Experimental Investigation of High-Burning-Rate Composite Solid Propellants

B ECAUSE motor thrust depends on the product of the exposed burning surface area and the propellant burning rate, high burning rates provide the ability to generate high thrust levelswithout the need for elaborate grain designs that provide high-burning surface areas. Complex, high-surface-area grain designs suffer from the fact that they reduce volumetric loading fraction (the ratio of propellant volume to chamber volume), which is one of the leading reasons for missiles’ size and weight. In addition, complex grain designs are more subject to mechanical failure under the harsh operational conditions of high-thrust/high-acceleration missiles. For these reasons, high-burning-rate propellants provide an enabling technology for a number of important missions. Ideally, alternatives can be identified that lie within the hazards class 1.3 propellant classification such that handling is less involved and the devices are inherently safer to personnel working with the system. It is well known that burning rates increase as particle sizes are decreased because the flame zone lies closer to the propellant surface where it provides more subsurface heating which drives the overall process. Because ammonium perchlorate (AP) is the most common oxidizer used in today’s composite solid propellants, there are numerous studies [1–5] that have shown burning rates increase inversely with AP particle size. The well-known Beckstead et al. multiple-flamemodel predicted a substantial increase in burning rate as AP particle size was decreased [6]. For unimodal distributions, their results showed an asymptotic ‘premixed flame’ limit that was approached at AP particle sizes somewhat below 10 m. From a practical standpoint, propellant processing limits preclude the use of this variable to a significant extent. In addition, very fine AP powder is more hazardous; particle sizes less than 15 m are classified as explosive by the U.S. Department of Transportation. Despite these drawbacks, one of the most effective strategies to increase burning rate is to minimize AP particle size such that mix viscosity and the desired hazards classification are met. Aluminum powder is the most common fuel in composite solid propellants, and reduction of its particle size has similar effects in increasing burning rate as with AP. The effect of aluminum content and powder size on burning rate has been studied in the classical literature [7–9] and has been the subject of more recent literature [10,11] with the development of nano-aluminum (nAl). Nanoaluminum has been proven to considerably enhance propellant burning rates over coarse aluminum. Shalom et al. showed in [12] that replacement of half of a propellant formulation’s 18% (by weight) coarse aluminum with nano-aluminum or ultrafine aluminum (UFAl) increased the burning rate by 84%. Using a unimodal aluminum distribution to determine the effect of particle size, Galfetti et al. [13] showed that replacing all of the micron-size aluminum in an AP/Al/hydroxyl-terminated polybutadiene (HTPB) (68/15/17) propellant increased the burning rate by 100%. Dokhan et al. in [11] showed that the addition of UFAl increased burning rate and that a higher fine-to-coarse AP ratio, in addition to the UFAl, further enhanced burning rate. With only 20% of the aluminum loading being UFAl, Dokhan et al. [11] showed that a burning-rate increase of 160% was delivered by increasing the fine AP loading from 20 to 40% and that a further 38% increase in burning rate could be realized with a fine AP loading of 60%. In addition to the effect of the fine AP loading, Dokhan et al. [11,14] obtained results that showed, for bimodal cases, substantial burning-rate increases occur at UFAl loadings as small as 20% of the total aluminum loading. Burn-rate catalysts or ‘modifiers’ are another commonmethod for enhancing burning rates of composite solid propellants. Extensive research regarding burn-rate modifiers has been conducted since the second half of the 1960s [1,3,15–25]. It has been shown that Received 18 January 2012; revision received 23 April 2012; accepted for publication 28 April 2012. Copyright © 2012 by Timothy D. Manship, Stephen D. Heister, and Patrick T. O’Neil. Published by the American Institute of Aeronautics and Astronautics, Inc., with permission. Copies of this paper may be made for personal or internal use, on condition that the copier pay the $10.00 per-copy fee to the Copyright Clearance Center, Inc., 222 Rosewood Drive, Danvers, MA 01923; include the code 0748-4658/12 and $10.00 in correspondence with the CCC. ∗Graduate Student, School of Aeronautics and Astronautics, 500 Allison Road. Member AIAA. Professor, School of Aeronautics and Astronautics, 500 Allison Road. Associate Fellow AIAA. Graduate Student, School of Mechanical Engineering, 500 Allison Road. Member AIAA (Corresponding Author). JOURNAL OF PROPULSION AND POWER Vol. 28, No. 6, November–December 2012