1.4 LIQUID ROCKET PROPELLANTS

The term "liquid propellant" is used to define both liquid oxidizers (liquid oxygen, liquid fluorine, nitric acid, etc.) and liquid fuels (RP-1, alcohol, liquid hydrogen, etc.). In some cases additives are used (water, ferric chloride, etc.). The propellants furnish the energy and the working substance for the rocket engines. The selection of the propellants is one of the most important steps in the design of an engine. It greatly affects overall engine system performance as well as the design criteria for each engine component. The propellant selection in turn is influenced by price, supply, handling, and storage considerations.

Monopropellants

Liquid monopropellants may be either a mixture of oxidizer and combustible matter, or a single compound which can be decomposed with attendant heat release and gasification. A rocket monopropellant must be stable in a natural or controlled environment, yet should produce hot combustion or decomposition gases when pressurized, heated, or fed through a catalyst. A liquid monopropellant engine system usually does have the advantage of simplicity of tankage, feed plumbing, flow control, and injection. Unfortunately, most of the practical monopropellants, such as hydrogen peroxide ( $\mathrm{H}_{2} \mathrm{O}_{2}$ ), have a relatively low performance. Thus, they are mainly used as secondary power sources in rocket engine systems, such as for turbopump gas generators and auxiliary power drives, and for attitude and roll control jets. Certain highperformance monopropellants, such as methyl nitrate $\left(\mathrm{CH}_{3} \mathrm{NO}_{3}\right)$, are rather unstable and are considered unsafe for rocket applications. However, some monopropellants promising relatively high-performance and safer operational characteristics have been under development recently. If successful, these may effect wider application of liquid monopropellant engines.

Bipropellants

In a liquid bipropellant system, two different propellants are used, usually an oxidizer and a fuel. Separate tanks hold oxidizer and fuel which are not mixed until they reach the combustion chamber. Present-day liquid propellant rocket engines use bipropellants almost exclusively because they offer higher performance, combined with safer operation.

The combustion of many bipropellant combinations is initiated by ignition devices such as: (a) chemical pyrotechnic igniters, (b) electric spark plugs, (c) injection of a spontaneously ignitable liquid fuel or oxidizer ("pyrophoric fluid") ahead of the propellant proper, (d) a small combustor wherein ignition is started by devices (a) or (b), in turn starting the main chamber by the hot gas produced.

Other bipropellant combinations ignite spontaneously upon mixing. Those combinations are defined as hypergolics and permit greatly simplified ignition, but pose certain hazards. For instance, accidental mixing of the fuel and oxidizer due to tank and other hardware failures could cause a violent explosion. These hazards must be considered when designing an engine system using hypergolic propellants.

Cryogenic Propellants

Some liquid propellants are liquefied gases with a very low boiling point ( $-230^{\circ} \mathrm{F}$ to $-430^{\circ} \mathrm{F}$ ) at ambient pressure and a low critical temperature ( $10^{\circ} \mathrm{F}$ to $-400^{\circ} \mathrm{F}$ ). These propellants are defined as cryogenics. The most common cryogenic propellants for rocket applications are liquid oxygen ( $\mathrm{O}_{2}$ ), liquid hydrogen ( $\mathrm{H}_{2}$ ), liquid fluorine ( $\mathrm{F}_{2}$ ), and oxygen difluoride ( $\mathrm{OF}_{2}$ ), or mixtures of some of them. Cryogenic propellants pose storage and handling problems. Elaborate insulation must be provided in order to minimize losses due to boiloff, the complexity depending on storage period and type of cryogenic. Recently, novel insulating techniques have been under development which should greatly reduce these losses. Adequate venting systems are needed for the developed gases. Storage and handling equipment and their components are extremely sensitive to atmospheric or other moisture; even minute quantities may cause a jamming of, for instance, a valve. Likewise, the design criteria, including materials selection for engine systems using cryogenic propellants, must consider the very low temperatures involved. The mechanical design of engine components for cryogenic propellant applications will be discussed in subsequent chapters.

Storable Liquid Propellants

In contrast to the cryogenic propellants, certain other liquid propellants are stable over a reasonable range of temperature and pressure, and are sufficiently nonreactive with construction materials to permit storage in closed containers for periods of a year or more. These propellants are defined as storables. Storable liquid propellants permit almost instant readiness of the rocket engine and may result in greater reliability due to the absence of extremely low temperatures and the need to dispose of boiloff vapors. Their application to military vehicles as well as to the upper stages of space vehicles has increased significantly during recent years. The mechanical design of storable liquid engine components will be further discussed in subsequent chapters.

Additives for Liquid Rocket Propellants

Sometimes, additives are mixed into liquid propellants for one of the following reasons: (a) to improve cooling characteristics; (b) to depress freezing point; (c) to reduce corrosive effects; (d) to facilitate ignition; and (e) to stabilize combustion.

Optimum Mixture Ratio

A certain ratio of oxidizer weight to fuel weight in a bipropellant combustion chamber will usually yield a maximum performance value. This is defined as the optimum mixture ratio. As a rule, the optimum mixture ratio is richer in fuel than the stoichiometric mixture ratio, at which theoretically all the fuel is completely oxidized and the flame temperature is at a maximum. This is because a gas which is slightly richer in fuel tends to have a lower molecular weight. This results in a higher overall engine systems performance. The optimum mixture ratio of some propellant combinations shifts slightly with changes in chamber pressure. Also, in actual application the mixture ratio may be shifted away from the optimum value for one of the following reasons: (a) lower chamber temperature to stay within the temperature limitations of chamber construction material; (b) required coolant flow; (c) improved combustion stability.

Density Impulse

In addition to the overall system-oriented specific impulse which we thoroughly discussed in paragraph 1-3, a quantity called "density impulse" is an important propellant performance parameter. It is the expression for the total impulse delivered per unit volume of the propellant. It is defined as:

$$\begin{equation*} \text { Density impulse }=I_{s} \cdot d(\mathrm{sec}) \tag{1-45} \end{equation*}$$

wherein

$$\begin{align*} d & =\text { bulk density or propellant combination } \\ & \text { spec. weight } \\ d & =\frac{\left(I_{w}-1\right)}{\frac{r_{w}}{d_{0}}+\frac{1}{d_{f}}} \tag{1-46}\\ r_{W} & =\text { (oxidizer/fuel) weight mixture ratio } \\ d_{0} & =\text { bulk density of the oxidizer, spec. weight } \\ d_{f} & =\text { bulk density of the fuel, spec. weight } \end{align*}$$

The Selection of Liquid Rocket Propellants

When selecting a propellant or propellant combination for a specific application, it is well to realize that most propellants, in addition to their advantages, may have certain disadvantages. Thus, propellant selection usually includes some compromises. The more important and desirable propellant features are listed below. Order of importance may vary as a function of application. (1) High energy release per unit of propellant mass, combined with low molecular weight of the combustion or decomposi- tion gases, for high specific impulse. (2) Ease of ignition. (3) Stable combustion. (4) High density or high density impulse to minimize the size and weight of propellant tanks and feed system. (5) Ability to serve as an effective coolant for the thrust chamber (optimum combination of high specific heat, high thermal conductivity and high critical temperature). (6) Reasonably low vapor pressure at $160^{\circ} \mathrm{F}$ (a frequent specification value) for low tank weight and low net positive pump suction head requirement. (7) Low freezing point (preferably less than $-65^{\circ} \mathrm{F}$ ) to facilitate engine operation at low temperature. (8) Absence of corrosive effects; compatibility with engine construction materials. (9) For storables: good storability as assisted by a high boiling point (preferably above $160^{\circ} \mathrm{F}$ ), by items $6,7,8$ and by the resistance to deterioration during storage. (10) Low viscosity (preferably less than 10 cp down to $-65^{\circ} \mathrm{F}$ ) to minimize pressure drops through feed system and injector. (11) High thermal and shock stability to minimize explosion and fire hazard. (12) Low toxicity of raw propellants, their fumes, and their combustion products. (13) Low cost. (14) Availability.

Liquid Rocket Propellant Performance and Physical Properties

Detailed methods to calculate the performance for any given liquid propellant or propellant combination can be found in the standard combustion engineering or rocket propellant textbooks. For the theoretical calculations, it is generally assumed that the ideal conditions exist as described in section 1.2 (Gas Flow Processes) of this chapter. The prime objective of propellantperformance calculations is to derive the quantities $c^{*}, C_{f}$, and $l_{s}$ through evaluation of the flame or chamber temperature $\left(T_{c}\right)_{n s}$; of the gas mean molecular weight $\pi$; and of the specific heat ratio $\gamma$ for a given $\left(p_{c}\right)_{n s}, p_{e}$ and $p_{a}$. The chamber temperature can be calculated from the heat of the chemical reaction of the propellants and from the specific heat of the gases. In practice it has been found that actual test results are usually 5 to 12 percent lower than the theoretical values obtained from calculations.

In addition to the assumption of certain idealized gas conditions, the performance equations discussed assumed and employed certain singular values for the most important gas properties: $\gamma, \mathrm{m}, R,\left(T_{c}\right)_{\mathrm{n} \mathrm{s}}$. For basic design information requiring greater accuracy, more rigorous calculations frequently employing electronic computers are usually conducted by specialists in the field. These consider that the gas properties are not necessarily constant along the path of flow. Two basic approaches can be taken: Calculations based on the assumption of unchanging or "frozen" gas composition along the nozzle axis, or based on the assumption of shifting composition. The applicable literature frequently uses the term "equilibrium" instead of "composition."

In calculations based on frozen composition, it is assumed that no further chemical reactions take place in the gases after leaving the combustion chamber and entering the nozzle, and that the combustion products at $A_{e}$ are in the same relative proportion as they were at $A_{i}$. The remaining principal variables then are pressure and temperature at the various stations. Assuming different initial sets of mixture ratios, chamber pressures, and gas compositions, a typical set of calculations, probably involving successive approximations, may be conducted to determine the optimum values of, for instance mixture ratio, chamber length, expansion area ratio, and nozzle contour, for a given propellant combination and vehicle trajectory.

Calculations based on shifting composition take into account additional variations, mainly those of gas composition, as they result from, for instance, incomplete combustion, dissociation, and reassociation. These calculations are an attempt to consider more nearly the true physical processes. Due to their extreme complexity and unpredictability, however, the results are frequently no more reliable predictions of test results than those obtained from calculations assuming frozen composition.

Thus, it is probably a matter of preference which approach should be taken. It is noted that the theoretical data based on a shifting composition usually give values several percent higher than those based on a frozen one. Therefore, in presenting performance data, the assumption of the type of composition assumed must be specified. As a rule, the thrust chamber designer will be supplied with the basic parameters by departments specializing in this field. We need not, therefore, concern ourselves further with this matter.

Performance and physical properties of numerous important liquid monopropellants and bipropellants are given in tables 1-4 through 1-10.