Rocket propellant
The thrust produced can be calculated by multiplying the mass flow rate of the propellants by their exhaust velocity relative to the rocket (specific impulse).
As combustion takes place, the liquid propellant mass is converted into a huge volume of gas at high temperature and pressure.
This exhaust stream is ejected from the engine nozzle at high velocity, creating an opposing force that propels the rocket forward in accordance with Newton's laws of motion.
Rocket stages that fly through the atmosphere usually use lower performing, high molecular mass, high-density propellants due to the smaller and lighter tankage required.
These features plus simplicity and low cost make solid propellant rockets ideal for military and space applications.
Their simplicity also makes solid rockets a good choice whenever large amounts of thrust are needed and the cost is an issue.
[citation needed] A drawback to solid rockets is that they cannot be throttled in real time, although a programmed thrust schedule can be created by adjusting the interior propellant geometry.
Solid rockets can be vented to extinguish combustion or reverse thrust as a means of controlling range or accommodating stage separation.
Casting large amounts of propellant requires consistency and repeatability to avoid cracks and voids in the completed motor.
Solid fuel rockets are intolerant to cracks and voids and require post-processing such as X-ray scans to identify faults.
The mixture is formed as a thickened liquid and then cast into the correct shape and cured into a firm but flexible load-bearing solid.
A comparison of the highest specific impulses achieved with the various solid and liquid propellant combinations used in current launch vehicles is given in the article on solid-fuel rockets.
All solid-fueled ICBMs on both sides had three initial solid stages, and those with multiple independently targeted warheads had a precision maneuverable bus used to fine tune the trajectory of the re-entry vehicles.
Some gases, notably oxygen and nitrogen, may be able to be collected from the upper atmosphere, and transferred up to low Earth orbit for use in propellant depots at substantially reduced cost.
Storable oxidizers, such as nitric acid and nitrogen tetroxide, tend to be extremely toxic and highly reactive, while cryogenic propellants by definition must be stored at low temperature and can also have reactivity/toxicity issues.
Liquid-fueled rockets require potentially troublesome valves, seals, and turbopumps, which increase the cost of the launch vehicle.
This conversion happens in the time it takes for the propellants to flow from the combustion chamber through the engine throat and out the nozzle, usually on the order of one millisecond.
In fact, LOX/LH2 rockets are generally limited in how rich they run by the performance penalty of the mass of the extra hydrogen tankage instead of the underlying chemistry.
However, liquid hydrogen is extremely well suited to upper stage use where Isp is at a premium and thrust to weight ratios are less relevant.
This means that vehicles with dense-fueled booster stages reach orbit earlier, minimizing losses due to gravity drag and reducing the effective delta-v requirement.
The main engines burned a fuel-rich hydrogen and oxygen mixture, operating continuously throughout the launch but providing the majority of thrust at higher altitudes after SRB burnout.
The first, shared with solid rocket motors, is that the casing around the fuel grain must be built to withstand full combustion pressure and often extreme temperatures as well.
However, modern composite structures handle this problem well, and when used with nitrous oxide and a solid rubber propellant (HTPB), relatively small percentage of fuel is needed anyway, so the combustion chamber is not especially large.
Liquid propellants are generally mixed by the injector at the top of the combustion chamber, which directs many small swift-moving streams of fuel and oxidizer into one another.
Liquid-fueled rocket injector design has been studied at great length and still resists reliable performance prediction.
The mixing is not a well-controlled process and generally, quite a lot of propellant is left unburned,[11] which limits the efficiency of the motor.
Additionally, as the burn continues, the hole down the center of the grain (the 'port') widens and the mixture ratio tends to become more oxidizer rich.
There has recently been an increase in hybrid motor development for nonmilitary suborbital work: GOX (gaseous oxygen) was used as the oxidizer for the Buran program's orbital maneuvering system.
Thermal rockets use inert propellants of low molecular weight that are chemically compatible with the heating mechanism at high temperatures.
