : final handbook /

"June 1961."

Solid propellant rockets, an introductory handbook; second stage.

Includes the original work (p. 1-162) published separately in 1955.

Solid propellant combustion; analytical survey

"Based on Soviet open literature available at the Aerospace Technology Division and the Library of Congress. The survey covers the period from January 1962 through May 1966."

Propellant actuated devices.

Mode of access: Internet.

An Investigation of instabilities characteristic of liquid propellant rocket motors :

"February 1952."

Organizational, direct support and general support maintenance repair parts and special tools list including depot maintenance repair parts and special tools for propellant draining kit, guided missile, 10293020, 1450-00-140-3490, Lance Field Artillery Missile System.

Includes index.