Quote from: ClaytonBirchenough on 12/22/2013 08:35 PMAlright, I'm having some trouble calculating the ISP of certain rocket engines. As I understand, specific impulse can be calculated by dividing the total impulse of a rocket motor by its weight.You need to divide the total impulse by the weight of the propellant, not the weight of the entire motor.2008 N * s / (1.618 kg * 9.8 m/s^2) = 127 s.

Alright, I'm having some trouble calculating the ISP of certain rocket engines. As I understand, specific impulse can be calculated by dividing the total impulse of a rocket motor by its weight.

Is there a number that describes the efficiency of stages as a whole? ISP is for propellant combination and engine performance but does not factor in the dry mass of the stage needed to support the propellant and its associated engine...(Not sure if that made any sense... everything I said sounded weird )

Escape velocity is sqrt(2) times your circular orbital velocity - which makes your delta V (sqrt(2) -1) times orbital velocity. That holds if you only burn at or near perigee. (you keep the perigee, and raise apogee to infinity.) If you do continuous burn, then you're essentially trying to raise both apogee and perigee to infinity. That sees like you'd be doubling the total dV needed that way...

The delta-V for a very-low-thrust trajectory between two circular orbits is simply the difference in the orbital speeds of the two orbits. A "circular orbit at infinity" has an orbital speed of zero. If you're in a parking at, say, 7 km/s, then the low-thrust delta-V to escape is 7 km/s. ...To get to Mars, of course, you need to do more than just escape: you need a positive asymptotic speed. It's not obvious to me that there's a good simple approximation available here. If the acceleration is high enough that the delta-V for a Hohmann-like transfer from Earth's orbit to Mars's can be delivered in a couple of months, then I suppose you could just add the delta-V, recognizing that it will be a bit of an underestimate. If, on the other hand, thrust is so low that several orbits of the sun will be completed on the way to Mars, then low-thrust approximation works well, it's just that this time the relevant circular speeds are those of Earth and Mars about the sun. Most likely, though, neither of these limiting cases applies. At least they give you upper and lower bounds on the delta-V required.

However earth's orbit about the sun is a much more leisurely degree per day. (360 degrees/365 days). In this case [of a spacecraft in heliocentric orbit] we'd remain in a 60 degree neighborhood of perihelion for two months. 100 micro-gees * 2 months = 5 km/s.For heliocentric orbits, low thrust, high ISP engines might be able to enjoy some Oberth benefit. Once out of planetary wells, it seems to me we can have our ion cake and eat it too.

In a LOX/H2 engine, do the two fluids enter the combustion chamber via the injector plate as gasses or as liquids?

...even though at the same time everyone also understood that the RL-10 used "gas/gas" in operation and why....

Quote from: Avron on 01/19/2014 02:36 PMIn a LOX/H2 engine, do the two fluids enter the combustion chamber via the injector plate as gasses or as liquids? in rl-10 they are gases, acc. to this post by ranulfchttp://forum.nasaspaceflight.com/index.php?topic=30547.msg1125002#msg1125002Quote...even though at the same time everyone also understood that the RL-10 used "gas/gas" in operation and why....