Performance Results for the Advanced Materials Bipropellant Rocket (AMBR) Engine

Performance Characterization of the Advanced Materials Bipropellant Rocket (AMBR) engine was completed at Aerojet’s Redmond, Washington test facility in the summer of 2009. This project was funded by the NASA In-Space Propulsion Technology (ISPT) Project Office. The primary goal of this project was to maximize the specific impulse of a pressure fed, apogee class earth storable bi-propellant engine using nitrogen tetroxide (MON-3) oxidizer and hydrazine fuel. The secondary goal of the project was to take greater advantage of the high temperature capabilities of iridium/rhenium material used for the combustion chamber and nozzle. The first round of hot fire testing of the AMBR engine occurred in October of 2008; during which, the engine demonstrated a maximum specific impulse of 333.5 seconds at a mixture ratio of 1.1 and a thrust of 151-lbf (672 N). This operating thrust level at this mixture ratio was close to the thermal stability (loss of fuel film cooling) thrust limit of the engine. A second round of testing, including random vibration, shock and extended hot fire testing, was added to the program with the goal of bringing the AMBR engine to a Technology Readiness Level (TRL) of 6. Following successful random vibration and shock testing, the AMBR engine went through a second hot fire test series in the summer of 2009 to document the performance (thrust and mixture ratio) operating margins and to demonstrate a long duration burn. During this testing, the engine demonstrated a specific impulse of 333 seconds at a mixture ratio of 1.1 and a thrust of 141lbf (627 N) (i.e. the AMBR engine demonstrated high performance with margin). This performance also occurred during a successful 2,700 second long burn. This program also successfully demonstrated the secondary goal of hot fire operation at a 4000°F (2200°C) combustion chamber temperature.