Design and demonstration of a radioisojet.

Abstract

A joint NASA Atomic Energy Commission effort was undertaken to develop and evaluate the performance of a radioisotope-heated thruster for possible spacecraft application. A Pm-147 oxide capsule was used as the heat source in a nominal 20-mlb NH3 thruster. The radioisojet mechanical design features were found to be quite similar to those for the thermal storage resistance jet. This permitted a parallel development and test program where the electrically heated thruster was constructed to the same physical dimensions as the radioisotope version. Initial and final testing of the electrically heated thruster was performed at General Electric Company. An interim 2-week continuous test program was performed at Mound Laboratory using both heater versions. Thrust, propellant mass flow, and temperature profiles were measured during these tests. The nominal 60-w (thermal) source produced core temperatures in excess of 1650°F at measured thrust levels up to 22 mlb. The corresponding specific impulse values were found to be greater than 230 sec for single-pulse operation having on-times up to 5 sec with thruster initially at nonflow steady state. A specific impulse of 152 sec steady state was demonstrated for continuous flow operation at 13 mlb thrust. There was excellent agreement in measured thrust and specific impulse for the electrically heated and radioisotope-heated thrusters at a rated power input of 60 w. Data correlated very closely also as a function of duty cycle and propellant flow rate over the range from zero to 100% on-time (i.e., continuous flow). Discrepancies in measured core temperatures between the two thrusters were attributed to location of sensing elements with respect to the heat source. Application of the radioisojet to a number of missions has been considered. Long-life, low-thrust applications, for which the power penalty must be borne by the auxiliary propulsion system, are attractive in terms of weight, reliability, and cost.