AIAA 95-2640
DiboraneKO, Rockets for
Use in Mars Ascent Vehicles
Robert Zubrin
Martin Marietta Astronautics
Denver, Colorado
31st AIANASMEISAEIASEE
Joint Propulsion Conference and Exhibit
July 10-12,1995/San Diego, CA
For psrmlsrion to copy or republlsh, contad the Amerlcan lnstltute of Aeronautics and Astronautlcs
370 L'Enfant Promenade, S.W., Washlngton, D.C. 20024
AI AA-95-2640
Diborane/CO, Rockets for Use in Mars Ascent Vehicles
Robert Zubrin’
Martin Marietta Astronautics
PO Box 179
Denver CO 80201
303-971-9299 (phone) 303-977-3600 (FAX)
Abstract
A new propulsion concept for Mars Ascent vehicles is
presented which burns diborane (B2H6) with
indigenous Martian CO, in a bipropellant engine.
Analysis is presented which shows that the expected
Isp for such an engine is about 300 s with C02:B,HG
mixture ratios of 2.5:l. The advantage of this system
is that both propellants are storable in space and o n
Mars and the propellant acquisition system has low
power requirements and is extremely simple. Analysis
is presented which shows that the B,H$CO;, system
outperforms all non-ISPP propulsion systems for both
a Mars Sample Return and a Mars ballistic hopper
mission, including those with much higher engine
specific impulses. The B,H$CO,
system is
outperformed on a mass basis by a chemical ISPP
system producing CH,/O, from hydrogen feedstock
and
Martian
CO,
using
a
combined
SabatierlElectrolysis-Reverse water gas shift cycle,
but the comparative simplicity of the B2HdCO2system
may make il the preferred option nevertheless. It is
therefore recommended that experiments be
undertaken to demonstrate the feasibility of B,H$CO,
engines.
lntroductlon
One possible means of simplifying an in-situ
propellant production (ISPP) system for Mars
applications is to find a fuel that can burn in pure CO,.
If such a fuel can be found, then the ISPP system
reducesto a C02 acquisition device - eliminating the
need
for
chemical
reactors,
condensers,
electrolysers, desiccant beds, and cryogenic
refrigerators, with a concomitant reduction in both
mass and power of the ISPP system by about 80%
In the past, some authors’.’ have considered burning
metals, such as magnesium, in CO,. The problems
with such systems are that the specific impulse
attained is only about 190 seconds and either a
hybrid or solidlliquid slurry engine must be
developed. Alternatively, it has been proposed3 that
c
raw CO, be used as a propellant fluid in a thermal
rocket. This however requires the use of a nuclear
reactor to provide the energy to heat the fluid.
Employing supercritical CO, as propellant without
substantial heating has also been proposed4, but this
results in a specific impulse of about 70 seconds,
which is too low for most applications.
In this paper we propose a new solution, that of
burning diborane, BH
,,
with CO, in a relatively
conventional bipropellant chemical engine.
Diborane is a high-energy fuel that has sometimes
been considered in the past for use in rocket
vehicles. Using LOX as an oxidizer, diborane
combustion can produce an Isp of about 41 0 s. B,H,
is stable, has a density of about 0.4 that of water, and
boils at -90 Cat 1 bar pressure. Its critical temperature
however is 9 C, and would be a storable on Mars (-50
C) under about 5 bars of pressure.
b
Diborane can react with CO, in accordance wilh:
B2HG+ 2C02 = 8 2 0 3 + H20 + CH,
AH=196 kcallmole
+C
(1)
With a AH of 196 kcallmole, this reaction has an
enthalpy of 7097 kJlkg, which translates into an ideal
Isp (infinite expansion, complete combustion) of 384
seconds.
In the case of incomplete combustion more typical of
rocket engines with finite nozzles and high
temperature exhausts, the reaction will occur in
accordance with:
BzHG + 2C02 = B203 + CO + 3Hz + C
AH = 147 kcal/mole
(2)
The ideal theoretical Isp of this reaction is 333
seconds.
These results strongly suggest that a real Isp of over
300 s can be attained in actual CO,/B,H, engines.
t Senior Member A I M
Copyright 0 1995 by Martin Marietta Corp. Published by the American Institute of Aeronautics and Astronautics, Inc. with
permission.
b
ISDof B2H6 Burnina i n C02
-
320
300 -
8
-
-X-
Clapplsp
Shafirovitch Isp
Schnackel Isn
Clapp -500 psi
Shafirovitch- 147 psi
Schnackel
1000 psi
-
Expansion-600
u)
U
C
al
u)
-a
VI
2201
200
0
1
2
3
4
5
6
7
8
C02/B2H6 Mass Mixture Ratio
Fig. 1 . Specific impulse of B,H$CO,
u
rockets. At a mixture ratio of 2.53, an Isp of 300 s can be attained
It will be noted that the mixture ratio in the above
chemical equation is 88:28, or 3.14 CO, to B,H, by
weight. This means that a rocket using such a system
would only have to bring 24% of its ascent propellant
from Earth; the remaining 76% could come from the
Martian atmosphere.
Because much of the product of reactions (1) and (2)
is in the condensed phase, it is difficult to accurately
estimate the rocket engine performance via simple
calculation. I therefore asked three different analysts
operating with different codes to estimate the
performance of a diboranelC0, engine. The analysts
were Jeff Schnackel. or Martin Marietta, Captain
Mitchell Clapp, of the USAF Phillips Lab, and Dr.
Evgeny Shafirovitch, of the Institute of Structural
Macrokinetics in Moscow. Schnackel assumed a
chamber pressure of 1000 psi, Clapp 500 psi, and
Shafirovitch 147 psi. Schnackel and Clapp both used
nozzle expansions of 600; Shafirovitch used an
expansion of 288. The results given by both
Schnackel and Clapp include a degradation factor of
0.95 to account for various engine Isp losses beyond
that calculated for finite expansion, Shafirovitch used
0.9. All three analysts assumed equilibrium flow. The
results of the code runs of all three analysts are given
in fig. I .
Schnackel did not report chemical composition of the
exhaust products, but Clapp and Shafirovitch did. In
both cases, the exhaust products were found to
largely resemble those predicted by equation (2),
with, however, small amounts of water and trace
amounts of CH, present.
In addition to these equilibrium chemistry runs, Clapp
did one calculation using a 2-D kinetics code. That
calculation, using a chamber pressure of 500 psi, a
nozzle expansion ratio of 600, and a mixture ratio o f
2.5:1, produced an engine Isp of 302.6 seconds, in
good agreement with the estimates given by the
degraded 1-D equilibrium results shown in Fig. 1 .
It can be seen that Isp peaks at close to 320 seconds
for CO;JB,H, mass mixture ratios between 1 and 1.5.
However the advantage of using this propellant
combination is maximized Ias large a portion of the
total propellant as possible is GO2, since that
component can be obtained readily on Mars. The
optimum mixture ratio is thus a compromise between
that which maximizes Isp and that which maximizes
propellant mass leveraging. In what follows, we derive
a figure of merit for estimating this optimum mixture
ratio.
I is clear that in such cases large values of L are
optimal, since c only affects the result in linear
fashion. Conversely, if the AV is large, c becomes
very important and smaller mixture ratios may be
optimal.
We define L as the "leverage factor."
L =Total Propellant UsedIPropellant from Earth (3)
The "L" value of a conventional Mars ascent vehicle
using propellant of terrestrial origin is thus always
equal to 1.0. A B,H$CO, rocket operating with a
mixture ratio of 2.5 has L=3.5. The mass ratio of a
rocket is given by:
M= mass ratio = eXp(AV/C)
In Figure 2 we present graphs of the wewdry mass
ratio for three different missions. One is a is a Mars
ascent vehicle (MAV) which flies to a low Mars orbit for
a total AV of 4.2 kMs. The second is a MAV which
flies to a highly elliptical Mars orbit for a total AV of 5.4
k N s . The final case is a MAV which flies from the
surface to direct Trans Earth Injection,for a total AV of
6.4 kMs. In all cases, Schnackel estimates for
B,H$CO, Isp are used.
(4)
where c is the exhaust velocity. The quantity (M-1) is
the ratio of propellant mass to vehicle dry mass. For
our mission analysis preliminary figure of merit we
therefore propose:
Z = (M -1)IL
-
For purposes of comparison, the weWdry ratios for
NTO/MMH propulsion (320 s Isp, L=l) are also
shown. It can be seen that the BH
,,
system
transported weWdry ratio is lower than that of the
NTOIMMH option by about a factor of 3. ll can also be
seen that, minimizing the transported weVdry ratio, a
mixture ratio of about 3 : l appears to be optimal, as
this minimizes 2 for the high A V missions and is as
good as the higher mixture ratios for the low A V
missions. For much lower AV missions, however such
(5)
The factor, 2, is just the vehicle's transported wetidry
mass ratio (i.e. the ratio of the transported propellant
mass to the vehicle's dry mass).
For AV's much smaller than the exhaust velocity,
equation (5) reduces to:
i/
TransDorted WetlDrv Ratio of B2H6IC02 MAV's
-
I \
4
-I\
NTOiMMH 5.4 kmis
-NTOiMMH
I
4.2 kmis
t l Z n b I t i U Z b.4 KmlS
B2H6lC02 5.4 kmls
BZH6lC02 4.2 km/s
I
0
1
2
3
4
5
6
C02182H6 Mass Mixture Ratio
Fig. 2. Transported wet/dry ratios for
B2H6propulsion. A mixture ratio of 3 : l appears optimal.
3
L
-
as a 1 kWs (50 km range) ballistic hopper,
be outgassed at high pressure during the next day by
heating it, with an estimated average daytime power
requirement of about 50 W being required. The
outgassed COz would be allowed to flow into another
vessel maintained at Mars ambient temperatures
where it would condense and liquefy. As the sorption
pumps absorb COz in preference to nitrogen and
argon, only a small amount of these Mars atmosphere
minority gas constituents would show up in the
condenser tank.(Mars atmosphere is 95% CO,, 2.7%
NZ,1.6% Ar, plus traces) As they would not liquefy
under Mars temperatures, they could then be
disposed of by periodic venting.
still higher
mixture ratios would be more optimal. On the other
hand, the use of Z as a figure of merii is
methodologically slightly biased towards high mixture
ratiollow Isp systems (since the inert weight of
tankage is neglected, as is the mass of the COz
acquisition system which would increase with greater
CO, acquisition requirements). If such factors are
taken into consideration the optimal mixture ratio for
the high AV missions could be as low as 2.53.
Propulsion System Design Considerations
-
One attractive feature of the B,H$CO, propulsion
system is that the combustion temperature is low, o n
the order of 2000 C (for a mixture ratio of 2.73, 500
psi chamber pressure, Clapp’s calculation). This is
about 1000 C less than a typical HJOz rocket. A
stainless steel combustion chamber should thus
certainly be possible. On the other hand, some of the
products are in the condensed phase. This could
cause erosion of engine materials, notably at the
throat. Ablative coating may therefore be necessary.
Part of the exhaust is elemental carbon. This could
cause coking. However, as in the case of LOX/RP
engines, such coking could be just the thing needed
to provide the engine with a protective layer. Further
analysis or experiment is necessary to determine how
much ablative protection is thus required.
Mission Analysis
In fig. 3 we show the results of mission analysis for
performing a Mars Sample Return mission using
various propulsion options, but a common set of
assumptions for the mass of avionics systems, sample
return capsules, science payload, and aerosheii and
lander masses as a percentage of landed cargo.
Shown in competition are options employing
conventional NTOlMMH bipropellants (lsp=320 s ) ,
Chlorine Pentafloride (CPF, lsp=360 s), and CH,/O,
(Isp =380 s), with ail propellant coming from Earth; and
options which produce all or part of their return
propellant on Mars, including CH,/O,-ISRU (lsp=380
s), and COJB,H, (Isp =300 s at a CO,:B,H,
mixture
ratio of 2.5:l). Ail five mission considered are
performed with a single vehicle the returns directly
from the Martian surface to Earth; i.e. no orbiter or
Mars orbit rendezvous is used. It can be seen that
with the given assumption of an Earth Return Vehicle
(ERV) consisting of the sample return capsule and all
trans-Earth cruise system avionics with a total mass of
50 kg, that the Conventional option using the
NTO/MMH propulsion system cannot be performed
with a Delta launch, and the option using the CPF
propulsion system is probably too heavy (Le. virtually
no launch margin) for a DeHa launch as well. If
terrestrial CH,/O, is used about 14% launch margin is
available, which is barely acceptable. On the other
hand, if CO&,H,
propulsion using indigenous CO,
for propellant is used, the launch margin triples to
43%, while the CH,/O, mission utilizing ISRU has
100% launch margin. Thus we see that despite its
higher power requirements, a CHJO, system
producing its propellant by the Sabatier/Electrolysis
(SE) cycle does offer the highest mass leverage.
However this is a case where second highest may be
best, as the ISRU system required by the COJB,H,
system is much simpler.
Diborane and CO, are both storable as liquids at Mars
ambient temperatures under about 80 psi of
pressure. As the propellant combination is relatively
dense, this is suggestive that a simple pressure fed
engine system m y be optimal. The pressure could
be created autogenously simply by raising the tanks
to about 0 C, or r could be supplied externally by
compressed helium gas. Given the low combustion
temperatures and the small engine sizes under
consideration for Mars Sample Return mission
applications (on the order of 100 - 200 ibf of thrust)
regenerative cooling of the nozzle should be
unnecessary.
C 0 2 Acquisition System
-
The COz acquisition system used by the vehicle
would be an activated carbon sorption pump such as
that recently demonstrated by Martin Marietta under
its in-situ propellant production IR&D. Such sorption
beds can absorb up to 40% of their weight in CO,
under Mars nighttime temperature condition?. Thus,
to acquire 1 kg of CO, per night would require a
sorption bed with a sorbent mass of 2.5 kg. This could
4
Mars Sample Return Mission ODtions
L
1200
1000
....-...D e b - T W
Capability
0
0
800
€I
$I
v1
v)
a
600
E3
I
-i
K4
400
TMI Backpack
Aeroshell
Decelerator
Lander Struct.
P/L Contingency
Science
ERV
ISRU plant (kg)
Power(kg)
Stages
Prop from Earth
200
n
the CH4/0, and HdO, ISPP systems operate similarly,
i.e. propellant is acquired or produced "one hop at a
time." The B,H$CO, system operates at a mixture ratio
of 3:1, giving it a net propellant leverage of 4. The
CH4/0,
ISPP
system
uses
a combined
Sabatier/Electrolysis-Reverse Water Gas shifl ISPP
system, allowing % to leverage hydrogen transported
from Earth into CH,/O, with a leverage of 18. The
HJO, ISPP option uses a Reverse Water Gas
Shift/Electrolysis Cycle to produce oxygen, which it
burns with q brought from Earth at a mixture ratio of
6 : l , giving it a net propellant leverage of 7.
In fig. 4 we show the mass that must be delivered to
Mars to fly a Mars ballistic hopper mission which visits
several sites in succession. It is assumed that the
basic mass of the hopper, excluding tanks and
engines is 50 kg, with the tanks and engines
comprising 10% of the propellant mass. In addition,
the options which use chemical ISPP systems
(CH,/O, ISPP and HJO, ISPP) have a 40 kg combined
ISPP piant and power unit that they must fly around
with, while the B,HdCO, option has a simpler 10 kg
combined sorption pump/power unit that it must take
with it when it hops. The AV for each hop is 2000 mis.
which includes 1650 m/s for ascent and 350 mis to
land after aerodynamic deceleration using vehicle
drag. Assuming 20% gravity losses during ascent and
no extension of range via gliding after re-entry, this is
sufficient to ailow the vehicle to traverse 500 km
during each hop. In the analysis shown it is assumed
that the NTO/MMH system has an ISP of 320 s, the
CPF has an Isp of 360 s, the CH4/0, has an isp of 380
s,the B,H$CO, has an Isp of 300 s and the HJO, has
system acquires the
an Isp of 450 s. The B,H,/CO,
C02 needed for each hop just before that hop, and
It can be seen that after about 3 hops, the required
transported mass of all the non-ISPP system grows
exponentially to infinity. The BJHdCO, system mass
is still reasonable after 6 hops - it takes about 8 hops
for this option's mass to get out of control. The CH,/O,
and HJO, ISPP options offer even more atlractive
mass leverage. However, these options require
storing liquid hydrogen not only during both transMars cruise, as would be needed for a sample return
mission, but also for an extended period (order of
i/
5
-0-
NTOlMMH
7
CPF
-C
0)
1,
2
'+
CH4102
HZO2
1650 m/s DV Ascent
350 mls descent
C
0
0
m
s
U
B
0
I//
C
0
-1
-
/
BZH6iC02
>
<
I
f -
--+0
1
2
3
4
5
HZ02lSPP
CH4/02 ISPP
6
Number of 500 km Hops
Fig. 4 Mass of Mars Ballistic Hoppers as a function of the number of hops undertaken
4
several hundred days) on the surface of Mars during
the subsequent hopping mission. This imposes
technical challenges that may make these higher
performing options unattractive. If so, the simpler
B,HdC02 option would emerge as a real winner.
chemical reactors, water condensers, electrolysis
units, dryers, and cryogenic refrigerators. Thus while
the propellant mass leverage of the B2HdC02system
is nowhere near as great as an S/E system producing
CHJO, from imported hydrogen (380 s Isp with and
propellant leveraging of 18:1), the mass and power
requirement of the propellant production system is
likely to be less by about a factor of 5. Moreover, both
the required imported fluid (B2H6) used by the
B,HdCO, system and the acquired propellant system
are storables. In comparison, the SE system requires
transporting hard cryogenic hydrogen to Mars and
storing soft cryogenic CH,/O,
on Mars for an
extended period of time. The simplicity of the
propellant acquisition system used by the B2H$C0,
system also argues for reliability, increasing the
probability of both mission success and acceptance
of the technology by the mission planning
community.
Conclusion
The advantages of B,HdCO, propulsion are clear.
Compared with conventional bipropellant transported
from Earth, less than 113 the propellant mass needs
to be delivered to Mars to allow the accomplishment
of a Mars sample return mission. For low AV missions,
such as a Mars ballistic hopper, the rocket equation
becomes linear, and an "effective Isp" of the
propulsion system may be calculating by multiplying
the engine Isp by the propellant mass leveraging.
With an Isp of 240 at a mixture ratio of 5:l (L=6), the
effective Isp of a B,HdC02 hopper is 1440 Seconds,
endowing it with capabilities for multiple hops that are
simply unattainable without using indigenous
propellants. Compared to other systems for using
indigenous propellants on Mars, the B,H$CO, system
has the advantage of offering the possibility of radical
simplification. Only a C 0 2 acquisition and
compression system is required, eliminating all
For these reasons it is recommended that the
B2H,/C02 propulsion system be investigated further.
As a first step, it is recommended that experiments be
undertaken to demonstrate the fundamental
feasibility of this technology by building and firing a
small B,H,/CO, engine.
6
Acknowledgment
I wish to thank Captain Mitchell 8.Clapp of the US Air
Force Phillips Lab, Jeff Schnackel of Marlin Marietta,
and Dr. Evgeny Shaiirovitch, or the Institute of
Structural Macrokinetics in Moscow for their
calculations in support of this paper.
References
1. S. Yuasa, and H. Isoda, "Carbon Dioxide Breathing
Propulsion for a Mars Airplane," AlAA 89-2863, 25th
AIANASME Joint Propulsion Conference, Monterey,
CA, July 1989.
2. E. Shafirovitch, A. Shiryaev, and U. Goldshleger,
"Magnesium and Carbon Dioxide: A Rocket
Propellant for Mars Missions," Journal of Propulsion
and Power, Vol9, No. 2, March-April 1993.
3. R. Zubrin, "Nuclear Rockets Using Indigenous
Martian Propellants," AlAA
89-2768,
25th
AIANASME Joint Propulsion Conference, Monterey,
CA, July 1989.
4. D. Pettit, "A Carbon Dioxide Powered Rocket for
Use on Mars," AAS 87-264, Case for Mars 111, Boulder
CO, July 1987
5. R. Zubrin, S. Price, L. Mason, and L. Clark, "Mars
Sample Return with In-Situ Resource Utilization An
End to End Demonstration of a Full-Scale Mars In-Situ
Propellant Production Unit," Report to NASA JSC,
Jan 13, 1995.