A Miniaturized, Green, End-Burning, and Sandwich Hybrid Propulsion System

Conventional hybrid rocket motors with thrust levels greater than 5 N rely on forced convection within the boundary layer as the primary heat transfer mechanism for fuel vaporization. For hybrid rockets with thrust levels of less than 5 N, oxidizer mass flow levels are sufficiently small that the rate of convective heat transfer is significantly reduced; and radiative heat transfer dominates the fuel vaporization mechanism. Radiative heating is a concern when implementing traditional hybrid rocket core-burn fuel grain designs for systems with thrust levels suitable for small satellites and CubeSats. Radiative heating induces a fuel-rich burn, which leads to low combustion efficiency and nozzle clogging. This paper presents a novel idea of using radiative heating in the design of a hybrid propulsion system suitable for CubeSats and small satellites. Additionally, this paper presents the test results of two fuel grain designs: an end-burning fuel grain design and a "sandwich " fuel grain design. This paper presents the test results of a 0.5 N end-burn system that produces a vacuum Isp of 162 s with a rise time of 1.1 s and a minimum impulse bit of 0.8 N.s. This paper also presents the test results of a 1.0 N end-burn system that produces a vacuum Isp of 136 s with a rise time of 2.2 s and a minimum impulse bit of 1.4 N.s. In addition, the test results for a 1 N sandwich system that produces a vacuum Isp of 170 s with a rise time of 0.3 s and a minimum impulse bit of 0.6 N.s are presented. Acrylonitrile butadiene styrene, polyvinyl chloride, Nylon-12, and an acrylic plastic known as polymethyl methacrylate were used as fuel; and gaseous oxygen was used as the oxidizer during these testing campaigns. These propellants provide several advantages, including benign handling properties, simplified plumbing, and greater burn efficiency over traditional monopropellant hydrazine.