Engineering topics
Publications and source records attributed to Kuenzly, J. D..
The engine has a conventional body and nozzle configuration. The monopropellant fuel is fed into the thruster with dual injection tubes via an injector shell with dual spray jets. The spray jets are positioned generally opposed to each other. A heater screen pack combination thermally decomposes the fuel after injection into the combustion chamber of the thruster.
Noncatalytic (thermal-decompositon) hydrazine thrusters can operate in both the pulsing and steady-state modes to meet the propulsive requirements of long-life spacecraft. The thermal decomposition mode yields higher specific impulse than is characteristic of catalytic thrusters at similar thrust levels. This performance gain is the result of higher temperature operation and a lower fraction of ammonia dissociation. Some life limiting factors of catalytic thrusters are eliminated.
A 333 mN electrothermal thruster designed to use MIL-grade hydrazine was demonstrated to be suitable for operation with low freezing point monopropellants containing hydrazine azide, monomethylhydrazine, unsymmetrical-dimethylhydrazine and ammonia. The steady-state specific impulse was greater than 200 sec for all propellants. The pulsed-mode specific impulse for an azide blend exceeded 175 sec for pulse widths greater than 50 msec; propellants containing carbonaceous species delivered 175 sec pulsed-mode specific impulses for pulse widths greater than 100 msec. Longer thrust chamber residence times were required for the carbonaceous propellants; the original thruster design was modified by increasing the characteristic chamber length and screen packing density. Specific recommendations were made for the work required to design and develop flight worthy thrusters, including methods to increase propellant dispersal at injection, thruster geometry changes to reduce holding power levels and methods to initiate the rapid decomposition of the carbonaceous propellants.
An electrothermal thruster designed for operation with MIL-grade hydrazine is suitable for operation with propellants having lower freezing points. These propellants are 76% hydrazine - 24% hydrazine azide, Aerozine-50, 50% hydrazine - 50% monomethylydrazine, and a TRW-formulated mixture of 35% hydrazine - 50% monomethylhydrane - 15% ammonia. A steady-state specific impulse of 200 sec was exceeded by all propellants. A pulse-mode value of 175 sec specific impulse was exceeded by the azide blend for pulse widths greater than 50 ms and was met by the carbonaceous propellants for pulse widths greater than 100 ms. Longer residence times were required for the carbonaceous propellants; the original thruster design was modified by increasing the characteristic chamber length and density of screen packing. A substantial amount of thermal energy must be supplied to initiate decomposition of propellants containing unsymmetrical-dimethylhydrazine and monomethylhydrazine. The rate controlling factor appeared to be the endothermic removal of methyl radicals.
The feasibility of operating small thrust level electrothermal thrusters with monopropellants other than MIL-grade hydrazine was studied. Analytical study, design, and fabrication of demonstration thrusters was performed, and an evaluation test program was initiated to evaluate monopropellants with freezing points lower than MIL-grade hydrazine, and to determine their applicability to electrothermal thrusters for spacecraft attitude control. Five demonstration thrusters were fabricated to determine the feasibility of operation with monomethylhydrazine, Aerozine-50, 77 percent hydrazine-23 percent hydrazine azide, and a mixture of hydrazine monopropellants consisting of 35 percent hydrazine-50 percent monomethylhydrazine-15 percent ammonia. The present thruster is designed to produce a steady-state thrust level of 0.344 N at 1.724 x 1 million N/M sq feed pressure. Vacuum specific impulse goals were set at 1961 N-s/kg steady-state and 1716 N-s/kg pulsed-mode.
The feasibility of operating small thrust level electrothermal thrusters is determined with monopropellants other than MIL-grade hydrazine. The work scope includes analytical study, design and fabrication of demonstration thrusters, and an evaluation test program where monopropellants with freezing points lower than MIL-grade hydrazine are evaluated and characterized to determine their applicability to electrothermal thrusters for spacecraft attitude control. Results of propellant chemistry studies and performance analyses indicated that the most promising candidate monopropellants to be investigated are monomethylhydrazine, Aerozine-50, 77% hydrazine-23% hydrazine azide blend, and TRW formulated mixed hydrazine monopropellant (MHM) consisting of 35% hydrazine-50% monomethylhydrazine-15% ammonia.