Mars Gashopper
Final Report on NASA Contract No.: NAS3-00074
Prepared By:
Robert Zubrin, Principal Investigator
Brian Frankie
Mark Caviezel
Gary Snyder
Dean Speith
Gilbert Chew
Frank Tarzian
Brian Birnbaum
Andrew Martin
Pioneer Astronautics
11111 W. 8th Avenue, Unit A
Lakewood, CO 80215
Approved By:
____________________________
Dr. Robert Zubrin
President
Pioneer Astronautics
8 June, 2000
FORWARD
This report was prepared by Pioneer Astronautics in Lakewood, Colorado, and
contains the results of a National Aeronautics and Space Administration (NASA) Small
Business Innovative Research (SBIR) Phase I study for the Jet Propulsion Laboratory
(JPL), as part of contract NAS3-00074, “Mars Gashopper”
SBIR RIGHTS NOTICE (MAR 1994)
These SBIR data are furnished with SBIR rights under Contract No. NAS3-00074. For a
period of 4 years after acceptance of all items to be delivered under this contract, the
Government agrees to use these data for Government purposes only, and they shall not be
disclosed outside the Government (including disclosure for procurement purposes) during
such period without permission of the Contractor, except that, subject to the foregoing
use and disclosure prohibitions, such data may be disclosed for use by support
Contractors. After the aforesaid 4-year period, the Government has a royalty-free license
to use, and to authorize others to use on its behalf, these data for Government purposes,
but is relieved of all disclosure prohibitions and assumes no liability for unauthorized use
of these data by third parties. This notice shall be affixed to any reproductions of these
data, in whole or in part.
MARS GASHOPPER
NASA SBIR CONTRACT # NAS3-00074
PROJECT SUMMARY
8 June 2000
The object of this NASA Phase I SBIR program was to assess the performance of a
Mars Gas Hopper, or “gashopper.” The gashopper is a novel concept for propulsion of a
robust Mars surface hopper vehicle which utilizes indigenous CO2 propellant to provide
Mars exploration with greatly enhanced mobility. The gashopper will acquire CO2 gas
from the Martian atmosphere, and store it in liquid form at a pressure of about 10 bar.
When enough CO2 is stored to make a substantial ballistic trajectory hop to another Mars
site of interest, the CO2 propellant tank will be moderately heated to raise it to 70 bar.
The propellant then runs through a hot pellet bed to form high temperature gas that is
expanded through a nozzle to produce thrust. The gashopper uses its CO2 propulsion
system for major liftoff, attitude control, and landing propulsive burn(s), as required.
Unlike chemical rockets, the gashopper’s exhaust will not contaminate the landing site
with organics or water. The gashopper has a potential flight range of 5 to 50 kilometers. It
can fly over terrain impassible to rovers, imaging as it flies, land to reconnoiter a remote
location, and then fly again. Thus, it offers unique capabilities for Mars surface
exploration.
During the Phase I program, Pioneer Astronautics designed and tested several
gashopper propulsion systems. The systems included the cold gas thruster, which uses
unheated liquid CO2 through the thrust nozzle, and three different hot thrusters. The hot
thrusters included the water heated thruster, which uses hot pressurized water (between
34 and 102 bar), the electrically heated thruster, which uses an electric current
(approximately 344 Amps), and the heated bed thruster, which uses preheated graphite
rods or a hot pellet bed to preheat the CO2 before it reaches the thruster.
Based upon the results of the above experimental setups, Pioneer Astronautics
believes that the use of the heated pellet bed thruster is the most efficient method of
operation of the gashopper. Following the test stand experiments, a flight unit was
designed, constructed and tested using the heated bed thruster idea. A 20 lb lift gaseous
He balloon was attached to the flight unit to partially counter its terrestrial weight
(approximately 50 lb). The flight unit provided an approximate thrust of 50 lbf, an Isp
over 80 seconds, and reached an apogee of about 40 meters.
The Mars gashopper clearly has the ability to be an enabling technology for Mars
exploration, as it will greatly expand the potential for long distance mobility on Mars.
Report on the Design and Operation of a Mars Gashopper
Robert Zubrin, Brian Frankie, Dean Speith, Gilbert Chew, Mark Caviezel,
Andrew Martin, Frank Tarzian, Gary Snyder, Brian Birnbaum
Pioneer Astronautics
11111 W. 8th Avenue, Unit A
Lakewood, CO 80215
303-980-0890
Introduction
This report describes work accomplished on the Mars gashopper project, contract
number NAS3-00074. The gashopper is a novel concept for propulsion of a robust Mars
surface hopper vehicle which utilizes indigenous CO2 propellant to provide Mars
exploration with greatly enhanced mobility. The gashopper provides a distinct advantage
over other surface mobility options because it utilizes indigenous Martian resources as its
propellant and because it is not confined to remaining on the ground, thereby allowing it
to easily travel long distances over rough terrain that would be impassible for rovers.
This project involved the design, construction and testing of several different
gashopper units, with each providing a different level of performance. The gashopper
design included one unheated CO2 rocket and three heated CO2 rockets. The three
options explored for preheating the CO2 prior to the nozzle were by hot water, using an
electrical current, and using a preheated packed pellet bed. In addition, a flight unit was
designed, constructed and tested using the preheated packed pellet bed. The flight unit
produced approximately 50 lbf thrust and 80 seconds of Isp. Assisted by a 20 lb lift
balloon to provide stability, it flew to an altitude of 40 meters and achieved a semi-soft
landing when its engine was restarted by radio control during descent.
Analysis shows that near-term gashoppers utilizing conventional materials could
achieve hopping ranges of tens of kilometers. Such systems could also perform aerial
imaging during flight and descent, thereby providing context information for surface
exploration wherever it chose to land. In addition, the gashopper could provide an ideal
test-bed for gaining the large experience base needed to perfect the automated landing
hazard avoidance systems that will be needed for the Mars Sample Return mission.
In short, our analysis and experiments to date indicate that Mars gashoppers are
feasible, and would offer unique enabling capabilities for Mars exploration.
Technical Background
The Mars Gas Hopper, or “gashopper,” is a novel concept for propulsion of a robust
Mars surface hopper vehicle which utilizes indigenous CO2 propellant to provide Mars
exploration with greatly enhanced mobility. The gashopper will acquire CO2 gas from
the Martian atmosphere, and store it in liquid form at a pressure of about 10 bar. When
enough CO2 is stored to make a substantial ballistic trajectory hop to another Mars site of
4
interest, the CO2 propellant tank will be moderately heated to raise it to 70 bar. The
propellant is then run through a hot pellet bed to form high temperature gas that is
expanded through a nozzle to produce thrust. The gashopper uses its CO2 propulsion
system for major liftoff, attitude control, and landing propulsive burn(s), as required.
Unlike chemical rockets, the gashopper’s exhaust will not contaminate the landing site
with organics or water. The gashopper has a potential flight range of 5 to 50 kilometers. It
can fly over terrain impassible to rovers, imaging as it flies, land to reconnoiter a remote
location, and then fly again. Thus, it offers unique capabilities for Mars surface
exploration.
Background
Mars is a very big place, with wildly varying terrain and enormous rises and dips in
surface topography. The development of controllable mobile systems capable of rapid
maneuvering across long distances in this environment would be of great value for the
exploration of Mars. Presently, mobile robotic technology for Mars missions is limited to
surface rovers, which are incapable of negotiating the chasms, descending into the
canyons, or climbing the mountains of Mars. While effective in their own sphere, as
demonstrated by the recent NASA/JPL Mars Pathfinder mission, such systems are limited
to exploring the region in close proximity of a lander site. While more capable rovers,
such as the Athena, are currently in development with travel capabilities on the order of
100 meters per day, such systems will still be limited to travel over comparatively mild
ground, with boulder fields, steep slopes, cliffs, canyons, and mountains all forming
impassible barriers to their movement. Furthermore, once it has left the landing area, such
a long range ground rover would no longer have context imagery available from its
descent vehicle to support its traverse, and context photos to support its movement could
only be supplied by much lower resolution images taken by satellites. In contrast, a flight
vehicle operating on Mars would be unrestricted by the Red Planet’s rough terrain, and
would be able to supply its own context photographs taken during descent towards each
new site it was sent to survey. Thus it is clear that systems for enabling true long range
exploration on Mars will have to be capable of flight. While airplanes, balloons, and even
helicopters have been proposed for such a role, the thinness of the Martian air combined
with the high speed of Mars’ high altitude winds strongly indicate that a controllable long
distance flight vehicle will most likely have to be rocket propelled.
Unfortunately, however, if the rocket vehicle so employed has to bring all of its
required propellant from Earth, the effect on the rocket equation caused by cumulative
V’s associated with performing a series of surface to surface ballistic hops rapidly
drives the vehicle’s propellant requirements to infinity. Indeed, the only way a viable
Mars Ballistic Hopping Vehicle (MBHV) can be designed is if it can derive at least a
significant fraction of its propellant from local Martian resources after each landing.
Several concepts for such a MBHV have been proposed in the past. For example, in
1989, Zubrin (Ref. 1) proposed using a nuclear reactor to heat raw native CO2 rocket
propellant. Such a vehicle would have unlimited mobility on Mars, as an infinite supply
of readily accessible propellant would be available to support its operation anywhere on
the Red Planet. However the nuclear reactor required to drive such a system is currently
unavailable.
5
An alternative option is to use a chemical Mars in-situ propellant production (ISPP)
systems (such as that described in Zubrin, references 3, 4, and 5) to produce a chemical
bipropellant such as CH4/O2 by combining a small amount of imported hydrogen with
Martian CO2. Such systems offer the potential of very high performance (specific
impulses of 375 s, propellant mass leveraging of 18:1), but the ISPP system required are
fairly complex and generally require transporting cryogenic hydrogen feedstock to (and
in this application, around) Mars. Furthermore, the production of chemical bipropellants
requires a lot of power, and since the landers are quite power limited, this implies that a
long period (typically many months) will be required on the surface between hops to
allow the MBHV time to produce its required propellant.
While in-situ produced bipropellants or nuclear heated CO2 may be necessary to
enable Mars hoppers with global reach, for shorter flights, ranging from 5 to 100 km, a
much simpler low-tech reusable rocket system suggests itself. We call such a system,
which uses raw CO2 heated to moderate temperatures (300 to 1000 K) to generate a low
specific impulse rocket exhaust, a Mars Gas Hopper, or “gashopper.”
In the gashopper concept, a small robotic flight vehicle, possibly equipped with a
microrover for local surface exploration is landed on Mars. Then using an activated
carbon sorption pump it acquires CO2 from the Martian atmosphere, storing it in liquid
form in a tank at a pressure of about 150 psi. Then, when flight to another location is
desired, solar power is used to warm the CO2 to about 30 c, thereby increasing its
pressure to close to 1000 psi. Once the warming of the CO2 is complete, a valve is
opened allowing the warm liquid to be into a bed of pellets that had been preheated to
perhaps 700 C. the hot pellet bed flashes the liquid CO2 to gas, and heats it to the
temperature of the bed. The hot gas is then expelled out a rocket nozzle to produce thrust
for takeoff. The same system can also be used to provide gas for RCS systems to control
the vehicle during flight. While the specific impulse of such a warm gas system is only in
the neighborhood of 100 s, this level of performance is sufficient to allow the vehicle to
fly 10 km or more, and still have enough propulsive capability for a soft landing.
Alternatively, instead of heating the propellant gas in a pellet bed, the gas can could be
warmed to a super-critical state in a carbon-overwrapped “supertank, and then expanded
directly out a rocket nozzle to produce thrust. Another alternative is to heat the CO2 in
the engine chamber using energy stored in a electrical batteries or flywheels. Depending
upon the technology assumptions employed, such a “supertank” and “electrical”
gashoppers may offer flight performance comparable to that of the hot pellet-bed
gashopper described above.
Whatever propulsion option chosen, after landing, additional CO2 can be acquired
from the atmosphere, enabling another flight. Thus the gashopper offers Mars scientists a
new tool, which can provide them with greatly enhanced mobility through the use of a
ballistic medium range flight vehicle that can refuel itself each time it lands. While the 5
to 50 km single flight range offered by the gashopper does not represent true global
mobility, repeated flights would allow for very long distances to be traveled. Moreover,
the vehicle would be capable jumping into canyons then flying out of them again, flight
to the tops of mountains, hopping over crevasses, and performing numerous other
maneuvers that are far beyond the capabilities of foreseeable surface rovers.
6
Despite its high molecular weight, carbon dioxide is clearly the propellant of choice
for such system because of its prevalence in the Martian atmosphere. The atmosphere of
Mars consists of 95.0% carbon dioxide, 2.7% nitrogen, and 1.6% argon. It can be
obtained by pumping the Martian atmosphere into a tank. At a typical Martian
temperature of 233 K, carbon dioxide liquefies under a pressure of 10 bars. Under these
conditions, assuming an ideal isothermal compression process, liquid CO2 can be
manufactured for an energy cost of just 84 W-h per kg. On a CO2 propelled gashopper,
this power could be provided by either a solar photovoltaic or solar dynamic concentrator
electrical power generation system. Unfortunately, currently available off-the-shelf
roughing pumps are only about 2% efficient, but even at that rate, a compressor driven
with 180 W could acquire 50 grams of CO2 per hour, or 0.5 kg per 10-hour day. (This is
the actual performance of a Vacuubrand P-79103-50 diaphram pump.) Alternatively, the
pumping power could be generated simply by using thermal energy from an RTG or solar
system source to heat a zeolite or activated carbon sorption pump to outgas CO2 which in
batch fashion is alternately adsorbed into the bed when its temperature is allowed to drop
to Mars ambient conditions. This method of acquiring and compressing CO2 on Mars
was demonstrated under simulated conditions on Earth by R. Zubrin in 1994 (Zubrin
1994), as part of a system to generate in-situ methane/oxygen chemical propellant on
Mars using the Sabatier/electrolysis (S/E) system. About 50 W (thermal) is required to
produce about 500 grams of CO2 per day using this technique. Still another option is to
acquire the CO2 out of the Martian atmosphere by freezing, as was demonstrated by
Frankie and Zubrin in a laboratory simulation performed in 1998 (Frankie and Zubrin,
1998). Using this system, about 30 W (electric) is required to produce 0.5 kg/day of CO2.
It should be noted that the sorption pump or freezer power requirements for these systems
were on the order of 10% the total power required for the synthesis of the chemical
bipropellants produced by the units they were feeding. This sharply illustrates the order
of magnitude propellant-production power-economies that can be achieved if raw
indigenous volatiles can be used as rocket propellants in place of chemical fuel/oxidizer
combinations that must be synthesized. Using this power-economy advantage, a CO2
driven ballistic gashopper would be able to fuel itself and carry out a program of 10 or
more hops in the time that a chemical rocket driven Mars ballistic hopper required to
acquire the propellant needed for a single hop. The CO2 acquisition system is simpler,
lighter, and more reliable than a chemical rocket indigenous propellant manufacturing
unit as well. Liquid CO2 has a density 1.16 times that of water and is eminently storable
under Martian conditions. It is these considerations, despite the modest specific impulse
performance possible with CO2, that makes a CO2 propelled gashopper an attractive
option to consider.
In addition, the use of CO2 as a propellant is uniquely necessary for astrobiology,
because CO2 exhaust will not contaminate the landing sites with organic molecules or
water, as any conventional bi-propellant propulsion system would. To explore the
potential of a simple gashopper concepts to meet emerging Mars surface exploration
requirements, Pioneer Astronautics performed a Phase I SBIR study for Jet Propulsion
Lab which included both analytical studies of gashopper vehicle systems and trades, and
the design, construction, and test firing of a variety of gashopper engine types. In
7
addition, a prototype gashopper flight vehicle using the hot pellet-bed rocket system
(PBRS) was constructed, and flown on a 100 m rocket-propelled hop.
The results of these studies and experiments strongly indicate that the gashopper is
feasible and that it could offer unique and enabling capabilities for long-range mobility
on Mars. A summary of these results is presented below.
Results of Phase I Program
The Phase I program consisted of general analysis, systems studies, and experimental
demonstrations. We discuss each of these in turn.
A. General Analysis
The first performance metric to be calculated in assessing the potential of the
gashopper is the specific impulse of the rocket system. This was calculated using the
Phillips Laboratory One-Dimensional Equilibrium (ODE) analysis program. Assuming a
chamber pressure of 1000 psi and a nozzle expansion ratio of 400, the ideal vacuum
specific impulse performance of the CO2 thruster as a function of chamber temperature is
displayed in Figure 1. One would expect delivered rocket system performance to be
between 90 to 96 percent of that shown in Figure 1.
In fig 1, the temperature values of greatest interest for near-term gashoppers are those
between 1000 and 1500 K, as pellet-bed engines made of conventional materials (i.e.
steel) can readily be constructed capable of supporting such operation. Beyond 1500 C,
unconventional approaches, such as graphite or carbide construction become necessary,
with the material in contact with the hot CO2 being limited to high temperature oxides.
Heating such a pellet bed also requires more power, however, which restricts such
systems to gashoppers equipped with RTG power sources In contrast, as we shall see,
photovoltaics are adequate to support the operation of the 1000 to 1500 K gashoppers.
300
250
Isp (s)
200
150
100
50
0
0
500
1000
1500
2000
2500
Temperature (K)
Figure 1. CO2 Thruster Ideal Vacuum Specific Impulse as a Function of Chamber
Temperature.
8
3000
Assuming rocket performance shown in Figure 1, super-tank gashopper velocity
capability and hopper range are shown as a function chamber temperature and mass ratio
in Figures 2 and 3, respectively. Depending upon the size of the payload carried and the
structural and electronic technologies employed, near-term gashoppers should be
buildable with mass ratios ranging from 1.3 to 2.0. It can be seen that for a near-term
gashopper operating at 1000 K, this implies delta-V capabilities ranging from 350 m/s
(for mass ratio= 1.3) to 950 m/s (for mass ratio= 2.0)
Gashopper Delta V as a Function of Mass Ratio
1200
Delta V (m/s)
1000
800
600
400
200
0
1.0
1.1
1.2
1.3
1.4
1.5
1.6
1.7
1.8
1.9
2.0
Vehicle Mass Ratio
Chamber Temp = 800 K
Chamber Temp = 1000 K
Chamber Temp = 1200 K
Chamber Temp = 1400 K
Fig. 2 Near-term Gashopper Delta-V as a function of Mass Ratio
Assuming 15% gravity losses and taking no credit for partial aerodynamic deceleration of
the vehicle (i.e. the full delta-V required for soft landing is supplied by the rocket
propulsion system), We see that this implies near-term gashopper flight ranges from 6 km
to over 50 km.
9
Gashopper Range per Hop as a Function of Mass Ratio
70
Range (km)
60
50
40
30
20
10
0
1.0
1.1
1.2
1.3
1.4
1.5
1.6
1.7
1.8
1.9
2.0
Mass Ratio
Chamber Temp = 800 K
Chamber Temp = 1000 K
Chamber Temp = 1200 K
Chamber Temp = 1400 K
Fig. 3 Near-term Gashopper flight range as a function of mass ratio.
Other key observations are that for a given mass ratio value, gashopper velocity and
range capabilities are not strong functions of chamber temperature.. These results
indicate that gashopper tank subsystem weights will be an important influence on
performance, but that CO2 gas temperatures can be quite moderate, say1,000 K, to get
good gashopper performance. These observations imply that common engineering
materials, such as stainless steel or titanium, can easily meet the design requirements of a
CO2 propelled gashopper. No advanced, high-temperature materials will be required to
develop such a system.
B. System Trades
Four primary types of gashoppers were considered as part of the present study. These
were;
* The Cold Gashopper: In this option cold liquid CO2 is stored, and then flashed to a 2
phase mixture in the rocket chamber, which is then expanded out a nozzle to produce
thrust. This is the simplest gashopper concept, and can attain the highest propellant mass
fraction. However, even if the propellant tank is heated to 30 C, (CO2 goes supercritical
at 31 C) specific impulse is limited to about 35 s (on Mars, about 25 s when tested at
Denver ambient conditions.) With this specific impulse , and a mass ratio of 2.0, the
gashopper would be limited to flight ranges of about 3 km. While modest, this
performance is enough to be of some interest for mission application, especially in view
of the attractive system simplicity. Cold Gashoppers were therefore included as part of
10
the experimental Phase I test program. However, higher specific impulse is required if
true long-range mobility is to be achieved.
* The Supertank Gashopper. As mentioned above, when CO2 is heated above 31 C, it
goes supercritical, with a resulting dramatic increase in propellant pressure for storage at
a given temperature. However, in recent years, various manufacturers have developed
“supertanks” consisting of graphite overwrapped aluminum, that have extremely high
strength to weight ratios. A state of the art supertank with a favorable (largely cylindrical)
geometry can achieve a metric of performance given in the form of pressure X
volume/weight (PV/W, in English units) of about 1 million inches. Thus if we consider a
20 liter tank (1220 cubic inches), pressurized to 5000 psi, such a tank would have a
predicted weight of 6.1 lbs, or 2.77 kg. If the CO2 were heated to 500 K, the tank could
hold 7.4 kg of CO2. This is not too bad a tank fraction, but at 500 K, the specific impulse
would be limited to about 75 s. If the tank were heated to 1000 K, a specific impulse of ~
130 could be attained, but only 3.7 kg of propellant could be stored in the tank. This low
propellant mass fraction would lead to low vehicle performance. Actually, however, at
1000 K, the supertank would not preserve its high strength, so even the 3.7 kg storage
estimate presented above is unrealistically optimistic. Thus, in fact, supertank gashoppers
are limited to the 500 K range. Assuming a vehicle mass ratio of 1.5, the 75 s specific
impulse of such a system would give it a hopping range of about 4.5 km. This is an
improvement on the cold gashopper, but the large propellant tanks needed (triple the size
needed by other concepts) pose issues, as this entire volume must be heated. Because
better performance is available from other concepts, and because supertanks rated for
even 500 K are not available off-the-shelf for SBIR-budget class costs, this concept was
not selected for experimental demonstration during Phase I.
* The Electrical Gashopper. In this concept, energy is stored in batteries, and then used to
heat the propellant while it is flowing through the chamber. Essentially a very high
powered resistojet using stored energy, this concept is fairly simple, and has the
advantage that the energy required for the thrust maneuver can be stored up over any
length of time, thereby reducing power requirements relative to other concepts, such as
the pellet bed system discussed below. However, the rate of power release required of the
batteries is very high. For example, consider a gashopper with a thrust of 150 lbf (668 N),
a reasonable level as it would give a 80 kg wet-mass gashopper a takeoff thrust/weight
ratio of 2.25 on Mars. Assuming a specific impulse of 130 s (~700 C), this means a mass
flow of 0.52 kg/s. If the propellant has been preheated in its tank to 30 C, 911 kJ/kg of
heating would be required to heat the propellant in the engine. At a flow rate of 0.52 kg/s,
this implies a battery discharge power of 474 kilowatts. It is unclear whether regenerative
batteries could be made that would support this kind of discharge on a repeated basis.
Also, it should be noted that current secondary batteries, such as lithium-ion only have a
storage capacity of about 125 W-hr/kg (450 kJ/kg). Thus 2 kg of batteries would be
needed to heat each kg of CO2 to 700 C. This would be fatal to the goal of achieving an
attractive vehicle mass ratio. If the propellant temperature is dropped to 600 K (327 C),
then only 468 kJ would be needed to heat each kg of CO2, and the battery and propellant
masses would be equal. When payload and other system masses are taken into account,
this would probably allow an electrical vehicle to be constructed with a mass ratio of
11
about 1.35 and a specific impulse of about 90 s. This would give it a flight range of about
3.5 km. This is not particularly attractive, as it is nearly equaled by the much simpler
Cold Gashopper. However, battery and flywheel technologies are currently in
development that could provide double to triple the power density available from current
lithium-ion batteries. These could increase the flight range to 7 to 10 km, which would be
a substantial improvement over the cold and supertank gashoppers. Thus, while posing
some issues, the electrical gashopper also offers promise. It was therefore selected for
experimental examination during Phase I.
* The Hot Bed Gashopper. In this concept the energy required to flash the liquid CO2 is
stored in the form of a bed of material which is heated to high temperature in advance of
the thrusting maneuver. If the material has a high melting temperature, such as graphite,
boron, or beryllium, the bed would take the form of a mass of pellets. We thus sometimes
also call this concept the pellet bed, or particle bed Gashopper. As a variant however,
materials with low melting points but high specific heats, such as lithium or aluminum
can also be used. Alternative concepts in which these metals are clad in steel, copper or
titanium tubes and then inserted in the engine as a cluster (like pencils in a can) with gas
flow channels formed by the interstices between the cladding tubes are also of
considerable interest. In either case, the CO2 in the storage tank is first warmed to 30 C,
which also pressurizes it to about 1000 psi. The liquid is then made to flow either through
the pellet bed of the bundle of liquid metal containing cladding tubes, during which time
it is heated to the temperature of the bed. The high temperature, high pressure gas is then
expanded out a rocket nozzle to produce thrust.
This concept has a number of advantages over the electrical gashopper. In the first place,
rate of energy transfer from the bed to the gas (which may be a show-stopper for the
electrical gashopper) is simply not an issue. When the gas flows through the hot pellet
bed, power is transferred in direct proportionality to the gas flow rate, ands that’s all there
is to it. The system is thus simple to start, run, and to throttle. The system is also likely to
be much more robust for repeated operation than batteries operating at high discharge
rates. Most importantly, however, the energy density available from a hot bed is much
higher than that possible with state of the art batteries. For example, if heated to 1000 K,
beds of boron, beryllium, or lithium will store 1160, 1700, or 2700 kJ/kg, respectively.
This is the equivalent performance, respectively, of that which would be offered
hypothetical rechargeable batteries that could provide 322 W-hr/kg, 472 W-hr/kg, or 750
W-hr-kg.. If the bed is heated to higher than 1000 K, its decisive energy storage
advantage increase even more. Thus near-term hot bed gashoppers operating at
temperatures ranging from 1000 K to 1400 K should be achievable with mass ratios of
about 1.5. This would allow flight ranges from 13 km (at 1000 K) to 23 km (at 1400 K).
If lightweight avionics are employed, the mass ratio could be probably increased to 1.6,
and these ranges would grow to 18 to 30 km.
An issue with the hot bed gashopper is the need to store a large amount of energy in
thermal form without it leaking to the environment faster than the spacecraft power
system can supply. As shown below, however, out analysis shows that for practical
gashopper systems of interest, that this problem can be well-handled for modest mass
impact by enclosing the engine vessel in a vacuum jacket containing multi-foil insulation.
12
Thus, because of its simplicity, potential high reliability, and high performance, the hotbed gashopper was selected as the gashopper baseline, and made the subject of most of
the Phase I programs analysis and experimental work.
Gashopper Design Concepts
An illustration showing a possible conceptual gashopper configuration is shown
in fig. 4..shown is the science payload and avionics, a reaction control system, a “pump”
(the intake for the freezer of sorption pump), and a set of folding solar panels which in
their stowed position also serve as the flight fairing. The engine is below the tank,
somewhat concealed in this drawing by the solar panels. It should be noted that there is
only one tank; the CO2 serves as its own pressurant.
Figure 4. The Gashopper Concept.
Hot Particle Bed Gashopper Propulsion System Concept. The hot particle bed
gashopper propulsion system concept is depicted in Figure 5. This gashopper stores CO2
from the absorption pump/separation unit as a liquid in a thermal insulated low-pressure
tank. As previously mentioned, CO2 can easily be stored as a liquid on Mars at a
pressure of 10 bar. This pressure can be raised to whatever higher value is desired for
engine operation by a small amount of tank warming. The high-pressure, liquid CO2 is
then passed through a simple, high-efficiency, hot particle bed or pellet bed heat
exchanger, where it is heated and gasified. At the exit of the heat exchanger, a
converging-diverging supersonic nozzle is attached from which the moderatetemperature, high-pressure CO2 gas is exhausted producing thrust. Before the heat
exchanger, upstream of the main control valve, a small amount of high-pressure liquid
CO2 is teed off and passed through a small tube heat exchanger, which is located on the
13
outer periphery of the internal region of the particle bed heat exchanger. This liquid CO2
is gasified and used for the RCS.
The hot particle bed heat exchanger draws substantially from past particle bed nuclear
reactor technology work (Powell, et al. 1988). The hot particle bed heat exchanger
consist of an annulus pack bed which consist of small spherical particles made of a high
heat-capacity material such as boron, beryllium, or special forms of graphite, see Figure
6. These particles are held in place by outer and inner concentric surface screens, known
as frits. These frits are sized such as to let the CO2 flow radially through them with little
pressure drop, but they are small enough to hold the small particles in place. These frits
can be made from a high-temperature metal alloy or carbon-carbon material. The top and
bottom of the packed bed annulus are closed with a solid, high-temperature metal alloy.
Because of the high surface area-to-volume associated with the small particle, high heat
transfer is ensured. In the hot particle bed gashopper heat exchanger a number of hightemperature material rods are located in the particle bed. They are electrically resistive
heated when the gashopper is at a Mars landing site to heat the bed. Power is supplied
these rods from the vehicles solar electric generation system.
During gashopper flight operation, the power is turned off and the CO2 is fed to hot
particle bed heat exchanger, as previously mentioned. The CO2 liquid enters from the
forward region of the heat exchanger, where it is manifold to the outer radial region of
the hot packed bed, see Figure 6. The liquid CO2 then flow radially through packed bed
where it is heated, by taking the energy from the hot particles, and the liquid gasifies.
The moderate-temperature, high-pressure CO2 gas is then collected in the annulus core
region and expanded through the nozzle. The technology needed to develop the packed
bed and the associated gashopper, is well with in the state-of-the-art.
14
Figure 5. The Hot Particle Bed Gashopper Concept.
Figure 6. The Hot Particle Bed Heat Exchanger Design.
15
A variant on the hot particle bed gashopper described above is to use a hot pellet bed,
with the primary distinction being that pellets (~1 to 2 mm diameter) are larger and easier
to handle, and impose a smaller pressure drop on the flow. In consequence, instead of
radial frits, pellet beds can be constructed consisting simply of engine tubes with a
screened off plenum at each end: the CO2 gas entering at one end and exiting at the other.
Such systems are much simpler to construct than radial-frit particle beds. Particle beds
offer better heat transfer, of course, but if metal pellets with high thermal conductivities
are employed, the heat transfer remains adequate in a pellet bed to get almost all the heat
out anyway. For this reason, the primary subject of experimentation of hot-bed rockets
developed and fired during the Phase I experimental program were systems of the pellet
bed type.
Materials for Hot Bed Heat Storage
One of the most critical parameters governing the potential performance of a hot bed
gashopper is the heat capacity of the storage material. Hot gaseous CO2 has a specific
heat of about 1.24 kJ/kg-K. in comparison, the specific heats of a variety of high heat
capacity substances is shown in Fig 7. To first approximation, the amount of bed material
needed to heat a given quantity of CO2 to any temperature will be given by the inverse of
the relationship between the materials (average) specific heat and that of CO2. Thus, for
example, if we are operating with a initial bed temperature of 1000 K, we see that the
average specific heat of boron between 300 K and 1000 k is about 1.7 kJ/kg K. Thus it
would take about 1.24/1.7 = 0.73 kg of boron to store enough energy to heat 1 kg of CO2
to this temperature.
Fig 7: Specific Heat -- Cp
4.5
Li
4.0
3.5
Be
kJ/ kg K
3.0
Al
Li
Be
B
2.5
B
2.0
1.5
Al
1.0
0.5
0.0
400
600
800
1000
1200
1400
T(K)
Fig. 7 Specific heats of selected hot bed materials as a function of temperature.
16
We may also note here in passing the possibility of using graphite as a heat storage
material. Specific heats for graphite are given by various sources as ranging from 0.7 to 2
kJ/kg-K, with the difference being apparently dependant not merely on the source, but on
the structure of the graphite considered. At the high end of this estimate range, graphite is
comparable to beryllium (and much cheaper). We order some graphite rods from lab
supply houses during the Phase I, and found there specific heat to be near 0.7 kJ/kg-K.
Aside from their disappointing low heat capacity, these rods used in a bundled rod
configuration in a test engine performed reasonable well, suffering no erosion by hot
CO2 when operating at about 950 K. If high –specific heat graphite can be found and
validated, it could offer a low-cost option with performance comparable to that presented
by beryllium in these charts.
The above approximation ignores;
a)The residual heat left in the bed.
b)The extra heat available from the engine case of other material present(cladding
tubes for Li or Al systems, for example.)
c) the latent heat of fusion from metals, such as lithium or aluminum, which are
brought to molten form.
To a fair extent, considerations (a) and (b) above cancel each other. In either a wellarranged particle or pellet bed, almost all the heat put in the bed can be gotten out. This is
because when the cool CO2 hits the bed, it initially cools the region it first hits nearly all
the way to 30 C (the CO2's’ininitialemperature). This remains true as the “cold front”
moves down the bed as its heat is exhausted, and only at the very end of the burn, when
the cold front has reached the inner frit (of the particle bed) or bottom (of the pellet bed)
will the chamber temperature drop below what it was at the start. If we choose to end the
burn there (to maintain Isp) a small amount of residual heat in the bed will be left
untapped. If we are willing to accept a loss of Isp at the end of the burn, however, all the
heat can be taken out. But to the extent we are unwilling to let Isp drop to get every scrap
of heat out of the bed, the difference can be largely made up from heat stored in other
engine materials (such as the steel or titanium engine case, etc.) Thus, with reasonable
accuracy, (a) and (b) above can be expected to cancel.
Consideration c, however, offers a significant benefit to aluminum and lithium bed
engines. This can be seen in figure 8, where we have integrated the total heat stored in
various media at various temperatures. For purposes of reference, it takes about 930 kJ to
heat CO2 (from 303 K, or 30 C, its initial condition leaving the storage tank) to 1000 K
and 50 bar. Comparing this to the data in Fig. 8, we see that at 1000 k, for example,
beryllium will store about 1800 kJ per kg at this temperature. Therefore 1 kg of CO 2 can
be heated to this temperature by about 0.5 kg of beryllium. The sharp jag in the lithium
and aluminum lines shown in fig 8 are a result of the phase changes of these materials
which occur at the temperatures indicated. This effect is particularly beneficial to
aluminum if operation is planned at 1000 K, increasing aluminum’s 1000 K heat capacity
by about 60%.
17
Fig 8: H eat C apacity -- C p  T + H f
 T = T - 3 0 0 K
5000
4500
4000
kJ / kg
3500
Al
Li
3000
Li
Be
2500
Be
2000
B
1500
Al
B
1000
500
0
400
600
800
1000
1200
1400
T (K )
Fig. 8. Heat capacity of various metals, with 300 K as the zero point. For comparison,
note that heating 50 bar CO2 to 1000 K requires 930 kJ/kg.
. It can be seen from figs 7 and 8 that if heat capacity per unit mass is the only
consideration, then lithium beats all other contenders hands down. However,
unfortunately lithium is not very dense, having a specific gravity of 0.53. Density is
important, because the denser the hot bed material is, the smaller the engine can be and
therefore the easier it will be to heat it. In fig. 9, we show the results if the heat capacity
per kg is multiplied by the material density, thereby providing heat capacity per liter. It
can be seen that in this case beryllium is the best, closely followed by boron. Lithium,
however is the worst, with a heat capacity per liter less than half that of beryllium. A
lithium engine would thus require twice the volume of engine casing as a beryllium or
boron one. Moreover, since the lithium would be molten, it would have to be contained in
cladding tubes made of copper, brass, or steel, all of which have poor heat capacity per
mass, and thereby negating much of the lithium’s mass advantage. Moreover, the bundled
cladding tube system required by the lithium engine is will offer poorer heat transfer
characteristics to that offered by a beryllium or boron pellet bed.
18
Fig 9: Heat Capacity -- Cp  T r+H f r
kJ / Liter
T=T-300K
5000
4500
4000
3500
3000
2500
2000
1500
1000
500
0
Al
Li
Be
B
B
Be
Al
Li
400
600
800
1000
1200
1400
T (K)
Fig. 9. Heat Capacity per volume of Candidate Hot Bed materials.
It is thus concluded that beryllium or boron pellet beds offer the best option for
gashopper propulsion.. Beryllium melts at 1550 K, boron at 2570 C. It can be seen that
throughout its range, beryllium is superior to boron in both heat per kg (significantly) and
heat per liter (slightly). For gashoppers operating above beryllium’s melting point, boron
provides the best option.
Heating the Pellet Bed
A key issue for the successful; implementation of the hot bed concept is our ability to
contain a hot vessel with a minimum of heat leak. This is critical, because unless heatleak can be minimized, a large power source will be needed to heat the bed quickly. On
the other hand, if the heat –leak is very slow, we can take days to heat the bed up, and the
power required can be quite low. As near term gashoppers will probably have be powered
with photovoltaics, achieving a low heat leak in the engine is essential.
We choose to analyze this problem by considering a reasonable design point of a
gashopper operating at 1000 K, employing a 12 liter engine vessel and having an
available 100 W for heating during the day and nothing for heating at night.
The 12 liter tank to be insulated could be spherical with an inside diameter (ID) of 28.4
cm (11.2 inches) and therefore a surface area of 2533 cm2. Or the tank could be of a
barrel design with many different length to diameter ratios. Consider a barrel design with
a cylindrical section length L’ = D the tank diameter. Assume the domed ends have a
19
height of D/(22). Then D=21.8 cm (8.6 inches) and the total tank length will be 37.1 cm
(14.6 inches), with a surface area of 2568 cm2. This is close enough to the spherical tank
area that the spherical area will be used for subsequent MLI analyses.
The heat loss rate q from the insulated surface that faces ambient is given by
q =  F A2 (T24 – Ta4)
(1)
where  is the Stefan-Boltzmann constant, F the radiation interchange factor equal to
surface emissivity s in this application, and A2 the outside radiating surface area. We
will assume the ambient temperature Ta is much less than the outside vehicle surface
temperature T2 as a worst case heat loss rate. Rearranging, the temperature T2 can be
calculated in terms of allowable heat losses:
T2 =
4
q
(2)
  s A2
A 5 watt heat loss, or 5% of the nominal heating rate of 100 watts, will give a T2=242K (33oC), whereas a 10 watt heat loss, will give a T2=288K (15oC), assuming an exterior
surface emissivity (s) of 0.1. The exterior surface should be gold plated with an s = 0.01
to 0.02 at beginning of life (BOL). The assumed value of 0.1 assumes a greatly degraded
and dusty gold surface.
The heat loss through the MLI blanket must equal this heat loss. Heat transfer through an
MLI blanket can be estimated assuming radiative heat transfer through N layers (Wolfe
and Zissis, 1978):
q =  eff A2 (T14 - T24)
(3)
eff = [1/(1/1 + 1/2 - 1)][1/(N+1)]
(4)
Equations (3) and (4) can be rearranged to give the required eff for a given heat transfer
rate q,
eff = q/[ A2 (T14 - T24)]
(5)
and the required number of layers of emissivity 1=2 to give this eff:
N~
1 /  eff
(6)
2 / 1 - 1
Equation (5) is shown in Figure 10 for =0.05 and compared to aluminized Mylar MLI
test data. In general, one cannot obtain the theoretical value because of other heat losses
20
through the blanket, most notably from conduction through the spacers (netting) that
keeps the aluminized surfaces from touching each other. Test data seem to indicate that
three times as many layers are required to achieve the theoretical eff to compensate for
this effect at low effective emittances, but this depends greatly on spacer material as well
as how tightly the blanket is held together, as for example by sewing. Some have
attempted to back out an effective thermal conductivity based upon test data, but the
problem is more fundamentally radiative as given by Equation (4).
Figure 10. Effective Emittance for an MLI Blanket (Wolfe and Zissis, 1978)
Internal surfaces are likely to remain clean and retain their emittance. Vacuum deposited
aluminum (VDA) has an =0.03 at room temperature which will increase to 0.05 at
elevated temperatures. Vacuum deposited silver has an =0.02 at room temperature
which is likely to also increase to 0.05 at the higher temperature limits of silver. We will
use 0.05 as the worst case inner surface emissivity, and will use doubly silvered inner
layers. Since this MLI will be inside a vacuum jacket, the silver will not be exposed to
ambient prior to Earth launch and will not oxidize. Gold has the lowest room temperature
emittance of all metals and is not corrosive; however, its emittance increases much more
rapidly with temperature.
Table 1 shows the required number of MLI layers for a 5 watt heat loss rate and inner
surface emissivities of 0.05. For a 700oC heated bed, the required eff of ~4x10-4 is
difficult to achieve in practice, but possible in the laboratory. An allowable heat loss rate
of 10 watts results in a more believable eff as shown in Table 2. A 600oC heated bed
requires an eff ~10-3, and a 700oC heated bed requires an eff ~ 8 x 10-4. The latter can
theoretically be achieved using 33 layers, and we have degraded this number by a factor
of three, from 33 to 100 layers, to be more consistent with test data (Wolfe and Zissis,
21
1978; Larson and Wertz, 1992). Recall this 10% heat loss is a maximum just prior to
gashopper launch on Mars.
Table 1. Number of MLI Layers Required for 5 Watt Heat Loss
T1, oC (K)
N
3N
eff
-4
600 (873)
5.9 x 10
43
129
700 (973)
3.8 x 10-4
67
200
-4
800 (1073)
2.6 x 10
99
296
900 (1173)
1.8 x 10-4
141 424
Table 2. Number of MLI Layers Required for 10 Watt Heat Loss
T1, oC (K)
N
3N
eff
-3
600 (873)
1.2 x 10
22
66
700 (973)
7.7 x 10-4
33
100
-4
800 (1073)
5.2 x 10
49
148
900 (1173)
3.6 x 10-4
71
211
Figure 11 shows the heating of the pellet bed tank assuming a constant heat input of 100
watts over 10 hours, followed by 14.4 hours of no heat input at night. The 100 layer MLI
1000
o
Bed Temperature, C
800
600
400
100 watt input
12 liter vessel
13 kg boron pellets
200
7.7x10 MLI effective emittance
Radiating to deep space
0
-4
0
1
2
3
4
5
6
Days, Martian
7
8
9
10
Figure 11. Heating of the Pellet Bed by 100 watts on the Martian Surface
blanket is predicted to perform quite well, reaching 700oC around noon by the fifth day.
The temperature increases about 200 Co for the first two days while the boron specific
heat is lower; overnight cooling amounts to only a few degrees assuming a worst case of
radiating heat to deep space instead of the Martian surface. The pellet bed temperature
exceeds the 700oC goal during the fifth day. It cooled by 11 Co the previous night, and if
left for another night would cool by 16 Co. This system could easily exceed 900oC if left
for a sixth day, but this is the limit for the silver coatings on the MLI and a higher
temperature tank material would be required to go further.
22
Assuming the temperature distribution falls linearly inside the 10 watt heat loss 100 layer
MLI blanket, from 700oC to 15oC (288 K), it will then fall to 600oC after 15 layers,
400oC after 44 layers, and 150oC after 80 layers. We can use aluminum foil up to ~600oC
(melts at 660oC), Kapton up to ~400oC, and Mylar up to ~150oC. A thermal analyzer
network will be developed in Phase II which accounts for different material layers and
variable emittance versus temperature.
The weight of this high temperature MLI blanket will be minimized by using multiple
materials. The inner surface will be at 700oC, which will melt aluminum, Kapton, and
Mylar. However, silver melts at 961oC, so doubly silvered titanium foil or stainless steel
foil can be used for the first 15 higher temperature layers. This temperature is not hot
enough to justify using tantalum, molybdenum, or tungsten foils which have been used in
higher temperature refractory ovens; additionally, the refractories have much higher
emittance and are much heavier. The next 30 layers can use aluminum foil or silvered
aluminum foil. The next 36 layers can be doubly silvered Kapton film (1/3 mil), and the
final 20 layers can be silvered or aluminized Mylar film (1/4 mil). Although it may be
possible to manufacture thinner metal foils, we have assumed using off-the-shelf 12
micron thick (0.5 mil) stainless steel foil (or 25 micron thick titanium foil) which is then
silvered on both sides. The weight breakdown of this MLI blanket follows:
15 layers of 12 micron (1/2 mil) stainless steel foil, silvered
30 layers of 12 micron (1/2 mil) aluminum foil
36 layers of 8 micron (1/3 mil) doubly silvered Kapton film
20 layers of 6 micron (1/4 mil) doubly silvered Mylar film
Total mass of foil/film
392 g
264 g
99 g
46 g
801 g
In addition, we will need spacers or scrim cloth. The Dacron netting frequently used in
MLI blankets is a low temperature material and only suitable for separating the outer 20
layers of Mylar. This netting only weighs 5.4 g/m 2, slightly less than the Mylar film. For
the higher temperature layers, fiberglass scrim is a good choice, but must be baked out
prior to assembly to remove the sizing from it which can contaminate the blanket. A warp
x fill of 20x10 gives a heavy fiberglass scrim of 60 g/m2, which can readily be relaxed to
10x10 to give 30 g/m2, comparable to a commercially available Nomex scrim. A 5x5 can
be custom woven to give an acceptable 7.5 g/cm2. Eighty layers would weigh 154g in
addition to the 28g for 20 layers of Dacron netting. Total high temperature blanket
weight, including spacers, will be ~1 kg.
Other options will also be examined in Phase II. The zirconia powders used by
Thermoelectron in their high temperature refractory blankets will be examined in place of
the fiberglass netting. Astroquartz (fused silica) fabric can also be used, and has been
used on Venus’ Magellan spacecraft, as a nuclear hard white blanket on defense
spacecraft, and for laser hardened blankets. An Astroquartz scrim cloth has also been
used as spacer material in these programs, and will be examined for possible use. As with
fiberglass, its sizing must be removed in an air furnace prior to blanket assembly.
Astroquartz has higher temperature capability than fiberglass that is probably not
23
required in a gashopper application. An Astroquartz sewing thread can be used to sew the
blanket together.
Additional tank materials will also be examined. The baseline Ti-6Al-4V pellet bed
spherical tank was sized at 0.250 cm (0.100 inches) thick to hold 7.1 MPa (1000 psia) at a
stress level of 0.2 GPa (28 ksi). It’s mass is 3 kg. Although the tensile strength of this
material is 1.2 GPa (170 ksi) at room temperature, it is ~0.6 GPa (~80 ksi) at 538oC
(1000oF), and extrapolated is perhaps 0.35 GPa (50 ksi) at 700oC. This is pushing it for
this material, although it only needs to hold pressure at temperature for seconds during
the gashopper flight and is probably therefore acceptable. Alternative tank materials
include stainless steel, although the tank would weigh twice as much, carbon-carbon, or a
graphite metal matrix. The carbon-polymer composites used so widely today, such as
graphite epoxy or graphite polycyanate, do not have the temperature capabilities required
for the gashopper. It is even possible that some of the older metal matrix materials, such
as graphite-titanium, where there are alternating layers of carbon fabric and titanium foil,
may work perfectly for this application while retaining higher strength than titanium at
temperature. The titanium becomes the matrix for this metal matrix composite,
manufactured at high temperature and pressure, although not nearly as extreme as that for
carbon-carbon.
In conclusion, our analysis shows that lightweight gashopper engines are possible which
can be heated to 700 C and beyond with power levels that are realistic for photovoltaic
powered vehicles. It should be noted that if RTG’s are available, the heating job becomes
much easier. RTG’s are only 5% efficient, so that even a 20 W electric RTG puts out
about 400 W of thermal power, round the clock. Using the waste heat from such a
system, the example gashopper engine described above could be heated to 700 C in 11
hours.
Power system
With an estimated heat leak of 10 W, even a 40 W (average 10-hour daytime) power
supply would be able to heat the bed within 20 days. A larger power requirement is that
of the refrigerator, which needs 60 W to acquire 1 kg a day of CO2 from the Martian
atmosphere by freezing. Since these activities do not need to be done simultaneously, and
other activity such as communication (which is assumed to be done through an orbiter) is
occasional and low power in any case, the power requirement of the gashopper is
estimated to be 60 W average daytime, and perhaps 5 W at night. Assuming that the 14.4
hours of night power has to be stored during the day, and taking into account battery
losses, we find a need to generate an average of 70 W for ten hours per day (i.e. 700 Whours per sol.). The solar flux on Mars is about 500 W/m2, so a 15% efficient
photovoltaic panel that is normal to the sunlight can generate 75 W/m2. Such a 1 m2
would have to track the sun, and would have an estimated mass of 3 kg. the tracking
mechanism would have an estimated mass of 1 kg, bringing the total to 4 kg.
An alternative solution is to use ultra-lightweight fixed panels that do not track. ABLE
Engineering has developed a technology called ultraflex, with a mass of 1.4 kg/m2, and a
power generation capability at Mars of 64 W/m2. . Employed in a fixed configuration,
24
these would have their power output degraded by a factor of 2 due to cosine losses. Three
square meters of this type of panel would mass 4.2 kg and provide 96 watts average 12hour daytime power, or 1152 watt-hours. This gives 64% margin against the 700 W-hour
daily power requirement. If Martina weather conditions caused the solar flux to drop so
much that this margin was insufficient, however, it would not mean mission failure.
Rather, since nearly all the power is needed for activities such as propellant acquisition of
engine heating that can be deferred if necessary, a drop in solar flux simply means that
the next hop will be postponed until more power is available.
The tracking system and the non-tracking ultraflex system weigh about the same.
However, the non-tracking ultraflex system is simpler to implement and has been brought
to an advanced state of development as the chosen power system for the Mars 2001
lander program. Some further development would be needed to provide the system with
suitable mechanisms to be restowed after each deployment in preparation for flight, but
this does not appear to be a fundamental problem. We therefore baseline the ultraflex
system for the gashopper.
Autogenous Pressurization
Part of the attractiveness of the gashopper propulsion system owes to the fact that the
CO2 propellant can be made to pressurize itself, so that no additional tank of pressurant is
required. Of course, as the tank blows itself down, CO2 gas must be produced from the
liquid reservoir, and the need to provide the heat of vaporization for this gas causes the
liquid temperature to drop. As the liquid temperature drops, its vapor pressure also
declines, and this causes the tank pressure to fall as well. If the tank pressure is being
used to pressurize the engine directly (i.e. without any regulation) this would also cause
thrust levels to fall in proportion.
In order to model these effects, a computer program was written which simulated the
blowdown of the tank in volumetric increments of 0.1% of volume at a time (i.e. 1000
steps to empty the tank). In fig 12, we show the results of this program, which in this case
assumed a tank initially completely filled with liquid at 27 C and a pressure of 62 bar.
The tank was then made to blow down until it was empty of liquid.
25
Temperature/Pressure/Liquid Volume Fraction vs. Mass Fraction Removed from the Tank
100
Temperature (C) / Pressure (bar) / Liquid Fraction (%)
90
80
70
60
50
40
30
20
10
0
0
0.1
0.2
0.3
0.4
0.5
0.6
0.7
0.8
0.9
1
Mass Fraction of CO2 removed from the tank
Pressure (bar)
Temperature (C)
Liquid Fraction
Fig. 12. Autogenous Blowdown of Gashopper Tank
It can be seen that in the course of this blowdown, the tank pressure fell from 62 Bar (911
psi) to 34 bar (500 psi). Thus without regulation final thrust, used to land, would be about
55% of the thrust available on takeoff.
There are several ways the issues raised by these results can be dealt with.
(a) They can be accepted as is, and used as the basis of performance of the engine. Since
the vehicle is lighter at landing, less thrust is needed in any case. If necessary to provide
adequate landing thrust, takeoff thrust can simply be over designed. This is not hard to
do. Since the pellet bed can transfer heat at almost any power level, increasing the takeoff
thrust simply requires making the throat a bit wider.
(b) Thrust can be kept constant throughout the burn by regulating the fluid released to the
engine down from whatever the tank pressure is to a constant level, say 450 psi. In that
case, the engine would never know that the tank pressure had fallen at all. In this case, the
throat and nozzle area would have to be expanded by a factor of 911/450 to produce the
same takeoff thrust as the previous example, but the thrust would be constant throughout
the burn (unless we chose to throttle further.) As an additional benefit, the engine vessel
would only have to be rated to 450 psi, and thus could be built lighter.
26
(c.) As a third approach, tank pressure could be maintained by employing a small subset
of the net flow to take heat from the pellet bed or engine case, and recycle it back into the
propellant tank to keep temperature and pressure there high. This is the highest
performance option, but complicates the system somewhat.
All of the above options are feasible. A selection of the optimal choice will be made on
the basis of further analysis performed during Phase II. For our purposes here, however,
the key thing to note is the conclusion that autogenous pressurization will work for the
gashopper, and thus no engine turbines, pumps, or ancillary pressuri9zzation systems are
needed.
Conceptual Gashopper Design Point
Based on the above analysis, we propose the following as a design point for a first
generation gashopper.
Table 3. Near-Term Gashopper
CO2 propellant
30 kg
CO2 tanks
3 kg (30 liters, graphite overwrapped, rated to 1500 psi at 32 C.)
Beryllium Bed
15.5 kg (12 liters, 70% packing fraction, operating at 1000 K)
Engine Case
3 kg (titanium)
Engine insulation
1 kg (multifoils)
Other engine parts
1kg vacuum jacket, various fittings.)
Refrigerator
3 kg (requires 60 W to acquire 1 kg/day propellant.)
Solar panels
4.5 kg (ABLE ultraflex fixed array, 45 W peak per kg on Mars)
Flight Avionics
4 kg
Structure
10 kg
Payload
10 kg
Total Mass
85 kg wet / 55 kg dry
Mass ratio =1.545
With an estimated specific impulse of 127 s, the vehicle above would be able to generate
a delta V of 541 m/s. Assuming 15% gravity losses on both takeoff and landing, and
taking no credit for aerodynamic effects that would actually help decelerate the vehicle,
this system would have a hopping range of 14.3 km. If we assume that aerodynamic
effects could be used to slow the vehicle from a maximum speed of 300 m/s down to 160
m/s, then 65% of the available delta-V could be used for ascent, and only 35% would be
needed to land. In that case, vehicle range would increase to 24 km.
Using the refrigerator to acquire CO2 at a rate of 1 kg per day would allow the vehicle to
hop at a rate of once per 30 days. Traveling 24 km per hop, this would give it an average
traveling speed of 800 m/s, approximately 8 times faster than that anticipated for the
Athena rover. Moreover the mobility of the system would be unimpaired by intervening
chasms, canyons, boulder fields, craters, mountains, ridges, or any other terrain obstacle.
In addition, it would be able to image or perform other remote sensing measurements as it
flew, thereby providing high resolution aerial context information to support surface
exploration wherever it chose to go.
27
In short, a gashopper such as that outlined above would add dramatic new capabilities to
our available set of tools for Mars exploration.
And with an engine temperature limited to 1000 K, it’s just the beginning.
Issues and Observations
It will be observed that the gashopper’s mode of travel requires repeated takeoffs and
landings on Mars. It is no secret that our ability to perform the latter reliably has recently
become a matter of considerable concern. With respect to this issue, the following can be
stated;
(a) As far as the propulsion system itself is concerned, the hot pellet bed gashopper is one
of the simplest rocket propulsion systems ever designed. It requires only one propellant,
no separate pumps or pressurants, and cannot create explosion hazards through propellant
leaks (i.e. Mars Observer.) It has no moving parts, cannot overheat, and can be readily
throttled with a single valve. Thus it is likely to be about as reliable as a rocket propulsion
system can possibly be.
(b) This then leaves the issue of achieving a safe landing despite the uncertainties
associated with the rough Martian terrain. In order to deal with this, an automated
onboard hazard avoidance system is necessary. We do not propose to develop such
systems in the course of an SBIR Phase II. However, it is recognized in the Mars program
that the development of such systems is necessary if we are ever to perform such highcost/high value missions as the Mars Sample Return, and a significant program has been
proposed at Jet Propulsion Lab to develop precisely such technology. That being the case,
the gashopper not only benefits from such a program, it itself will enormously benefit
that program by providing it with a test vehicle on Mars.
The point needs to be emphasized. With multiple spacecraft involved, the Mars Sample
Return will be a highly complex and expensive mission, possibly involving international
collaboration with France or other partners. We certainly would not want to test out our
automated hazard avoidance system on that mission for the first time. Lower cost Mars
landing missions could be used to test the hazard avoidance system prior to the MSR
mission, but how many times? Once? Twice? We only fly Mars landers at most once
every two years. That’s not going to give us much of a data base. But if we send a
gashopper to Mars, we can test the hazard avoidance system repeatedly across a variety
of terrain types, weather, and illumination conditions. Thus, in addition to its exploration
value, the gashopper provides the ideal engineering test bed that is critically necessary to
perfect the automated surface hazard avoidance system that the lander for the MSR
mission will absolutely require.
Phase I Experimental Program
The goals of the Phase I experimental program were to look at engineering issues
associated with the actual hardware gashopper hardware, and to test the hardware to gain
28
experience and gather performance data from the hardware testing.
will flow directly into future gashopper activities.
Experience gained
The Unheated CO2 Rocket
The simplest experiment to perform is that of expanding unheated liquid CO2 from a
storage tank out a nozzle. This concept avoids the complexity of adding heat to the CO2
flow. A schematic of the experimental concept is shown below in Figure 13. CO2 is
stored in a high-pressure cylinder. Liquid and gas are in equilibrium; the pressure in the
bottle is fixed by the temperature and the saturated state of the CO2. When the remotely
actuated valve V1 is opened, the gas pressure pushes liquid out through the “dip tube” and
into the system plumbing. A turbine flow meter measures the volumetric flow rate of the
CO2; with the density of the CO2 fixed by the upstream temperature, the mass flow rate
can be found. The throttling valve reduces the pressure of the CO2 stream from the high
initial pressure (associated with saturation conditions at the starting temperature in the
tank) down to the desired chamber pressure. Note that, without heating, the temperature
downstream of this valve will be much lower than that upstream of the valve. A
thermocouple and pressure transducer measure the temperature and pressure,
respectively, in the thrust chamber.
T1
CO2
Tank
F
M
V1
T2
P1
Rocket
Nozzle
V2 P2
Symbols
T1 Temperature upstream of flow meter
P1 Pressure upstream of flow meter
FM Flow meter
V1 Remotely actuated valve
V2 Throttling valve
T2 Temperature in thrust chamber
P2 Pressure in thrust chamber
Figure 13. Schematic of Unheated CO2 Rocket Experiment.
Construction of the Laboratory Demonstrator
Following the design schematic shown in Figure 13, construction of a laboratory test
stand and rocket nozzle was completed. The specifications of the rocket nozzle are
shown in Figure 14. A photograph of the test stand as built is shown in Figure 15, and
Figure 16 shows a closeup view of the rocket nozzle as installed on the stand. Although
this test stand was used initially to test the cold-gas version of the CO2 gashopper rocket,
the test stand was also reconfigured to test the electrically heated and hot-bed gashopper
rockets as well.
29
Figure 14. Specifications of Cold Gashopper Thrust Chamber/Nozzle.
Figure 15 (left). Mars Gashopper Cold-Gas Demonstrator Test Stand.
Figure 16 (right). Closeup External View of Unheated Gashopper Nozzle.
A short series of experiments was performed to characterize the system for an
assortment of mass flows. Figure 17 is a photograph of the test stand during one such
experiment, showing the exhaust plume. Figures 18 and 19 show the thrust, calculated
specific impulse, system pressures, and mass flow rates. These early tests used short run
times of approximately 5 seconds. Note that the calculated values for specific impulse
are less accurate at the start of propellant flow, since finite time is required for the turbine
flow meter to spin up. The results of these experiments are summarized in Figure 20,
which shows the measured pressures in the thrust chamber, the generated thrusts, the
rates of flow of CO2, and the calculated specific impulses for various settings of the
metering valve (which throttles the CO2 flow). Note that, as the metering valve was
opened (increasing the valve coefficient Cv), the flow rate, chamber pressure, and thrust
all increased, but, contrary to expectations, the specific impulse decreased. This result
30
may indicate that some of the CO2 was expelled as liquid, rather than as a solid, so that
some enthalpy was unavailable to be converted into kinetic energy. In addition, the flow
rates were less than expected, resulting in lower chamber pressures and thrusts than
desired.
20
50
18
45
16
40
14
35
12
30
10
25
8
20
Load Cell
Specific Impulse
6
Specific Impulse (s)
Thrust (lbf)
Figure 17. Unheated Gashopper Experiment. The nozzle is at the center of the picture
and the plume extends to the right.
15
4
10
2
5
0
0
0
5
10
15
20
25
30
35
40
45
Time (s)
Figure 18. Thrust and Specific Impulse for Unheated CO2 Gashopper Thruster
Demonstrator, 7 Turns Open on Throttle Valve.
31
900
0.3
800
0.25
700
0.2
P tank outlet
P chamber
DP
Flow Meter
500
400
300
0.15
Flow Rate (kg/s)
Pressure (psig)
600
0.1
200
0.05
100
0
0
20
22
24
26
28
30
32
34
Time (s)
Figure 19. System Pressures and Flow Rates for Unheated CO2 Gashopper Thruster
Demonstrator, 7 Turns Open on Throttle Valve.
25
1
0.8
Peak Pchamb (bar)
Peak Thrust (lbf)
0.7
Peak Specific Impulse (s)
Peak Mass Flow Rate (kg/s)
15
0.6
0.5
10
0.4
0.3
5
Mass Flow Rate (kg/s)
Pressure (bar), Thrust (lbf), Spec. Impulse (s)
0.9
20
0.2
0.1
0
0
0
0.1
0.2
0.3
0.4
0.5
0.6
0.7
0.8
Valve Coefficient (Cv)
Figure 20. Thrust Chamber Pressure, Thrust, Specific Impulse, and Mass Flow Rate for
Various Metering Valve Settings on the Unheated Gashopper Test Stand.
Based upon these results, Pioneer theorized that the plumbing may be too restrictive for
the vapor/liquid mixture downstream of the metering valve, and therefore decided to
replace it with a larger metering valve. Typical results from this setup are shown below
in Figures 21 and 22, with the metering valve open one full turn (approximate Cv = 1.2).
As expected, substantially greater flow was measured. The thrust did not increase
proportionately, resulting in a lower measured Isp. It should be noted however, that in
this set up the metering valve is downstream of the flow meter. When it is opened wider,
it causes the static pressure in the line at the position of the flow meter to drop. This can
cause partial boiling, resulting in two-phase flow, which in turn will cause the flow meter
to give an artificially high reading (it reads volumetric flow.) If the flow meter
exaggerates the mass flow, an inaccurately low specific impulse will be calculated, since
specific impulse is thrust/mass flow. In all probability, the specific impulse of this
thruster was in the 25 to 28 s range, just as in the previous tests. This conclusion is
supported not only by the previous test data, but by the known enthalpy of the fluid
which is what determines the specific impulse in any case.
32
25
50
Load Cell
45
Specific Impulse
20
40
15
30
25
10
20
Specific Impulse (s)
Thrust (lbf)
35
15
5
10
5
0
0
20
22
24
26
28
30
32
34
36
38
40
Time (s)
800
0.8
700
0.7
600
0.6
500
0.5
400
0.4
300
0.3
P tank outlet
P chamber
DP
Flow Meter
200
100
Flow Rate (kg/s)
Pressure (psig)
Figure 21. Unheated Gashopper Experiment. Thrust and Specific Impulse for Unheated
CO2 Gashopper Thruster Demonstrator, 1 Turn Open on New Throttle Valve.
0.2
0.1
0
0
20
22
24
26
28
30
32
34
36
38
40
Time
Figure 22. System Pressures and Flow Rates for Unheated CO2 Gashopper Thruster
Demonstrator, 1 Turn Open on New Throttle Valve.
33
Electric Heated Thruster Test
An experiment was devised to preheat liquid CO2 near the critical point in a propellant
tank, then extract the liquid through a heated tube propellant feed line which converts the
CO2 to gas for expulsion through a rocket thruster. The experimental setup is shown in
Figure 23. The black aluminum scuba tank was preheated using heat tape. After passing
through a flow valve and flowmeter, the liquid flowed through a 1.5 meter, 0.635 cm OD
(0.250 inch), 0.051 cm (0.020 inch) wall 316SS heated tube, before entering the thrust
chamber. The tube was heated by passing battery current through the pipe via large
copper electrodes separated by 1 meter, and insulated by fiberglass covered with black
polyethylene foam pipe insulation. Three car batteries in series provided 29.6v across the
pipe’s 85 m resistance under load. The current draw was 344 amps for a total heating
rate of 10.2 kW, which decreased slightly during the 15 second firing. Pressure
transducers and thermocouples were fitted into the thrust chamber.
Figure 23. Electrically Heated Tube Setup for CO2 Thruster
Figure 24 shows a thruster firing two days earlier, on May 9, when the propellant tank
was not properly preheated, and the resultant exhaust plume was primarily liquid. The
batteries were sized to only provide enough heat to convert CO2(l) at 30oC to CO2(g) at
41oC prior to entering the thrust chamber. On the morning of May 11, we were able to get
a good liquid load, and reasonable preheating of the propellant tank, but were not
instrumented to get a good liquid flow temperature in this Phase I experiment.
34
Figure 24. Firing of the Electrically Heated Tube on May 9 (Insufficient Preheat)
The measured thrust, inferred Isp, inferred flowrate, temperatures and pressures are
shown in Figures 25-28 for the May 11 firing. The thrust was 10 lbf higher than expected,
probably because of uncertainties in predicted liquid densities. The flowmeter measures
volumetric flowrate calibrated to liquid flowrate, and because we did not have an
accurate flow temperature measurement (which will be properly instrumented in Phase
II), we have a large uncertainty in flow density and hence mass flowrate shown in Figure
26. The CO2 liquid density varies dramatically near the critical point of 31.1oC as shown
in Figure 28, and we will assume for now that the liquid flow was near the critical point.
Doing so results in a specific impulse of 53 seconds near the end of the firing as shown in
Figure 29, which is close to the theoretical value of 56 seconds predicted.
30
Heated Tube CO Thruster
2
Thrust, lb
f
25
29.6v, 344a, 10.2 kW
5/11/2000
20
15
10
5
0
6020
6025
6030
6035
Time, s
Figure 25. Thrust Profile of the Electrically Heated Thruster
35
6040
6045
0.5
Flow @ critical temperature
Flow @ room temperature
Flowrate, kg/s
0.4
0.3
0.2
0.1
0
6020
6025
6030
6035
6040
6045
Time, s
Figure 26. Inferred Flowrate of the Electrically Heated Thruster
Liquid Density, kg/m
3
1000
800
600
400
Liquid is under pressure,
increasing to critical pressure of 72.45 atm
at critical temperature of 31.11 C.
200
0
0
5
10
15
20
Liquid Temperature,
25
o
30
35
C
Figure 27. Density of Liquid CO2 Near the Critical Point
ISP, seconds, ground level
60
50
40
30
Heated Tube CO Thruster
20
10
0
6020
2
29.6v, 344a, 10.2 kW
5/11/2000
Assumes tank flow @ 31 C
6025
6030
6035
6040
6045
Time, s
Figure 28. Inferred Specific Impulse of Electrically Heated Thruster
Two of the measured temperatures are shown in Figure 29. The thermocouple on the
Electrically heated tube was taped directly at the end of it under insulation, prior to the
electrode closest to the thrust chamber shown in Figure 23. It reaches nearly 130oC until
the steady state flow cools it to approximately 30oC. The chamber temperature
thermocouple was clearly not inserted far enough into the thrust chamber to obtain a good
reading.
36
140
Temperatures, deg C
120
Chamber Temperature
Electrocuted Tube
100
80
60
40
20
0
-20
6020
6025
6030
6035
6040
6045
Time, s
Figure 29. Temperature Profiles (Chamber Temperature is Probably Not Representative)
Figure 30 shows the supply tank pressure and the thrust chamber pressure. The black
propellant tank was preheated over an hour until pressure read 1210 psig. Unfortunately,
there was residual air in the tank during filling, so this cannot be correlated directly to
temperature. Future work requires more complete instrumentation and development of
precise CO2 liquid fill techniques, including propellant weight measurement. The
chamber pressure is well behaved and approximately 130-140 psi below the tank
pressure.
1400
Tank Pressure
Chamber Pressure
1200
Pr 1000
es
sur 800
es,
psi 600
g
400
200
0
6020
6025
6030
6035
6040
6045
Time, s
Figure 29. Tank and Chamber Pressures of the Electrically Heated Thruster
Water Heated Thruster Test
An alternative to using electricity to directly heat the carbon dioxide flow tube is to have
a heated medium that transfers heat to the flowing stream. The first of these
configurations attempted was a water heated thruster. In this configuration, a one liter
steel vessel was filled with approximately 750 ml of deionized water, wrapped with heat
tape, and covered with insulation. The 3/8” carbon dioxide flow tube penetrated directly
through the middle of the steel water vessel. The water vessel had both temperature and
pressure monitors to record the state of the water. After the water was heated above one
bar, a valve on the top of the vessel was opened to allow trapped air to escape and ensure
37
that the vessel was filled with a pure water/steam mixture. The reason a two phase
mixture was used was to enhance the heat transfer; although water holds significantly
more energy per unit volume than steam, condensation of steam on the cold CO2 flow
tube, and natural convection between the steam and water phases, could allow
significantly more energy to be transferred to the carbon dioxide per unit time.
Figure 31: This is the water heated CO2 thruster undergoing testing. The insulated and
foil wrapped water vessel is at the left of the photograph. Carbon dioxide flowed directly
from the water vessel through a length of flexline to the engine.
Several runs were made, with the steam at ambient pressure, 500 psig, 750 psig, 1000
psig, and 1500 psig. Unfortunately, in every case when the carbon dioxide started to
flow, the chamber temperature dropped almost immediately to 0 degrees C or slightly
below. This indicates that the heat transfer rate from the water into the carbon dioxide
flow tube was insufficient, and ice actually began to build up on the tube exterior, fixing
the carbon dioxide enthalpy in the thruster. With this fixed enthalpy, the performance of
the water heated carbon dioxide thruster was limited, with specific impulses measured in
the range of 35 s. Although techniques for enhancing heat transfer could be examined,
other options for heating the carbon dioxide appeared more promising and the water
heated thruster was dropped.
Hot bed Engines with 40 lbf thruster
To experimentally evaluate and test hot bed options, a one-liter stainless steel sample
cylinder was prepared as a test article. The input of the bottle is ½” NPT threaded input,
and the output is ½” NPT, which fed to the thruster section. A new thruster was
designed for 40lbf (178 N) thrust when using CO2 at 44.8 bar (650 psi) input and
exhausting into Denver area ambient atmospheric pressure, 0.82 bar (12 psi). The throat
38
diameter is 5.79 mm and the exit diameter is 14.8 mm. The divergence angle is 15. The
hot bed container is heated with two 627 W heat tapes, and insulated with about 5 cm of
high temperature and fiberglass insulation, then over-wrapped with aluminum foil.
Hot bed Engine Vertical Test Stand and Instrumentation:
The hot bed engine bottle has an internal and an external thermocouple, and there is a
thermocouple in the chamber (i.e. the area immediately upstream of the thruster throat).
There are pressure transducers for measuring chamber pressure and CO2 tank pressure, as
well as a flow meter and a load cell. The load cell was calibrated with known-mass
weights to ensure accuracy. Accurate and reproducible measurements of load cell,
pressure transducers, thermocouples were read by the LabVIEW data acquisition system.
The test stand was built up of 2”x4” lumber on a 1m x 2m folding table. Instrumentation
wires were routed off the test rig at the fulcrum point so as to not impart false loads into
the test stand.
A ‘protoflight’ version of a gashopper vehicle was also built. This system utilizes a
fiberwound bottle for CO2, stainless steel plumbing, an electrically operated propellant
valve, and a heavy stainless steel sample bottle to hold the hot media. A customfabricated stainless steel screen at the bottom of the bottle ensures that the hot particle
bed media is not discharged through the rocket thruster during operation.
39
Figure 32: Working on the hot bed vertical test stand. The fiberwound CO2 bottle is at
top, the yellow remote controlled valve in between, and the foil wrapped hot bed engine
at bottom.
Figure 33: Firing the hot alumina particle engine on the test stand. The visible CO 2
exhaust shows we have utilized nearly all the heat capacity of the bed – hot exhaust is
completely invisible.
40
Figure 34: The protoflight version of the gashopper chained to the floor for tethered static
tests.
Figure 35: The protoflight version of the gas hopper undergoing flight/hover test. The
initial mass of the hopper is 51 lbs, and the engine thrust is rated 40 lbf. A 20 lbf. assist
from a counter balance allowed flight and provided upward guidance to avoid upsetting
the gashopper.
Hotbed test series:
The first candidate hotbed material was alumina pellets, commercially available in 3mm
x 8 mm cylindrical form. The alumina pellets were measured to have an aggregate
specific density of 0.48. Thus, the one liter hoke bottle was filled with 0.48 kg of
alumina. Using 30 C as our reference temperature, we can load 290 kJ into the alumina
pellets. This corresponds to the energy required to heat 0.31 kg CO2 liquid to 700C.
Time averaged Isp measurements were only about 35 seconds with the alumina hot bed.
The best explanation is that the alumina does not transfer heat to the CO2 fast enough to
be useful for the hotbed rocket engine. Figure 33 shows a test firing of the alumina filled
engine; Figure 36 shows the thrust and specific impulse achieved during an alumina
hotbed test run.
41
Isp & Thrust vs. Time Run 33c Alumina Pellets
70
Isp (sec) and Thrust (lbf)
60
50
40
Thrust
Isp
30
20
10
0
172
174
176
178
180
182
184
186
Time (sec)
Figure 36: Thrust and specific impulse of the alumina hotbed experiment. The alumina
did not average significantly more than about 35 seconds Isp during the major portion of
the firing.
To improve on the performance of the engine, the alumina pellets were replaced with
graphite rods. Commercially available rods with 6 mm diameter were procured and cut
to appropriate lengths to fit into the hotbed bottle. We were able to load 0.65 kg of
graphite rods into the hotbed bottle. Graphite has a wide range of reported specific heats,
but a conservative estimate is 754 J/(kg K). Using 30 C as our reference temperature, we
can load 328kJ into this hotbed media at 700 C. We achieved only about 50 seconds Isp
when using the graphite as the hotbed media. This was less than expected, most likely
because of heat transfer rate limitations. Additionally, it is theorized that the shape of the
graphite rods allows the flow to channel. The channeling allows some cool, unheated
CO2 to ‘sneak past’ the graphite rod media without heating to desired operating
temperature, thus lowering the average temperature of the exhaust. Finally, there is the
possibility of reaction between graphite and CO2 at elevated temperatures via the
Boudouard reaction, which would endothermically produce carbon monoxide and lower
the temperature of the exhaust gas. However, inspection of the graphite rods after use
revealed no degradation after being heated to over 650 C and exposed to high
temperature gas followed by quench with liquid CO2.
42
Figure 37: Graphite rods after being heated to over 650C and exposed to high
temperature CO2 during engine testing. No degradation was apparent.
Copper Coated Steel BB’s
In the next experiment, copper coated steel BB’s were used. We loaded 4.84 kg of BB’s
into the hotbed container. When heated to 700 C, 1459 kJ is available as heat for CO 2
propulsion. Good correlation between the load cell thrust data, and thrust deduced from
chamber pressure readings was attained during test runs. Results from the test run are
shown in Figure 38. Specific impulse attained with this set up was over 80 seconds,
undoubtedly due to the good heat transfer characteristics between the 4.5 mm diameter
copper coated steel BB’s and the working fluid. Chamber pressures as high as 54 bar
(780 psi) have been recorded with the BB’s and enable the engine to operate at 20% more
thrust than was initially planned during the design phase. There was uncertainty about
how much pressure drop would be manifested while flowing CO2 through different media
types in the hot bed. In this case, the pressure drop is lower than estimated, and the CO2
exhaust is underexpanded at the exit.
This indicates a loss in efficiency, but is
acceptable for this phase of experimentation. The BB-loaded hotbed showed the highest
performance, and despite its high weight, was selected for the next round of activity:
flight test. Although steel BB’s are too heavy for consideration for a Mars system, they
simulate well the performance of boron pellets that have similar thermal conductivity,
volumetric heat capacity, thermal diffusivity, but weigh only 1/4th as much.
43
C O 2 G a s H o pp e r R o c k e t E n g in e P e rf or m a n c e
B B h ot be d te s t
9 0.0
8 0.0
Spec ific Im pulse (se cond s)
and T hrust (lbf)
7 0.0
6 0.0
5 0.0
Is p ( te m p b a se d )
T h r u s t ( p r e s su re b a s e d )
4 0.0
3 0.0
2 0.0
1 0.0
0.0
0
2
4
6
8
10
12
T im e ( se c)
Figure 38: The results of the copper coated steel BB hotbed test. Low pressure drop
improved thrust over the nominal 40 lbf, and Isp reached a peak above 80 seconds.
Flight Test Series:
The hot particle bed CO2 propulsion unit showed enough utility and compactness to be
considered for integration into a crude protoflight system. Our first action was to remove
the flowmeter and pressure transducers from the test-configured propulsion set up, then
to attach three aluminum legs to the fiber wound bottle to approximate a lander
configuration. The all-up weight of the hardware including liquid CO2 propellant is 23
kg (51 lbs). With a little over 50 lbs (213 N) thrust available, we needed a little extra lift
to conduct a flight test. We counterbalanced the protoflight gas hopper with a 20 lb. lift
bungee cord. The hopper was constrained to the floor with three chains and flight to ~0.3
m altitude was accomplished. Due to the stabilizing effect of the bungee cord lift, the
flight was stable and lasted about 6 seconds, limited by propellant depletion. Figure 35
shows this tethered static test.
A second flight was made with a longer floor tether - about 1.5 meters – and a different
counterbalance scheme. Unfortunately, the counterbalance failed and the gas hopper
flailed about the lab and broke a leg. It was rebuilt in a short time and prepared for the
non-constrained flight test.
The non-constrained flight test featured our proto-flight unit with a 20 lb. lift helium
weather balloon, and is shown in Figure 39. The electrically operated valve is rather
44
slow (1.5 seconds lag time) to operate and the radio remote control that we added to it
was an additional small delay in control time. Our plan was for a 7 second boost,
followed by apogee and a hopefully well-timed retro thrust for a DC-X style soft landing.
The boost and ascent went well, with peak altitude estimated at over 40 meters. The
vehicle was then allowed to descend, and the engine was successfully restarted to avoid
an obstacle in the landing zone before landing. The vehicle then ascended to about 10
meters, and then was allowed to descend again, touching down at low descent velocity.
However, because of cross-wind induced horizontal motion it dragged and tipped on
impact, which again damaged one of the landing legs. Irrespective of the unsatisfactory
landings, the fact remains that the gashopper has been built, demonstrated on a test stand
and successfully integrated into a simple flying vehicle and flown under Phase I
activities. The fact that the engine is so simple that it could be brought from the test
stand to radio-controlled flight within so short a time speaks very well for the likelihood
that simple, reliable systems can be developed for Mars missions.
45
Figure 39: The balloon assisted gashopper flight. Left: CO2 gashopper before liftoff.
Fully loaded, the gashopper weighs 51 lbs, and requires a little help from a helium
weather balloon providing 20lbs of lift. Center: Liftoff! The gashopper ascends rapidly
on about 50 lbs thrust. Right: The CO2 gashopper reached an estimated 40 meter apogee,
then started downward towards a planned DC-X style landing.
46
Conclusions
1.The Gashopper can provide surface mobility of a type that is not obtainable with
any extant Mars exploration system
2. The Gashopper can travel at an average speed two orders of magnitude faster
than existing rovers and one order of magnitude faster than projected rovers.
3. The Gashopper is not constrained by the terrain barriers limiting rover access.
It can cross chasms, canyons, boulder fields, craters, and fly up mountains. Such travel is
impossible for rovers..
4.A gashopper can do aerial photography and other kinds of remote sensing when
it flies. It thus can create aerial context photography for itself at each new site. In
contrast, rovers lose context when they leave their landing areas.
5. In contrast to balloons, gashoppers can be directed to fly precisely to a
designated location. They can land repeatedly. In contrast, successful balloon landings at
more than one site are problematical, and directing them to a succession of particular
sites is impossible.
6. The gashopper provides an ideal tool for perfecting the automated hazard
avoidance systems that will be needed for the Mars Sample Return mission. No other
system can offer as much landing experience for as little cost or time.
7. Several gashopper concepts may be possible. The Cold gashopper is the
simplest. The hot bed gashopper offers the best performance.
8. Hot bed gashoppers can be built with existing materials that can achieve single
hop ranges on the order of 20 km, with repeated hops every 30 days.
9. A pellet bed consisting of beryllium pellets is a leading candidate for a hot bed
gashopper engine. Such pellets can be well simulated in developmental engines using
brass.
10. A hot pellet bed gashopper is a simple engine which can be built without
stressing its materials. It is thus likely to be very reliable..
11. Hot pellet bed gashopper engines can be readily built with thrust and Isp
performance adequate to meet the requirements of robotic Mars exploration.
12. Hot pellet bed gashoppers can be autogenously pressurized. Such vehicles can
therefore be built with a single propellant tank. No pressurants or pumps are needed.
13. Hot pellet bed gashopper engines can be heated to flight temperature with
power sources that are realistic for robotic Mars exploration vehicles employing
photovoltaic power.
14. The simplicity and reliability of hot pellet bed engines have potentially
important commercial applications
15. There is thus a substantial foundation to show that gashopper technology is
feasible and offers extensive potential benefits to both NASA and the nation. Further
work to develop the technology is clearly warranted.
47
References
Blackledge, M., D. G. Pelaccio, J. W. Lewis, and F. J. Perry (1993), “Application of SpaceBased Interceptor Propulsion Technology to Satellites and Interplanetary Vehicles,” AIAA 932119, 29th AIAA/SAE/ASME/ ASEE Joint Propulsion Conference and Exhibit, Monterey, CA,
June 28-30, 1993.
Pettit, D. R. (1989), “A Carbon Dioxide Powered Rocket for Use on Mars,” AAS 87-264,
The Case for Mars III: Strategies for Exploration – Technical, Volume 75, Science and
Technology Series, American Astronautical Society Publication, 1989.
Powell, J., et al. (1988), “Particle Bed Reactor Orbit Transfer Vehicle Concept,” AFAL-TR88-014, Air Force Astronautics Laboratory Report, 1988.
Rapp, D., P. B. Karlmann, D. L. Clark, and C. M. Carr (1997), “Absorption Compressor for
Acquisition and Compression of Atmospheric CO2 on Mars,” AIAA 97-2763, 33rd
AIAA/SAE/ASME/ ASEE Joint Propulsion Conference and Exhibit, Seattle, WA, July 6-9, 1997.
Shafirovitch, E., A. Shiryaev, and U. Goldshleger, “Magnesium and Carbon Dioxide: A
Rocket Propellant for Mars Missions,” Journal of Propulsion and Power, Vol 9, No. 2, MarchApril 1993.
Shafirovitch, E. and U. Goldshleger, “Mars Multi-Sample Return Mission,” Journal of the
British Interplanetary Society, Vol. 48, No. 7, July 1995.
Wiley Larson and James Wertz, eds., (1992), Space Mission Analysis and Design,
Microcosm, Inc., Torrance, CA.
William Wolfe and George Zissis, eds., (1978), The Infrared Handbook, Environmental
Research Institute of Michigan.
R. Zubrin, "Nuclear Thermal Rockets Using Indigenous Martian Propellants," AIAA-892768, AIAA/ASME 25th Joint Propulsion Conference, Monterey, CA, July, 1989..
Zubrin, R., S, Price, L. Mason, and L. Clark (1994), “Report on the Construction and
Operation of a Mars In-Situ Propellant Production Plant,” AIAA-94-2844, 30th AIAA Joint
Propulsion Conference, Indianapolis, IN, June, 1994.
Zubrin, R., S, Price, L. Mason, and L. Clark (1995a), “An End to End Demonstration of Mars
In-Situ Propellant Production,” AIAA-95-2798, 31st AIAA/ASME Joint Propulsion Conference,
San Diego, CA, July 10-12, 1995.
Zubrin, R., “Diborane/ CO2 Rockets for Use in Mars Ascent Vehicle,” Journal of the British
Interplanetary Society, Vol. 48, No. 9, September 1995.
Zubrin R. and B. Frankie, “Mars Atmospheric Carbon Dioxide Freezer,” Final Report on
NASA Contract NAS9-99020, Presented to NASA JSC June 14, 1999,
48
Form Approved
OMB No. 0704-0188
REPORT DOCUMENTATION PAGE
Public reporting burden for this collection of information is estimated to average 1 hour per response, including the time for reviewing instructions, searching existing data sources, gathering
and maintaining the data needed, and completing and reviewing the collection of information. Send comments regarding this burden estimate or any other aspect of this collection of
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1. AGENCY USE ONLY (Leave blank)
4.
2. REPORT DATE
June 8, 2000
3. REPORT TYPE AND DATES COVERED
SBIR Final Report; 12/10/99 – 06/09/00
TITLE AND SUBTITLE
Mars Gashopper
5.
FUNDING NUMBERS
Contract No.
6.
AUTHORS
Robert M. Zubrin, Brian Frankie, Dean Speith, Mark Caviezel, Andrew Martin
Brian Birnbaum, Frank Tarzian, Gilbert Chew, Gary Snyder
7.
PERFORMING ORGANIZATION NAME(S) AND ADDRESS(ES)
NAS 3-00074
8.
PERFORMING ORGANIZATION
REPORT NUMBER
Pioneer Astronautics
11111 W. 8th Avenue, Unit A
Lakewood, CO 80215
9.
PA-GH-1
SPONSORING/MONITORING AGENCY NAME(S) AND ADDRESS(ES)
10. SPONSORING/MONITORING AGENCY
REPORT NUMBER
NASA Jet Propulsion Laboratory
California Institute of Technology
4800 Oak Grove Dr
Pasadena, CA 91109
11. SUPPLEMENTARY NOTES
12a. DISTRIBUTION/AVAILABILITY STATEMENT
12b. DISTRIBUTION CODE
For general distribution
13. ABSTRACT (Maximum 200 words)
The Mars Gas Hopper, or “gashopper,” is a novel concept for propulsion of a robust Mars surface hopper vehicle
which utilizes indigenous CO2 propellant to provide Mars exploration with greatly enhanced mobility. The gashopper
will acquire CO2 gas from the Martian atmosphere, and store it in liquid form at a pressure of about 10 bar. When
enough CO2 is stored to make a substantial ballistic trajectory hop to another Mars site of interest, the CO2
propellant tank will be moderately heated to raise it to 70 bar. The propellant is then run through a hot pellet bed to
form high temperature gas that is expanded through a nozzle to produce thrust. The gashopper uses its CO2
propulsion system for major liftoff, attitude control, and landing propulsive burn(s), as required. Unlike chemical
rockets, the gashopper’s exhaust will not contaminate the landing site with organics or water. The gashopper has a
potential flight range of 5 to 50 kilometers. It can fly over terrain impassible to rovers, imaging as it flies, land to
reconnoiter a remote location, and then fly again. Thus, it offers unique capabilities for Mars surface exploration.
Pioneer Astronautics proposes to demonstrate the feasibility of the gashopper.
14. SUBJECT TERMS:
15. NUMBER OF PAGES
47
Mars Exploration; Surface Mobility; In Situ Fuel Vehicles; Ballistic Flight Vehicles; Thermal Rockets
16. PRICE CODE
17. SECURITY CLASSIFICATION
OF REPORT
Unclassified
NSN 7540-01-280-5500
18. SECURITY CLASSIFICATION
OF THIS PAGE
Unclassified
19. SECURITY CLASSIFICATION
OF ABSTRACT
Unclassified
Computer Generated
20. LIMITATION OF ABSTRACT
UL
STANDARD FORM 298 (Rev 2-89)
Prescribed by ANSI Std 239-18
298-10
49