Session
Technical Session IX: Propulsion
Location
Utah State University, Logan, UT
Abstract
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 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 potential concern when implementing traditional hybrid rocket core burn fuel grain designs for systems with low thrust levels adequate for small satellites and CubeSats. Radiative heating causes the system to be fuel rich leading to inefficient combustion and nozzle clogging. This paper presents a novel idea of using this radiative heating phenomenon in the design of a hybrid propulsion system suitable for CubeSats and small satellites. This paper presents the test results of two fuel grain designs, the first being an end burning design and the second a “sandwich” fuel grain design. ABS, PVC, Nylon-12 and acrylic known as PMMA were used as fuel and gaseous oxygen (GOX) was used as the oxidizer during this testing campaign. These propellants provide several advantages including: benign handling properties, simplified plumbing, and greater burn efficiency over traditional monopropellant hydrazine.
A Miniaturized Green End-Burning Hybrid Propulsion System for CubeSats
Utah State University, Logan, UT
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 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 potential concern when implementing traditional hybrid rocket core burn fuel grain designs for systems with low thrust levels adequate for small satellites and CubeSats. Radiative heating causes the system to be fuel rich leading to inefficient combustion and nozzle clogging. This paper presents a novel idea of using this radiative heating phenomenon in the design of a hybrid propulsion system suitable for CubeSats and small satellites. This paper presents the test results of two fuel grain designs, the first being an end burning design and the second a “sandwich” fuel grain design. ABS, PVC, Nylon-12 and acrylic known as PMMA were used as fuel and gaseous oxygen (GOX) was used as the oxidizer during this testing campaign. These propellants provide several advantages including: benign handling properties, simplified plumbing, and greater burn efficiency over traditional monopropellant hydrazine.
