Experiments concerning the ballistic characterization of several nanoaluminum (nAl) powders are reported. Most studies were performed with laboratory composite solid rocket propellants based on ammonium perchlorate as oxidizer and hydroxyl-terminated polybutadiene as inert binder. The ultimate objective is to understand the flame structure of differently metallized formulations and improve their specific impulse efficiency by mitigating the two-phase losses. Ballistic results confirm, for increasing nAl mass fraction or decreasing nAl size, higher steady burning rates with essentially the same pressure sensitivity and reduced average size of condensed combustion products. However, aggregation and agglomeration phenomena near the burning surface appear noticeably different for microaluminum (μAl) and nAl powders. By contrasting the associated flame structures, a particle-laden flame zone with a sensibly reduced particle size is disclosed in the case of nAl. Propellant microstructure is considered the main controlling factor. A way to predict the incipient agglomerate size for μAl propellants is proposed and verified by testing several additional ammonium perchlorate/hydroxyl-terminated polybutadiene/aluminum formulations of industrial manufacture.

Microstructure Effects in Aluminized Solid Rocket Propellants

Activated aluminum powders for space propulsion

Experimental investigation of the combustion products in an aluminised solid propellant

Metal fuels are commonly used as high-energy ingredients in both solid and hybrid rocket propulsion systems. However, aggregation and agglomeration effects can sensibly modify ballistic properties and anyhow reduce the delivered specific impulse with respect to the computed ideal value. A survey of possible effects is offered, indicating widely different flame structures. A good understanding of the relevant burning phenomena opens the way to improved system performance by reducing combustion and gasdynamic expansion losses.

Aggregation and Incipient Agglomeration in Metallized Solid Propellants and Solid Fuels for Rocket Propulsion

Burning of Metallized Composite Solid Rocket Propellants: Toward Nanometric Fuel Size

A survey is offered of the present status of microaluminized propellants industrially used worldwide in most space applications, and also new directions are pointed out making profitable use of the nanoaluminized propellants meanwhile tested in advanced international laboratories. The survey is limited to formulations with inert binders. Different industrialand research-type of solid rocket propellants, mainly but not only, of the family AP/Al/HTPB were experimentally analyzed at the Space Propulsion Laboratory of Politecnico di Milano. In general, they feature the same nominal composition but implement different grain size distributions of the oxidizer or metal fuel. The basic properties of all formulations were compared to that of a standard 68/18/14 propellant already certified for flight. In order to test their possible eligibility in terms of performance for future launchers, steady ballistic properties in terms of burning rate and flame structure were studied by a variety of experimental techniques ranging from high-speed and high resolution digital video recording to chemo-physical analyses of the solid combustion residues. Surprising differences were revealed in the flame structure of apparently similar formulations. In turn, this allowed shedding new light on the complex phenomena governing metallized solid rocket propellant burning and thus acquiring improved prediction capabilities of motor performance. Nomenclature Roman symbols Al = metallic aluminum Al2p = aluminum oxide content in XPS analysis Al2O3/Al = aluminum oxidation efficiency Alox = aluminum oxide at.% = atomic concentration C1s = carbon content in XPS analysis C2Cl4 = inert liquid Cl2p = chlorine content in XPS analysis Dp = particle diameter, μm D43 = mass mean diameter, μm DAP = mean ammonium perchlorate grain size, μm Is = specific impulse, s Is,ideal = ideal (adiabatic thermochemical equilibrium) specific impulse, s n = steady burning rate pressure sensitivity O1s = oxide content in XPS analysis p = pressure, bar qr = radiant flux, W/cm rb = steady burning rate, mm/s tign = ignition time, ms (t*)cc = average residence time in combustion chamber, ms (Vcc/Ab)1/2 = combustion chamber free volume / burning area ratio at half burning time, m Greek symbols ΔIs = specific impulse loss, s ηc* = efficiency of characteristic velocity, % ηcs = efficiency of thrust coefficient, % ηIs = efficiency of specific impulse, % ρcc = combustion chamber gas mixture density, g cm ρp = propellant density, g cm α-Al2O3 = stable phase alumina or Corundum Abbreviations AP = Ammonium Perchlorate CCPs = Condensed Combustion Products HTPB = Hydroxyl-Terminated Polybutadiene MOPs = Micrometric Oxide Particles SEM = Scanning Electron Microscopy SOPs = Smoke Oxide Particles TMD = Theoretical Maximum Density XPS = X-Ray Photoelectron Microscopy XRD = X-Ray Diffraction 1

L.T. De Luca

Papers

Burning of Aluminized Solid Rocket Propellants: from Micrometric to Nanometric Fuel Size