Ok. So when they list thrust for these commercial/NASA/Military rockets, is this the same as "Max. Thrust" they list on the motor details of our amateur motors?
The "thrust" listed on the NAR motor designations is not MAXIMUM thrust but AVERAGE thrust... and there is quite a bit of "fudging" allowed in that designation as well...
Basically, though, generally speaking, the higher the number, the higher the "kick" off the pad (usually when maximum thrust occurs, but it depends on the grain geometry of the motor).
As someone else said, if you know the burn duration of the motor, and multiply it by the average thrust, you'll get a rough approximation of the total impulse of the motor in newton/seconds.
Without seeing the thrust/time curve, however, you're really in the dark as to what the motor is actually best suited for... heavy, draggy rockets need a hard "kick" (thrust spike) at liftoff and early in flight to accelerate and get it up to a speed high enough to be stable. Then the thrust can taper off to a lower "average thrust" until motor burnout to continue accelerating. A lighter, more streamlined rocket will go higher if that impulse is spread out over a longer burn duration, however, and it doesn't require the hard kick to get it up and moving... in fact it can be counterproductive from a performance standpoint (especially if it causes the rocket to shred or something, plus drag squares as velocity doubles). That's why you don't just go for maximum thrust over everything when choosing motors.
The same thing basically applies in "real" space rockets... you need enough pure raw thrust to give you at least a 1.15-1.2 thrust/weight ratio at liftoff (for most NASA rockets, 1.3 T/W ratio is "safer"). Propellant density and stage size/weight is also important design criteria, as maximum specific impulse is nowhere near as important for liftoff and first stage acceleration as maximum thrust. That's why most rockets use either solid fuels or dense liquid fuels (kerosene/oxygen, or hypergolic propellants) for first stages. Once the rocket is out of the lower atmosphere and accelerated to the point where the second stage takes over, specific impulse starts to become more important... thrust is less of a concern, so long as thrust levels maintain sufficient acceleration to prevent excessive gravity losses. The higher the specific impulse of the propellant combination used on the second stage, the less propellant required for a given payload, and thus lower dry weight of the second stage. Second stages are FAR more sensitive to excessive dry weight than first stages-- for instance, you have to save about 10 pounds of weight on the first stage of most rockets to get an additional pound of payload to orbit... on second stages, it's much less... if you shave off about 1.1-1.2 pounds of weight from the stage you get an additional pound of payload to orbit. Third stages (on a three stage rocket) it's basically a 1:1 ratio-- every pound of dry weight you save, you can carry an additional pound of payload. SO, usually on second stages, it's beneficial to choose the highest performing (most energetic) propellant combination you can get, which is typically hydrogen/oxygen. It requires larger tank sizes, meaning a heavier stage, but it's more than made up for by the additional specific impulse from the propellant.
For third stages and in-space propulsion stages, specific impulse is KING... thrust is virtually unimportant (other than its effects on the mission timeline and such). That's why liquid hydrogen is the propellant of choice for most uses. This is also where electric propulsion really shines, where the ISP is typically DOUBLE that of the best liquid hydrogen rocket engines. The only problem there is, these types of engines have such low thrust levels that they take a LONG time to accelerate the payload to the required velocities. You could send a an unmanned probe to the Moon with a small amount of propellant and a highly efficient electric thruster, but it will have to 'spiral out' over time, making multiple passes through the Van Allen Radiation Belts, so you have to "harden" your spacecraft to deal with that repeated radiation dose. That's also why you wouldn't want to use such a low thrust high efficiency engine for a manned mission-- taking a month or two to accelerate out to the Moon and making multiple passes through the Van Allen Belts isn't a good idea with a manned spacecraft, for obvious reasons. The only factor where thrust becomes an issue for in-space propulsion is you need enough thrust to prevent excessive gravity losses or maximize benefits like the Oberth Effect, which is sort of the free "gravity assist" you get by burning your rocket engine to raise the orbit when the vehicle is approaching perigee of the orbit and the velocity is the highest.
Anyway, that's why the Saturn V was configured the way it was-- kerosene and oxygen are dense propellants with good ISP (far better than typical solid propellants) and produce high thrust in large rocket engines like F-1. A hydrogen first stage actually would have a higher specific impulse, but it would have been at least half again, if not TWICE the size of the kerosene powered S-IC stage. The specific impulse gain would not have made up for the additional mass of such a mammoth first stage. On the other hand, the additional specific impulse of liquid hydrogen and oxygen was beneficial to maximize payload on the second stage, despite the larger tankage. The S-II was about a million pounds fully fueled and the five J-2 engines only generated about 1.1 million pounds of thrust altogether, so basically at staging the T/W ratio was only about 1:1, maybe even less when one figures in the weight of the lunar stack and S-IVB above it! Of course it rapidly burns off propellant and as the stage/stack gets lighter, the T/W ratio increases. The S-IVB third stage of course used hydrogen/oxygen burned in a single J-2 engine of about 220,000 lbs thrust.
Later! OL JR
