James R. Datler
Continental Oil Company Ponca City, Oklahoma
Rocket Propulsion: The Chemical Challenge
This discussion will focus on liquid propellant rockets, presently of great importance. These are generally bipropellant systems of relatively low density which possess higher specific impulses, I,,, than are attainable with solid propellant systems. They give short burn-out times and generally are characterized by a simpler chemistry than are solid propellants. Greater energy and vehicle flexibility are achieved, however, a t the cost of a certain amount of difficulty in handling-particularly in the cryogenic oxidants such as liquid oxygen-ozone/ oxygen mixtures or fluorine. The structural weight and the complexity of liquid propellant rockets are also greater. The One-Dimensional Fast-Burning Rocket I n this section, we shall discuss the equation of motion of a one-dimensional rocket and develop some formulas for use in comparing vehicle performances and relating them to chemical parameters. The equation of motion for rockets is generally derived under the assumption that the engine exhaust velocity, that is, the velocity of the rocket-ejected species, is constant. This gives (1) for the equation of motion
+
MJ(II M*+I). The inert masses are assumed to drop away with zero velocity with respect to each burnout stage. The equation for the velocity of an n-stage rocket utilizing the same fuel in each stage will then be
where M,, is the payload of the final (nth)stage. It may not be clear that the term
is greater than M/m, but Leitman (4) has very elegantly done the proof and has further shown that the velocity will be optimum if each stage loading-ratio in equation (4) is the same. By assuming identical fuel-loading ratios and letting the 2, 3, . . .nth stages be the payloads for the 1,2,. . .(n - 1)" stages, each working stage has a full/ burnout weight ratio of M/m. Ideal staging of this kind gives us, instead of (2), the family
where c is the effective nozzle exit velocity of burned material and u is the attainable velocity at burnout when the loaded mass is M and the "burnt-out" mass is m. The last term, gat,, is usually negligible if t,, the burn-out time, is short. The gravitational acceleration a t the surface of the earth, go, is a constant. Since the effectivevelocity c is related to the I,, according to c = I,,ga (t), equation ( 1 ) becomes
where the subscripts to the v's refer to the number of stages. Analogous to (3) and (51, the height attainable by staged rockets is the family
v = I& ln(M/m)
where, again, the subscripts refer to the number of stages. Figure 1 shows the orbital height attainable as a function of the I,, for staged rockets. All stages are assumed to be 80% fuel loaded. Equations (5) and (6) indicate that, with enough staging, u approaches infinity and the required I,, approaches zero. This does not represent a realizable vehicle, and the problem has been discussed by Wertheimer (5).
(2)
with got, assumed negligible. By solving for the kinetic energy 1/2mu2,a t burnout it has been shown (8)that the orbital height attainable for such a vehicle is given by h
=
1/~go.,[ln(M/m)121,,a
(3)
where the orbital injection energy is assumed small. It is clear that the orbital height attainable in an unstaged rocket is very strongly dependent on the I,,. I n order to discuss the attainable orbital height of staged rockets, we consider the following model. Each stage of weight, M I , has a payload which is made up of the inert mass of the ithstage, I,,and the total mass of the (i I)thstage. Equation (1) becomes vi = c log
+
Presented at the Ninth Tetrasectional Meeting, American Chemical Society, Oklahoma Division, March 16, 1963.
58 / Journal of Chemical Education
----The
cover
1.5 X 10' pounds of thrust streems from F-l rocket engine during aeries of test at NASA's high thrust test area, Edwards, California. The F-l is the most powerful engine in the United States. In the first stage of the Saturn V booster, fixF-1'swill launch the first manned exploration to the moon. Photo hy Rocketdyne, a Division of North American Aviation.
.mading Different Fuels
I n the previous section, u7edeveloped the equation of notion and predictive scheme based on "ideal" loading .atios for staged rockets with the same fuel. Here we ihall develop an idealized scheme for calculating the yelocity attained when loading different fuels. We :onsider only the velocity increment when n stages are mrned out in a constant gravitational field. For implicity, we again assume identical loading ratios; tnd as before, we assume that deadweight drops off with zero relative velocity. The equation for the velocity increment with n stages is which can be rewritten as
On the basis of a given set of boundary conditions, principally the mass ratio M / m , we can calculate the velocity after burning out n stages. For purposes of discussion, suppose we have a system with the lift-off stage loaded with RP-1, kerosene, and LOX, liquid oxygen, the second stage loaded with a gaseous fission nuclear engine, and the third loaded with a plasma-jet or ion engine of very high I,,. If we take representative specific impulses to be 300, 1000, and 1500 seconds each, the velocity increment for a three-stage vehicle is If the rocket loading to SOY0 fuel, then ( M / m ) is 5 and the velocity increment is 143,000 f t sec-'. For three RP-1/LOX stages, the velocity increment is just 46,000 ft sec-'--not quite enough for a lunar landing. I n Table 1, we have a listing of the velocity requirements for various missions. I n Figure 2, we have a plot of the velocity attainable as a function of M / m , the mass ratio of each stage, for various values of 2 Is,,(i). Table 1.
Some Velocity Requirements
Miwision
Approximate velocity, .. (ft/seo)
Low Orbit Escape Hieh Orbit. Venus Probes L U Landing ~ Solar System Escape Venus Round Trip
30,000 42,000 45,000 50;000 60,000 60,000 to 100,000
Chemistry in the Rocket Engine
I
'
i
100
I
200
I
Ya
4scmoa
I
M .
I
Ya
I
Lm
Figure 1. Orbitol height attoinoble as a function of specific impulse. The curves refer to one., two., and three-stage vehicles, all loaded 80% with Fuel.
I M l m ) . THE W S 5 R l T i O
Figure 2. The mission velocity 0 %a function of the m a s ratio M/m. The lobeied curve, refer to three-stage vehicles, all having identical mots ratios in each +age and burning different fuels.
A rocket vehicle moves by virtue of the physical reaction generated when the materials (fuel) burned in the rocket engine are ejected rearward. The rocket represents a straightforward application of Newton's Third Law of Motion. The rocket engine itself is just a device for heating or burning the fuel gases to high temperature and then directing their expansion in such a way that the mass flow is unidirectional. This simply means that the chemical energy of the fuel is converted to the kinetic energy of the reaction products. The rocket nozzle is just the artifact that directs the mass flow so that unidirectional ejection, and hence thrust, is obtained. In Figure 3, typical rocket nozzle profiles are illustrated. These are all deLaual nozzles, so called because of their eonuergent-diuergent shapes. There is an ideal design, and it is found in the conical nozzle; however use of this is precluded by length and physical design parameters. The "egg-cup" or bell-shape, which gives more radical expansion regimes, is used instead. I t is generally assumed that the chemical energy available, the enthalpy H, is converted to translational energy of the exhaust gases by means of an isentropic expansion in such a nozzle. The conservation of energy over the short internal distance inside the nozzle takes the form where T is the kinetic energy, '/%n d , of the ejected species. The enthalpy change d H is given by Volume 41, Number 2, February 1964
/
59
where V, P, ff,, and S are the thermodynamic volume, pressure, and temperature in the chamber, and the entropy. Equation (11) becomes dT
=
-VdP-ad8
(12)
and integrating from inlet to exit AT=-
S',
V ~ P a.
j-r
ds
I,, (13)
and it is clear that the kinetic energy mill be a maximum when f d S = 0, or (ATmax= AH)ds=o
(14)
This imposing of isentropic flow means that o n l ~ the translational energy of the combustion products is available for thrust. I n isentropic expansion, the internal degrees of freedom of the gases contributes nothing to the specific impulse.
Figure 3.
specific heat ratio C , / C , = 7. The value of the specific impulse is quite sensitive to several design variables. These are the maximum chamber temperature and the chamber pressure. I n addition, the pressure recovery ratio P,/P, is of extreme importance. The specific impulse is sometimes related to engine design (8)through a set of efficiencyparameters, as in =
dnx(2JH7/g~)L'9
(16)
where the q,, vt, and qr are the conversion, the combustion, and the kinetic energy efficiency of the rocket engine, and H, is the enthalpy of the combustion reaction. The specific impulse is related to the net thrust in the following way. The specific impulse is c.'go, and the thrust is given by P = ZUp~/gO + ( P . - P).l (17) where @, is the rate of propellant flow in pounds, A is the rocket-nozzle effective cross section, and PC and Poare the exit pressure and ambient pressure. If the rocket nozzle is ideal and expansion of the combusted fuel is complete, then (P, - Po)= 0, so that equation (17) becomes: F = t&,I,* (IS) and the net thrust is seen to depend on the specific impulse and rate of propellant flow. I n the case of chemical rockets in which the propellant flow rate of a particular engine design is restricted to a small range, i t is seen that net increases in the total thrust for such an engine can come about only through improvements in I,,, since from (18)
Typisol nozzle shapes.
The Nafure of fhe Specific Impulse
From the previous sections, it is clear that the specific impulse figure plays a commanding part in the attainment of useful rocket performance, particularly for long-range missions. For this reason, it is worthwhile to dwell on the nature of that parameter. The specific impulse represents the amount of thrust obtained in burning unit mass of fuel in unit time and is expressed in seconds. It is an implicit function of rocket engine design and very dependent on the rate of burning of the fuel. Improvements in rocket engine design have led to reevaluations of I,,'s over the years. I n 1958 the I,, for Hs/LOX was given as 335 seconds (6). I n 1947 it was given as 364 seconds (7), and in 1961 it was given as 376 seconds (8). A large variety of expressions are available for the specific impulse. All are related to the enthalpy of combustion. The most general expression is just
where H , and H , are the enthalpy of combustion and of the products a t exit, and J is the heat/work conversion constant. This expression ignores the chemical complexity of the combustion and energy conversion process. The I,, can be shown to be a function of the average molecular weight of the ejected species, the chamber pressure, the chamber temperature, and the 60
/
Journol of Chemical Education
Figure 4. the
Dependence of I.* on chamber pressure, P.. for H2/LOX ond goroline/LOX is shown.
The variation of
Explicitly, the specific impulse can be shown (9) to be given by
illustrating the strong dependence of I,, on the various chemical parameters. The 7 in equation (21) is just the ratio of specific heats for the exhaust gases, C,/C,. The dependence of I,, on the molecular weight, MW;
the chamber temperature, 8,; and the ratio (PJP,) is shown in Figures 4, 5, and 6. The way in which the chemistry of reaction is controlling can be seen by considering a simple fuel such as kerosene (RP-1) and liquid oxygen (LOX). Refined hydrocarbons such as RP-1 burn such that the molecular weight is near a value of 22, and the exhaust is rich in H20, CO, and COz. An attempt to increase the temperature may simply result in more dissociation with no net increase in translational energy
Figure 5. Dependence of I.. on chomber temperature b.. tion of the I,* for Hz/LOX and gamline/LOX is 3hown.
The "aria-
The Combustion Producfs
The reaction products generally spend only an extremely short time within the combustion chamber. They undergo reaction and are heated until, within the chamber, the pressure goes to very high values (-1000 psia) and the temperature rises to a high value. Outside the combustion chamber, the hot gases undergo isentropic expansion to the ambient pressure. This takes place in the nozde which is a carefully engineered divergent shape, such that, a t a certain rate of mass flow, zh, and chamber pressure, Po, the gases entering a t the chamber end will expand to ambient pressure smoothly. The requirement that the pressure be a t the ambient value when the combustion products leave the nozzle is a necessary one. If the pressure of the hot gases heyond the nozde were greater than ambient, then the subsequent expansion would he isotropic, leading to no net physical reaction for the vehicle. As it has been remarked-the rocket engine is just a device to direct the mass flow. The processes that go on inside the combustion chamber are important. Because the reaction products are generally water and its dissociation products, carbon oxides, nitrogen oxides and the like, there are certain deleterious processes that can take place. These are ionization, dissociation, and excitement of internal degrees of freedom. Toble 2. Specific Impulses for Some Propellant Combinations (chamber pressure = 500 psio) Fuel
Oxidizer
Gasoline Hydrazine Gasoline Aniline Ammonia Alcohol Gasoline Hydrazine Hydrogen Ammonia Hydraaine Hvdroeen
Hydrogen peroxide Hydrogen peroxide Nitric acid Nitric acid Nitric acid Oxygen Oxygen Oxygen Oxygen Fluorine Fluorine Fluorine
Table 3. Species
Figure 6. Dependence of I.* on overage molecular weight. tion of Lp for H9/LOX and goroline/LOX ir shown.
r
CYK)
1 . 2 0 2938 1.22 2860 1.23 3116 3088 i . 2 4 2599 1.22 3344 1.24 3460 1.25 3238 1.26 2755 1.33 4268 1.33 4666 1.33 3088
Idsec) 248 262 240 235 237 259 264 280 364 306 316 373
Energies o f Dissociotion ond Ionization
D (Kcal/mole)
I (Keal/mole)
The voria-
and, therefore, no real increase in measured chamber temperature. If through nozzle design PJP, is raised, dissociation may decrease with a net loss due to increased molecular weight values. An increase in oxidant flow may actually raise 8,but may result in more of the heavy oxygen-rich material such as CO or C02, and there may be energy losses due to vibrational excitation. The complement of characteristic parameters is surely governed by the chemistly of the situation, and improvements in performance must come as a result in improving fuels and increasing their inherent energy contents.
lonizafion and Dissociafion Processes
The temperature a t which rocket engine combustion chambers operate is quite high. I n Table 2 some operating parameters for several fuel/oxidizer combinations are listed. I n Table 3 we have listed the dissociation and ionization energies of some combustion products. Figure 7 is a graphical illustration of how dissociation can be deleterious. I n the combustion reaction, species such as HzO and COz are formed. These in turn can be broken down in high-temperature fields to form H, Volume 41, Number 2, February 1964
/ 61
H+, CO and 0. As the temperature and pressures rise, ionization phenomena such as
we rearrange equation (27) and take the antilog, we find
H=H++e-
and K , is expotentially dependent on both AH and AS. The dependence on A S is, however, controlling due to the positive nature of the entropy increase. The increasing values of K,, with $, are reflected in Figure 8, where per cent dissociation of several combustion species is plotted as a function of temperature.
OH = OH+
+ e-
(22)
and dissociation reactions such as
K,
= e -AH/RJ.~M/R
(28)
absorb energy from the available source of chemical energy. While the thermal energy rises monotonically with it,, the available chemical energy H(it,) decreases at higher temperatures due to the processes shown in (22) and (23). This can be seen more clearly if we consider, say, a water molecule a t some temperature slightly above its decompensation temperature (in that particular pressure environment). Its kinetic energy is that it has by virtue of the enthalpy of the reaction producing it, or TE,
= H(9J
(24)
If the water molecule now undergoes dissociation as in equation (2.7) the kinetic energy is given by TH THO = H ( 9 d - D H ~ O (25) where Dnaois the energy of dissociation of the water. The kinetic energy invested in the water molecule is greater before dissociation by the amount D",O. The chemical energy imparted to the hot gas has been reduced, with a loss in ultimate rocket engine performance.
+
WIY,
-
Figure 7. The chemical energy as a function of temperoturer. Curves I and 2 reprerent two readion schemer r h w e chemical energy falls off at high temperoture. The diogonol line represents the monotonically increasing thermal energy, and 9,. is the operating temperature of o rocket engine.
The available chemical energy invested in the hot gases leaving the chamber and entering the expansion nozzle is given by the enthalpy H($,). In the chamber it is approximately true that AG
= AH
- 9,AS
(26)
so that the equilibrium of equation (23) is dependent on the Gibb's free energy AG, and is governed by the expression -R3 lnK,
= AO = AH
- ffAS,
(27)
where K, is the equilibrium constant. While a is a relatively pressure and temperature insensitive function, A S for a reaction such as (23) is monotonically increasing since there is always a positive increase in entropy when dissociation occurs. If 62
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Journal of Chemical Education
Figure 8. Dissociation of typical rocket combustion species as o function of temperature.
Internal Degrees of Freedom-Equipartition
I n the high-temperature fields that exist in the rocket engine, the excitation of the internal dcgrees of freedom of polyatoms can serve as an energy sink. The excitation of vibrational and rotational modes in polyatoms and diatoms serves no useful purpose since it robs the engine system of useful heat and adds nothing to the translational energies of the ejected species. The deleterious effect that the excitement of internal degrees of freedom can cause may best be understood by recalling that diatoms and polyatoms undergo several kinds of motion, and therefore, may possess an energy corresponding to each kind of motion. The principle of equipartition assumes these energies to be additive, so that the total energy of a polyatoms is the sum of its kinetic energy, its vibrational energy, and its rotational energy. Only the kinetic energy described as always by three spatial coordinates, or degrees of freedom, can contribute to rocket thrust. Any energy absorbed into vibrational or rotational motion is separate from the translational energy, and lost. A triatom such as HsO needs nine spatial coordinates to describe it. It has nine degrees of freedom. Since it clearly has three translational degrees of freedom, it has six internal (vibration and rotation) degrees of freedom. Associated with each principal moment of inertia is a rotational degree of freedom, leaving us with three vibrational degrees of freedom. In a fully excited molecule, though, all degrees of freedom do not enjoy equal thermal weight. Each translational degree has an average energy of k 9 where k is Boltzmann's constant. Each rotational degree of freedom k$, but each vihrational has an average energy of mode has an average energy of kit. Therefore, for H20, the average energy is 6/43 per molecule, only 3/z kt9 of which is invested in translation, or useful to a rocket engine.
I n Table 4, we have the fundamental frequencies, expressed as wave numbers of rocket combustion species, shown along with the temperature that corresponds to those fundamental frequencies. These temperatures, calculated from the fundamental frequency by multiplying by 1.438 cm-degrees-the so-called second radiation constant-represent the temperature a t which the first vibrational mode, above the ground state, will become excited. We can say, in a simple-minded way, that a species a t this average temperature will absorb energy. This energy, however, is not translational, but vibrational, and is lost to the engine. Most of the temperatures in Table 4 are quite high, and a comparison with the chamber temperatures listed in Table 2 shows that most rocket engine chambers are a t lower temperatures than those corresponding to the vibrational fundamental. There will be, however, distributional effects which lead to there being more vibrationally excited species present than a linear approximation mould lead us to expect. Within each vibrational level there is a band of rotational levels which can lead to the absorption of energy for rotational promotion. Generally these are such closely spaced levels that promotion causes only a slight energy absorption. I n a vibrational level, promotion to higher rotational energy levels amounts to only a tenth to one-hundredth of a vibration excitation. Table 4.
Fundamental Frequencies and Their Temperature
It is important that we do not confuse the idea of equipartition with that of promotion from low-lying energy levels to higher ones. For example, an excited polyatom can have its vibrational, rotational and translation degrees of freedom excited, and short of dissociation is still able to absorb energy into its internal degrees of freedom. What the equipartition principle refers to is the essentially independent and additive character of the various degrees of freedom and their energies. Promotion from low-lying levels to higher excited levels refers to the energy absorbed by a species in a particular degree of freedom. The vibrational degrees of freedom do not contribute much to the heat capacity of a species until its temperature is near that temperature that corresponds to the vibration fundamental. This is usually called the vibrational characteristic temperature. Equilibrium or Frozen Flaw
At the present time most specific impulse calculations are based on one or two assumptions concerning the states of the internal degrees of freedom. These two hypotheses have to do with the vibrational excitation of the exhaust species. These hypotheses are the following:
Frozen Flow: The chemical composition of the combustion products are jixed throughout the expansion process in the nozzle. Equilibrium Flow: The chemical composition is equilibrated all along the profile of B in the nozzle. The one hypothesis-frozen flow-is tantamount to the assumption of a constant temperature throughout the length of the nozzle, while the second hypothesisequilibrium flow-is roughly the same as assuming a decreasing &profile along the nozzle. Calculations based on equilibrium flow lead to better abstraction of heat into the translational energy of the rocket ejected species. I n Table 5 below we have some 1,;s calculated from frozen and equilibrium flow compared. Table 5.
1s ;,
in Frozen Flaw and Equilibrium Flaw (P, = 1000)
Fuel
Oxidizer
Frozen
Equilibrium
RP-l Hydrasine Ammonia RP-1 Hydrasine Hydrogen Ammonia. Hydraeine Hvdroeen
Hydrogen peroxide Hydrogen peroxide Nitric acid Oxygen Oxygen Oxygen Fluorine Fluorine Fhmrine
266 27i 255 286 30 1 388 330 333
272 282 260 301 313 391 359 364
3%8
4111
The Chemical Challenge
The challenge inherent in all this is obvious. It is clear that the chemistry of the combustion process prescribes a limit to the available energies from chemical fuels. The means to improvement are just the better utilization of present fuels. At present the hydrogen/fluorine system is the best bipropellant combination, giving specific impulses of over 400 see. The chemical problems, then, can be characterized by reflecting on the energy richness of the H2/Fz system. Better energy will be available by raising the hydrogen and/or the fluorine contcnt of present propellant systems. Consider the hydrocarbon fuels such as JP-4 and RP-1. They are relatively carbon rich, leading t o molecular weight values near 20-22. Increasing the content of lighter hydrocarbons leads to better molecular weight figures, since the C/H ratio is smaller. Hydrocarbon fuels are not very susceptible to improvement by burning with fluorine as the oxidant. When the hydrocarbons are burned with fluorine, stable C-F compounds can be formed, leading to decreases in I,, due to higher molecular weights. Now, oxygen as LOX is a cryogenic fluid boiling at - 183°C. This leads to severe insulation problems in liquid propellant boosters. A RP-1/LOX booster or rocket must have the fuel insulated from the oxidant so that there are no flow problems. Here is an area of research. We need to develop combinations of good oxidizers with readily available fuels that can get around the structural weight problems posed by the need for insulation or separate tankage. Fluorine-rich oxidants may he a partial answer. Oxygen difluoride, for example, boils higher than LOX and is a more energetic oxidizer. I t suffers in that its cost (10) is so very high (-$400/ton). An improveVolume 41, Number 2, February 1964
/
63
ment in the production costs of OFz would certainly give a boost to its possible utilization in rocket fuels. Research on other Ox-Fz compounds holds high-energy promise. 03F2(If), for example, shows promise both as an igniter and exotic oxidant. The ideal rocket fuel would be a self-igniting monopropellant that is readily storable a t room temperature. At present the best chemical fuel is hydrogen (bp, -251°C), and the best chemical oxidant is fluorine (bp, - 188°C.). They are far from ideal, but they are the best available. They give molecular weights near nine, and I,, g400 secs. They are both excessively dangerous to handle and require complex rocket hardware. The means to improved use of these and other hydrogen and fluorine rich fuels are: 1. 2. 3. 4.
Increased stability Increased safety in handling Lighter insulations for vehicles Better engine materials
I n the realm of the new multiengined boosters, another problem shows itself. It is necessary that all engines light simultaneously. There certainly is room for a great deal of research in ignition reliability. I n the Saturn, for example, there are eight nozzles that must light-off a t once. The success of Saturn launches does not contraindicate ignition studies! On the whole, rocket propulsion challenges the chemist to better understand the chemical bond and the thermal properties of materials. Better materials are needed to store fuels, with long-term stability; and better insulations are needed to insure complete fuel usage. It may be possible that sophisticated additives may be put into fuels to permit safer, more stable use. Perhaps there are even methods of varying the boiling points so that compounds now incompatible would become monopropellants. While the energy available from such a system must be the same as before, the savings in rocket weight would be very great, since the inert masses (structural weight) are cut drastically. Burning of fuel takes place after injection and
breakup of the fuel and oxidant streams into droplets. The actual combustion is a sophisticated process of heat and mass transfer in a small spherical system. The nature of this process, near the point of injection and downstream of the injector, challenges the chemist to understand the problems associated with surface tension and liquid viscosities a t high pressures. If the chemist could find a means to decrease the time of droplet evaporation, the combustion efficiency could certainly be raised. Present day engines do not get all the heat into the exhausted species. If the combustion efficiency could be improved, perhaps by lowering the surface tension of the fuel droplets, then we could expect better performing engines. The challenges inherent in rocket propulsion are many; and all are interesting. They represent research in areas of organic and physical chemistry. They represent research into the basic character of matter and materials a t low temperatures, high temperatures, and a t elevated pressures. The rapidity of rocket propulsion research promises to greatly enrich our understanding of materials, fast kmetics, and combustion. I t is an exciting wide-open field for the chemist to get into. Literature Cited
(1) RUTHERFORD, D. E., "Classical Mechanics," 2nd ed., Interscience Publishers, Inc., New York, 1957, p. 82 ti. (2) BARRERE,M., JAUMOT~E, A,, DEVEUBEKE, B. F., AND VANDENKERCAOVE, J., "Rocket Propulsion," Elsevier Publishing Co., Amsterdam, 1960, p. 119. J. R., Am. J . Phys., 30,770 (1962). (3) DAFLER, G., Am. J. Phys., 26,28 (1958). (4) LEPPMAN, (5) WERTAEIMER, A,, Am. J. Phys., 25, 385 (1957). G.P., "Rocket Propulsion Elements," John Wiley (6) SUTTON, and Sons, Inc., New York, 1949. (7) TOEMEY, J. F., I d . and Eng. Chm., 49,1339 (1957). H.C., Chm. Eng. Pmg. S w p . , 57, (33), l(1960). (8) RODEAN, (9) BARRERE,M., et al., C h m . Eng. Pmg. Symp., 57 (33) 62 (1960). (10) KIT, B., AND EVERED,D . S., "Rocket Propellant Handbook," The Macmillan Go., New York, 1960, p: 97. (11) MCGEE,H.A,, JR., AND MARTIN,W. J., C~yogates,2, (5). l(1962).
1964 Dexter Award Nominees Sought The Division of History of Chemistry of the American Chenkal Society is now asking for nominees to he considered for the 1964 Dexter Award in the history of chemistry, which is administered by the division. The award consists of a plaque and a check for $1000. The 1963 Dexter Award was won hy Dr. Douglas McKie, head of the Department of History and Philosophy of Science, University College, London. The presentation of the Award wae made at ameeting of the Chemical Society in London, October 24,1963. Information, in duplicate, should be sent before March 10, 1964, to Dr. Sidney M. Edelstein, Secretary, Division of History of Chemistry, ACS, Dexter Chemical Corp., 845 Edgewater Rd., Bronx 59, N. Y.
64 / journal o f Chemical Education
