Figure 4.1 shows a block diagram and a photo depicting the basic components of a rocket engine. To begin with, the engine needs some form of propellant. This includes both fuel and oxidizer. The main energy that will be converted to propulsion energy is stored in the propellant if it is a combustion-type engine. If the engine is simply a thermal engine, then the energy could be stored electrically or in nuclear fissile material. In the purely thermal engines, a heat source is used to heat an exhaust gas. The exhaust gas is practically inert and might be something as simple as water. In these cases, the propel- lant is simply a means to convert the heat energy into propulsive energy.
But, in most typical modern rocket engines, the heat is generated through a chemical reaction between the propellant chemicals. The fuel and the oxi- dizer are typically mixed together in an exothermic reaction. An exothermic reaction is defined as a reaction where chemical bonds are broken with less energy required than that needed to make the bonds. The excess energy is released as heat. A more simple definition is that an exothermic reaction is any reaction that releases heat. A very pertinent example of such a reaction
is the mixing of liquid hydrogen, H2, with liquid oxygen, O2. The chemical reaction is as follows:
2H2 + O2 + Ignition Heat → H2O + Excess Heat. (4.1) A spark for ignition on the left side of the equation enables the burning of the liquids together to produce water and a large amount of heat as the byproducts of the reaction. As can be seen in Equation 4.1, there is a proper mixing ratio of liquid hydrogen to liquid oxygen. Two diatomic hydrogen molecules per one diatomic oxygen need to be in the chamber for an effi- cient use of the propellants. This ratio is known as the stoichiometric ratio and is what we all became familiar with in high school chemistry when the teachers had us balancing chemical equations. Don’t be misled, though. Even though there are twice as many diatomic hydrogen molecules needed in the mix doesn’t mean that there is twice as much of it by mass. Recall that the molecular weight of hydrogen is much less than that of oxygen.
The reaction in Equation 4.1 is the one that occurs inside the engines of the Space Shuttle. The Space Shuttle Main Engines (SSMEs) react liquid hydrogen and liquid oxygen together to generate the thrust that drives the rocket into space. With each mole of liquid oxygen burned, 483.6 kJ of heat are produced. The SSMEs burn about 500,000 kg of O2 during launch. Using the molecular weight of diatomic oxygen, we can find the number of moles burned as
Released energy is converted to exhaust mass to generate thrust
Inert gas or liquid Energy from nuclear or electric source transfers energy to propellant Stored energy in propellant released Ignition transfers energy to propellant Fuel & oxidizer Fuel Oxidizer Ignition energy added & stored energy released in combustion chamber Propellants FigurE 4.1
# , . , , moles kg kg mole mole = 500 000 = 0 032 / 15 625 000 ss. (4.2)
Multiplying the number of moles by the heat generated per mole gives the total heat, ΔH, released to be
∆H=
(
15 625 000, , moles)(
483 6. KJ mole/)
=7 556 25, , 00 000, KJ≈7 6. TJ. (4.3) That is quite a bit of heat, indeed.The calculation above shows us that the exothermic reaction within a rocket engine releases a tremendous amount of heat energy, which in turn heats up the remaining gas products. In the case of the SSMEs, the combus- tion byproduct is water as shown in Equation 4.1. As these products (water vapor) get superheated inside the combustion chamber, they are forced out of the rear of the engine and are accelerated by a nozzle as they exit. Once they reach the exit of the nozzle at extremely high exhaust velocities, the result is a net reaction force against the rocket following the law of conserva- tion of momentum and Newton’s Laws (as discussed in Chapter 3).
In some instances, as with the SSMEs, the fuel and oxidizer need an igniter to spark the reaction. Simply mixing the propellant fluids isn’t enough to start the reaction, therefore, energy is added to the system. As the SSMEs prepare to fire, they use sparkplugs to ignite an internal “blowtorch” of hydrogen and oxygen, which blows the flame through the rest of the combustion chamber. Once the reaction is started it will continue to burn as long as there is propel- lant flow. Often, people confuse the sparks they see flying across the bottom of the SSMEs just before launch as the igniters. These sparks are used to keep any excess propellant gas from pooling in dangerous quantities underneath the engines. The spark shower keeps any propellant clouds ignited before they have time to pool.
In some engines no igniter is needed, such as in the Space Shuttle orbital maneuvering system (OMS) thrusters. Those smaller rocket engines imple- ment a single engine based on the Apollo Service Module’s Service Propulsion System. The engine uses monomethylhydrazine (MMH) for fuel and nitro- gen tetroxide (N2O4) for oxidizer. When the two propellants are mixed, they are volatile enough to spark the reaction without an external ignition source. A self-starting reaction like this is called hypergolic. The advantages of using hypergolic systems are fairly obvious. The mechanical systems are much less complex. The combustion rate of a hypergolic engine can be controlled by two flow control valves: one to control the fuel and one the oxidizer. Another advantage to hypergolic propellants is that large explosive quantities can’t gather in one place. This is because the two compounds are volatile with each other and as they come into contact they start to burn. A disadvantage of hypergolic systems is that they typically have a significantly lower Isp than nonhypergolic ones.
Once the propellants are mixed and reacted within the combustion cham- ber of the rocket, they expand and the force of the combustion is redirected out of the chamber through a nozzle. The simplest description of a rocket nozzle is that it is a component of a rocket (or an air-breathing engine like a jet) that produces thrust by the redirection and acceleration of exhaust gases. The nozzle converts the thermal energy of the chemical reactions in the com- bustion chamber (or the heated gas in a nuclear thermal engine) into kinetic energy through thermodynamic expansion by directing the kinetic energy vector along the axis of the rocket’s flight path, which is in line with the nozzle axis. Figure 4.2 shows the J-2X engine that was evolved from the ones used on the Saturn IV upper stage and will be used on the Ares-I upper stage and the RS-68, which is used on the Delta IV and also will be used as the main engines of the Ares V.