space resources RICHARD M. LAWRENCE WILLIAM H. BOWMAN'
for teachers
Ball Stab University Muncio, Indiano 47306
Chemical Rocket Propellarrts It is natural that man, in his present stage of technological development, is sending vehicles on missions into space. These vehicles must be powered with rocket engines that produce thrust by expelling propellants a t high velocities. In the case of chemical rocket engines the ejected propellants are the products of highly exothermic chemical reactions. Rocket scientists have the problem of selecting combinations of combustion chamber geometry and propellants to optimize the performance of these engines. Chemical rockets are classified as solid or liquid based on the physical state of the stored propellants. Typical solid propellants are heterogeneous mixtures of an elastomeric material and an oxidizer such as ammonium perchlorate, or homogeneous gelatinized colloidal mixtures of nitrocellulose, nitroglycerine, or diethylene glycol dinitrate. Other substances are added to these mixtures to increase the efficiency of combustion and to improve the physical properties of the propellant. The solid propellant is packed or molded inside a rocket case into the general shape of a cylinder with a hollow core. To fire the rocket, the surface of the propellant along the hollow core is brought to a sufficiently high temperature to sustain combustion following ignition. Ignition is accomplished by detonating a small pyrotechnic device near the charge or by introducing a small amount of another substance that combusts by itself or with the charge. Burning then occurs along the hollow core with the remainder of the charge acting as the combustion chamber. Some of the heat produced melts, sublimes, and boils the reactants near the surface of the charge, and the vapors react as they are swept along in the gas stream. The rate of combustion per unit of surface area of the charge is dependent on the pressure in the combustion This article is one of a series of articles based on resource units in LAWRENCE, R. M., AND BOWMAN, W. H., "Space ltesources for Teachers: Chemistry," NASA EP-87, 1971, available through the Superintendent of Documents, Government Printing Office, Washington, D. C. 20402. Present address: Laboratory of Biochemical Pharmacology, National Institute of Arthritis and Metabolic Diseases, Bethesda, Md. 20014. Some liquid propellant rockets derive their thrust from the thermal decomposition uf a single substance. The herdware design of these liquid monopropellant rockets is essentially the same as for one of the propellants of the hipropellant type.
chamber and on the propellant temperature. The pressure in the combustion chamber in turn is related to the amount of material burning per unit time and to the size of the exhaust opening or throat. The surface temperature of the charge is determined by its specific heat and thermal conductivity. The rate of penetration of heat into the charge is regulated by additives, and charges typically are consumed at the rate of 0.1 to 5 cm/sec in the absence of extreme erosive effects. Solid-propellant rockets have the advantages of simplicity and reliability along with the disadvantage that the solid propellant may decompose or react prior to launching. A solid-propellant rocket is stored a t a sufficiently low temperature to retard premature reaction or decomposition of its charge. Homogeneous propellants are also very sensitive to shock. Solidpropellant rockets have the further disadvantage that once the combustion process is initiated, it cannot be throttled easily but continues until all of the charge is consumed. Liquid propellants play a prominent role in our space efforts. Rocket engines using liquid propellants have the advantage that they can be throttled because the propellant storage tanks and combustion chamber are separated. The complex systems required for the storage and movement of liquid propellants, however, impose penalties of decreased reliability and increased mass. Furthermore, many liquid propellants must be handled and stored a t very low temperatures. Liquid nitrogen tetroxide and aerosine, a liquid mixture of hydrazine and unsymmetrical dimethyl hydrazine, were the propellants used in the Titan I1 boosters to launch the Gemini series of manned satellites. Liquid hydrogen and liquid oxygen are used to power the 200,000-lb thrnst 5-2 engines in the second and third stages of the Saturn V-Apollo. For combustion in a liquid-propellant engine the liquid fuel and oxidant are pumped or forced to the combustion chamber (Fierure Commonly the liquid fuel is c i r c u ~ a t e d ' t h ~ o uthe ~ h tubular walls of chamber and nozEle thus warming the the fuel to the proper temperature for ignition and protecting these structures of the engine from the effects of high temperatures. The liquid fuel and oxidant are metered, atomized, and mixed upon injection into the combustion Some propellants such as NIOa and NzH4 bite On mixing ( h ~ ~ e r g O l i c ) , while other propellants such as Hz and 0%must be Volume 48, Number 5, Moy 1971
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PRORLUNT
COMBUSTION
NOZZlE THROAT
Figure 1. Componenh of a typical liquid bipmpollont rocket. Fuel and oxidant are pumped from their storage tank. to tho combustion chamber. Circulation of the fuel through the walls of tho combustion chamber and nozzlecoolsthere drudurer.
ignited initially by a glow plug, a spark plug, a pyrotechnic igniter, or the injection of a small amount of hypergolic fuel. I n the combustion chamber the temperature of the atomized droplets of propellants rises rapidly. Vaporization occurs a t their surfaces and the droplets typically exist for only several microseconds as they are accelerated to the speed of the surrounding gas stream. Reaction of the gaseous fuel and oxidant occurs sufficiently downstream from the fuel injectors to prevent convective heating and subsequent erosion of the fuel injector plates. The combustion chamber must be of sufficient size to essentially complete the processes of atomizing, mixing, igniting, and burning the propellants by the time the gases enter the exhaust nozzle. The area of the nozzle throat determines the pressure in the combustion chamber and consequently the speed of the reaction and the required length of the combustion chamber. These and other parameters have no optimum values but are selected for a given set of conditions including the nature of the propellants, external pressure, and the thrust expected from the rocket engine. To gain further insight into some of the criteria used in the selection of chemical rocket propellants it is beneficial to first consider some quantitative aspects of rocket propulsion. The purpose of a rocket engine is to provide the accelerating force or thrust needed to place a space vehicle into a planetary orbit or an interplanetary trajectory. The acceleration produced is directly proportional to the thrust of the engine which in turn is directly proportional to the amount of propellant ejected per unit time and to the directed velocity of the ejected propellant. Large acceleration is achieved by expelling large amounts of propellant a t high speeds from a space vehicle having the smallest possible mass. While rocket engines are rated in terms of thrust, propellants are rated in terms of specific impulse. Specific impulse, I,,, is the ratio of the thrust produced to the mass of propellant expelled per unit time and is expressed in the units lb/(lb/sec) or lb-se~/lb.~The importance of this quantity is that it is directly proportional to the velocity of the ejected propellants and to the acceleration produced by the rocket engine. Because the directed velocity of the ejected propellant is directly proportional to the random velocities of the gases in the combustion chamber which in turn are direotly proportional to the square root of the ratio of the temperature in the combustion chamber, T,,to the average particle mass, m, it follows that (1)
336 / Journal of Chemical Education
Furthermore, because the temperature in the combustion chamber is nearly proportional to the heat of reaction, AH, of the propellants, we have (8)
Thus propellants are favored that undergo highly exothermic reactions and yield products of small molecular mass under the reaction conditions. To illustrate this relationship, consider the reaction of some potential elemental rocket fuels with oxygen. The heat released per unit mass of oxide formed is plotted as a function of the atomic number of the elemental fuel in Figure 2. The units of the ordinate
",
L 0
I
20 I\lOMlC NUMBER
10
I
30
Figure 2. Reaction energy of elements with oxygen. The heat released per pound of oxide formed is seen to vary periodicdly with atomic number. Maxima occur for the elements beryllium, aluminum, and rsondium.
reflect the importance of minimizing the mass of the propellant in maximizing the acceleration of a rocket. As can be noted in the figure, the most exothermic reactions are those involving the elements hydrogen, lithium, beryllium, boron, magnesium, aluminum, silicon, and scandium. The oxides of these elements other than hydrogen have high boiling or sublimation points, however, and would tend to aggregate in the gas flow of a rocket engine and even condense on the combustion chamber walls. Use of these elemental fuels would give highly exothermic reactions but would yield low values of specific impulse because of the large effective mass of the reaction products (eqn. (2)). In practice, oxygen and fluorine are used as oxidants, and fuels typically include elemental hydrogen and compounds containing boron, carbon, nitrogen, and hydrogen. The resulting combustion temperatures are in the range of 4500'-700O0F. The average mass of the combustion products depends on the nature of the fuel-oxidant pair, the reaction conditions, and the thermal stabilities of the products. Fuel-oxidant combinations are favored that at equilibrium yield gaseous products of small molecular mass and appropriately low thermal stabilities. Low thermal stability of the equilibrium products reduces the heat released in the combustion chamber but may result in an overall increase in the value of specific impulse because of a greater percentage reduction of the average molecular mass in the rocket exhaust. I n addition, the average molecular mass of the ejected propellant often can be reduced by injecting excess fuel into the combustion chamber. Although less heat In the literature the mass of rocket propellants commonly is given in ~ounds. It follows that the units of specific impulse are conventional pounds force divided by the mass flow rate (Ib/sec).
380-
-
O,-h
Figure 3. Performance of represent.. tiye rocket propellants 141.
360- F,-H,
2
% $
340-
OiH,
320 -
is released under fuel-rich conditions, the reduction of the 300average mass of the exhaust 3 280 - F,-B,Y products is even greater. For V instance, for use in rocketry 260 - 0,-N,H,lCHJ2 a typical weight ratio of $ oxygen to hydrogen is 4:l. 240 - NP4-N9Hd HNO,.NQ-N,H,(CH,), Under these fuel-rich condi220tions the hydrogen-oxygen reaction a t high temperature can be written approximately as (3) ",
$-ii
Similar equations apply to other fuel-oxidant combimations. Assuming the composition of the propellants ejected from the nozzle of the rocket engine is the same as that occurring in the combustion chamber, sccalled frozen equilibrium, the values of specific impulse
of typical chemical propellants range from 200 to 450 lb-sec/lb (Fig. 3). Although much information about propellant performance can be determined through combustion experiments, such trials are very costly and time consuming. A more efficient and revealing method of predicting the performance of propellants involves the use of statistical thermodynamic methods. Data have been compiled for more than 20 elements and 200 combinations of the elements from which quantities such as specific impulse can be calculated. Literature Cited RATTIN,E. J., in "Astronautics for Saience Teachera" (Editor: MBIW NEB. J. C.)John Wiley 6r Sons, In=.,New York, 1965, p. 258. V m n a o ~FR&NzH.. ~, J. CXBM, EDUO., 46.140 (1968). Bunnows. M. C.. "Exploring in Aerospace Rocketry," NASA Lewis Research Center, 1967, Chapter 4. ROTHAOCX. A. M., in "Current Resesrch in Astronsutical Soienoes" (Editor: B R O O LL. ~), Pergamon Press. New York. 1961, Vol. 6 , pp. 914310
(5) "Propulsion for Deepspace."NASA EP-41.1969.29 pp. (6) Humen, M. W.."Thruat into Space," Holt. Rinehsrt, and Winston. Ine.,N e v York. 1966,224 pp. (7) B ~ o x * w ,R. 8.,F. E. B e ~ m s B. . J. CLARK.AND F. J. ZEGEZNIK in, "Conferenoe on Selected Technology for the Petroleum Industry." NASASP-5053, 1966. pp. 5-23.
Volume 48, Number 5, May 1971
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