Diehl, John M

Ordin, Paul M

Rowe, William H

English

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RESEARCH MEMORANDUM 
INVESTIGATION OF THE DIBORANE - HYDROGEN PEROXIDE 
PROPELLANT COMBINATION 
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By William H. Rowe, Paul M. Ordin 
and John M. Diehl 
Flight Propulsion Research Laboratory 
Cleveland, Ohio 
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NA llONAl ADVISORY COMMITTEE 
FOR AERONAUTICS 
WASHINGTON 
May 28, 1948 
https://ntrs.nasa.gov/search.jsp?R=19930085775 2020-06-17T16:41:55+00:00Z
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NACA RM No. E7K07 
NATIONAL ADVISORY COMMITTEE FOR AERONAUI'ICS 
RESEARCH MEMORANDUM 
INVESTIGATION OF THE DIBORANE - HYDROGEN PEROXIDE 
PROPELLANT COMBINATION 
By William H. Rowe, Paul M. Ordin 
and Jobn M. Diehl 
SUMMARY 
1'wo runs were made with liquid diborane and hydrogen peroxide 
in a small-scale experimental rocket engine. After tbe first run} 
,.bich was very erratic, several modifications were made to the 
apparatus and test procedure. In the second run combustion started 
immediately and continued smoothly for tbe 12 seconds of operation. 
The diborane - hydrogen peroxide weight ratio investigated was 0.27 
(stoichiometric, 0.136). The maximum experimental specific impulse 
attained during the run ,res 211 pound-seconds per pound for a 
combustion-chamber pressure of 237 pounds per square inch gage. 
This experimental specific impulse was 81 percent of the theoretical 
specific impulse for the same mixture ratio and combustion-chamber 
pressure. 
INTRODUCTION 
The use of diborane (B2H6 ) as a rocket fuel is of current 
interest because of the large amount of energy per unit weight 
available from its reaction .n tb oxidants. Tbe thermodynamic prop-
erties and gaseous equilibriums of boron and tbe oxides of boron 
have been determined by the National Bureau of Standards (refer-
ence 1). Considerable work in connection with the chemistry of 
boron and boron compounds has been conducted by Schlesinger and 
associates at the University of Chicago. In addition, the inves-
tigation to determine many of the physical] chemical, and physio-
logical properties of diborane is being sponsored by the Naval 
Research Lab ora tory, \{asbington, D. C. Few experimental investi-
gations, however, of the performance of diborane as a rocket fuel 
have been conducted. 
The experimental results of two runs conducted at the NACA 
Cleveland laboratory using 95- percent liquid diborane and 87-percent 
2 CONFIDENl'IAL NACA RM No. E7K07 
bydrogen peroxide in a IOO-pound thrust rocket engine are presented 
herein. Also presented are theoretical values of specific impulse, 
reaction-chamber temperature, and exhaust temperature for the oper-
ating conditions of one of the runs. 
APPARATUS AND PROCEDURE 
A diagrammatic sketch of the 100-pound thrust rooket unit 
illustrating the propellant (catalyst, fuel, and oxidant) systems 
and the gas pressurizing systems used for pumping the propellants 
is shown in figure 1. The entire apparatus was mounted on a stand. 
that was pivoted to permit horizontal motion. 
The oxidant (87-percent hydrogen peroxide) and catalyst 
(44-percent calcium-permanganate solution) were contained in 
stainless-steel tanks. A chrome-molybdenum steel tank (fig. 2) 
was used to hold the fuel (95-percent diborane BzH6)' Tbe fuel 
tanlc vTaS equipped ion th a stainless-steel bursting disk and high-
pressure stainless-steel hand valves that were packed with Teflon, 
a polytetrafluoroethylene plastic. Check valves were installed in 
each of the propellant lines. The fuel t8nk and lines were refrig-
erated with dry ice to maintain the diborane (boiling point, 
-134.50 F at 1 atmosphere) in a liquid state up to the injector 
plate. 
The oxidant and catalyst supply tanks were pressurized with 
water-pumped nitrogen gas from a common nitrogen supply and the 
fuel supply tank was pressurized with oxygen-free helium. The gas 
pressure to each of the propellant tanks was controlled with two-
stage preSS1lre regulation. 
A detailed sketch of the 100-pound-thrust rocket engine is 
shown in figure 3. The engine comprises a stainless-steel injector 
plate containing six solid-jet injector nozzles, a thick-walled 
1 
copper combustion chamber with an inside diameter of 24 inches and 
a length of 18~ inches, and a convergent-divergent copper exhaust 
nozzle designed to provide complete expansion at a pressure ratio 
of 20.4. The two diborane injector nozzles were located diamet-
rically opposite to each other and produced impinging jets meeting 
apprOximately 3/8 inch upstream of the point of tropingement of the 
four hydrogen-peroxide injector nozzles. The hydrogen-peroxide 
injector nozzles were located 900 apart on a radius outside that 
of the fuel nozzles. The catalyst was injected through a hole in 
CONFIJ)ENTIAL 
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NACA RM No. E7K07 CONFIDENTIAL 3 
the center of the injector plate. The inner walls of the combustion 
chamber and exhaust nozzle were chrome-plated to resist corrosion 
and erosion. The ratio of the combustion-chamber volume (up to the 
tr~oat of the eXhaust nozzle) to the cross-sectional area of the 
exhaust nozzle throat was 255 inches. 
Thrust was measured by means of two strain gages mounted on a 
bar subjected to bending by the engine thrust. (See fig. 1.) The 
two s ages were connected in a resistance bridge circuit of a modi-
fied, continuous-recording, self-balancing potentiometer, which 
traced a thrust-time relation. Propellant comsumption was deter-
mined by differences in weights before and after the run. Specific 
impulse, defined as the thrust per unit rate of propellant consump-
tion, was obtained by integrating the thrust-time diagram with a 
planimeter and dividing by the propellant consumption. Combustion 
pressure was measured by a Bourdon-type pressure recorder. 
Propellant-supply pressures were recorded by taking motion pictures 
of the Bourdon gages. Color motion pictures were taken of the 
exhaust flame. 
Prior to operation, the oxidant and catalyst systems were 
loaded and the diborane feed system was flushed with oxygen-free 
helium. The diborane tank was then removed from dry-ice storage, 
weighed, placed in a drY-ice-alcohol bath, and connected to the 
feed system. 
For the first run, the catalyst (44-percent solution of calcium 
permanganate) was injected first; after a delay of about 2 seconds, 
the injection of 87-percent hydrogen peroxide followed. After 
another delay of approximately 2 seconds, the liquid diborane was 
admitted and the catalyst flow immediately shut off. A fuel-
oxidant weight ratio of 0.67 was selected, which would permit a 
burning time of approximately 5 seconds for the quantity of diborane 
provided. 
For the second run, several changes were made to the apparatus 
and the ' firing procedure: (1) The type of air-operated diborane 
control valve was changed; (2) the catalyst injection was omitted; 
(3) nozzle sizes were changed for a fuel-oxidant weight ratio 
of 0.27 (stoichiometric, 0.136), which would permit a burning time 
of approximately 12 seconds for the quantity of diborane provided; 
(4) the motor was heated by a hand torch just prior to the loading 
of the diborane to facilitate the initial decomposition of the 
hydrogen peroxide; and (5) the diborane and hydrogen peroxide were 
injected simultaneously into the preheated engine. 
CONFIDENTIAL 
4 CONFIDENTIAL NACA RM No. E7K07 
RESULTS AND DISCUSSION 
Poor performance and erratic burning were obtained in the ~irst 
run and data required to evaluate the performance of the propellant 
combination could not be obtained. The white smoky appearance, 
typical of straight peroxide rune, changed to a clear green flame 
when the diborane was injected; but the flame varied in intensity 
and burning pieces of molten boron or boron compounds were observed 
to drop from the exhaust nozzle during the run. Although for the 
intended fuel-oxidant weight ratio of 0.67 the burning time should 
have been only 5 seconds, there was no indication of cessation of 
burning at the end of 25 seconds at which time the propellant 
systems were shut off. Smoke issued ~rom the exhaust nozzle and 
burned upon contact with the air for about l~ hours after the pro-
pellant systems were shut off. 
After the first run, the exhaust nozzle was completely clogged 
with a hard brown solid and the inside of the rocket engine was 
completely coated with a soft, white, fluffy deposit apprOximately 
1/4 inch thick under which was a layer of a dark brown hard sub-
stance (fig. 4). The hydrogen-peroxide injectors were clean but 
the tips of the diborane injectors were filled with a solid white 
deposit. The lines, the air-operated valve, and the check valve 
of the diborane system had a very fine yellowish-white coating on 
the inside surfaces. The erratic operation was attributed to a . 
reduced and an intermittent flow of diborane, which may have been 
caused by decomposition of the diborane in the injectors or by 
faulty operation o~ the air-operated control valve, or both. 
In the second run, in which the diborane and hydrogen peroxide 
were injected Simultaneously, immediate and smooth combustion 
occurred. The exhaust ~lame was a brilliant green with many lumi-
nous particles clearly recorded by the color motion pictures. An 
enlargement o~ a frame from the motion pictures is presented in 
figure 5. The luminous particles in the exhaust flame are prob-
ably deposits Originally formed on the engine walls and products 
o~ combustion condensed or solidi~ied by the decrease in tempera-
ture during the expansion process through the exhaust nozzle. 
Thrust and combustion pressure increased progressively during the 
12 seconds o~ operation, followed by slig!1t afterburn1ng, which 
occurred for about 5 seconds. 
Experimental thrust and combustion-chamber pressure, propel-
lant flow ra~es, and experimental and theoretical values of specific 
impulse for the second run are plotted against running time in 
CONFIDENTIAL 
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NACA EM No. E7K07 CONFIDENTIAL 5 
figure 6. Thrust and combustion-chamber pressure increase rapidly 
for the first l~ seconds, after which time both curves have a 
marked tendency to approach a level. The propellant-flow-rate 
curves were plotted from data of the combustion-chamber pressure, 
injection pressures, total yropellant consumptions, and total time. 
The propellant flow rates were high at the start of the run because 
of the high differential pressure across the injection nozzles due 
to the low combustion-chamber pressure; but after the first 
1 12 seconds, these flow rates tended to level off at lower steady 
values for the duration of the run. 
A corresponding change is seen in the experimental specific-
impulse curve obtained from the thrust and total-propellant-
consumption data. During the first second of the run, the values 
of thrust and combustion-chamber pressure were low, and the propel-
lant consumption rate was high; consequently, the specific impulse 
was low. As the combustion-chamber pressure and thrust increased 
and the total propellant consumption rate dropped, the specific 
impulse rose. Values for the theoretical specific-impulse C1.l.rve 
were calculated for the experimental mixture ratio and combustion 
pressures using the data of references 1 and 2. It was assumed 
that all products of reaction were ideal gases and that no recom-
bination or condensation occurred during expansion. The maximum 
experimental specific impulse value of 211 pound-seconds per pound 
is apprOximately 81 percent of the maximum theoretical value of 
262-pound-seconds per pound for the mixture ratio investigated. As 
is to be expected, the low combustion and nozzle efficiencies of 
the first part of the run are made evident from a comparison of the 
specific-impulse curves. 
A contributing factor to the initial low performance is the 
poor mixing and combustion resulting from the use of only two 
solid-jet injector nozzles for the diborane and four for the hydro-
gen peroxide. With better injection, starting and over-all effi-
ciency could be improved. It should be noted that the mixture 
ratio investigated is not the optimum for this~uel-oxidant com-
bination. 
The performance data for the second run are summarized in 
the following table: 
CONFIDENTIAL 
r---
6 CONFIDENTIAL NACA EM No. E7K07 
Diborane - hydrogen peroxide weight ratio • • • • • • • • 0.27 
Maximum thrust, ( lb ) •••••••• • • • • • • • • • • • 68 
Maximum combustion-chamber pressure, (lb/sq in. gage) • 237 
Maximum specific impulse, (lb-sec/lb) • • • • • • • • • 211 
Reaction temperature (calculated for maximum specific 
impulse), 0]' • • • • • • • • • • • • • • • • • • 4995 
Exhaust temperature (calculated for maximum specific 
impulse), <>:F • • •••• • • •• • • • • • • 2972 
As a result of the second run, the walls of the rocket engine 
were coated with a 1/8- to 1/16-inch gray-to-dark brown, hard layer 
of boron and boron oxides (fig. 7). At the injection end of the 
engine and for the first 3 or 4 inches downstream, the deposit was 
rough and uneven in thickness and gray in color. On the remainder 
of the engine up to the exhaust-nozzle throat, which was clean, the 
coating gradually became dark brown, smooth, and more even in thick-
DeSS with many gas pockets and a surface glaze. The smooth, glazed 
coating, especially in the Vicinity of the convergent section of 
the exhaust nozzle, was probably formed by heat and flowing gases 
during the second run. 
SUMMARY OF RESULTS 
From two runs using approximately 95-percent liquid diborane 
and 87-percent hydrogen peroxide in an experimental lOO-pound thrust 
rocket engine, the following results were obtained: 
1. During the first run, in which hydrogen peroxide and a 
catalyst were lnjected prior to the diborane, very erratic combus-
tion attributable to a faulty supply of diborane was produced. 
2. For the second run, in which the catalyst was omitted and 
hydrogen peroxide and diborane were injected Simultaneously, 
immediate and smooth combustion was obtained for the duration of 
the run (12 sec). The diborane - hydrogen peroxide weight ratio was 
0.27 (stoichiometric, 0.136), the maximum combustion-chamber pres-
sure was 237 pounds per squa.re inch gage, and the specific impulse 
was 211 pound-seconds per pound, which is 81 percent of the theo-
retical specific impulse for these operating conditions. 
CONFIDENTIAL 
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NACA RM No. E7K07 CONFIDENTIAL 7 
3. The engine was much cleaner after the second run than after 
the first. For the second run, a layer o~ boron and. boron oxide 
1/16 to 1/8 inch thick was deposited on most of the engine surfaces 
with only a negligible amount of deposit on the exhaust-nozzle 
throat. 
Flight Propulsion Research Laboratory, 
National Advisory Committee for Aeronautics, 
Cleveland, Ohio. 
l'{Elli'ERENCES 
1. Wacker, Paul F., Wooley, Harold W., and Fair, Myron F.: Thenno-
aynamic Properties and Caseous E~uilibria of Boron, Oxygen and 
the Oxides of Boron. Tech. Rep., Nat. Bur. Standards, Heat 
and. Power Div., Jan. 25, 1945. (Bur. Aero., Navy Dept.) 
2. Hirschfelder, J. 0., McClure, F. T., Curtiss, C. F., and Osborne, 
D. W.: Thermodynamic Properties of Propellant Gases. NDRC 
Rep. No. A-116, Nov. 23, 1942. (~ension and Revision of 
NDBO Rep. No. A-48 (OSRD No. 547).) (Abs. in Bib. Sci. and 
Ind. Reps., vol. 2, no. 10, Sept. 6, 1946, p. 743, PB 28531. ) 
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Investigation of the Diborane - Hydrogen Peroxide Propellant Combination

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