Three-Dimensional Printing of “Green” Fuels for Low-Cost Small Spacecraft Propulsion Systems

Abstract

This paper details the advantages of employing modern additive manufacturing methods to fabricate hybrid rocket fuels for intrinsically safe and green small spacecraft propulsion systems. Using additive manufacturing overcomes multiple issues frequently associated with hybrid propulsion, including poor volumetric efficiency, system ignitability, and low fuel regression rates. When certain three-dimensionally printed thermoplastics are subjected to a high-voltage low-wattage charge, electrostatic arcing along the surface pyrolizes a small amount of material that, with the introduction of an oxidizer, "seeds" combustion and produces immediate and reliable ignition. Thermoplastic fuel grains can be printed with port shapes that enhance burn properties and increase volumetric efficiencies. Embedded helical fuel ports significantly increase regression rates. The presented test results from several prototype systems using gaseous oxygen and printed acrylonitrile butadiene styrene demonstrate the various advantages of additive manufacturing, including low-power ignition, regression rate enhancement, and system scalability. The test results from both ambient and vacuum tests of a 25 N flight-weight small spacecraft thruster are presented. Multiple burn tests allowed statistical characterization of ignition timing and burn-to-burn thrust, as well as total impulse consistency. The test results demonstrating specific impulse values exceeding 295 s are presented. When fully developed, this propulsion technology has the potential for "drop-in" replacement of many hydrazine-based propulsion applications. = propellant mass flow rate, kg∕s N = number of fuel grain helical loops n = fuel regression rate fit exponential factor O∕F = oxidizer-to-fuel ratio P = helical pitch distance, cm P R = helical pitch ratio P 0 = combustion chamber pressure, kPa p exit = nozzle exit pressure, kPa R g = gas constant for combustion products, J∕kg ⋅ K r = instantaneous longitudinal mean of the fuel port radius, cm r 0 = initial fuel port radius, cm _ r = instantaneous longitudinal mean of the fuel port regression rate, cm S = internal fuel port run length, cm T 0 = combustor flame temperature, K t = time, s γ = ratio of specific heats ΔM fuel = consumed fuel mass, g η = combustion efficiency μ = sample mean ρ fuel = fuel material density, g∕cm 3 σ = sample standard deviation τ rise = motor ignition time, ms

Catalyst development for an arc-ignited hydrogen peroxide/ABS hybrid rocket system

“…3 Previously, the authors of this report have experienced considerable success with small spacecraft systems using gaseous oxygen (GOX) and 3-D printed thermoplastics like ABS (acrylonitrile-butadienestyrene) as the propellants. 4,5 However, GOX was rejected for this upper-stage application due to its low density, required storage pressures, and significant potential for fire hazards when stored in significant quantities at high pressures. Due to its high density, hydrogen peroxide (H 2 O 2 ) was considered to be very promising oxidizer for this application.…”

“…6 The thruster system was adapted from previous "Green" hybrid systems designed and tested at USU. 4 For this prototype motor, the hydrogen peroxide was directly injected into the fuel grain port using a commercial hollow-cone 50-degree nozzle to atomize the oxidizer. Using Utah State's patented Arc-Ignition system, the ABS was pyrolyzed by directing a high voltage through a 3D printed ignitor cap made of ABS.…”

“…The small port configuration produced the greatest pressure drop and overall motor burn stability; however, it suffered from a significantly reduced burn lifetime compared to the larger injector. 4 Anon. Dimension 1200es, Durability Meets Affordability.…”

An Overview of Cube-Satellite Propulsion Technologies and Trends

“…Additionally, a system using AF-M315E propellant will not likely encounter design issues and failure modes associated with control of mixture ratio of propellant, vapor diffusion and reaction, and oxidizer flow decay [27]. A major disadvantage of AF-M315E is that it is difficult to ignite due to its Ionic Liquid (IL) (high water content) form [26,28]. Experiments have been conducted with multiple catalyst systems to augment its ignitability, but room temperature ignition does not currently exist and the preheating process can consume large amount of energy (up to 15,000 J).…”

“…Experiments have been conducted with multiple catalyst systems to augment its ignitability, but room temperature ignition does not currently exist and the preheating process can consume large amount of energy (up to 15,000 J). These requirements on the energy consumption impose severe restrictions on spacecraft and even more so on CubeSats [28].…”