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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)[reply]

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)[reply]

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)[reply]
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)[reply]
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