luke strawwalker
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Because liquid propellants have MUCH more energy than solid propellants and are MUCH more efficient. The main motivator in large rockets (IE "real" space cargo rockets) is SPECIFIC IMPULSE (ISP). This is basically a measure of fuel economy, but it goes beyond that... it has to do with how much energy the propellants can produce and how well that can be converted into propulsive effort... Thrust is only PART of the equation... ISP quickly becomes MUCH more important than thrust as the rocket ascends. High (maximum possible) thrust is only important basically at liftoff, as the rocket needs sufficient thrust to get up and moving and "punch through" the dense lower atmosphere (fighting gravity and aerodynamic drag and accelerating the rocket). As the rocket burns off propellant, it accelerates faster and faster, because the rocket engines have to push a lighter and lighter vehicle... Gee loads start building, and the rocket fairly soon either has to throttle down the engines, or shut down one or more of them to keep the acceleration gee forces in check. This can be done in several ways; for solids it has to be "designed into" the motor by using propellants with different burn rates in different parts of the motor, or by the design of the core cavity, or both. Some liquid engines, like SSME, throttle down, but this complicates the engine design since it has to be DESIGNED to be able to do this. Other engines like F-1 couldn't throttle, so the Saturn V merely shut down the center engine partway through the first stage burn to minimize the gee loads on the rocket. Once the rocket is through Max-Q (maximum aerodynamic drag and pressure on the rocket) it's mainly fighting gravity, as air drag starts falling off. At this point, ISP starts becoming more and more important to how the rocket performs, and how much payload it can lift. Usually about this time, the first stage burns out (be it solid or liquid) and the second stage takes over.
For the second stage, usually the engines have larger expansion ratios and larger engine nozzles, to make better use of the propellants at higher altitudes, since usually by the time the rocket stages, the engines are then operating basically in a vacuum for all intents and purposes. (Using larger nozzles at ground level isn't an option, as they'd be over-expanded and suffer flow separation inside the nozzle from the air pressure pushing inward on the rocket exhaust flow, creating turbulence. Likewise, nozzles optimized for sea-level operation are under-expanded in a vacuum and don't capture as much of the propulsive energy as they could from the rocket exhaust, which expands much more at altitude than at sea-level. SSME's used on the shuttle are a tradeoff between underexpansion and overexpansion; they're over-expanded at sea level and under-expanded in a vacuum). Note also that the thrust/weight ratio doesn't have to be anywhere NEAR as high on "real" space rockets as on our model rockets... thrust/weight ratios on liquid propellant large space rockets are usually in the 1.10:1 to 1.25:1 ratio. Typically thrust/weight ratios (T/W) is about 1:1 on second stages, maybe a little higher. Going too low on the T/W ratio actually loses energy and efficiency to gravity losses on ascent (remember the force of gravity is pulling everything down with an acceleration of 32.2 ft/sec squared, so if the T/W is less than 1:1 on an upper stage (flying perfectly vertical, which most do not-- that's where proper trajectory modeling and good guidance comes in) will actually DECELERATE until the T/W>1:1. SO, thrust isn't AS important, because the rocket is already MUCH lighter (since the first stage and it's propellant weight is long gone) and is getting lighter all the time, since it's constantly burning off propellants. ISP, on the other hand, becomes MUCH more important! Once a vehicle is in space (in orbit or in deep space on some kind of transfer trajectory) ISP is KING, and thrust is relatively unimportant... That's why upper stages (like the S-IVB on Saturn V, or Centaur on Atlas) can have relatively low-thrust engines compared to their weight (including payload)-- at that point, ISP is the most important, because the lower the ISP, the more propellant it takes to do the same amount of work, and the larger and heavier the tanks have to be to hold that propellant, and the stronger the stage structures have to be to handle the loads... which all equates to LESS DELIVERABLE PAYLOAD! (Plus, it "back-ripples" down the design of the rocket; larger upper stages require larger first and second stages to lift them, larger (or more) booster rocket motors or larger liquid engines to lift it, etc.) In fact, in-space propulsion stages are INCREDIBLY sensitive to ISP, because each additional pound of weight required for propellants and stage dry mass (tanks and structures) subtracts 1:1 from payload... IOW, every additional pound of weight in the in-space stage equals a pound less payload you can carry! (the lower the stage, the less sensitive the stage is to weight, and therefore ISP, from a payload capabilities standpoint... IIRC, on Saturn V, shaving a pound of weight off the third stage got you an extra pound of payload to the moon-- for the second stage, you'd have to shave off about 3-4 pounds to get an extra pound of payload to the moon, and you'd have to shave off like 11 pounds on the first stage to get an extra pound of payload... This is why having solid rockets on the FIRST stage (or booster rockets) isn't particularly critical, but you DO NOT want solids on upper stages (if you can help it!)
ISP isn't JUST determined by how much energy the propellants produce when they burn (how energetic the reaction is) but ALSO by the weight of the byproduct gases in the rocket exhaust. A lighter byproduct molecule from the burned propellants, like a lightweight car with a big motor in it, can accelerate MUCH faster and to a higher speed than a heavy car can with the same motor. That's part of the "theoretical ISP" of different propellant combinations... hydrogen and oxygen burning together give one of the highest theoretical ISP's available, because it produces a LOT of energy when burned, and the byproduct molecules can accelerate easily to higher velocities. Hydrocarbon based propellants like kerosene, on the other hand, have carbon atoms, which are heavy, in their exhausts, and thus accelerate less and are somewhat less energetic and less efficient in combustion, meaning lower theoretical ISP. Propellant combinations of various storable propellants are either even less energetic, or less efficient in combustion, therefore accelerating their exhaust gases less and resulting in less ISP. Solid propellants have the LOWEST ISP of ANY of the propellants... their combustion is usually pretty inefficient, the byproduct gases are heavier and slower to accelerate, and thus produce less energy per pound of propellant than other propellant combinations.
There are other factors to propellant choices than just thrust and ISP, important as they are. Part of it deals with the complexity of the system, the requirements of the system, and the characteristics of the propellants themselves. For instance, solid propellant rockets are EXTREMELY heavy... they must have their propellants mixed and poured or cast in a special factory, and be moved FULLY FUELED and ready to burn, from the factory to the launching pad. Liquid propellant rockets are transported EMPTY and are therefore MUCH lighter and safer than solids. All this can have HUGE effects on your supporting infrastructure. For instance, the VAB at KSC in Florida, where the Saturn V's were assembled, used to have offices inside with large staffs. When shuttle replaced Saturn, with its large SRB's, those offices had to be put elsewhere, due to the safety dangers of handling the large SRB's. The crawlers, which moved huge Saturn V's empty (but complete with cargo) were stressed to near the limits hauling the extremely heavy shuttle SRB's fully fueled to the launching pads, so much so that putting the cargo into the shuttle was done at the pads (well, that's PART of the reason anyway). Making an even LARGER rocket than Shuttle using even larger (or more) SRB's requires an all-new six-truck crawler, and even IT will be nearly maxed out... so the rocket won't be able to grow much in the future for additional power when it's needed. Liquid rockets, being moved empty, don't have these limitations (not anywhere near as limited anyway) and thus can EASILY grow in the future with larger stages, longer stages, or add-on liquid (or even solid) boosters if necessary later on.
The propellant choice also greatly depends on the demands and operating conditions of the system. For instance, say manuevering engines, like the SPS on the Apollo, the LMDE (descent engine) on the LM, or the LMAE (ascent engine) on the Lunar Module, or the OMS engines on the shuttle... or RCS thrusters on Apollo, the LM, or Shuttle-- these engines need to be EXTREMELY reliable, and as simple as possible (for lighter weight, but mostly for reliability); their absolute efficiency isn't as important, so a lower ISP is less important than high reliability and simplicity. That's why these engines use storable liquid propellants like hydrazine, nitrogen tetroxide, hydrogen peroxide, etc... things that, while highly toxic, can remain liquids without boiling off in space like cryogenic liquid propellants like liquid hydrogen and liquid oxygen do, and which are usually HYPERGOLIC, meaning they burst into flame and combust on contact with each other, eliminating the need for ignitors or other devices to start the engine up when needed. Sometimes it's a tradeoff, because the LM could have hauled MUCH MUCH more cargo had it been designed to use LH2 (liquid hydrogen) and LOX, BUT, the problems associated with storage and handling of these propellants in deep space during the three day transit to the moon (and accounting for propellant boiloff, which must be vented to prevent the tanks from exploding from the excess pressure, which means you have a more complex system and have to carry more fuel than you need to account for boiloff) would have made the already incredibly complex LM design HUGELY more complex... so the choice was made to just use lower-efficiency storable propellants that were much simpler and safer... but which couldn't deliver as much performance or as much cargo to the moon's surface... but which also didn't need a complex storage system or complicated engine-startup to get the astronauts back off the moon, either... the LMAE actually had a backup system in case the main engine start system didn't work, which basically just popped the valves open to the engine, allowing it to start instantly-- can't do that with a hydrogen engine!
A lot of times you'll see solid "kick motors" used to boost satellites or space probes to their final transfer orbits or trajectories... this is where the simplicity of a solid motor outweighs its natural inefficiency. That's why Von Braun used clusters of Sargeant solid rocket motors to inject the first US satellite Explorer I into orbit in 1958, and why some satellite launchers still use small final "kick motors" even today. Of course, for LARGE propulsion jobs like accelerating a lightweight vehicle over a long period of time, the incredibly high ISP benefits of ion engines, Hall-effect electric thrusters, solar-electric propulsion, nuclear-electric propulsion, and such REALLY shine because of the incredibly efficient usage of the propellants, meaning only relatively very small amounts of propellants are required (and even then, some propellants are more efficient than others, and some tradeoffs come into play... different propellants include LH2, ammonia, argon, etc...) The problem with these extremely efficient engines is, their thrust is TINY compared to chemical rocket engines burning propellants and expelling them out the back... so the acceleration force they can provide is tiny, though they can operate continuously for MONTHS and over time, that tiny acceleration adds up! They don't scale well, though, which limits what they can be used for... also, it limits the size of the spacecraft you can use them on... for instance, a manned spacecraft needs to quickly traverse the Van Allen Radiation Belts surrounding Earth, so a high-thrust chemical engine is needed to quickly accelerate the vehicle to the moon (or elsewhere). An unmanned cargo vehicle, lander, or spacecraft probe, however, can take MONTHS spiralling out to the moon or elsewhere in the solar system under the low but incredibly efficient thrust of ion or electric propulsion, so long as the electronics are designed to handle the radiation. The ultimate in high-thrust, high-efficiency (high ISP) rocket engines are nuclear thermal propulsion, which has been successfully developed and tested in the US and Soviet Union (our NERVA engines of the 60's were successful but canceled). The hazards of radiation contamination make them unsuitable for use in Earth's atmosphere, and the risks of a rocket malfunction spreading radioactive materials in the event the rocket blew up during launch draws the ire of the environmentalists, plus nuclear propulsion in space is a political hot-potato due to the treaties we have in place against nuclear proliferation in space...
Hope this clears it up for you!
Later! OL JR
