Condensed phase reactions in solid propellants

The paper investigates the cause for the difference between differential scanning calorimetric results and mass spectrometric studies on polystyrene (PS) ammonium perchlorate (AP) propellants as related to the method of preparation of the propellant and the difference in experimental conditions by the use of mass spectrometry. Sufficient time is given for the product sublimates to interact with each other and attain equilibrium. It is shown that the propellant decomposition is a nonadditive phenomenon and that even a physical mixture of AP and PS does not yield additive decomposition products of its components. Results on the identification of a yellow compound containing chlorine in the bulk of the propellant suggest a condensed phase reaction. The occurrence of the reaction in the porous condensed phase of the propellant may explain the larger exothermicity of the propellant compared to the additive heats of decomposition of its components.

Thermal Decomposition Studies on Ammonium Perchlorate—Polybutadiene Composite Propellant

The thermal decomposition of ammonium perchlorate based solid composite propellant using carboxyl terminated polybutadiene as binder has been studied employing thermogravimetry and differential thermal analysis techniques. The thermal decomposition characteristics of the propellant have been found to be quite similar to those of pure ammonium perchlorate with activation energy, 32 Kcal/mole and 60 Kcal/mole respectively in the low and high temperature regions. The effect of the sample size and shape on the thermal decomposition has also been evaluated.

Storage stability of solid rocket fuels. 2. Effect of oxidizer loadings

Ageing behaviour, leading to ballistic changes, has been studied as a function of oxidizer loading in polystyrene/ammonium perchlorate solid-propellants. The ageing studies were carried out at 100 °C in air. Change in burning rate decreased as the oxidizer loading increased from 75 to 80%. The change in thermal decomposition rates both at 230 and 260 °C also decreased as the oxidizer loading in the propellants increased. The shapes of the plots of the changes in burning rate and thermal decomposition rate (230 and 260 °C) at different storage times for different oxidizer-loaded propellants seem to be exactly similar. These results lead to the conclusion that the thermal decomposition of the propellant may be responsible for bringing about the ballistic changes during the ageing process. Infrared studies of the binder portion of the aged propellant indicate that peroxide formation takes place during the course of ageing and that peroxide formation for a particular storage time and temperature increases as the loading decreases.

Combustion and thermal decomposition characteristics of composite solid propellants

The role of thermal decomposition of the binder and the oxidiser in the thermal decomposition, ageing and combustion of composite solid-propellants has been investigated. The present study shows that the burning rate and ageing of polystyrene and ammonium perchlorate propellant are related to the thermal decomposition of the propellant itself and ammonium perchlorate.

Storage stability of solid rocket fuel. 1. Effect of temperature

Ageing behaviour of polystyrene (PS)/ammonium perchlorate (AP) propellent leading to ballistic changes has been studied. It follows a zero-order kinetic law. Ageing behaviour leading to change in burning rate ( ) in the temperature range of 60–200 ° C was found to remain the same. The dependence of the change of the average thermal decomposition (TD) rate at 230 and 260°C on the change in burning rate for the propellant aged at 100 ° C in air suggests that the slow TD of the propellant is the cause of ageing. The safe-life (for a pre-assigned burning-rate change limit) at 25 ° C in air has been calculated as a function of the rate of change.

Role of aluminum-magnesium alloys in humid aging of composite solid propellants

The humid aging of composite propellants containing a terpolymer of polybutadiene, acrylic acid, and acrylonitrile (PBAN) as a binder has been studied as a function of aging temperature, relative humidity, and aging time. Three composite types - AP-PBAN, AP-Al-PBAN, and AP-(Al-Mg) alloy- PBAN - have been studied. The burning rates of all three propellant types were unaffected by aging. The calorimetric values of composites containing aluminum-magnesium alloy decreased on aging, and the lattice parameter of the alloy decreased to a value close to that of aluminum. Water absorption in all of the samples increased with increases in the temperature, relative humidity, and aging time. The compression strength of the nonmetalized and aluminized samples decreased on aging, whereas that of the composites containing the alloy increased. The latter effect has been traced to reaction of residual carboxyl groups on the polymer chains with magnesium, leading to cross-linking. The reaction between the -COOH groups and magnesium has been proved using infrared spectroscopy. (Author)

Effect of Storage Temperatures on the Mechanical Properties of the Composite Solid Propellants

Accelerated ageing studies for three composite propellant formulations, namely polystyrene (PS)/ ammonium perchlorate (AP), polymethylmethacrylate (PMMA)/AP and poly phenol formaldehyde (PPF)/AP have been carried out in the temperature range of 55-125°C. Measurements of the ultimate compression strength (Uc) and isothermal decomposition rate (TD rate) were monitored as a function of storage time and temperature. The change in Uc was found to be linearly dependent on the change in TD rate irrespective of the propellant systems. Analysis of the results further revealed that the cause of ageing for both Uc and burning rate (r) is the thermal decomposition of the propellant. The safe-life for the change in mechanical properties was found to be higher compared to the change in r for PS and PMMA based propellants.

A review of hybrid rockets: present status and future potential

This paper discusses the potential of the hybrid rocket engine as a viable and attractive mode of propulsion for both space vehicles and missiles. Research and development work on this engine in other countries is presented and evaluated. The various advantages of a hybrid engine over solid and liquid engines and its problems are highlighted. It has been argued that because of the low technology needed in the development of the hybrid system, it constitutes a cost-and-time-effective propulsion system for several applications in space programmes as well as weapon systems. In support of this conclusion, experience on the developmental studies of a variable thrust 100 kg engine is presented. Some future possibilities for hybrid propulsion systems are cited.

Kinetics of propellant decomposition

Equations for solid-state decompositions which are controlled by the phase-boundary movement and nucleation have been examined using ammonium perchlorate/polystyrene propellant decomposition at 503 K and 533 K. It was found that 3 different equations governed by the nucleation process show a good fit of data at these temperatures. However, the best fit was obtained for the following Avrami-Erofeev equation, [-In (1 - α]1/4=kt.

Some observations on the ignition of composite solid propellants

Heat-up times derived from studies on the ignition characteristics of a few model composite solid propellants, containing polystyrene, carboxy-terminated polybutadiene, plasticised polyvinyl chloride and polyphenol formaldehyde as binders, show that they are directly proportional to the mass of the sample and inversely proportional to the hear flux. Propellant weight-loss prior to ignition and high pressure ignition temperature data on the propellants, ammonium per chlorate, and binders show that the ignition is governed by the gasification of the binder pyrolysis products. The activation energy for the gasification of the pyrolysed polymer products corresponds to their ignition behaviour suggesting that propellant ignition is controlled by the binder.

Universal behaviour in erosive burning of solid propellants

This paper considers the extensive data and correlations on the erosive burning of solid propellants. A relatively simple nondimensional relationship between the ratio of the actual to nonerosive burn rate (eta) and a quantity g, which is the product of g(0)-the ratio of free stream mass flux to the mass flux from the surface for nonerosive condition-and Re-0(m), where Re-0 is the Reynolds number based on the nonerosive mass flux of the propellant and port diameter, is shown to correlate most data within the accuracies of the experiments with m = -0.125. This shows the above relationship to account for the effects of pressure, aluminum, even up to a proportion of 17%, burn rate catalysts, and motor size. It is concluded that the suggested correlation between eta and g may be adopted universally for most practical propellants. (C) 1997 by The Combustion Institute.

Modeling re-ignition and chuffing in solid rocket motors

A nonlinear model is developed to numerically simulate dynamic combustion inside a solid rocket motor chamber. Using this model, the phenomena of re-ignition and chuffing are investigated under low-L* conditions. The model consists of two separate submodels (coupled to each other), one for unsteady burning of propellant and the other for unsteady conservation of mass and energy within the chamber. The latter yields instantaneous pressure and temperature within the chamber. The instantaneous burning rate is calculated using a one-dimensional, nonlinear, transient gas-phase model previously developed by the authors. The results presented in this paper show that the model predicts not only the critical L*, but also the various regimes of L*-instabihty. Specifically, the results exhibit (1) amplifying pressure oscillations leading to extinction, and (2) re-ignition after a dormant period following extinction. The re-ignition could be observed only when a radiation heat flux (from the combustion chamber to the propellant surface) was included. Certain high-frequency oscillations, possibly due to intrinsic instability, are observed when the pressure overshoots during re-ignition. At very low values of initial L*, successive cycles of extinction/reignition displaying typical characteristics of chuffing are predicted. Variations of the chuffing frequency and the thickness of propellant burned off during a chuff with L* are found to be qualitatively the same as that reported from experimental observations.

Magnetic and dielectric interactions in nano zinc ferrite powder: Prepared by self-sustainable propellant chemistry technique

The structural, magnetic and dielectric properties of nano zinc ferrite prepared by the propellant chemistry technique are studied. The PXRD measurement at room temperature reveal that the compound is in cubic spinel phase, belong to the space group Fd (3) over barm. The unit cell parameters have been estimated from Rietveld refinement. The calculated force constants from FTIR spectrum corresponding to octahedral and tetrahedral sites at 375 and 542 cm(-1) are 6.61 x 10(2) and 3.77 x 10(2) N m(-1) respectively; these values are slightly higher compared to the other ferrite systems. Magnetic hysteresis and EPR spectra show superparamagnetic property nearly to room temperature due to comparison values between magnetic anisotropy energy and the thermal energy. The calculated values of saturation magnetization, remenant magnetization, coercive field and magnetic moment supports for the existence of multi domain particles in the sample. The temperature dependent magnetic field shows the spin freezing state at 30 K and the blocking temperature at above room temperature. The frequency dependent dielectric interactions show the variation of dielectric constant, dielectric loss and impedance as similar to other ferrite systems. The AC conductivity in the prepared sample is due to the presence of electrons, holes and polarons. The synthesized material is suitable for nano-electronics and biomedical applications. (C) 2014 Elsevier B.V. All rights reserved.

Gd1.96-xYxEu0.04O3 (x=0.0, 0.49, 0.98, 1.47, 1.96 mol%) nanophosphors: Propellant combustion synthesis, structural and luminescence studies

Gd1.96-xYxEu0.04O3 (x = 0.0, 0.49, 0.98, 1.47, 1.96 mol%) nanophosphors were synthesized by propellant combustion method at low temperature (400 degrees C). The powder X-ray diffraction patterns of as formed Gd1.96Eu0.04O3 showed monoclinic phase, however with the addition of yttria it transforms from monoclinic to pure cubic phase. The porous nature increases with increase of yttria content. The particle size was estimated from Scherrer's and W-H plots which was found to be in the range 30-40 nm. These results were in well agreement with transmission electron microscopy studies. The optical band gap energies estimated were found to be in the range 5.32-5.49 eV. PL emission was recorded under 305 nm excitation show an intense emission peak at 611 nm along with other emission peaks at 582, 641 nm. These emission peaks were attributed to the transition of D-5(0) —> F-7(J) (J = 0, 1, 2, 3) of Eu3+ ions. It was observed that PL intensity increases with increase of Y content up to x = 0.98 and thereafter intensity decreases. CIE color co-ordinates indicates that at x = 1.47 an intense red bright color can be achieved, which could find a promising application in flat panel displays. The cubic and monoclinic phases show different thermoluminescence glow peak values measured under identical conditions. The response of the cubic phase to the applied dose showed good linearity, negligible fading, and simple glow curve structure than monoclinic phase indicating that suitability of this phosphor in dosimetric applications. (C) 2014 Elsevier B.V. All rights reserved.