Performance Characteristics and limitations of Che ical' Propellants ARTHUR J. STOSICK Jet Propulsion laborafory, California lnsfitute of Technology, Pasadena 3, Calif.
The general characteristics and limitations o f chemical propellants are examined from two fundamental and general points o f view. The thermodynamic requirements are reviewed and the fact i s stressed that ultimate limits are nearly approached b y present systems. The second major consideration concerns the fact that combustion characteristics o f liquid propellants depend primarily on heat and mass transfer processes, rather than on chemical details. Logistic considerations are o f dominant importance in selecting liquid propellants. In the case o f solid propellants many opportunities for useful chemical approaches to improved propellants are clear.
T
HE broad interest in matters pertaining t o rocket missiles is clear from the frequent accounts in newspapers and mag% zines. Not only do we find factual n e m releases, of present and future missiles, but various manufacturers adorn their advertisements with pictures of missiles which, in some instances, they may have helped to design or build. Because of this broad interest, a symposium dealing with propellants is timely. T o provide a background for the specialized papers to follow the major performance characteristics of propellants and some of the intrinsic limitations of chemical propellants are discussed. Following the pattern of an earlier analysis the discussion is limited to propellants containing only the elements carbon, hydrogen, oxygen, and nitrogen, but the arguments are readily extended to include other elements. Horrever, the main conclusions are not greatly altered. Specific examples of computations of propellant performance have been omitted in order not t o direct attention from the underlying principles. The computation procedures used in these early calculations, as well as later improvements, have since been described by other authors both in journal articles and in textbooks. These well known publications also contain many examples of specific numerical results, both theoretical and experimental.
ments based on conservation of momentum lead to a relation betn-een the thrust, the jet velocity, and the mass flom- rate of discharged fluid. To calculate the thrust, the mass flom- rate and the jet velocity must be known. The flow rate is determined by the design parameters of the rocket system; the jet velocity is primarily determined by the propellant choice, and secondarily by operating conditions, and is readily computed by simple thermodynamic procedures, complicated numerically only by the necessity t o consider chemical equilibria in multicomponent systems. The basic thermodynamic equations are:
Reversible adiabatic flow Perfect gas with constant heat capacity q = O
AS =
sz - SI = 0
Thermodynamic Considerations A rocket derives its thrust bv conducting a chemical reaction which produces hot gases inside a chamber equipped with an escape nozzle or orifice. Thrust is produced because of the excess pressure inside the chamber as compared to the outside pressure. This excess pressure drops from a high value near the closed end to a zero value a t the exit of a welI-designed nozzle. Thus, a t the closed end of the combustion chamber there will be a pressure several hundred t o several thousand pounds per square inch in excess of the external pressure. At the end of the nozzle, the static pressure inside the nozzle matches that of the outside atmosphere. This pressure gradient in the chamber and nozzle causes the gas to accelerate to a very high velocity. If, in Figure 1, the force contributions arising from this pressure field are added, the result is the value of the thrust. At this point it becomes clear that unless gases are produced there mill be no thrust; this excludes reactions which produce no gases or only small amounts of gases. Simple physical argu722
- AH =
-(
H~ H , ) = A (kinetic energy)
u1 = 50 ft./sec.
