2818
INDUSTRIAL AND ENGINEERING CHEMISTRY
The data obtained indicate that two or more competing mechanisms of ignition are present when the fuels investigated ignite at atmospheric pressure. ACKNOWLEDGMENT
The authors wish to express gratitude to C. J. Vogt for helpful criticism, valuable advice, and sustained interest in this investigation. LITERATURE CITED
Standards On and Lubricants,” D 286-30, pp. 49-50, Philadelphia, American Society for Testing Materials, 1942. (2) Ibid., D 613-41T, PP. 190-7. (l)
Vol. 43, No. 12
(3) Boerlage, 0. D., and Broeee, J. J., “The Science of Petroleum,” Vol. IV, pp. 2894-2909, London, Oxford University Press, 1938. (4) Bogen, J. S., and Wilson, G. C., U.O.P. Booklet No. 260, Chicago, Universal Oil Products Co., 1944. ( 5 ) Boodberg, A. A., thesis, University of California, 1940. (6) Jost, W., “Explosion and Combustion Processes in Gases,’’ p. 436, New York, McGraw-Hill Book Co., 1946, translated by H. 0. Croft. (7) Marder. M., “Motorkraftstoffe,”Erster Band, pp. 347-9, Berlin, Springer Verlsg, 1942 (lithoprinted by Edwards Brothers. Inc., 1945.) (8) Peden, R. C., thesis, University of California, 1938. (9) Starkman, E., Trans. Am. Inst. Chem. Engrs., 42, KO.1, 107-20 (1946). RECEIVED October 30, 1950.
ADIABATIC FLAME TEMPERATURE IN JET M WILLIAM S. MCEWAN AND SOL SKOLNIK’ Chemistry Division,
U.S. N a v a l
O r d n a n c e Test Station, China Lake, Calif.
T h e adiabatic flame temperature is of interest principally in obtaining the specific impulse of a propellant. In order bo obtain values of specific impulse, a knowledge of the average molecular weights and ratio of the specific heats of the products of combustion also is required. Each of these factors involves the computation of the relative composition of the products of combustion; hence, any discussion of flame temperature or impulse calculation principally amounts to a consideration of nieans and methods of obtaining the products of combustion. Several systematic procedures for performing this calculation are briefly reviewed and evaluated. The effects of flame temperature of propellants with different enthalpies are given. These calculations indicate that propellants with low enthalpies (about -875 calories per gram) yield flame temperatures which are essentially independent of pressure, while the flame temperatures of propellants with en thalpies of the order of -640 calories per gram are pressuredependent.
PRECISE knowledge of the adiabatic flame temperature of a rocket propellant actually is of little practical importance by itself. However, it is an important parameter in the calculation of the theoretical performance of a propellant when combined with the effective molecular weight and the ratio of specific heats of the products of combustion. Since the desired effect from a rocket propellant is the production of impulse, this value of impulse serves as a convenient means of rating different propellant compositions. This parameter is defined as the force per unit weight of propellant integrated over the burning time
A
I s p = F / w f dt
(1)
where I,, is specific impulse, F is force, and w is weight of propellant. The derivation of I s , in terms of thermodynamic quantities is based on elementary principles of thermo- and hydrodynamic laws and the following assumptions: 1. The combustion and expansion processes are adiabatic 2. The products of combustion are in thermal equilibrium P
Present address, U. 5. Naval Powder Faotory, Indian Head, Md.
3. The flow through the nozzle is streamlined flow 4. The velocity of the gases in the combustion chamber relative to the rocket motor is zero
If force is defined simply by
F = -mdu/dt
(2)
where m is mass, v is velocity, and t is time, after integrating between limits ( t = 0 when v = 0) and dividing by the gravitational constant, g,
Zap = - v / g
=
Ft/gm
(3) is obtained. On the basis of the assumption that the velocity of gases in the combustion chamber relative to the rocket motor is zero, the sum of the kinetic energy and enthalpy of a unit mass of fluid is constant. There is thus obtained for a unit mass of gas He f ’ / z d = He f ’/zv:
(4)
where H is the enthalpy, 2. is the velocity, and subscripts c and e indicate the state in the chamber and after expansion, respectively. Since vc is assumed to be zero, HP
-He
=: 1/22):
(5)
and Since
where T is the combustion chamber temperature and T,is the exhaust temperature, the flame temperature assumes a practical importance. The exhaust temperature can be expressed readily in terms of flame temperature, T,, chamber pressure, P,, exhaust pressure, Pc,and ratio of specific heats of combustion products r-1
T, = T,(P,/P,) 7-
(8)
by making the assumptions that: the products of combustion behave like perfect gases; y is independent of temperature; and no changes in composition occur during the expansion (frozen
I N D U S T R I A L AND E N G I N E E R I N G CHEMISTRY
~
~
Figure 1.
l
~
HVORODEN EQUILIBRIA CIRCUITS
OXVOLN EQUILIBRIA CIRCUITS
Wheatstone Bridge Analog of the Mass Action
Law
equilibrium). If Equations 7 and 8 are substituted in Equation 6, the following expression for specific impulse is obtained
Is,
9.324
li
[
1"'
T02yJm(y - 1) 1 - (Pe/Po)
(9)
COMPUTATION OF PRODUCTS OF COMBUSTION
The calculation of the flame temperature is made by equating the enthalpy of the propellant composition to the enthalpy of the products of combustion
where P , is the partial pressure of the ith component. Therefore, the knowledge of the composition of the gaseous products of combustion is of primary importance in the computation of both the adiabatic flame temperature and the specific impulse. Any general treatment of the calculation of either of these parameters really requires a discussion of the procedures for obtaining the composition in terms of partial pressures or mole fractions of the gaseous species assumed to be present in appreciable quantities. The species present, of course, depend on the temperature and pressure as well as the elements composing the unburned material. However. for the carbon, hydrogen, oxygen, and nitrogen (CHON) system, generally ten species are assumed to be present in significant concentrations. These species are carbon monoxide, carbon dioxide, oxygen, nitric oxide, nitrogen, water, hydrogen, atomic hydrogen, atomic oxygen, and hydroxyl radicals. The relative amounts of these species are governed by the total pressure, the conservation of mass requirements, and the mass-action relationships. The solution of this general problem, since it involves a number of equations, several being nonlinear, equal to the number of molecular species present, often becomes awkward and time consuming. At the present time there are available almost as many systems for the solution of these equations as there are individuals who have done work in this field. It is apparent that some systematic method of procedure must be developed or obtained if one insists on working in this field. The authors have developed two methods of their own. An attempt, therefore, will be made to give and somewhat assess the relative merits of a number of representative methods which have appeared in the literature. CALCULATION OF PRODUCTS OF COMBUSTION Edse ( 9 ) was one of the first to publish an organized stepwise procedure for calculating the composition of the products of combustion of the generalized carbon-hydrogen-oxygen-nitrogen propellant. His method involved the selection .of a reasonable value for the ratio of carbon dioxide to carbon monoxide from a
2819
graph based on the oxygen balance and total carbon. This ratio was introduced into the prescribed procedure and the partial pressures of the 10 components were computed. From these partial pressures the atom ratios were calculated and these values were compared with the known stoichiometric atom ratios. After repeating this procedure twice, the value of the carbon dioxide-carbon monoxide ratio then could be corrected by interpolation in a direction to bring the computed atom ratios, usually the oxygen to hydrogen ratio, into coincidence with the known stoichiometric atom ratio. The difficulties with this procedure were that it involved a recourse to graphs in order to procure the entering values of the carbon dioxidecarbon monoxide ratio, and that it was necessary to compute all 10 partial pressures in order to sum up the computed atom l l l ~ ratios after each iteration. At the U. S. Naval Ordnance Test Station the necessity arose of computing large numbers of these problems for the preparation of Molier charts for a variety of liquid and solid propellant compositions. Since Edse's method could not be readily adapted to automatic mechanical computers such as those of the International Business Machine Corp., this method was replaced by one (6) in which all the constants appearing in the equations were separated from the variables and classified according t o temperaLure dependence, mixture dependence, and temperature-mixture dependence. These constants then could be calculated once and tabulated. The atom ratios could be calculated directly as a function of only the carbon dioxide-carbon monoxide ratio and the constants without the necessity of first computing all 10 partial pressures. Starting values for this ratio could be obtained by the machine from the stoichiometric ratios and two sets of arbitrary constants dependent on temper1' ' '?H2 ature and pressure. Adjustments between iterations were then made by linear interpolation from the carbon dioxide-carbon monoxide ratio and the known and computed atomic ratios. The partial pressures were computed only after the final iteration, Since this system required no graphs or judgment by the operator, it proved to be readily adaptable to IBM calculating equipment. Brinkley of the U. S. Bureau of Mines ( I ) has contributed greatly to the solution of the general problem by showing the nature of the governing equations for any complex system and by showing how to select a consistent set of chemical equilibria with which to operate. The latter i s accomplished essentially by providing rules for dividing the molecular species into two groups consisting Figure 2. Examples OF of independent and derived comM o d i F i c e t i o n OF Wheatstone Bridge to ponents. In his method explicit Produce Square Roots expressions for the derived comor Squares ponents are obtained in terms of the independent components and D o t t e d lines indicate mechanical coupling so as to the mass-action constants. T h e equate rerlrtanser
2820
INDUSTRIAL AND ENGINEERING CHEMISTRY
explicit expressions for the independent components are obtained in terms of the derived components and the conservation of mass relationships. Solution is obtained by substituting values for the independent components, calculated by assuming the partial pressures of the derived components are zero, into the mass-action relationships for the derived components. These values then are substituted into the equation for the conservation of mass and new values for the independent components are obtained. This procedure is repeated until the differences in composition caused by suecessive iterations become sufficiently small.
CARBON
VOl. 43, No. 12
functions. This method is claimed to converge extremely rapidly. Hottel, Williams, and Satterfield of the Massachusetts Institute of Technology (3)have attempted to make the task of others easier by providing a number of generalized thermodynamic charts prepared for a considerable number of carbonhydrogen-oxygen-nitrogen mixtures a t 300 and 14.7 pounds per square inch absolute. Directions are given for obtaining values of thermodynamic functions a t pressures other than the standard ones. Specific impulses calculated by means of these charts are claimed to be nithin 2% of values computed by other procedures. Although the outlined methods are more or less systematic and are usable by individuals who have only little knowledge of chemistry and mathematics, these methods are time consuming and laborious. These methods are especially irkSERIES SET some vhen the requirement is for the ca~culation of the performance characteristics of a few odd propellant compositions. This is not an infrequent requirement at most activities engaged in propellant research and development. In order to minimize the amount of tedious effort involved in these calculations, hIcI3wan and Skolnik ( 6 ) developed an analog computer for the determination of the products of combustion of
--CARBON
SERIES SET
OXYGEN CARBON SERIES SET
Figure 3. Series Summation of Resistances to Give Analog of Atom Sum and Total Pressure
0-c
Double lines indicate mechanical coupling
The Bureau of Nines has used this system to compile by means of the Electronic Super-Computer, located at National Ballistics Laboratory, Aberdeen, Md.> the compositions of an enormous number of problems covering a wide range of temperature, pressure, and compositions. The results of these calculations are not tabulated but they are kept in the form of IBM cards. Inquiries about a specific problem may be addressed to the Bureau of Mines, Pittsburgh, who then will attempt to provide the interpolated values from the data on the cards. Just how long this takes, the cost, or the backlog of work is not known. Kreiger and White of the Rand Corp. ( 4 ) have published a general method similar in concept to that of Brinkley's, but differing from it to the extent that the expressions for the minor constituents have been linearized in terms of the free energy change for the dissociation reactions which define the minor components and the logarithms of the major components and the chamber pressure. The iteration procedure for this system is said to be slow to converge when the oxidizer-fuel ratio is c ose to stoichiometric. Sachsel, Mantis, and Bell of Battelle bIemorial Institute ('7) have published a method, specifically for the oxidation of aluminum borohydride but generally applicable. They used the equilibrium constants in logarithmic form to w i t e expressions for the four atomic summations in terms of four of the 14 variables. Two arbitrary trial values were given to two of the four variables; the four permutations of these were substituted into two of the four atomic summation expressions. The resulting equations were solved for values of the other two variables. These values then were used to solve the remaining two atomic summation functions. Finally a graphical interpolation is made betmTeen the four initial values and the resulting values of the two atomic summation functions, so as to obtain zero for the two
-CARBON
SERIES SET
0 - C SERIES SET SHORTED
H-C
ROGEN SERIES SET -CAREON
020;AND
SERIES SET
0 SHORTED OUT
NITROGEN SERIES SET
N-C
Figure 4.
Analog for Conservation of Mass Relationships h are as in empirical formula OarCpN-,Hh
a, 8 , Y, and
carbon-hydrogen-oxygen-nitrogen systems. The basic principles utilized in the design of this computer are thase of interacting wheatstone bridges. Because all the reactions encountered in this system can be represented by equilibria involving three species, the six independent mass-action equations can be arranged as six wheatstone bridges shown in Figure 1. Since some of the components appear as one-half powers and others are squared, bridge circuits were devised to perform squaring and rooting operations. Examples of these circuits are shown in Figure 2. The molecular species are represented as series resistances whose sum is proportional to the total pressure of the system, Figure 3.
\
INDUSTRIAL AND ENGINEERING CHEMISTRY
December 1951 Table
1. Gas Composition at Different Temperatures and Pressures for Three Propellants Solid' Propellanta
Pressure, Atmospheres 1
Composition, Mole % ' 08 0 H NO N¶ OH H20 H, 2750 0.3290 0.1718 0.0042 0.0031 0.0178 0.0021 0.1375 0.0177 0.2491 0.0678 3000 0.3448 0.1304 0.0163 0.0153 0.0422 0.0056 0.1287 0.0405 0.2048 0.0714 3250 0.3554 0.0786 0.0296 0.0452 0.0892 0,0094 0.1152 0,0632 0.1367 0.0772 lo 2750 0.3258 0.1847 0.0006 0.0003 0.0056 0.0007 0.1409 0.0060 0.2682 0.0672 3000 0.3321 0.1712 0.0030 0.0021 0.0128 0.0025 0.1380 0.0167 0.2561 0:0655 3250 0.3440 0.1429 0.0104 0.0085 0.0265 0.0059 0.1318 0.0353 0.2265 0.0681 20 2750 0.3256 0.1861 0.0003 0.0002 0.0040 0.0005 0.1414 0.0043 0.2706 0.0672 3000 0.3306 0.1763 0.0016 0.0011 0.0090 0.0018 0,1394 0.0122 0.2631 0.0651 3250 0.3387 0.1559 0.0064 0.0047 0.0185 0,0047 0.1348 0.0271 0.2424 0.0660 40 2750 0.3255 0.1870 0.0001 0.0001 0.0028 0.0004 0.1417 0.0031 0.2723 0.0673 3000 0.3295 0.1797 0.0008 0.0006 0.0064 0,0013 0.1403 0.0088 0.2680 0.0648 3250 0.3362 0.1653 0.0036 0.0025 0.0129 0,0036 0.1370 0.0203 0.2540 0.0646 Liquid Propellanta 1 2500 0.0234 0.0667 0.0164 0.0020 0.0043 0.0057 0.02469 0.0164 0.5886 0.0298 2750 0.0396 0.0460 0,0236 0.0069 0.0109 0.0065 0.2334 0.0366 0.5263 0.0626 10 2500 0.0190 0.0723 0.0025 0.0003 0.0011 0.0000 0.2617 0.0067 0.6185 0.0229 2750 0.0252 0,0652 0,0109 0,0015 0,0041 0,0046 0.2476 0.0167 0.5900 0.0332 3000 0.0376 0,0493 0,0204 0,0054 0.0121 0,0085 0.2379, 0.0392 0.5308 0.0562 20 2500 0.0134 0.0789 0.0031 0.0002 0.0007 0.0012 0.2575 0.0060 0.6229 0.0159 2750 0.0198 0.0722 0.0105 0,0011 0.0023 0.0046 0.2516 0.0157 0.5984 0.0226 3000 0.0330 0.0568 0.0167 0.0034 0.0072 0.0078 0.2438 0.0328 0.5564 0.0409 40 2500 0.0114 0,0809 0,0043 0,0002 0.0006 0.0011 0.2562 0.0041 0.6261 0.0156 2750 0.0183 0.0728 0.0087 0,0007 0.0016 0.0038 0.2522 0.0116 0.6084 0.0233 3000 0.0300 0.0603 0.0138 0.0022 0.0071 0.0000 0.2420 0.0250 0.5665 0.0400 Liquid Propellantc 1 2500 0.1647 0,1407 0,0009 0,0006 0.0070 0.0006 0.1625 0.0076 0.4374 0.0779 2750 0.1715 0.1234 0.0164 0,0043 0.0197 0.0034 0.1552 0.0254 0.4001 0.0808 lo 2500 0.1608 0 1493 0.0000 0 0006 0.0015 0.0000 0.1602 0.0024 0.4482 0 0738 2750 0 1643 0'1445 0 0016 0'0006 0.0059 0.0013 0.1595 0.0098 0.4419 0:0704 3000 0:1649 0:1307 0:0072 0:0031 0.0135 0.0041 0.1576 0.0269 0.4230 0.0722 2o 2500 0.1603 0.1514 0.0002 0 0000 0.0014 0 0012 0.1617 0.0199 0.4575 0.0622 2750 0.1706 0.1399 0.0007 0:0002 0.0039 0:0012 0.1611 0.0072 0.4556 0.0620 3000 0.1698 0.1325 0,0038 0.0016 0.0098 0.0028 0.1562 0.0186 0.4294 0.0742 40 2500 0.1465 0.1575 0.0001 0,0000 0.0009 0.0001 0.1646 0.0016 0.4693 0.0635 2700 0.1594 0,1409 0.0004 0.0001 0.0027 0.0006 0.1619 0.0047 0.4592 0.0635 3000 0.1567 0.1455 0,0026 0.0009 0.0063 0.0025 0.1623 0.0142 0.4464 0.0624 a JPN composed (8) of 52.2% nitrocellulose (12.25% N),43% nitroglycerine, 3.0% diethylphthalate, 0.6% diphenylakine, 1.25% potassium nitrate a n d 0 1% nigrosine dye (added). Enthalpy = -368. b Composed (by weight) of 13.00% methanol, 20.15% hydrazine, 3.71% water, and 62.85% nitric acid. Enthalpy = -876. -644. c Composed (by weight) of 19.075% octane and 80.925% nitric acid. Enthalpy
Tfmz.,
cox
co
-
The mole fraction, therefore, is equal to the resistance of a given component divided by the total resistance. The three conservation of mass equations also are represented by wheatstone bridge circuits; two arms of each of these are the stoichiometrical ratios calculated from the empirical formula. These circuits are shown in Figure 4. The potentiometers representing massaction constants are calibrated in terms of temperature intervals. The proper scale is selected by switching the correct ked-resistor into the rooting circuits. This arrangement allows the utilization of the maximum resistance which in effect gives a precision that is independent of the pressure involved. With the proper switching arrangement the iterative process is completely mechanical and even can be made automatic. The details of construction and operations of this computer are described in reference (6). Table I gives the composition of product gases a t different pressures for three typical propellant compositions. F L A M E TEMPERATURES
40
In
W
30
a W
U. u)
0
F
4
E W
2 20 W In
After the composition of the combustion products has been determined by any of the several methods outlined, the calculation of the flame temperature is a relatively simple operation. The value for Q of Equation 10 is obtained by adding to the enthalpy of the propellant at 300" K., the sum of the enthalpies of sufficient carbon dioxide, water, oxygen, and nitrogen
a
10
4
Figure 5. Flame Temperatures at Different Pressures of, Propellant Composition with Different Enthalpies Enthalpy in calories per gram
A. B.
C.
-360 -644 -015
0,
2821 from 0' to 300' K. to make up a unit mass of propellant. Linear interpolation between the two computed temperatureenthalpy points to obtain the temperature a t which the enthalpy of the product gases equals the enthalpy of the unburned propellant gives the adiabatic flame temperature, Table I1 gives the flame temperatures of a series of propellant compositionswith different enthalpies. Consistent with expectations, the propellants with the highest enthalpies exhibit the highest flame temperatures a t a given pressure. However, the flame temperatures of the hot propellants are markedly pressure dependent, while those of the cool propellants are almost independent of pressure. For example, the flame temperature of a relatively hot propellant (-368 calories per gram) changes by 9.0% in the pressure range of 40 to 1 atmospheres The flame temperature of a propellant with an enthalpy of -875 calories per gram changes only 4.5% in the same pressure range. This effect of pressure on the flame temperature is illustrated in Figure 5 for propellants with different enthalpies.
INDUSTRIAL AND ENGINEERING CHEMISTRY
2822
Flame Temperatures of Propellant Compositions with Different Enthalpies
Table II.
Pressure, Atmosvheres
Specific Enthalpy, Calories per Gram -368 - 644 -876 -368 - 644
x.
CONCLUSION
2817 2689 2461 3014 2847 2544 3056 2906 2661 3099 2974 2578
-368 - 644 -672 - 368 - 644 -875
The adiabatic flame temperature of a propellant is an important parameter in the calculation of theoretical performance characteristics. The computation of the flame temperature requires a knowledge of the composition of the product gases which may be determined by one of several systematic mathematical procedures, The flame temperature is but little dependent on pressure for propellants with low enthalpies but it is dependent markedly for propellant of high enthalpies. The effect of ernpirical formulas of the same enthalpy on flame temperature is not, predictable without a knovledge of the composition of the product gases. ACKNOWLEDGMENT
111.
Flame Temperatures as Function of Pressure for Different Propellant Compositiona of Assumed Equal Enthalpies Pressure, Atmospheres Propellant Flame Temperature, O K. 2915 2881 2838 2500 2137 3000 2905 2876 2502 2139
A
10
B C
D
E A B C
20
D E
Propellant
C
H
A B C
0.1170
0,4006 0.2787 0.4712 0.3004 0.3247
D E
0,2086 0,0341 0,2451 0,2497
0 0.3573 0.3967 0.3039 0.3452 0.3349
The large number of calculations involved in obtaining the data given in the tables mere performed by Jane McCuistion. LITERATURE CITED
(1: BrinMey, S. R., and Kandiner, H. J., IXD.ENG.CHEM.,42, 860 (1950).
(2) Edse, R., Air Tech. Service C o m m a n d , PiogressRegt., No. IRE-47 (1946).
Atom % a
compositions for which equal enthalpies were assumed. The importance of a knowledge of the composition of product gase-, are emphasized by these results.
Flame Temperature,
- 875
Table
Vol. 43, No. 12
S
0.1260 0.1159 0,1913 0.1093 0.0935
(3) Hottel, H. C., Williams, 6. C., and Satterfield, C. N., “Thermodynamic C h a r t s for Combustion Processes, P a r t I,”New York. John Wiley & Sons, 1949. (4)Kreiger, F. J., and K h i t e , W.B., Project Rand Rept., RA15 055 (October 1942).
(5) McEwan, W. S.,NavOrd Rept. 1239, U. 3. Yaval Ordnance Test Sta. (July 1960). (6) McEwan, W. S.,and Skolnik, 801, Ren. Sei. Instrument, 22, No. 3, 125-32 (1951).
( 7 ) Sachsel, G. E”., Bell, J. C., and X a n t i s , M. E., “Third Symposium
If compositions having the same enthalpy hut differing in empirical formulas are considered, then no simple correlation between flame temperature and composition appears possible. Table I11 lists the calculated flame temperature of five different
on Combustion, Flame and Explosive Phenomena,” p. 620, Baltimore. Williams R- milkins Co., 1949. (8) Wimpress, A. K.,“Internal Ballistics of Solid-Fuel Rockets,” p. 4,New York, McGraw-Hill Book Co., 1950. RECEIVED July 3, 1951.
E. S. STARKMAN’, A. G. CATTANEO, AND S. H. MCALLISTER Shell Development Co., Emeryville, Calif.
To
investigate gas turbine combustion problems separate from the complexities of compressor and turbine, a small combustion chamber was designed to incorporate the principles of contemporary full scale burner tubes. More specifically, i t was used to study the influence of fuel and operating factors on the carbon deposits found in combustion chamber liners. The carbon deposition tendencies of paraffinic, naphthenic, and aromatic fuels increased in that order. For a gven ratio of carbon to hydrogen, deposits decrease as volatility increases, more generally increasing with specific gravity of the fuel. Deposits are influenced by operating conditions. Increased air temperature, increased air-fuel ratio, and decreased pressure minimize them. For a given combustion chamber, normally encountered operational ranges of air temperature and pressure influence the amount of carbon as much as changes in fuel composition. 1 Present address, Department of Mechanical Engineering, University of California. Berkeley, Calif.
Besides providing a laboratory test for carbon deposits, this work resulted in the derivation of a formula by which the carbon deposition tendency of a fuel can be predicted from its characteristics.
URING the development stage of aircraft gas turbine engines, kerosene has generally been used a s the fuel, Choice of kerosene was prompted by a number of reasons, of which low vapor pressure was the primary one (1). It had not been ascertained from an operational standpoint whether fuels of the kerosene type were really the most desirable, and a survey in laboratory apparatus of various potential gas turbine fuel components appeared necessary. Same of the desired qualities wer? apparent from piston engine experience (%). One aspect of performance which could not suitably be evaluated from piston engine data, because of the difference in mode of combustion, was the tendency for coking or carbon formation in the gas turbine combustor. Carbon formation is objectionable for many reasons, among which are low combustion efficiency, local hot spots on the
