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American Institute of Aeronautics and Astronautics
1
Analysis and Prediction of Dual-Mode Chemical and
Electric Ionic Liquid Propulsion Performance
Brian R. Donius* and Joshua L. Rovey†
Missouri University of Science and Technology, Rolla, Missouri, 65401
(formerly University of Missouri-Rolla)
An analytical and numerical investigation of the performance of a dual-mode
propulsion system using ionic liquids is presented. Chemical bi-propellant
performance of select propellants is determined using Chemical Equilibrium with
Applications. Comparison of predicted specific impulse of ionic liquids with
hydrazine and UDMH show that the ionic liquid propellants have 3-12% lower
specific impulse when paired with nitrogen tetroxide. However, when paired with
hydroxlammonium nitrate, the specific impulse is comparable. Density Impulse for
ionic liquids is found to be superior due to their higher density. Analytical
investigation of an electrospray electric propulsion system shows that some ionic
liquids are capable of operating in a purely ionic regime, providing very high
specific impulse (~ 6000 sec). The predicted chemical and electric performance data
will be used to quantify mass savings for representative dual-mode propulsion
missions.
Nomenclature
CEA = Chemical Equilibrium with Applications
CPIA = Chemical Propulsion Information Agency
ε = Nozzle expansion ratio
e = Fundamental charge
g0 = Gravitational constant
HAN = Hydroxlammonium Nitrate
ILs = Ionic liquids
IRFNA = Inhibited Red Fuming Nitric Acid
ISP = Specific Impulse
K = Ratio of specific heats
= Mass flow rate of ion i
MW = Molecular weight
NTO = Nitrogen Tertroxide
P2/P1 = Ratio of nozzle exit pressure to combustion chamber pressure
Pc = Combustion Chamber Pressure
q/m = Charge to mass ratio
R’ = Universal gas constant
Tm = Melting temperature
UDMH = Unsymmetrical Dimethylhydrazine
V = Voltage
Ve = Exit velocity
WFNA = White Fuming Nitric Acid
* Graduate Research Assistant, Aerospace Plasma Laboratory, Mechanical & Aerospace Engineering, 210
Toomey Hall, 400 W. 13th St., Student Member AIAA.
† Assistant Professor of Aerospace Engineering, Mechanical & Aerospace Engineering, 292D Toomey
Hall, 400 W. 13th St., Senior Member AIAA.
48th AIAA Aerospace Sciences Meeting Including the New Horizons Forum and Aerospace Exposition
4 - 7 January 2010, Orlando, Florida AIAA 2010-1328
Copyright © 2010 by the American Institute of Aeronautics and Astronautics, Inc. All rights reserved.
American Institute of Aeronautics and Astronautics
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I. Introduction
OMBINATION of high-thrust chemical and high-specific impulse electric propulsion into a single
dual-mode system has the potential to greatly enhance spacecraft mission capability. Ideally, the
combined system would share both hardware and propellant to provide the greatest reduction in system
mass and maximum spacecraft flexibility. In the following sections we describe and review the current
state-of-the-art in dual-mode propulsion and ionic liquids. We then quantify the anticipated chemical and
electric thruster performance of popular ionic liquids. Finally, we summarize predicted propulsion
performance and identify orbit maneuvers of interest that will be used in future work to quantify the
benefits of a dual-mode system.
A. Ionic Liquids
An ionic liquid (IL) is either an organic or non-organic salt in a molten (liquid) state. Because of its
molten state, the cation and anion of the salt dissociate, but the overall liquid remains quasi-neutral. All
salts will obtain this state if heated to the proper temperature, but there is a sub-group known as room
temperature ionic liquids that can remain liquid at or below 293 K. Although known since 1914, recent
developments in chemistry have allowed the number of known ILs to reach well into the hundreds.1 The
exact mechanism for this molten salt behavior has yet to be identified, making the prediction of IL
properties difficult. However, in general, common properties of ILs are high conductivity, viscosity, and
negligible vapor pressure.
Ionic liquids are being considered as replacements for traditional explosives and rocket propellants2,3
and volatile industrial solvents in chemical processing. Although thought of as universally benign by most
professionals, a joint publication recently sought to highlight the combustibility of many ILs as they
approach decomposition temperature.4 This work also pointed out that combustion of several ILs becomes
drastically more vigorous when sprayed rather than combusted as a pooled sample. Other articles have
suggested that IL hydroxlammonium nitrate (HAN) may be used as a substitute for hydrazine as a
monopropellant.5-7 These analyses show that HAN performance both in pure and mixed forms is on the
same order as hydrazine. Lastly, hypergolic behavior for several ILs has been reported when combined
with traditional space storable oxidizers.8-10
Ionic liquids have also been investigated as electrospray propellants.11-13 Ionic liquids for electrospray
application were initially investigated because of their low vapor pressure. Previous electrospray liquids
had relatively high vapor pressure and would boil off the emitter. This interest evolved to include multiple
studies that explored plume emission of electrospray ionic liquids. Results showed that a spray of an ionic
liquid can approach a purely ionic regime (PIR) of emission similar to that found in field emission electric
propulsion.
B. Dual-Mode Propulsion
The main goal of a dual-mode propulsion system is to reduce spacecraft mass and enhance flexibility
through the use of common system resources. Examples of common resources are either hardware or
propellant. The modes of propulsion can be characterized as the tasks to be performed by the propulsion
system during a given mission. These include but are not limited to high thrust orbit transfer maneuvers,
low thrust station keeping, low thrust orbital transitions, and precision low thrust pulses for attitude control.
Although the dual-mode concept has been passed around the academic community for quite some time,
focused research has been nearly nonexistent.
One of the few instances where dual-mode propulsion has been applied was the Mars Global
Surveyor.14 However, this system consisted of two chemical systems, a bi- and mono-propellant thruster.
The system was developed on a tight budget and a rushed schedule dictated by the loss of the Mars
Observer craft. The system consisted of a common fuel bi-propellant thruster and catalyst monopropellant
thruster (Fig. 1). The bi-propellant thruster was used in conjunction with aero-braking to bring the probe
into orbit around Mars; while the monopropellant system was used for attitude control. The common
propellant to both thrusters was hydrazine.14 This common propellant allowed for the integration of three
identical tanks that were readily available, reducing costs. Dual-mode propulsion was selected for this
mission not based on improved performance, but out of the desire to speed up development, cut costs, and
ease integration issues due to the smaller size of the Delta II launch vehicle selected for the mission.
Although the MGS does represent a dual-mode system its use of hypergolic propellants and catalyst
C
American Institute of Aeronautics and Astronautics
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monopropellant thrusters leaves little room for improved performance, it is therefore necessary to examine
new concepts for dual-mode propulsion.
II. Chemical Propulsion Analysis
The NASA Glenn Research Center, Gordon-McBride Chemical Equilibrium with Applications code is
used to analyze chemical combustion and rocket performance of multiple bi-propellant combinations.15
Specifically, ionic liquid fuels and oxidizers are paired with conventional fuel and oxidizer to determine
rocket performance.
A. Investigated Fuels
Ten different ionic liquids were selected based on available literature data. To evaluate the capabilities
of an ionic liquid fuel in a chemical propulsion system, it was first necessary to determine fundamental
properties, such as heat of formation, melting point, and density. Because investigation of combustion
properties of ionic liquids has only recently begun to receive attention, limited knowledge regarding heat of
formation of these liquids has been determined. This led to the selection of 1-Butyl-3-methylimidazolium
Dicyanamide and 1-Butyl-3-methylpyrrolidinium Dicyanamide as two of the IL fuels. Additionally, 1-
Butyl-3-methylimidazolium Dicyanamide has been reported to have hypergolic reactions with standard
storable oxidizers.10 The heat of formation of both Dicyanamide IL’s was determined through calorimeter
tests and calculations using the quantum chemistry program Gaussian 3.16-17 Reported results include a
difference of approximately 1% between the calculation and experimental data. Heats of formation for
eight energetic ILs based on 5-aminotetrazolate have also been determined.18 To allow for a comparison
with current space storable propellants, hydrazine and UDMH were also evaluated. Hydrazine has the
greatest performance of all traditional storable fuels, but is highly toxic and notoriously unstable. UDMH
has only slightly reduced performance from hydrazine, but a larger storable temperature range and greater
stability. All thermodynamic data for these propellants can be found in Table 1.
Figure 1. Mars Global Surveyor schematic showing dual-mode
bi- and mono-propellant hydrazine thrusters14
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B. Investigated Oxidizers
Four oxidizers were selected for the combustion analysis. One of the oxidizers, hydroxylammonium
nitrate (HAN), is an ionic liquid. The other three are common to current state-of-the-art chemical rocket
systems. Specifically, nitrogen tertroxide (NTO), white fuming nitric acid (WFNA), and inhibited red
fuming nitric acid (IRFNA) are considered. NTO is a highly toxic, storable, space oxidizer with extensive
flight heritage and in most reactions provides the best performance. WFNA is essentially pure nitric acid
doped with a small percentage of hydrofluoric acid to permit storage in a variety of container materials. In
general, WFNA has reduced performance compared to NTO, but, aside from being corrosive, it is relatively
benign. In terms of percentage by mass, IRFNA is 83% HNO3, 14% N2O4, 2.4% H2O, and 0.6% HF. In
general, IRFNA usually has the same performance as WFNA, but has a far reduced melting point. The
ionic liquid HAN was selected because of afore mentioned interest in HAN-based monopropellant systems.
For the combustion analysis, HAN is considered as an oxidizer in a pure liquid state at 316.05 K.
C. Performance Criteria and Simulations
In this project the measures of performance selected for investigation are specific impulse (ISP), density
impulse, and storability. Specific impulse is defined as the thrust per unit weight flow rate of propellant
and represents how efficiently a system uses propellant. It is determined from Eq. (1)19 for ideal flows.
The pressure ratio is determined by solving the transcendental equation Eq. (2) by assuming an area ratio
(ε). Density impulse takes into account how easily the oxidizer-fuel combination can be stored. The
storability of the propellant for this project is qualitatively described as the need for additional heating or
cooling of the propellant to maintain liquid phase in the storage tanks. This is important to reducing strain
on the power system of the satellite and to prevent excessive propellant loss due to boil off. The storability
for this study was quantitatively defined as the ratio of the melting point temperature of hydrazine to the
melting point temperature of the ionic liquid Eq. (3). With this description, a value greater than one
signifies that a propellant is at least as storable as hydrazine, while a value less than one signifies that
additional heating of the propellant is likely to be required. To determine these measures of performance
the chemical composition and thermodynamic state of the exhaust stream from the rocket combustion
chamber must be known.
Table 1. Thermo-chemical data for fuel and oxidizers investigated in this study
Fuel # Fuels Name Formula
Hf(KJ/mol) Density (g/cm3)Melting Point (K)
- Hydrazine
N2H4- 1.01 274.69
- UDMH
C2H8N2- 0.79 216
1 1-Butyl-3-methylimidazolium Dicyanamide
C10H15N5363.40 1.06 267
2 1-Butyl-3-methylpyrrolidinium Dicyanamide
C11H20N4218.90 1.02 223
3 Hydrazinium 5-aminotetrazolate
CH7N7383.60 1.48 398
4 Guanidinium 5-aminotetrazolate
C2H8N8205.40 1.54 399
5 Aminoguanidinium 5-aminotetrazolate
C2H9N9302.30 1.51 366
6 Guanylguanidinium 5-aminotetrazolate
C3H10N10 306.90 1.41 414
7 4-Amino-1H-1,2,4-triazolium 5-aminotetrazolate
C3H7N9565.00 1.62 387
8 4-Amino-1-methyl-1,2,4-triazolium 5-aminotetrazolate
C4H9N9546.00 1.46 249
9 4-Amino-1-ethyl-1,2,4-triazolium 5-aminotetrazolate
C5H11N9523.40 1.39 235
10 1,5-Diamino-4-methyl-1,2,3,4-tetrazolium 5-aminotetrazolate
C3H9N11 655.10 1.57 444
Oxidizers
Nitrogen Tertroxide (NTO)
N2O4- 1.44 261.95
White Fuming Nitric Acid (WFNA)
HNO3- 1.33 231.6
Inhibited Red Fuming Nitric Acid (IRFNA)
83% HNO3
+14% N2O4
+ 2.4% H2O + .6% HF
-
1.59 216
Hydroxlammonium nitrate (HAN)
NH3OHNO3-79.68 1.83 316.05
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(1)
(2)
(3)
The NASA Chemical Equilibrium with Applications code was used to determine the performance of
each propellant combination. This program has been under continual development since the late 1950’s
and offers users the capability to determine equilibrium composition and adiabatic flame temperature for
any reaction.15 A recent addition to the program also allows the user to define a new fuel given the heat of
formation and molecular composition. A series of 1536 simulations were performed using CEA by varying
Pc, equivalence ratio, and propellant combinations for a fixed expansion ratio (ε) of 40.
D. Results
The chamber pressure was varied between 150 psia and 600 psia because these are typical levels in on-
orbit engines.20 As can be seen in Fig. 2, the variation of engine performance with pressure is minimal.
This is to be expected because these relatively small pressures are not capable of greatly effecting the
product species dissociation. Therefore all subsequent analyses will be restricted to the 300 psia case.
For the 48 propellant combinations, the mixture ratio was varied between 0.6 and 1.3 in order to find
peak performance for each combination. An equivalence ratio of unity (stoichiometric) represents the point
of complete combustion. Generally it was found that peak performance rested far to the right of the
stoichiometric condition. This is a common result for rocket performance because as excess fuel is
introduced to the system, molecular weight decreases faster than chamber temperature. A summary of the
results for all oxidizer fuel combinations can be seen in Table 2a and Table 2b. All performance shown in
Tables 2a and 2b is scaled with respect to the peak performance of hydrazine when combined with NTO
(333, s and 406, g-s/cm3). For example, UDMH-NTO has an ISP that is 97% of that obtained with
hydrazine-NTO.
Figure 2.Variation of ISP with change in pressure for a sample ionic liquid
260
265
270
275
280
285
290
295
300
0.6 0.8 1.0 1.2
ISP, s
Equivalence Ratio, -
150psi
300psi
450psi
600psi
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Results show that hydrazine and UDMH provide higher ISP than each investigated ionic liquid fuel,
regardless of the oxidizer. However, simulations also show that hydrazine and UDHM combustion with
HAN oxidizer provides higher ISP than when combusted with NTO. With NTO oxidizer, the ionic liquid
fuels have 3-12% lower ISP compared to hydrazine and UDMH. NTO out performed IRFNA and WFNA
for any given propellant with the singular exception of 1-Butyl-3-methylimidazolium Dicyanamide, which
resulted in only a small gain of five seconds ISP. IRFNA and WFNA have almost identical performance for
any given propellant combination. When combined with HAN for ideal combustion, the performance of all
fuels improved. That being said no IL-HAN combination resulted in superior performance to baseline
hydrazine-NTO. However, the prediction of superior performance of HAN to NTO warrants further
investigation.
In the area of density impulse the ionic liquids show superior performance to the traditional propellants,
especially when combined with IRFNA and HAN. This is due to the greater density of the ionic liquids in
comparison to that of hydrazine or UDMH. Density impulse takes into account how easily the oxidizer-fuel
combination can be stored. This would seem to offer these propellant combinations a niche of small volume
budgeted systems with proportionally high mass budgets.
Analysis indicates that storability of ionic liquids is a major difficulty. Only four propellants showed
equal or superior storability than hydrazine (highlighted in green). As a result the remaining six propellants
do not represent storable space propellants and are dropped from consideration in the remainder of the
Table 2b. Summary of chemical propulsion – IRFNA and WFNA oxidizer
Fuel Specific Impulse (-) Density Impulse (-) Specific Impulse (-) Density Impulse (-)
Hydrazine 1.00 1.00 0.96 0.92
UDMH 0.93 0.85 0.93 0.79
1 0.90 0.96 0.90 0.88
2 0.90 0.95 0.90 0.87
3 0.93 1.18 0.93 1.07
4 0.88 1.13 0.88 1.00
5 0.89 1.14 0.89 1.02
6 0.88 1.09 0.88 0.98
7 0.90 1.20 0.90 1.08
8 0.90 1.14 0.90 1.03
9 0.90 1.11 0.90 1.01
10 0.91 1.19 0.91 1.07
IRFNA
WFNA
Table 2a. Summary of chemical propulsion – NTO and HAN oxidizer
Fuel Specific Impulse (-) Density Impulse (-) Specific Impulse (-) Density Impulse (-) Storability (-)
Hydrazine 1.00 1.00 1.03 1.14 1.00
UDMH 0.97 0.85 1.01 0.97 1.27
1 0.88 0.91 0.99 1.13 1.03
2 0.94 0.95 0.99 1.11 1.23
3 0.96 1.16 1.00 1.36 0.69
4 0.90 1.10 0.96 1.35 0.69
5 0.92 1.11 0.97 1.34 0.75
6 0.91 1.07 0.97 1.29 0.66
7 0.93 1.16 0.98 1.40 0.71
8 0.93 1.12 0.98 1.33 1.10
9 0.93 1.09 0.98 1.31 1.17
10 0.94 1.16 0.98 1.38 0.62
NTO
HAN
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study. With a storability factor of 0.87, HAN does not represent a storable propellant, but as it is
substantially closer to hydrazine than any other ionic liquid that failed the criteria it will be carried forward
in the analysis.
E. Error Analysis
Although CEA has been held in high regard for its predictive capabilities, it is still necessary to assess
the accuracy of the performance predictions. To quantify the error in performance predicted by the code, a
series of test cases were used to compare the CEA output with actual engine performance for systems found
in the Chemical Propulsion Information Agency (CPIA) engine manual.20 Given the mixture ratio,
expansion ratio, and chamber pressure, the predicted ISP for equilibrium flow was compared to the actual
performance (Fig. 3). Results show that as thrust level increases, the accuracy of the simulation increases.
For engines with thrust levels less than 100 lbf, the values predicted by CEA for ISP were found to be on
average 15.6% higher than that of test data, with a standard deviation of 3.7%. To ensure a more
conservative estimate of ISP for a real system in future studies, a value of ISP 19.3% below the predicted
values will be used representing a one standard deviation factor of safety.
III. Electrical Performance
A dual-mode propulsion system with common propellant must have an electric propulsion system that
can operate with a combustible fuel. In this investigation, the focus is on ionic liquids. Electrospray electric
propulsion systems have been operated with ionic liquids and are the focus of the electric performance
analysis. Specifically, we use test results of ionic liquids similar to those proposed in this investigation to
predict electric propulsion performance of an electrospray thruster operating with a combustible ionic
liquid propellant.
A. Methods of Modeling Electrospray Propulsion
The traditional electrospray consists of emission of fine droplets from a micro jet that forms at the apex
of an electro-fluidic structure known as a Taylor cone.21 This cone is formed from the application of a start-
up potential between a capillary containing a conducting fluid and an extractor grid downstream. Once
formed, the applied voltage on the Taylor can be varied quite substantially, even to values below the start-
up potential. Excessive voltages can result in the deformation of the cone and can even result in multiple
secondary emission points. A stable Taylor cone emits droplets and ions in proportions dependent on the
flow rate of propellant. Higher flow rates result in predominantly droplet emission, while minimum flow
rate yields a nearly pure ion regime for some ionic liquids.
To quantify the performance of an electrospray propulsion system, initially two models for traditional
solution based sprays were investigated.21-22 Both models are empirically based on a series of dope polar
solvent solutions. These models take a series of electrochemical properties and a proportionality curve
between emitted current and dielectric constant to predict the specific impulse, thrust, and droplet size.
Unfortunately, these two models are not valid for use with ionic liquids for various reasons. First there is
disagreement between the two methods as to the functional form of the emitted current and dielectric
Figure 3. CEA error in ISP based on engine performance data from CPIA
engine manual as a function thrust
0
5
10
15
20
25
110 100 1000 10000 100000 1000000
%Error
log(Thrust, lbf)
MMH NTO
UDMH NTO
UDMH IRFNA
Aerozine NTO
H2 LOX
American Institute of Aeronautics and Astronautics
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constant. Further analysis by Chen23 suggests that the results of this curve may depend greatly on mobility
of ions present in the fluid. Since ionic liquids typically are composed of complex ions it can be expected
that ion mobility’s for ILs will be very different from the conventional additives of polar solvent solutions.
Second, the typical values of ionic liquid dielectric constant fall at the very edge of the range investigated
in both models. Lastly, neither model is capable of predicting results near the pure ionic regime, which is of
most interest to this study.
Although no formal method exists for modeling whether an ionic liquid can produce a pure ionic
regime, emission close to PIR has been reported for several ionic liquids.13,24 It appears that a general
requirement is that the fluid possess high conductivity and surface tension and a relatively low viscosity.25
For all but the most viscous fluids, a regime close to PIR has been induced by heating the fluid.18 The
properties of Fuel 1 (1.139 S/m and 0.0466 N/m respectively)26,27 show that it falls into the reported range
of conductivity and surface tension of the other fluids reported to have obtained PIR. Unfortunately,
electrochemical data for HAN and Fuels 2, 8, and 9 are unavailable in the published literature. Therefore
the two 5-aminotetrazolate based ILs were neglected outright. While it can be noted that if Fuel 2 could
undergo PIR electrospray its performance would be virtually identical to that of Fuel 1 because of their
similar molecular composition. HAN was maintained because it represents a known ionic liquid
monopropellant and any resultant dual-mode system would be of high interest. Because fuel 1 has
properties similar to those ILs that have shown PIR, 13,24,26,27 we assume it will operate in PIR with similar
flowrates.
Emission in the PIR consists of pure ions and ions traveling with clusters of N number of neutral pairs
(degree of solvation). A value from literature is then needed to approximate this value for the percentage of
the mass flow that is pure ions and is ions of different degrees of solvation. This percentage can be
determined from time of flight curves for IL’s that have operated in PIR. This time of flight curve is
produced experimentally by cutting flow rate to the emitter and measuring the current emitted as a function
of time. The ions will arrive at the collector in order of least to greatest mass (higher solvation). Assuming
an ideal collapse of emission a typical curve would resemble Fig. 4. The successive step down in current
represents the final arrival time of a particular species and assuming that all emission strikes the collector
would also represent the steady current contribution of that species. Given the current contribution of a
species the mass flow rate of that species would be given by Eq. (4), where m is the mass of one molecule
of the species and e is the fundamental charge. The total mass flow rate of the jet would be found by
summation of the component mass flow rates.
Using this technique and actual time of flight curves13 the average total mass flow rate for PIR emission
ionic liquids becomes 1.2x10-12, g/s/emitter. An examination of the TOF curves shows that typically only
the first two ion states are present in the emission in any discernable quantity. The averaged result is a 40%
emission of pure ions and a 60% emission of the 1st solvated state. In reality the time of flight curve will
vary from the ideal because of collisions of molecules and non-finite cessation of emission from the
emitter. Based on the available data the potential associated with PIR emission for a typical ionic liquid is
1760 V.13
(4)
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B. Results
Applying ideal equations for a charged mass in a potential field the specific impulse for the system
would be 6526s for HAN and 4511s for 1-Butyl-3-methylimidazolium Dicyanamide. Using the power
level and efficiency of the BHT-200 (200 W, 43.5%) results in a system thrust of 2.7 mN and 4.0 mN
respectively for the propellants. Comparing these results with the performance of the BHT-20028 (1390 s,
12.8 mN) shows a trade off in thrust for higher propellant efficiency when using an electrospray in PIR. If
emitters of 50 μm inner diameter and 150 μm outer diameter were spaced 50 μm apart the system would
consist of 34,798 needles and occupy a circular area with a diameter of 5.3 cm for HAN. A comparison of
this arrangement with the space charge-limit shows that it would be operating at only 3-5% of the
maximum.
IV. Dual-mode Benefits and Future Work
A. Dual-mode Benefits
A typical spacecraft thrust history is normally dictated years before it reaches the launch pad. Systems
are often equipped only to perform station keeping and attitude control after separation, with a rare few
having the ability to raise themselves to their final orbit. In a dynamically changing world and ever rising
threat of on-orbit debris, the need for an operationally responsive space technology is increasing important.
A dual-mode system will allow an on orbit asset to change orbit as the need arises rather than predicting the
event. If a rapid orbit change is required the system can respond with a chemical high thrust burn and then
gradually return to its primary orbit in a high efficiency electrospray mode. Also, in the event the asset is
never called upon to perform such radical maneuvers, the system will be able to maintain its primary orbit
for an extended period using the ionic liquid propellant purely in the electrospray mode.
B. Future Work
Using the performance predicted in this paper (shown in Table 3), a series of orbital maneuvers will be
considered, including debris avoidance, orbit raising, and inclination adjustment. The goal is to determine
and quantify any mass savings associated with an ionic liquid based dual-mode system. Accurately
predicting the weight of individual system components such as power processing units and fuel tanks will
be of utmost importance.
Figure 4. Example time of flight trace for nearly pure ionic emission
0
50
100
150
200
250
300
010 20 30 40 50 60 70 80 90 100
Current, nA
Time, ms
I(Cation+)
I(Cation+ (CationAnion)1)
I(Cation+ (CationAnion)2)
I(Cation+ (CationAnion)n)
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V. Conclusion
The assessment of chemical propulsion performance for a series of ionic liquids has been determined
through use of CEA. It was found that similar performance to traditional storable propellant combinations
may be possible if ionic liquids are teamed with the HAN oxidizer. Performance of HAN and 1-Butyl-3-
methylimidazolium Dicyanamide operating in the pure ionic regime of electrospray emission has been
approximated and has been found to operate at much higher specific impulses when compared to a standard
Hall thruster. Given these performance predictions and yet to be determined system weights a series of
mission thrust profiles will be simulated to determine how much flexibility would be gained and what
potential mass saving may be incurred by use of a dual-mode system.
References
1CRC Handbook of Chemistry and Physics, 88th ed., Taylor and Francis Group, LLC, 2008, Section 6 Fluid
Properties, Sub. Ionic Liquids
2Zube, D., Wucherer, E., Reed, B.,“ Evaluation of Han-Based Propellant Blends,” 39th AIAA Joint Propulsion
Conference, AIAA 2003-4643 Huntsville Alabama 2003.
3Boatz, J., Gordon, M., Voth, G., Hammes-Schiffer, S.,“ Design of Energetic Ionic Liquids,” hpcmp-ugc, pp.196-
200, 2008 DoD HPCMP Users Group Conference, 2008
4Smiglak, M. et al.,“ Combustible ionic liquids by design: is laboratory safety another ionic liquid myth?,”, Chem
Commun. 2006
5Amariei, D., et al.,“ Influence of the fuel on the thermal and catalytic decompositions of ionic liquid
monopropellants,” 41st AIAA Joint Propulsion Conference, AIAA 2005-3980 Tucson Arizona 2005.
6Chang, Y., Kuo, K.,“ Assessment of Combustion Characteristics and Mechanism of a HAN-Based Liquid
Monopropellant,” 37th Joint Propulsion Conference, AIAA 2001-3272 Salt Lake City Utah 2001.
7Meng, M., Khare, P., Risha, G., Yetter, R.,Yang, V., “Decomposition and Ignition of HAN-Based
Monopropellants by Electrolysis,” 47th AIAA Aerospace Sciences Meeting, 2009-451, Orlando, Florida, 2009
8Schneider, S. et al.,” Ionic Liquids as Hypergolic Fuels,” Energy & Fuels 2008, 22, 2871-2872
9Gao, H., Joo, Y., Twamley, B., Shreeve, J.,”Hypergolic Ionic Liquids with 2-2,Dialkyltiazanium Cation,”
Chemistry International 2009, 48, 1-5
10Dambach, E.,” An Investigation into the hypergolicity of Dicyanamide-Based Ionic Liquid Fuels with Common
Oxidizers,” USAF public release
11Lozano, P., Sanchez, M.,“ Efficiency Estimation of EMI-BF4 Ionic Liquid Thrusters,” 41st AIAA Joint
Propulsion Conference, AIAA 2005-4388 Tucson Arizona 2005.
Table 3. Propulsion Systems for Dual-mode Benefits Study
*Chemical thrust assumed for a small spacecraft
Propellant Isp (s) Thrust (N)* mdot (kg/s) Type
N2H4-NTO 270 10 3.7E-02 Bipropellant
N2H4-HAN 277 10 3.6E-02 Bipropellant
BIMDAC-NTO 239 10 4.2E-02 Bipropellant
BIMDAC-HAN 267 10 3.7E-02 Bipropellant
N2H4 227 10 4.4E-02 Monopropellant
HAN 218 10 4.6E-02 Monopropellant
Propellant Isp (s) Thrust (mN) mdot (g/s) Type
Xenon 1390 12.8 9.4E-07 Hall
HAN 6526 2.7 4.2E-08 Electrospray
BIMDCA 4511 4.5 8.8E-08 Electrospray
Electric Propulsion (200 W)
Chemical Propulsion
American Institute of Aeronautics and Astronautics
11
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