Talk:Arcjet rocket

Any information on the thrust produced by the Arcjet? Mass of the thruster and its rate of propellant usage would permit comparison with other types of thrusters. Andrew Swallow 08:19, 20 September 2007 (UTC)

According to http://rocket.itsc.uah.edu/u/cassibj/Arcjet%20system.htm Isp varies from around 350s to a theoretical maximum of 1500s. I don't know about thrust, but apparently it's significantly higher than with conventional ion thrusters or VASIMR. —Preceding unsigned comment added by 213.84.101.176 (talk) 20:49, 15 February 2009 (UTC)

The page you link to suggests propellant flow rates of 0.00002-0.0002 kg/s, at maximum Isp=1500 this translates to a little under 3 N of thrust (Isp * g * flow_rate). Eridane (talk) 10:35, 14 February 2010 (UTC)

In general, kinetic energy power in the exhaust is power = exhaust velocity * thrust / 2. (Isp = exhaust velocity / g). Vacuum-optimized rocket nozzles are fairly good at turning thermal energy into exhaust kinetic energy, often exceeding 80% efficiency. If using hydrazine propellant, that would supply some energy beyond the electrical input; ammonia would sap some of it. Consider a hypothetical 2kW (electrical input) hydrazine engine with 600s Isp. Hydrazine has a monopropellant Isp of ~230s in vacuum, so the engine is getting 15% of its energy from the hydrazine ((230/600)^2), the rest electrically. Exhaust power = 2kW / 0.85 * 0.8 = 1.88kW. Thrust is 0.64N, propellant consumption is 0.11 g/s. Hopefully that's clear enough you can do the same calculation on other numbers.Evand (talk) 14:41, 11 February 2016 (UTC)

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