Artifact: Rocket Engine, MR-50F, Liquid Propellant, Mars Viking Lander Reaction Control

A Rocket Research Corporation (RRC) MR-50F monopropellant hydrazine thruster developed for application on the Viking Mars Spacecraft. This 8 pound thrust engine had a rated impulse of 38.69-N and a maximum duty cycle of 20,000 pulses (the thruster is capable of steady state firing for a total of 3,504 seconds). 12 of these engines were installed in groups of 3 on four quadrants of the Viking Lander’s protective Aeroshell Reaction Control System (to provide three-axis attitude control thrust from shortly after the Lander separated from the Orbiter until the aeroshell was jettisoned; as well as the de-orbit impulse required to alter the Lander’s trajectory for the Mars encounter). An additional 4 of these thrusters were installed on the Lander itself as part of the Terminal Descent System (the four MR-50F thrusters enabled roll control on the Lander and augmented the 3 Terminal Descent System main engines).

A description of the engine follows; refer to the diagram at the lower left for graphic presentation of the component locations discussed. Monopropellant hydrazine flows from a trim orifice through the thrust chamber vale and on through an all-welded single feed tube to a rigimesh dispersion element. Hydrazine is distributed from an integrated rigimesh spud through the upper 0.25-inch, 25-30 mesh Shell (Corporation) 405 catalyst bed. An intermediate bed plate with screens welded on each side separates the upper bed from the lower bed, which is 0.65 inches of 14-18 mesh Shell 405 catalyst. A lower bed plate with screen retains the catalyst. Decomposed hydrazine exists through a 40:1 expansion ratio RAO nozzle. A 1/16-inch OD x 0.010-inch wall tube is provided to measure thrust chamber pressure. The propellant valve is solenoid-operated and was manufactured by Parker Aircraft Company and was actuated by 19 watts @ 33 VDC.

The chamber and nozzle assembly consists of a catalyst chamber, nozzle and chamber pressure tap fitting. The chamber-nozzle body is fabricated from Haynes Alloy No. 25 one-piece machining with a cylindrical lower catalyst bed section and a straight 40:1 contoured nozzle (the choice of Haynes alloy was driven by its thermal properties at elevated temperature coupled with its easy machine-ability). The chamber section is designed with a minimum wall thickness of 0.020 inches. The lower end of the cylindrical portion of the chamber contains a step to which a diaphragm assembly is heliarc welded. The nozzle convergent zone is a 50-degree angle conical section blending into the 0.197 inch-diameter nozzle throat. The nozzle divergent section employs a Rao contoured nozzle.

The engine is enclosed in a shield for thermal control. At the aft end of the chamber, a diaphragm is welded between the chamber and heat shield to provide for lateral structural support, differential thermal expansion and for high thermal resistance to limit heat transfer from the chamber to the heat shield. The rocket engine assembly (REA) is mounted from the upper heat-shield flange. Total weight is 1.2 pounds.