ROCKET PROPELLANTS R. J . T H O M P S O N ,
JR., Rocketdync,
JtVocKET
Chemists contribute to rocket program primarily through development of propel I ants. Since W o r l d W a r I!, chemists h a v e been a b l e to increase the specific impulse of fuels 2 5 % ; another 5 0 % gain is theoretically attainable with chemicals
62
C&EN
JUNE
2 3,
1958
Canoga Park, Calif.
PROPELLED
MISSILES,
satel-
lites, a n d «pace vehicles n o w under development o r ia their research phases are highly sophisticated, extremely complex mechanisms. Contributions from practically every -field of physical science, engineering, and technology are required for their successful development, fabrication, and operation. T h e chemist contributes primarily through the selection, development, a n d evaluation of t h e rocket propellant. Through
are forced into die combustion chamber by high pressure gas, and pump fed systems. Solid propellants are a sub ject of intense interest to the chemist today, and remarkable improvements in solid propellant technology are being achieved. However, within the limits of this paper I propose to discuss mainly liquid propellants employed in large, p u m p fed rockets. P u m p fed, liquid propellant rocket engines are employed in most of the large missiles (IRBM and ICBM) and as the high thrust booster stage in the satellites which have been launched to date. Flying Chemical Plant
Testing on the ground to reach the sky
his concern with the propellant, the chemist soon finds himself involved in nearly every phase of the rocket power plant development. The unique characteristic of the rocket engine among prime movers is that it is a self-sufficient system, capable of operating independent of its environ ment. In fact, optimum performance ΑΒΟΛΕ: Rocketclyne tests three rocket en gines at its field propulsion lab
is attained in a vacuum. A fuel and oxidizer carried within the system are burned to produce hot gases. Ihe thermal energy is then converted to directed kinetic energy as the gases are expelled through a nozzle to produce thrust. Rockets are generally classified ac cording to whether they use liquid or solid propellants. The liquid propellant rockets are further subdivided into pres sure fed systems, in which the liquids
From the chemist's point of view, the rocket engine may b e conceived as a high throughput, lightweight flying chemical process plant in which quanti ties up to the order of a ton per second of liquid (or solid) reactants are con verted to hot, gaseous products at a precise, automatically controlled rate. The conversion jDrocess is governed by the familiar laws of thermodynamics, reaction kinetics, and mass and heat transfer. By quite straightforward ap plication of these principles, the chem ist is able to specify fairly accurately the desired properties of a rocket propel lant for a specific application. Then he embarks on a program of experimental evaluation of known materials and syn thesis of new compounds in a continual effort to approach the ideal specifica tions more closely. The principal functions of the rocket propellant are to provide a high speci fic impulse by near-stoichiometric com bustion in the thrust chamber, to pro vide suitable turbine chive gases by off-stoichiometric (generally fuel-rich) combustion in a gas generator, a n d to cool the thrust chamber with one of the reactants. It is often also desirable, though not essential, that one of the reactants be suitable for lubricating and cooling pump bearings and gears and for use as a hydraulic actuating fluid for valves. First and foremost, the rocket pro pellant must provide a large combustion energy release, together with suitable gaseous products. The rocket de signer's measure of propellant perform ance is the specific impulse, defined as the pounds of thrust delivered by a flow of one pound per second of gas through the nozzle. The specific im pulse is directly proportional to the exit gas velocity and, hence, to the square root of the exit kinetic energy. JUNE
2 3,
1950
C&EN
63
In thermodynamic terms, specific impulse is proportional to the square root of the enthalpy change which the gas undergoes in expanding isentropically through the nozzle. Two variables, the chamber and exit pressure, are under the control of the designer. T h e pressure ratio across the nozzle should be as large as possible for optimum performance, although it may be limited by other practical design considerations. Three other variables—combustion temperature, combustion gas molecular weight, and specific heat ratio—are inherent properties of the propellant. Their optimum values for any fuel and oxidized combination can in principle be precisely calculated from basic thermodynamic data. Unfortunately, these basic data, comprising enthalpies, entropies, heat capacities, and chemical equilibrium constants of all the reactants and products are not yet always known, particularly for some of the newer fuels. This is a very active area of propellant chemistry research. Since the specific heat ratio varies over a range of only about 109< for most of the propellants of interest, the specific impulse depends primarily on the ratio of combustion gas temperature to molecular weight, T../M. This important ratio accounts for some of the unique features of rocket propellants compared to fuels for air breathing engines. In an air breathing engine, the working gas is predominantly hot nitrogen. The over-all molecular weight is only slightly affected by the nature of t h e fuel. Therefore, one tries to maximize the heating value per pound or per gallon of the fuel. In the rocket, heating value of the total propellant (not just the fuel) and the molecular weight of t h e product gas are equally important. This has several significant implications. First, the relative performances of different rocket propellants need not b e in the same order as their combustion temperatures, and frequently a r e not. Second, the heat of combustion of a compound may by itself be a poor or misleading indication of its performance potential as a rocket fuel. Third, the importance of low molecular weight effectively limits the atomic composition of rocket fuels to t h e light elements in the first two rows of the periodic table. The presence of elements heavier than aluminum (atomic weight = 27) is generally detrimental. High density in the u n b u m e d propellant is an important factor in design 64
C&EN
JUNE
2 3.
1958
optimization. (This is the reason solid propellants are effective competitors to liquids, although they generally have lower specific impulses.) Theory indicates that a high hydrogen content in t h e propellant should be desirable. This is generally borne out in practice; most useful rocket fuels are relatively rich in hydrogen. T h e specific impulses of rocket propellants in operation, development, and research today range from about 200 to 400 pounds thrust per pound per second at sea level a n d somewhat higher at higher altitudes. T h e velocity attained by a rocket-powered vehicle is directly proportional to the specific impulse, and the maximum vertical range attainable is proportional to the square of t h e specific impulse. The horizontal range is also proportional to the square of t h e specific impulse over relatively short ranges. However, because of the curvature of t h e earth, the horizontal range increases much more rapidly with increasing specific impulse a t longer ranges. For example, at a 5000 nautical mile nominal range, a \c/c increase in specific impulse increases the range by approximately 17c in a typical case. F o r another example, consider four identical, single-stage rockets of about the best design practically attainable today, loaded with propellants having spécifie impulses of 200, 300, 400, and 500 pound-seconds per pound. The first will attain a maximum range of about 12O0 miles, the second about 4000 miles, t h e tbird will become an earth satellite, and the fourth will escape from the earth entirely. These figures are based on a simplified, approximate calculation a n d should not be taken too literally, b u t they do serve to indicate t h e importance of a high specific impulse.
Propellants and Variables Let's consider typical propellants and the three propellant variables in the specific impulse equation: combustion temperature, product molecular weight, and specific heat ratio. For example, optimum performance of an oxygengasoline system is attained with a b o u t 6 5 % of the oxygen required for stoichiometric combustion to carbon dioxide and water. The decreased temperature of the fuel-rich mixture is more than offset, up to this point, by the decreased p r o d u c t molecular weight. T h e high performance of hydrogen fuel systems results largely from the low molecular weights of their h y d r o gen-rich exhausts. Fluorine-hydrazine and fluorine-ammonia propellants, on the other hand, derive their high performance primarily from their high combustion temperatures. The desired simultaneous attainment of high temperature and low molecular weight has not been attained with any of these propellants. Indeed, it does not appear that the two can be attained simultaneously, since the lightest stable combustion products from high heat release reactions, water a n d hydrogen fluoride, already have molecular weights of 18 and 20. Carbon a n d the light metals will yield still higher molecular w e i g h t products. In general, the propellant should combine a lowheat of formation, high h e a t of combustion, high percentage of hydrogen, and high combustion efficiency. In order to realize in practice its full theoretical performance potential, t h e propellant must ignite smoothly and reliably and burn stably a n d efficiently. Stability and efficiency of ignition a n d combustion are dependent on both the propellant a n d t h e combustor design. An intensive fundamental and applied combustion research p r o g r a m involving
ROBERT J. THOMPSON, JR., manager of research a t Rocketdyne division of N o r t h American Aviation, has devoted a major portion of his career to propellant chemistry and work in thermodynamics, kinetics, and combustion processes. H e has been with North American since 1954, supervising propellant chemistry. Earlier, he was engaged in rocket system design as a senior research engineer at Bendix Aviation, as chief of research at M. VV. Kellogg, a n d as a senior research associate with Allegheny Ballistics L a b . A native of California, h e received his bachelor's degree in chemistry from U.C.L.A. in 1940 and his P h . D . in physical chemistry from the University of Rochester in 1946. H e served on NACA's subcommittee on rocket engines from 1951 to 1956 and now serves on its aircraft fuels subcommittee.
Specific Impulse Determines Racket Velocity a n d Range Rocket
Oxidizer
chemists, physicists, and design a n d development engineers has been i n progress throughout the rocket industry for several years. Considerable i n sight has been gained into the fundamentals of the combustion process. A t t h e same time, semiempirical design criteria have been developed for t h e attainment of smooth, efficient combustion in chambers of minimum size and weight. A s rocket engines increase in thrust and progress t o w a r d more energetic propellants, c o n t i n u e d combustion research is required to extract and utilize substantially all the energy theoretically available in the propellant. Turbine P o w e r T h e requirements on the propellant for turbine gas generation are superficially rather similar t o those already discussed for the thrust chamber: high available energy, low molecrilar weight, good ignition, a n d efficient, stable combustion characteristics. T h e significant difference is the temperature limitation imposed by t h e design operating t e m p e r a t u r e of the uneooled turbine. This relatively low temperature (about one third to one fourth the thrust c h a m b e r gas temperature) i s attained b y far off-stoichiometric combustion, i.e., employing one reactant as a diluent and coolant. Since the temperature is fixed b y t u r b i n e design considerations, the attainment of a low molecular weight is particularly impor-
Specific Impulse at 500
tant in this case. Both theory a n d experience show that lower molecular weight is generally a c h i e v e d by fuelrich operation. For optiimnm turbine design and performance, it is desirable that t h e turbine drive gas properties b e accurately known and highly r e p r o d u c i b l e . At the lower •temperatures characteristic of the turbirae gas, diemical equilibrium is generally" not attained in the few milliseconds of reaction time available. Therefore, ^thermodynamic calculations can no longer serve a s a reliable guide. T h e gas composition and properties are determined by a combination of chemical reaction, kinetics a n d mass and heat transfer processes. T h e chemical a n d physical properties of the propellant and the physical design of the gas generator coirkbu.sk> r are inextricably combined i n controlling the heterogeneous combustion process and determining the resultant gas properties. Experimental study of fuel-rich combustion is therefore an important segment of rocket ^propellant research. Turbine gas can also be generated b y exothermic· decomposition of an energyrich compound, such as hydrogen peroxide. TTTie whole subject of liquid monoprop«ellants, a s such compounds are called» which can decompose to p r o duce hot gases for turbine drive, p r i mary thru st, or for o t h e r applications, is fascinating to the chemist and a very active research field. Since it is somew h a t peri pheral to the main subject, I
P.S.I.A.
will n o t discuss it further here, however. At least one of the reactants must be suitable for cooling the thrust chamber. This coolant must have sufficient thermal stability t o withstand exposure to a liquid side wall temperature of 1000° F., or more, for a period of about one second. In particular, the coolant should not decompose pyrolytically to produce solids which m i g h t deposit on the heat transfer surface. Any deposits which tend to reduce the heat transfer rate may raise the wall t e m p e r a t u r e to the failure point. Assuming t h e thermal stability to b e a d e q u a t e , t h e effectiveness of the coolant depends primarily on several of its physical properties, namely, the heat capacity, thermal conductivity, density, and viscosity. T h e liquid film heat transfer coefficient varies directly as about t h e 0.8 power of the density, two-thirds p o w e r of thermal conductivity, a n d one-third power of heat capacity, and inversely as about the square root of viscosity. T h e suitability of a propellant as a coolant frequently cannot be accurately predicted because of a lack of adequate data on its chemical stability and physical properties u n d e r t h e high temperature, high pressure conditions of use and uncertainty regarding t h e actual heat rejection rates which will b e encountered. Research into the heat transfer characteristics of rocket motors and propellants is another active area of interest to the chemist and chemical engineer. JUNE
2 3,
1958
C&EN
65
Need a Real Emulsion Stabilizer? Use an ORZAN® F o r a stable, easy to use, economical emulsifier, one of the O R Z A N products (surface active ligriin sulfonates) may be the answer t o your emulsifying require ments. ASPHALT FOR ROADS: ORZAN S is a n excellent stabilizer for slow-break (SS-1 and SS-lh) asphalt emulsions. Used with a resin soap emulsifier, O R Z A N s enable the emulsion to pass the ASTM cement test. ASPHALT F O R INDUSTRY: OR Z A N Ρ can be used as the primary emul sifier for emulsions of asphalt to be used in sizing of insulation board and other fibrous materials. A unique property of O R Z A N Ρ is its ability to precipitate im mediately o n the fibers when alum is added. By its dispersing action O R Z A N Ρ also makes asphalt sizing go farther. Less is required to get the desired sizing effect. W A X E S , ROSINS, O I L S : One of the O R Z A N s will do the j o b ! Let us make recommendations based upon your par ticular system. C r o w n Zellerbach Corporation C h e m i c a l Products Division C a m a s 1, W a s h i n g t o n Our emulsion consists of in
- Send literature and
samples of ORZANs Literature only Name Firm Ad d r ess
\ΜΕΕΜΠΠΜΠΜ 66
C&EN
JUNE
2 3,
1958
Numerous other chemical and physi cal properties which affect the utility and operational characteristics of a propellant can be mentioned only briefly. Suitable materials of construc tion, both metals a n d nonmetals, must be selected for compatibility with t h e often highly reactive or corrosive pro pellants. High propellant density i s obviously desirable, since it will reduce the over-all missile size, as well as t h e size, weight, and power requirement o f the p u m p . Propellant density is sec ond in importance only to specific i m pulse in determining the performance capability of a given rocket power plant. T h e relative importance of spe cific impulse and density cannot b e stated in general b u t must be assessed for each individual application. U n fortunately, simultaneous attainment o f all these desirable properties is, to s a y the least, quite unlikely. The propel lant chemist, in cooperation with t h e design engineer, must reach t h e best compromise h e can among these often conflicting requirements. H e can then apply himself both to modifying t h e propellant to improve its less desirable properties and to developing better techniques for overcoming t h e problems introduced by these properties. Finally, specifications must b e p r e p a r e d and quality control analysis a n d test procedures established to ensure precise batch-to-batch reproducibility of propellant properties in conform ance with an established standard. T h e role of the chemist in rocket development can perhaps best b e illus trated by tracing the story of t h e development of hydrazine. During World W a r II, the Germans h a d ex perimented with literally hundreds o f possible n e w rocket propellants. Only two of these were outstanding, t w o chemicals virtually unknown in this country at the time except as laboratory curiosities: high strength hydrogen peroxide and anhydrous hydrazine. T h e Germans h a d learned to make and h a n d l e high strength peroxide a n d were using it routinely to p o w e r their V-2 rocket t u r b o p u m p and V-l missile launchers. T h e technology of higli strength peroxide was quickly acquired by the American rocket a n d chemical industry, and it has been used for several years to drive the h u g e turbop u m p of the Rocketdyne-built Red stone engine, as the oxidizer in several bipropellant aircraft rocket engines, a n d in other applications. Hydrazine was a different story. I t is corrosive, poisonous, and has a nasty
habit of exploding unexpectedly and with great violence w h e n mishandled. I t has uniquely valuable properties as a high energy rocket fuel but is prac tically worthless as a fuel for any other type of engine. The Germans had not advanced very far with hydrazine. They had m a d e small scale tests which showed its potential value as a high energy rocket fuel, b u t they had not manufactured, handled, or tested it on a large scale. The explosion hazard h a d not b e e n effectively dealt with, manufacturing processes were quite un economical, and the hydrazine itself was generally impure and not consistent in quality. The V-2 rocket burned alcohol as its fuel. The potential value of hydrazine in rockets was quickly recognized and small scale experimental testing started i n the United States more than 10 years ago. The expected high per formance w a s confirmed, but in all other respects t h e results were dis couraging. Manufacture of hydrazine of the required p u r i t y proved difficult, slow, and prohibitively expensive. T h e experimentalists were p l a g u e d with problems of corrosion, toxicity, and, most seriously, frequent explosions. Then, o u r engineering effort on rockets was small, funds w e r e severely limited, we were grappling with a mul titude of problems in learning a whole new technology. T h e further compli cation was simply b e y o n d the imme diate capability of the infant American rocket industry. Hydrazine was re luctantly s e t aside b y t h e engine de signers. T r i e Rocketdyne engine for t h e Redstone missile employed liquid oxygen and alcohol. So did the rocket engines for the Navy Viking missile, t h e X - 1 , X-2, and D-558 rocket powered aircraft, and several others. Later, the p o w e r of t h e rocket engine was in creased by substituting for alcohol the special grades of kerosine originally de veloped for supersonic turbojet aircraft. Hydrazine was not forgotten, how ever. Several rocket engine manufac turers, chemical companies, universi ties, and government laboratories con tinued to study t h e hydrazine molecule. Chemical modification of the molecule improved t h e physical properties and greatly decreased the explosion hazard without seriously compromising its es sential advantages. Gradually the chemists and chemical engineers learned the secrets of economical manu facture, corrosion resistant materials, p r o p e r handling, and the whole com plex technology of living comfortably
with a highly energetic new chemical. T h e various laboratories worked in friendly competition, b u t all basic data were freely a n d fully exchanged. Today a whole family of hydrazine derivatives and related compounds are known. One of these n e w derivatives, unsymmetrical dimethyl hydrazine ( U D M H ) , is now playing a vital role in t h e American satellite program. U D M H is t h e fuel employed with nitric acid in the second stage rocket of the V a n g u a r d satellite. It is also the principal component in the b l e n d e d fuel, code n a m e d "Hydyne," used in place of alcohol in the modified Redstone engine, first stage of the Jupiter-C missile which launched the Explorer satellites. This fuel substitution, with no physical change in the engine, very substantially increased the impulse and vertical range of the vehicle. In less than six months from the initial request for a h i g h e r performing fuel, the Hydyne development at Rocketdyne was carried through theoretical calculations, laboratory tests, and research scale ïocket tests to successful completion of full scale rocket engine evaluation tests. Such a speedy development would h a v e been quite impossible without the background of sound chemical research on hydrazine and U D M H . W e may expect to hear much more about the hydrazine family of fuels in the future. T h e full potential of the chemical rocket is still far from being realized in practice. Since the end of World W a r II, t h e specific impulses of operational rockets have increased only about 2 5 % for both liquid and solid propellants. This relatively modest performance increase, together with major improvements in design, fabrication, and operating techniques, have taken us from the approximately 150 mile range of the V-2 to the intercontinental missiles and satellites of today. Theory indicates that a further increase of about 50% in the specific impulse of chemical systems is attainable over the operational systems of today. T h e cnemist will need to apply all his skill, ingenuity, and equipment in achieving the advances in propellant technology which will reduce these theoretical improvements to practice a n d thus provide the jDropulsive power required for our future exploration of outer space. BASED O N a paper presented before the 35th Annual Meeting of the American Institute of Chemists in Los Angeles, April 10, 1958.
Here's how you can MEASURE OPTICAL PATH DIFFERENCE with the AO-Baker Interference Microscope
lm
First, as shown in the photomicrograph* above, the microscope analyzer was rotated until the background was brought to extinction. Readings were taken directly from the analyzer scale. Averaged settings resulted in reading of 70.4°.
2 . Next, the analyze χ was rotated until the nucleus ortli_eccll was brought to extinction. Averag*edsettings resulted in reading of 138.2°.
3 . T h e Optical Path Difference, in degrees, is twice the difference between the t w o readings:
O P D = 2 (138.2°-70.4°) = 135.6°; or O P D = ( ^ Q O ).54(5 =.2057 Microns. Optical path difference measurements can be made t o an optimum accuracy of 1/300 wavelength. This unique ability to measure opti
