MOST rocket propulsion is achieved through chemical propellants, where chemical energy is converted into the kinetic energy of hot gases that are expelled from the combustion chamber. A propellant is composed of two parts, a fuel that burns and an oxidiser that aids its burning. The chemical energy is first converted into heat energy through combustion. Because of the heat energy that is released, the gaseous products of the combustion expand. As the hot gases escape through the flared-up geometry of the nozzle at the aft end of the combustion chamber, they gain kinetic energy and exit at a very high velocity. The greater the chemical energy content of the propellant, the higher the exit velocity of the hot gases and the resulting thrust.
Rocket propellants come in solid, liquid and gaseous forms. As compared to solid and liquid propellants, gaseous propellants have to be compressed to very high pressures or cooled to very low temperatures in order to achieve a high density. Propellants that are usually gases at room temperature but become liquids when cooled to very low temperatures to achieve the high density are called cryogenic liquid propellants.
Cryogenics is the science and technology of temperatures below 120 Kelvin (−153° Celsius), the limit being defined by the boiling point of methane, a principal component of natural gas.
Thus, all cryogenic rocket engines are liquid engines but they should be distinguished from rocket engines that use earth-storable liquid propellants that are liquids at ordinary temperatures and can, therefore, be stored as liquids easily. The most common cryogenic propellants used in rocket engines are liquid hydrogen (LH2), which liquefies at −253° C, as the fuel and liquid oxygen (LOX), which liquefies at −183° C, as the oxidiser.
Cryogenic propellants are preferred as rocket propellants when rockets have to carry payloads of high mass because they have the greatest efficiency in terms of thrust generated. This efficiency is measured by what is called “specific impulse”. It is defined as the thrust generated per unit mass of propellant consumed per unit time or, equivalently, the rate of mass ejected from the rocket nozzle. It is measured in units of seconds.
High values of specific impulse are obtained from high exhaust gas temperature and from exhaust gas having very low (molecular) weight. To be efficient, therefore, a propellant should have a large heat of combustion to yield high temperatures, and the combustion products should contain light molecules made of elements such as hydrogen (the lightest), carbon and oxygen. Another important factor is the density of a propellant. A given weight of a dense propellant can be carried in a smaller, lighter tank than a low-density propellant of the same weight. The advantage of cryogenic propellants is that they are the most energetic and, therefore, have the highest specific impulse. The LOX+LH2 combustion yields the highest amount of total energy and the product of combustion is water vapour, with a low molecular weight.
Compared with the specific impulse of about 260 s of solid chemical propellants, 340 s of earth-storable liquid propellants such as hydrazine and dinitrogen tetroxide, 350-360 s of semi-cryogenic propellant mixtures such as LOX and kerosene, a cryogenic mixture of LOX+LH2 has a specific impulse of about 450 s.
This means that for a given amount of thrust required, the rocket needs to carry a lesser amount of cryogenic propellants, which directly translate into a higher weight of payload it can carry. For example, to produce one tonne of thrust, the Vikas liquid engine of ISRO’s PSLV burns 3.4 kilograms of propellant per second. A cryogenic will deliver the same thrust with only 2 kg of propellant per second. This translates into a higher payload capacity of the GSLV powered by a cryogenic engine. For an upper stage cryogenic engine, with every one second increase in the specific impulse, the payload gain is of the order of 15 kg.
Unlike other propellants, the optimum mixture ratio for LOX and LH2 is not necessarily that which will produce the maximum specific impulse. Because of the extremely low density of LH2, the propellant volume decreases significantly at higher mixture ratios. The maximum specific impulse typically occurs at a mixture ratio of around 3.5. However, by increasing the mixture ratio to, say, 5.5, the storage volume is reduced by one-fourth. This results in smaller propellant tanks, a lower vehicle mass, and less drag, which offset the loss in performance that comes with using the higher mixture ratio. In practice, most LOX/LH2 engines typically operate at mixture ratios of about 5 to 6. The specific impulse can also be increased by choosing an appropriate combustion cycle. The Russian cryogenic engine supplied to ISRO uses the complex “staged combustion cycle” (SCC) as against the simpler and more flexible “gas generator cycle” (GGC), because it increases the specific impulse.
In the SCC, the fuel LH2 is burnt with a little LOX in a pre-combustion chamber. The hot gases drive the high-spin turbo pumps. The exhaust gases are then injected into the combustion chamber along with some more LOX. In the GGC, on the other hand, the exhaust gases are ejected or wasted. The SCC results in a more energy-efficient engine with a specific impulse that is marginally more than what the GGC gives.
R. Ramachandran
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