<rockets>

A solid propellant is a monopropellant fuel-a single mixture of several chemicals (the oxidizing agent and the reducing agent or fuel). This fuel, as implied, is in its solid state and has a preformed or molded shape. The propellant grain, this interior shape of the core is an important factor in determining a rocket's performance. The variables determining grain-relative performance are core surface area and specific impulse.

Surface area is the amount of propellant exposed to interior combustion flames, existing in a direct relationship with thrust. An increase in surface area will increase thrust but will reduce burn-time since the propellant is being consumed at an accelerated rate. The optimal thrust is typically a constant one, which can be achieved by maintaining a constant surface area throughout the burn. Examples of constant surface area grain designs include: end burning, internal-core and outer-core burning, and internal star core burning. Various shapes are used for the optimization of grain-thrust relationships since some rockets may require an initially high thrust component for takeoff while a lower thrust will suffice its post-launch regressive thrust requirements. Such a compromise has implications as seen, but it optimizes overall rocket performance. Complicated grain core patterns, in controlling the exposed surface area of the rocket's fuel, often have parts coated with a non-flammable plastic (such as cellulose acetate). This coat prevents internal combustion flames from igniting that portion of fuel, ignited only later when the burn reaches the fuel directly.

Specific Impulse, the thrust per unit propellant burned each second, measures rocket performance and more specifically, internal thrust production a product of pressure and heat. Thrust in chemical rockets (inclusive of both solid and liquid fueled rockets, is a product of the hot and expanding gasses created in the combustion of an explosive fuel (a reduction-oxidation reaction). The degree of the fuel's explosive power coupled with the rate of combustion is the specific impulse. In designing the rocket's propellant grain specific impulse must be taken into account since it can be the difference betwixt a conflagration of failure (explosion), and a successfully optimized thrust producing rocket. If a propellant with a high specific impulse is used as the fuel for a rocket whose grain design offers a high surface area ratio, high amounts of thrust will ensue ignition. And if the engine grain casing cannot withstand the extreme pressure and temperature it will rupture and explode. Thus, the function involving the variables of both specific impulse and surface area must be considered in grain design.

The departure from the use of gunpowder to more powerful fuels (higher specific impulses) marks the development of modern solid fueled rockets. Once the chemistry behind rocket fuels (fuels provide their own "air" to burn) was discovered, scientists sought the evermore-powerful fuel, constantly approaching new limits. A composite propellant is a mechanically mixed combination of the oxidizer and the fuel. Some common solid oxidizers are: ammonium perchlorate (NH4-ClO4) and ammonium nitrate (NH4-KNO3), chemicals providing far more oxygen than potassium nitrate (KNO3), the oxidizing agent in gunpowder. These oxidizers are often mixed, in making composite propellants, with synthetic rubbers such as: polystyrenes, polysulfides, and polyurethanes. Another type of propellant is homogeneous where the oxidizer and the fuel are combined as one molecule. Propellants of this type often use a double-base (combination of two propellants) of nitrocellulose and nitroglycerin (C3H5(ONO2)3).

Advantages/Disadvantages: Solid fueled rockets are relatively simple rockets. This is their chief advantage, but it also has its drawbacks. Once a solid rocket is ignited it will consume the entirety of its fuel, without any option for shutoff or thrust adjustment. Another key disadvantage is the danger involved in the premixed fuels of monopropellant rockets. In the double-base homogeneous nitrocellulose-nitroglycerin propellant, for example the nitroglycerin is too unstable (sufficient shock will detonate) to use individually add thus a more stable propellant like nitrocellulose (a form of gunpowder) is added. Composite engines, having the fuel and oxidizer as separately mixed elements, are less sensitive to shock and therefore safer to use. A relatively low specific impulse limits the use of solid rockets when large amounts of thrust precondition. The Saturn V moon rocket used nearly 8 million pounds of thrust that would not have been feasible with the use of solid propellant, requiring a high specific impulse liquid propellant. The ease of storage of solid propellant rockets is another main advantage employing a high level use in the military. Some of these rockets are small missiles such as Honest John and Nike Hercules; others are large ballistic missiles such as Polaris, Sergeant, and Vanguard. Liquid propellants may offer better performance, but the difficulties in propellant storage and handling of liquids near absolute zero (0 degrees Kelvin) has limited their use unable to meet the stringent demands the military requires of its firepower.

As with conventional solid fuels rockets, liquid fueled rockets burn a fuel and an oxidizer. The apparent distinction is the liquid state of the fuel and the oxidizer. Several layers of complexity are added to this rather innocent looking point. The unfolding performed here will illuminate the necessity for this complexity.

There are two metal tanks holding the fuel and oxidizer respectively. Due to properties (discussed later) of these two liquids, they are typically loaded into their tanks just prior to launch. The separate tanks are necessary, for many liquid fuels burn upon contact. Upon a set launching sequence two valves (see figure 4.2) open, allowing the liquid, hitherto blocked, to flow down the pipe-work. If these valves simply opened allowing the liquid propellants to flow into the combustion chamber at their own leisure, a weak (if any at all)thrust production would incur as well as an unstable flow rate (leading to a unstable thrust rate). Two solutions have been devised to solve this problem: (1) a pressurized gas feed and (2) a turbopump feed.

The simpler of the two, the pressurized gas feed, adds a tank of high pressure gas to the propulsion system. The gas, an unreactive, inert, and light gas (such as helium), is held and regulated, under intense pressure, by a valve/regulator. The purpose of this gas is maintain a pressure forced flow of the liquid propellants, pushing them out, as one might expend liquids from a straw by blowing into it. As noted, more than a valve is needed to execute this operation in a rocket, thus the regulator controls the amount of gas flowing into each propellant tank. If the gas was controlled only by a valve, opened during the initial launch sequence, the gas would flow to form an equilibrium of pressures in the gas tank, the piping, and the propellant tanks. This is problem. Although the gas tank will be able to withstand the equilibrium pressure, the piping and the propellant tanks might not and the rupture ensuing will cause a conflagration of failure. One could use propellant tank able to bear such pressures but the mass of these tanks would be exorbitant. Thus, the regulator controls a flow that maintains a constant pressure within the propellant tanks--a situation solving the problem of fuel transfer. The constant force (pressure) exerted on the surface of the propellants will give a constantly regulated flow as they are pushed into the combustion chamber. The regulator functions to maintain these constant flows by adjusting the flow of the gas entering the propellant tanks. The gas flow must constantly be regulated; as pressure is fed into the propellant tanks, to compensate for the fuel leaving, pressure is removed from the gas storage tank. And this gas is progressively being sent into the propellant tanks, as more of this low-pressure gas is necessary to maintain a constant pressure within the propellant tanks. The pressure of a gas is indirectly related to the volume the gas occupies. This law explains how the pressure decreases in the gas supply tank (whose volume does not change), and how this action can maintain a constant pressure inside the propellant tanks (whose volumes increase, as the liquid propellants exit, with the influx the gas in effect replacing the fuel). Given this criteria the flow of propellants is ultimately controlled by the pressure the system is set to maintain. Thus, a high rate of propellant flow is achieved by simply increasing the set pressure of the system.

The second, and often preferred, solution to the fuel transfer problem is a turbopump. A turbopump is the same as regular pump in function and bypasses a gas-pressurized system by sucking out the propellants and accelerating them into the combustion chamber. The idea seems simple but the implementation of it not. The gas-pressurized method worked because great pressures could be easily stored in the gas storage tank, but in the turbopump model the pump has to do all the work. And energy to run the turbopump must generated. The large propellant tanks looming over the turbopump suggest a source of stored energy. To convert this chemically stored energy to productive pump energy a miniature rocket engine is added (yes, one is not enough). This small engine typically uses the same propellants as the main rocket but at a much lower thrust production due to decreased size. The exhaust (or thrust) of this engine beats down upon a turbine (a propeller-like disk with hundreds of blades), causing it to spin rapidly. This action converts the chemical energy into the mechanical energy the turbopump needs to operate. A shaft, connected to the rotating turbine, spreads in opposite directions to two additional turbines. The rotation of these turbines, controlled by a gear train (a set of gears) along the shaft, controls the flow of the propellants the spinning of the turbines induces. This configuration is analogous to a high power waterwheel accelerating the water in a stream, where the stream in the turbopump model is the piping that leads to the combustion chamber. Note that the three turbines in this model are enclosed and entirely isolated from one another connected only by the shaft. Also note that the blades on the on the outer two turbines (the propellant accelerating turbines) are both powered, via the same shaft, by the interior, rocket powered turbine.

Now that the propellants are being gushed into the combustion chamber we run into more complications. As the oxidizer and fuel are mixed and ignited inside the combustion chamber thrust is created. Ultimately this thrust will push the rocket upwards but while inside the thrust wants to push everywhere, even into the piping the propellant is coming out of. The intense pressure created in this converging section of the propulsion system must be accounted for in determining rate of propellant flow and combustion chamber shear strength. If the rate of propellant flow is to small the propellant will not be bale to enter the combustion chamber--a problem avoided by the proper use of a gas-pressurized or turbopump system. If the combustion chambers integrity cannot maintain burn-time pressures the engine will explode, thus high-strength (although heavy) steel or an alloyed metal (composed of several metals; lighter) or composite material is used. Another problem is the intense heat created by combustion of the fuel and oxidizer. This is usually solved by circulating the propellants around the exterior of the combustion chamber and nozzle. The propellants are (as will be seen in the following paragraph) extremely cold and they evaporate slightly, as the flow over the hot surface of the combustion chamber, absorbing some of the engine's heat. This evaporation actually has three effects: (1) as mentioned, evaporative cooling, (2) increase in propellant flow (from increase in total pressure of increased volume evaporated gas), and (3) catalytic (although temperature of the relatively adiabatic system might increase, the creation of more gaseous reactants (which burn more efficiently) will probably improve overall performance).

Liquid Oxygen is the most common oxidizer used. Other oxidizers used in liquid propellant rockets includeing: hydrogen peroxide (95%, H2O2), nitric acid (HNO3), and liquid fluorine. Of these choices liquid fluorine, given a control fuel, produces the highest specific impulse (amount of thrust per unit propellant). But due to difficulties in handling this corrosive element, and due to the high temperatures it burns at, liquid fluorine is rarely used in modern liquid fueled rockets. At STP (standard temperature, 25 degrees Celsius, and pressure, 1 ATM or 760 torr) oxygen and fluorine are gaseous elements. This state could be used, and combustion would occur, but the amount of gaseous oxygen or fluorine, storable in the oxidizer tank, would be insufficient in producing useful thrust. Thus, the temperatures of these gases are significantly reduced, thereby changing into a liquid state. And it is in this form that the oxidizers must be used. The reason for this is simply that the atoms of oxygen or fluorine are much closer to one another in the liquid state-the oxidizer is more concentrated and thus more useful. The liquid fuels often used include: liquid hydrogen, liquid ammonia (NH3), hydrazine (N2H4), and kerosene (hydrocarbon).

Advantages/Disadvantages: Liquid propellant rockets are the most powerful (in terms gross thrust) propulsion systems available. They are also among the most variable, that is to say, adjustable given a large array of valves and regulators to control and augment rocket performance. Unfortunately the last point makes liquid propellant rockets intricate and complex. Not that this scares away the designers (they are "rocket scientists") but what it does do is lower reliability. Figures 4 and 5 may look relatively simple, but you may have noted that many parts and pipes and wires that connect everything were omitted. A real modern liquid bipropellant engine has thousands of piping connections carrying various cooling, fueling, or lubricating fluids. Also the various sub-parts such as the turbopump or regulator consist of a separate vertigo of pipes, wires, control valves, temperature gauges and support struts. Given this myriad of parts, the chance of one integral function failing is large. Thus, many rockets are rated in terms of reliability--one of the reasons for the Titan series' popularity. As noted before the liquid oxygen is the most commonly used oxidizer, but it too has its drawbacks. To achieve the liquid state of this element, a temperature of -183 degrees Celsius must be obtained--conditions under which oxygen readily evaporates, losing a large sum of oxidizer just while loading. Nitric acid, another powerful oxidizer, contains 76% oxygen, is in its liquid state at STP, and has a high specific gravity--all great advantages. The latter point is a measurement similar to density and as it rises higher so to does the propellant's performance. But, nitric acid is hazardous in handling (mixture with water produces a strong acid) and produces harmful by-products in combustion with a fuel, thus its use is limited.