PROPELLANT CHOICE

Liquid rocket engines can burn a variety of oxidizer - fuel combinations, some of which are tabulated in Table I. Most of the propellant combinations listed are dangerous, toxic, and expensive. The amateur builder of rocket engines on the other hand, requires propellants that are readily available, reasonably safe and easy to handle, and inexpensive. Based on experience, ROCKETLAB recommends the use of gaseous oxygen as the oxidizer and a hydrocarbon liquid as the fuel. They give good performance, the combustion flame is readily visible, and their high combustion temperature presents an adequate design challenge to the amateur builder. The propellants are used in the Atlas missile and the Saturn space booster. In these systems, however, liquid rather than gaseous oxygen is used as the oxidizer.

TABLE I

Gaseous oxygen can be readily and inexpensively obtained in pressurized cylinders in almost any community because of its use in oxy-acetylene welding. With reasonable precautions, to be detailed later, the gas (and cylinder) is safe to handle for rocket test stand use. Gas pressures are easily regulated with commercial regulators and gas flow rate is easily controlled with commercially available valves.

Hydrocarbon fuels, such as gasoline and alcohol, are readily available in any community. Safety precautions are already known by most responsible individuals due to wide use of the fuels in internal combustion engines for automobiles and power equipment.

All subsequent sections of this publication will refer to, and assume, that the propellants to he used in amateur liquid-fuel rocket engines are gaseous oxygen and hydrocarbon fuel.

The flame temperature of hydrocarbon fuels burned in gaseous oxygen at various combustion chamber pressures is shown in Figure 3 for the stoichiometric mixture ratio. Mixture ratio is defined as the weight flow of oxidizer divided by the weight flow of fuel, or

where
Wo = lb of oxygen/sec
Wf = lb of fuel/sec

When a stoichiometric ratio is achieved just enough oxygen is present to chemically react with all the fuel; the highest flame temperature is achieved under these conditions. If a lower flame temperature is desired it is usually better to have more fuel present than oxidizer; this is known as burning "off-ratio" or "fuel-rich." This condition is less severe on the rocket engine than burning, at stoichiometric or oxygen-rich conditions.

Figure 4 indicates how the flame temperature varies when combustion chamber pressure is held at a constant value and the mixture ratio is allowed to vary.

Figure 3 Flame temperature versus chamber pressure at stoichiometric mixture ratio.

Figure 4 Flame temperature versus mixture ratio at constant chamber pressure.

The thrust developed per pound of total propellant burned per second is known as specific impulse and is defined as

Figure 5 indicates the maximum performance possible from hydrocarbon fuels burned with gaseous oxygen at various chamber pressures with the gas expanded to atmospheric pressure. This graph can be used to determine the propellant flow rate required to produce a certain thrust. Suppose you wish to design a rocket engine using gaseous oxygen/gasoline propellants to be burned at a chamber pressure of 200 psi with a thrust of 100 lbs. At these conditions the propellant performance, from Figure 5, is 244 lb of thrust per lb of propellant burned per second. Therefore

Figure 5 I_sp performance of hydrocarbon fuels with gaseous oxygen.
Since the maximum Isp mixture ratio (r) for oxygen/gasoline is 2.5, we have:

PROPELLANT PROPERTIES