MA SS AV"HjSF.T-r 41:
Modeling the Characteristics of PropulsiorSystems Providing Less Than 10 N Thrust
by
Thomas Michael Chiasson
B.S. Astronautics, United States Air Force Academy (2010)
Submitted to the Department of Aeronautics and Astronautics
in partial fulfillment of the requirements for the degree of
Master of Science in Aeronautics and Astronautics
at the
MASSACHUSETTS INSTITUTE OF TECHNOLOGY
June 2012
@ Massachusetts Institute of Technology 2012. All rights reserved.
A
Author ...................................
Department of Aeronautics and Astronautics
May 24, 2012
\F7
.....
Certified by.................
-Paulo C. Lozano
H.N. Slater Associate Professor of Aeronautics and Astronautics
Thesis Supervisor
/ /
Accepted by .......................
(
I
Eytan H. Modiano
Professor of Aeronautics and Astronautics
Chair, Graduate Program Committee
ARCHIVES
TF
2
4
Acknowledgments
I would first like to thank my advisor, Prof. Paulo Lozano, for his patient guidance
and advice during my two years at MIT. I couldn't have chosen a better lab or advisor.
I would also like to thank Jaime Ramirez and Jim Sisco at Aurora Flight Sciences
for including me on the project that led to this research and giving much needed
assistance on a program that turned out to be much bigger than I expected.
Special thanks goes to Tom Coles for his genius on every software program I could
ever want to use. If you were an app, I'd put you on my home screen.
As always, I am grateful to the U. S. Air Force and the U. S. taxpayer for providing
for my education. I hope I prove to be a good investment.
At the end of 20 years of education, it seems appropriate to thank the teachers
who have inspired and believed in me over the years: Mrs. Kilgore, Ms. Diaz, Ms.
Aragon, Coach Scott, Eric Hamp, Mr. Thomas, Lt Col Adelgren, Dr. Vergez, Zhou
Laoshi, Kevin Gibbons, Maj Hague, Dr. Brown, Col Lawrence, and Col Anthony, to
name just a few. I hope that someday I can do for someone else what each of you
have done for me.
Thanks to my GC brothers who convinced me to come here in the first place:
Tony, Joel, Daniel, Bruce, and countless others. Thanks to my family in Boston who
made sure I stayed: Mark, John, Nayoon, Chance, Adelina, Chi, Sofia, Megan, and
everyone at the Justice House of Prayer.
Thanks to my family, without whom I would never have achieved anything, and
especially to my parents, Mike and Louisa. You believed in me before I believed in
myself. I hope I can make you as proud of me as I am of you.
Finally, but most importantly, thanks to Abba Father, Jesus, and Holy Spirit. You
never cease to surprise and amaze me, and I know that it's always just the beginning
of eternity.
5
6
Contents
1
Introduction
15
2
Fundamentals of Space Propulsion
19
3
Predicting Relationships from Historical Data
23
3.1
Cold Gas Thrusters . . . . . . . . . .
. . . . . . . . . . . . . . . . .
26
3.2
Monopropellant Thrusters . . . . . .
. . . . . . . . . . . . . . . . .
30
3.3
Bipropellant Thrusters . . . . . . . .
. . . . . . . . . . . . . . . . .
36
3.4
Resistojets . . . . . . . . . . . . . . .
. . . . . . . . . . . . . . . . .
38
3.5
Arejets . . . . . . . . . . . . . . . . .
. . . . . . . . . . . . . . . . .
43
3.6
Ion Engines
. . . . . . . . . . . . . .
. . . . . . . . . . . . . . . . .
49
3.7
Hall Effect Thrusters . . . . . . . . .
. . . . . . . . . . . . . . . . .
51
3.8
Pulsed-Plasma Thrusters . . . . . . .
. . . . . . . . . . . . . . . . .
54
3.9
Vacuum Arc Thrusters . . . . . . . .
. . . . . . . . . . . . . . . . .
56
3.10 Field-Emission Electric Propulsion
. . . . . . . . . . . . . . . . .
56
3.11 Colloid Thrusters . . . . . . . . . . .
. . . . . . . . . . . . . . . . .
59
3.12 Electrospray Thrusters . . . . . . . .
. . . . . . . . . . . . . . . . .
61
4
Modeling Propellant Storage and Management
63
5
Validation
67
5.1
Aerojet Cubesat Propulsion Module . . . . . . . . . . . . . . . . . . .
67
5.2
BepiColombo Ion Engines
69
. . . . . . . . . . . . . . . . . . . . . . . .
7
5.3
University of Minnesota Shackleton Crater Reconnaisance Mission Bipropellant System
5.4
. . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
70
Clyde Space PPT Module . . . . . . . . . . . . . . . . . . . . . . . .
70
6
Applications
73
7
Conclusion
79
A Thruster Data
83
B Source Code
105
8
List of Figures
3-1
Historical data for xenon Hall thrusters.
. . . . . . . . . . . . . . . .
24
3-2
Historical data for hydrazine monopropellant thrusters. . . . . . . . .
25
3-3
Historical data for cold gas thrusters.
Most of the data is gathered
from using nitrogen gas, and the outlier is from using helium.....
27
3-4
Historical data for cold gas thrusters. . . . . . . . . . . . . . . . . . .
28
3-5
Historical data for cold gas thrusters. . . . . . . . . . . . . . . . . . .
28
3-6
Historical data for cold gas thrusters. . . . . . . . . . . . . . . . . . .
29
3-7
Historical data for cold gas thrusters, looking at just the low thrust
regim e. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
29
3-8
Historical data for hydrazine monopropellant thrusters. . . . . . . . .
31
3-9
Historical data for hydrazine monopropellant thrusters. . . . . . . . .
32
3-10 Historical data for hydrazine monopropellant thrusters. . . . . . . . .
32
3-11 Historical data for hydrazine monopropellant thrusters. . . . . . . . .
33
3-12 Historical data for hydrogen peroxide monopropellant thrusters. . . .
33
3-13 Historical data for hydrogen peroxide monopropellant t hrusters.
. . .
34
3-14 Historical data for hydrogen peroxide monopropellant t hrusters.
. . .
34
3-15 Historical data for hydrogen peroxide monopropellant t hrusters.
. . .
35
3-16 Historical data for bipropellant thrusters. . . . . . . . . . . . . . . . .
37
3-17 Historical data for bipropellant thrusters. . . . . . . . . . . . . . . . .
37
3-18 Historical data for hydrazine resistojets.
. . . . . . . . . . . . . . . .
39
3-19 Historical data for hydrazine resistojets.
. . . . . . . . . . . . . . . .
39
3-20 Historical data for hydrazine resistojets.
. . . . . . . . . . . . . . . .
40
3-21 Historical data for hydrazine resistojets.
. . . . . . . . . . . . . . . .
40
9
3-22 Historical data for water resistojets. . . . . . . . . . . . . . . .
41
3-23 Historical data for water resistojets. . . . . . . . . . . . . . . .
41
3-24 Historical data for water resistojets. . . . . . . . . . . . . . . .
42
3-25 Historical data for water resistojets. . . . . . . . . . . . . . . .
42
3-26 Historical data for arcjets using various propellants. . . . . . .
44
3-27 Historical data for arcjets using various propellants, zooming in on the
low thrust regime. . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
45
3-28 Historical data for ammonia arcjets. . . . . . . . . . . . . . . . . . . .
45
3-29 Historical data for ammonia arcjets. . . . . . . . . . . . . . . . . . . .
46
3-30 Historical data for hydrogen arcjets. . . . . . . . . . . . . . . . . . . .
46
3-31 Historical data for hydrogen arcjets. . . . . . . . . . . . . . . . . . . .
47
3-32 Historical data for hydrazine arcjets.
. .. . . .. . . . .. . .. . . .
47
3-33 Historical data for hydrazine arcjets.
. .. . .. . . . . .. . .. . . .
48
3-34 Historical data for hydrazine arcjets.
.. . .. . . . .. . . . .. . . .
48
3-35 Historical data for hydrazine arcjets.
.. . .. . . . .. . . .. . . ..
49
3-36 Historical data for xenon ion engines . . . . . . . . . . . . . . . . . . .
50
3-37 Historical data for xenon ion engines . . . . . . . . . . . . . . . . . . .
50
3-38 Historical data for xenon ion engines . . . . . . . . . . . . . . . . . . .
51
3-39 Historical data for xenon Hall thrust,ers . . . . . . . . . . . . . . . . .
52
3-40 Historical data for xenon Hall thrust ers . . . . . . . . . . . . . . . . .
53
3-41 Historical data for xenon Hall thrust ers . . . . . . . . . . . . . . . . .
53
3-42 Historical data for PPTs. ......
. . . .. . . .. . . .. . . . .. .
54
3-43 Historical data for PPTs. ......
. . .. . . .. . . . .. . . .. . .
55
3-44 Historical data for PPTs. . . . . . . . . . . . . . . . . . . . . . . . . .
55
3-45 Historical data for FEEPs. . . . . . . . . . . . . . . . . . . . . . . . .
57
3-46 Historical data for FEEPs. . . . . . . . . . . . . . . . . . . . . . . . .
58
3-47 Historical data for FEEPs. . . . . . . . . . . . . . . . . . . . . . . . .
58
3-48 Historical data for colloid thrusters. . . . . . . . . . . . . . . . . . . .
59
3-49 Historical data for colloid thrusters. . . . . . . . . . . . . . . . . . . .
60
3-50 Historical data for colloid thrusters. . . . . . . . . . . . . . . . . . . .
60
10
6-1
Results from application of a preliminary version of the program for
the System F6 project. . . . . . . . . . . . . . . . . . . . . . . . . . .
6-2
73
Using the program to compare propulsion options for a 100 kg spacecraft. Electrosprays are not compared. Mass fractions over 90% are
not shown. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
6-3
Using the program to compare propulsion options for a 100 kg spacecraft, including electrosprays. Mass fractions over 90% are not shown.
6-4
74
75
Using the program to compare propulsion options for a 10 kg spacecraft. Electrosprays are not compared. Mass fractions over 90% are
not shown. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
6-5
76
Using the program to compare propulsion options for a 10,000 kg spacecraft. Electrosprays are not compared. Mass fractions over 99% are
not shown. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
11
77
12
List of Tables
3.1
Thrusters included in the database. . . . . . . . . . . . . . . . . . . .
3.2
Specifications of VACCO MiPS butane thruster unit with integrated
propellant tank for cubesats.[38] . . . . . . . . . . . . . . . . . . . . .
3.3
24
30
Specifications of Alameda Applied Sciences Corporation pVAT for Illinois Observing NanoSatellite (ION).
. . . . . . . . . . . . . . . . . .
56
3.4
Specifications of MIT Space Propulsion Lab's electrospray arrays. . .
62
4.1
Choosing a tank material. . . . . . . . . . . . . . . . . . . . . . . . .
64
5.1
Comparison of results from program with actual propulsion systems. .
68
A. 1 Cold gas thrusters . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
83
A.2 Cold gas thrusters with liquid propellant (mass includes integrated tank) 85
A.3 Hydrazine monopropellant thrusters . . . . . . . . . . . . . . . . . . .
86
A.4 Hydrogen peroxide monopropellant thrusters . . . . . . . . . . . . . .
88
A.5 Bipropellant thrusters
. . . . . . . . . . . . . . . . . . . . . . . . . .
88
. . . . . . . . . . . . . . . . . . . . . . . . . . . . .
89
A.6
Water resistojets
A.7
Hydrazine resistojets
A.8
A.9
. . . . . . . . . . . . . . . . . . . . . . . . . . .
90
Ammonia arejets
. . . . . . . . . . . . . . . . . . . . . . . . . . . . .
90
Hydrogen arcjets
. . . . . . . . . . . . . . . . . . . . . . . . . . . . .
91
A.10 Hydrazine arcjets . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
91
A.11 Ion engines
. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
92
A.12 Hall Thrusters . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
94
A.13 Pulsed plasma thrusters
98
. . . . . . . . . . . . . . . . . . . . . . . . .
13
A.14 Vacuum arc thrusters . . . . . . . . . . . . . . . . . . . . . . . . . . .
101
A.15 Field emission electric propulsion
. . . . . . . . . . . . . . . . . . . .
102
. . . . . . . . . . . . . . . . . . . . . . . . . . . . .
103
A.16 Colloid thrusters
14
Chapter 1
Introduction
In spacecraft design, scaling equations are often used to estimate the mass of a system
for preliminary design purposes. Although well developed scaling equations have long
been used for space lift or orbit insertion propulsion systems, few if any can be applied
to the low thrust regime. [35} Thrusters providing less than 10 N of thrust are currently
the focus of the majority of research and development on space propulsion.[115] [361
These systems range from integrated tank, thruster, and power processing units for
cubesats, to interplanetary travel, to extremely fine precision control. They are even
more diverse in the means that they provide propulsion; the database that has been accumulated in the course of this research includes models (specific historical thrusters)
categorized into eleven distinct types of propulsion, electrical and chemical, using a
variety of propellants, with several other propulsion concepts for which not enough
data could be gathered to form equations.
However, the systems level designers are usually not interested in how the thruster
works, they care about what it will provide to their spacecraft in terms of thrust
and total velocity change and what it will cost in terms of mass, power, volume,
risk, or money. These numbers require an initial design phase in which the type
of propulsion is chosen and some initial calculations are done to estimate how large
the thruster, propellant storage, and (if necessary) electrical power system will be.
In scenarios when designers want to compare several different spacecraft options, of
which propulsion is only one term in a larger equation, this process can be tedious.
15
Perhaps the designer wants to run some simulations of several different options for
the space system in which the number of satellites and their masses are varying,
drastically changing the propulsion requirements, or perhaps some initial calculations
need to be run to compare the merits of one propulsion system against another to
make an initial design decision. The variety of propulsion options often makes these
situations like comparing apples and oranges. The fact that there is ongoing research
into improving and miniaturizing virtually every type of propulsion only makes these
decisions even more difficult, stressing the need for the most current data.[70][71]
This project was initially started to answer the seemingly simple question of how
spacing in a satellite cluster would cause the collision avoidance requirements to
change the mass fraction of their propulsion systems. A program that quickly and
reliably estimated the mass fraction of propulsion systems featuring monopropellant,
cold gas, and Hall thrusters based off required acceleration, total change in velocity
(AV),
and initial satellite mass was created as part of a parametric model that
would run several thousand scenarios. The program was then used to help simulate
scattering and regathering the cluster. Surprisingly, the results indicated that such
operations would be feasible even for satellites under 10 kg using dual high thrust
and low thrust monopropellant propulsion systems. Furthermore, since the program
was created from the ground up using historical data, there was already information
gathered on actual thrusters on the market that could conceivably be a part of such
a system.
Since then, the program has been expanded to include eighteen different propulsion and propellant options built off of data gathered from over 300 thruster models
ranging from 10 N to to 10 IN of maximum thrust. It can be run iteratively within a
larger program, or it can be used stand alone: any situation where a designer requires
a quick estimate of required power or mass of a propulsion system (including thruster,
propellant, propellant management, and power system) given the initial mass of the
spacecraft and the required AV and maximum acceleration. Data has also been gathered on specific impulse, efficiency, minimum impulse bit, and feed pressure, although
only specific impulse and feed pressure go into the mass calculations and efficiency
16
and minimum impulse bit could not be found for a few thrusters.
This thesis will describe how the program was created, including the database of
thruster specifications, equations that were developed from the database, equations
used to model the other parts of a propulsion system, a few necessary assumptions,
validation, and some interesting applications.
17
18
Chapter 2
Fundamentals of Space Propulsion
Before an examination of the program itself it would be useful to review some fundamentals of space propulsion. Space propulsion at the most basic level is the application of Newton's third law by expelling mass at a rate n and velocity c to propel a
spacecraft in the opposite direction. In propulsion, the resulting force is called thrust
(T).
T =m - c
(2.1)
While thrust relates to how quickly a spacecraft can change velocity, in orbital
mechanics it is often useful to describe maneuvers in terms of total velocity change
(AV) required to move the spacecraft from one orbit to another. These two variables,
thrust and AV, tend to dictate propulsion choice from a design standpoint. Large
spacecraft, short schedules, or rapid maneuvering will require higher thrust from a
propulsion system, whereas drag cancellation, non-elliptical orbits, and long mission
times and/or distances will require larger AV.[66]
AV is by definition dependent on the mass of propellant (mp)expelled from the
propulsion system. This relationship is described in the rocket equation, where ms/c
is the initial mass of the spacecraft including propellant and g is the acceleration of
19
9.81 m/s 2 due to gravity.
myro,
=
ms/c(1 - e~ V
"IP9)
(2.2)
As can be seen from this equation, specific impulse (I,,) is also a determining
factor of a spacecraft's propellant mass fraction. Ip describes the amount of thrust
derived from a propulsion system compared to the mass flow rate of its exhaust
(rh).
T
ISP =
T
rh - g
(2.3)
Although the concepts are different, the relationship between AV and I,, is analogous to that between distance and miles per gallon. Missions requiring high AV must
use propulsion systems with high I,p in order to keep the propellant mass fraction
from being prohibitively large. Unfortunately, propulsion systems with high I,
tend
to have low thrust to power ratios.
For the purposes of this thesis, power refers to the electric power input of a
propulsion system. All propulsion systems require some minimal amount of power
to operate valves and possibly heaters or pumps, but some propulsion systems use
electric power to create thrust, as opposed to depending on chemical reactions. The
thrust of these electric propulsion systems is directly dependent on the input power.
Increasing the power in turn requires a power system, including a power source
such as batteries and/or solar panels and a power processing unit. All spacecraft
require a power system regardless of propulsion, but electric propulsion systems can
have power requirements high enough to significantly increase the mass of a spacecraft's power system. In practice, this means that electric propulsion systems provide
limited thrust.
This brings up a fundamental distinction between electric and chemical propulsion. Chemical propulsion provides unmatched thrust, whereas no electric propulsion
system has ever been designed that can lift the weight of its own required power system off the surface of the planet. However, electric propulsion far surpasses chemical
propulsion in terms of I,p, resulting in less propellant and making them the natural
20
choice for high AV missions, such as interplanetary travel.
Another important factor in propulsion system design is the propellant management system. Depending on the propellant used (which also depends on the type of
propulsion), multiple tanks may be required. The pressurization of propellant tanks
is also important, and an optimal low mass balance must be found between large
thin-walled tanks and small thick-walled tanks. For some types of propulsion systems, thrust or I,
may be dependent on the feed pressure of the propellant into the
thruster.
21
22
Chapter 3
Predicting Relationships from
Historical Data
The foundation of the modeling process is a spreadsheet containing specifications of
all the models of thrusters providing 10 N or less of thrust that could be found. Data
has been collected on 385 specific thruster models. Best fit equations were then found
to represent the relationships between different system characteristics. For example,
Figure 3-1 shows the relationship of thrust to efficiency for Hall thrusters. Each point
represents data gathered from a real world Hall thruster, while the curve is a best fit
of the historical data, which is used to predict system characteristics in the modeling
process.
This approach is time intensive due to the difficulty of finding published specifications for many thrusters; however, it ensures that the modeling process is using
hard data gathered from real world systems to compare thruster options. Of course,
it does not guarantee that every thruster designed will also fit this historical data.
In some case, the data was so chaotic that it was difficult or impossible to find a relationship between certain system specifications, and a flat average was used instead
of an equation. Figure 3-2 gives an example of data that was just barely organized
enough to find an equation for. In this case, it was decided that although some of the
data points were drastically different from the best fit curve, using the equation of
the curve would still be preferable to taking a flat average. So although the historical
23
Table 3.1: Thrusters included in the database.
Thruster Number of Models
Cold Gas (gas)
Cold Gas (liq)
Monopropellant (H4N2)
Monopropellant (H202)
Bipropellant (liq)
Resistojet (H20)
Resistojet (H4N2)
Arcjet (NH3)
Arcjet (H2)
Arcjet (H4N2)
Arcjet (other)
Ion Engine (Xe)
Hall Thruster (Xe)
PPT
VAT (Cr)
PEEP
Colloid
32
3
32
7
9
9
7
7
5
11
4
37
81
63
4
18
13
Unused
43
Total
385
80
70
60
50
5-40
30
20
10
0
0
0.5
1
2
1.5
Thrust (N)
2.5
3
3.5
Figure 3-1: Historical data for xenon Hall thrusters.
24
1.4 i
1.2
1
-0.8
20.6
0.4
0.2
0
0
2
4
6
Thrust (N)
8
10
12
Figure 3-2: Historical data for hydrazine monopropellant thrusters.
data provides the best starting point for generating a quick initial estimate of the
characteristics of an undesigned system, it should be kept in mind that once such a
system is actually designed, its characteristics could be very different, for better or
for worse.
Special care was taken to ensure that the equations were most accurate for lower
levels of thrust, where slight differences in mass have a more serious impact on the
system as a whole. Sometimes the thrusters were divided into two different regimes
and separate equations were developed for both.
However, the low thrust end of
thruster technology is currently a focus of much research, meaning that low thrust
data is few and far between, has often been gathered from experimental models, and
is subject to change in the near future.[36] However, the data gathered is still more
useful that what can be predicted by extrapolating from larger thrusters. Low thrust
technology often uses innovative designs employing different technology and throws
out assumptions that are essential to traditional methods of thruster design, which
have been developed for the high thrust regime.[115]
25
It is important to stress that although this method uses real world data, it will not
necessarily output the specifications of an actual real world thruster. The results from
this method are representative of all of the thrusters researched and will probably not
be an exact match of any of them specifically. These results are better thought of as
the specifications of a theoretical thruster that could be built given what has been
built in the past. Some of the thrusters in the database have not yet been flown or
are engineering models of thrusters still in development, but they still provide useful
data points on what can be achieved. In light of this, special care was taken not to
propagate relationships beyond the limits of the data that has been gathered. If a
thruster providing less than 1 N of thrust has never been built, this program will not
try to conjure up what such a thruster might look like. But if thrusters providing 1
and 2 N have been built, this program will predict the characteristics of a thruster
that provides 1.5 N. On the other hand, if the required thrust is higher than that
of any existing model, then the program will divide the thrust until it is in range
of existing models and design a system that includes multiple thrusters but unified
propellant and power systems.
If the user wants to know what the options are in terms of thrusters that have
already been built, the database itself is still a very useful starting point. Past users
have been able to validate results of this program by finding thrusters in the database
that are on the market and provide even better performance than what the program
estimates.
In conclusion, while the relationships developed from this database of historical
data are not perfect, they are the most logical starting point for providing a solid
backbone of historical data to the model. As will be discussed, each propulsion type
provided its own unique set of challenges.
3.1
Cold Gas Thrusters
Cold gas thrusters are the simplest type of thruster. Pressurized gas is simply released
through a nozzle to create thrust. Specific impulse (I.,)
26
is largely a function of
propellant choice, and any kind of gas can be used. Lighter gases deliver higher I,,
at the cost of lower density, resulting in heavier tanks. Nitrogen gas is usually chosen
for its lack of reactivity and balance of relatively low molecular mass and relatively
high density, although helium has also been used, as can be seen in Figure 3-3.
160
140
*
120
100
C80
60*
40
20
0
0
2
4
6
8
10
Thrust (N)
Figure 3-3: Historical data for cold gas thrusters. Most of the data is gathered from
using nitrogen gas, and the outlier is from using helium.
Because of the low I,
provided by inert gas, propellant mass is usually pro-
hibitively large, making cold gas thrusters uncompetitive compared to other options
as a primary propulsion system. Usually cold gas thrusters are chosen for their simplicity and low cost, and little thought is given to optimizing the mass of a small
valve attached to a comparatively large propellant tank. The chaotic nature of the
data reflects the crude nature of this propulsion option.
However, the simplicity, flight heritage, and comparative safety of cold gas thrusters
have made them an attractive option for small, cheap university satellites that often
face safety restrictions due to being secondary missions on larger rockets.[70] Special
attention has been paid towards propellants that can be stored as a liquid and expand
27
0.4
0.35
0.3
0.25
0.2
0.15
0.1
0.05
I
0
**
*
0
2
4
6
8
10
Thrust (N)
Figure 3-4: Historical data for cold gas thrusters.
40
35
30
25
20
a:
15
10
5
0
0
2
4
6
8
Thrust (N)
Figure 3-5: Historical data for cold gas thrusters.
28
10
18
16
14
S12
e
10
1.1
8
wo 6
LL.
4
2
+
T
0
0
2
4
6
8
10
Thrust (N)
Figure 3-6: Historical data for cold gas thrusters.
0.8
-
0.7 0.6
-
2 0.5 -
0.4
y =0.0881n(x) +0.7413
0.3686
0.3R2=
0.2
0.1
0
0
0.02
0.04
0.06
Thrust (N)
0.08
0.1
0.12
Figure 3-7: Historical data for cold gas thrusters, looking at just the low thrust
regime.
29
into a gas during emission. Liquefied gas thruster packages with integrated propellant
tanks and nozzles have recently been developed specifically for cubesats.[71]
Most of the characteristics of cold gas thrusters were predicted by simply averaging
the data that was gathered, since no relationships could be found except for I,,, which
is a constant dependent on propellant choice. For feed pressure, which is the driving
factor of tank mass (a dominating part of propellant heavy cold gas systems), an
equation was developed specifically for the low thrust regime. Even this equation is
a poor fit of the data, but it is more accurate than a flat average across the entire
range of thrust.
The program offers three propellant options for cold gas thrusters: nitrogen, helium, or butane (stored as a liquid).
thrusters are identical except for the I,.
The characteristics for nitrogen and helium
The characteristics for butane thrusters are
gathered from VACCO's MiPS thruster shown in Table 3.1. (Only three models of
liquefied gas thrusters could be found, the data on this model was the most complete,
and it is currently on the market for use on cubesats.) All characteristics except for
feed pressure for low thrust helium and nitrogen thrusters are constant.
Propellant
Butane
Max Thrust (mN)
Mass (g)
I,, (s)
Min Impulse Bit (mNs)
55
456
65
0.25
Max Propellant Storage (g)
53
Table 3.2: Specifications of VACCO MiPS butane thruster unit with integrated propellant tank for cubesats.[38]
3.2
Monopropellant Thrusters
Monopropellant thrusters use the chemical decomposition of a single propellant, historically either hydrazine or hydrogen peroxide, to create thrust.
The propellant
flows through a heated catalytic bed in the combustion chamber to initiate the reaction.
Hydrazine delivers a higher specific impulse, but because it is toxic and
30
1.4
1.2
1
-0.8
20.6
0.4
0.2
0
0
2
4
6
Thrust (N)
8
10
12
Figure 3-8: Historical data for hydrazine monopropellant thrusters.
unstable, hydrogen peroxide is sometimes chosen instead. As with other chemical
thrusters, I., is largely a constant dependent on propellant choice, although some
models have decreased feed pressure at the expense of I., in order to optimize for
lighter tanks. Although monopropellant thrusters are certainly a chemical and not
an electric propulsion system, higher thrust does correlate to higher electrical power
requirements of the heated catalytic bed.[116]
Monopropellant thrusters are a natural choice for secondary thrusters or station
keeping on large systems. [5][35][90][30] Only one propellant tank is required, specific
impulse is substantially higher than that of cold gas thrusters, and electrical power requirements are minimal. Recent research has focused on miniaturizing this propulsion
option down to the micro- and nanosat range; however, utilization of monopropellant
thrusters in this regime is dependent on the miniaturizing of the associated valves.
If developed, monopropellant thrusters could provide an attractive option for a dual
propulsion system on small satellites by attaching high and low thrust thrusters to a
single propellant tank.
31
240
'
220
-
200
0
y = 206.21x 048
2
R = 0.5934
180 A
160 4
140
0
2
4
6
Thrust (N)
8
10
12
Figure 3-9: Historical data for hydrazine monopropellant thrusters.
4
3.5
3
$
2.5
2 2.
y = 0.0026eoc
lmm
2
R =0.7599
t.2
v 1.5
0.5
0
140
160
180
200
220
240
Isp (s)
Figure 3-10: Historical data for hydrazine monopropellant thrusters.
32
4035 30 25 -
0 20
y = 28.684x0-7943
R2 = 0.8878
20
10
5
0
0
0.2
0.4
0.6
0.8
Mass (kg)
1
1.2
1.4
Figure 3-11: Historical data for hydrazine monopropellant thrusters.
0.07
.
0.06
0.05
7
y = 0.0077e'.4 x
2
=0.5954
0.04 -R
20.03
0.02
0.01>
0
*
0.2
0.4
0.6
Thrust (N)
0.8
1
1.2
Figure 3-12: Historical data for hydrogen peroxide monopropellant thrusters.
33
160
-
150
y = 136.76x 0"" 9
R2 = 0.6005
140
-
130
120
110
100
0
0.2
0.4
0.6
Thrust (N)
0.8
1
1.2
Figure 3-13: Historical data for hydrogen peroxide monopropellant thrusters.
0.7 -r0.6 -
0.5
0.4
0 01
y = 0.1034e -.
0.8267
0.3 -R2=
0.2
0.1
0
80
100
120
140
160
Isp (s)
Figure 3-14: Historical data for hydrogen peroxide monopropellant thrusters.
34
2.5
2
y =0.4715el-3&"
0.9848
1.5
CLI
0.5
.-
tr
0
0
0.2
0.4
0.6
Thrust (N)
0.8
1
1.2
Figure 3-15: Historical data for hydrogen peroxide monopropellant thrusters.
Two separate sets of equations were developed for thrusters using hydrazine and
hydrogen peroxide. Although trends could be identified, the data is still relatively
scattered. This suggests that either other factors are at work in the design of monopropellant thrusters, or that they have not been optimized as much as they could be,
perhaps because I., is still low enough that propellant tank mass dominates the system. Theoretically, monopropellant thrusters could be miniaturized and integrated
with propellant storage as cold gas thrusters have been, although not to the same
extent, since the heat gradient caused by propellant decomposition becomes harder
to manage as mass and volume decrease. Issues with propellant storage and safety
would still prohibit their use on many small satellites, which tend to be launched
as a secondary payload on a rocket whose main customer would not welcome the
risk of a volatile propulsion system stored next to their larger and more expensive
spacecraft.[70] Despite these barriers, Aerojet has been developing a 1U hydrazine
thruster module for cubesats in accordance with most range safety requirements. [91]
35
3.3
Bipropellant Thrusters
Bipropellant thrusters, the powerhouses of space propulsion, use the chemical reaction
of two propellants to create thrust. These thrusters are unmatched in terms of thrust
to weight ratio and boast the highest I,
of any type of purely chemical propulsion.
However, they are also the most complicated chemical propulsion option, requiring
two propellant tanks and the associated equipment required for pressurization and
fuel delivery.
While often the only viable option for space lift and orbit insertion, the complexity of bipropellant thrusters does not lend itself to miniaturization and low thrust
applications. Thrusters that do provide less than 10 N are usually used as secondary
propulsion, and as the chaotic nature of the data suggests, have not been optimized for
mass. [27] [23][5] While there has been initial investigation into miniaturizing bipropellant thrusters using microelectricmechanical systems (MEMS) technology, propellant
storage is still a barrier to implementation. [71] [561
On the other end of the spectrum, electric propulsion often outperforms bipropellant thrusters on large satellites if the AV requirement is high. While bipropellant
thrusters have the highest I,
out of any chemical propulsion option, it is still far sur-
passed by most electric propulsion options. A high AV requirement often demands
a high I,, to keep the propellant and tank masses from becoming prohibitively large.
These factors restrict bipropellant thrusters to high thrust operations, most of which
are over 10 N and outside the scope of this research.
Although data from all available models was averaged to predict thruster characteristics, the program assumes monomethylhydrazine as fuel and nitrogen tetroxide
as oxidizer in its propellant storage calculations. Power requirements were only found
for one thruster and feed pressures were only found for two.
36
0.8
0.7
0.6
0.5
0.3
0.2
0.1
0
0
2
4
6
8
Thrust (N)
10
12
14
Figure 3-16: Historical data for bipropellant thrusters.
325
320
315
310
305
e
300
295
290
285
280
275
270
0
2
4
6
8
Thrust (N)
10
12
14
Figure 3-17: Historical data for bipropellant thrusters.
37
3.4
Resistojets
Resistojets improve the performance of cold gas thrusters by heating the propellant
using electric resistors.
This method raises the I,
of the thruster, decreasing the
tank and propellant masses, but adds the complexity and mass associated with the
electrical power requirements.
The first widely used resistojets were actually modified monopropellant thrusters,
and as a hybrid of chemical and electrical thrusters were called electrothermally augmented hydrazine thrusters (EHTs). The development of EHTs arose from the convergence of stringent mass limits and surplus power supply on communication satellites
that were already using monopropellant hydrazine thrusters. [35] [90] [30] However, resistojets can use inert gas as well. Often propellants that can be stored as a liquid
and easily vaporized are chosen, such as water or ammonia.
The heat augmentation process raises I.
up to a point and then levels out, mean-
ing that the specific impulse of resistojets is largely a function of propellant choice.
There is a relatively strong correlation between thrust and power, although low power
resistojets can have abnormally high thrust to power ratios due to the mechanical
thrust provided by propellant pressure.
Resistojets are often chosen instead of cold gas thrusters in cases where spare
electrical power is already available.
This similarity of application with cold gas
thrusters, combined with the greater variety of propellants used, makes the data
difficult to analyze.
Resistojets are also amiable to miniaturization for the same reasons as cold gas
thrusters, although limits are reached due to the difficulty of transferring heat to
the propellant within a smaller chamber. The increase in specific impulse and using
liquid propellant can keep mass low, however, this must be balanced against the
higher power requirements. On the high thrust end of the spectrum, resistojets are
limited by the material properties of the heating element, a barrier which is overcome
by arcjets.
Two sets of equations for resistojets have been developed, one for hydrazine and
38
1000
900
800
700
-600
d 500
400
2
y = 373.54x + 770.3x + 52.949
R = 0.7676
300
200
200
100
0
0
0.2
0.4
0.6
0.8
1
Thrust (N)
Figure 3-18: Historical data for hydrazine resistojets.
1
0.9
0.8
0.7
0.6
10.5
20.4
0.3
0.2
0.1
0
0
200
600
400
Power (W)
800
1000
Figure 3-19: Historical data for hydrazine resistojets.
39
320
300
280
y = 186.93x0.0s
R2 = 0.9485
VI
.260
240
220
200
0
200
400
600
Power (W)
800
1000
Figure 3-20: Historical data for hydrazine resistojets.
2.8
2.6
2.4
. 2.2
2
a.
y = 0.8191n(x) + 2.8825
R= 0.7511
1.8
z
1.6
1.4
1.2
1
0
0.2
0.4
0.6
0.8
Thrust (N)
Figure 3-21: Historical data for hydrazine resistojets.
40
1
1600
1400
1200
R2 = 0.9263
1000
i
800
600
400
200
0
0
0.05
0.1
0.15
0.2
0.25
Thrust (N)
0.3
0.35
0.4
Figure 3-22: Historical data for water resistojets.
3.5
3
2.5
no2
1.5
1
0.5
0
0
200
400
600
Power (W)
800
Figure 3-23: Historical data for water resistojets.
41
1000
240
220
200
180
n.
160
140
120
100
0
200
400
600
Power (W)
800
1000
Figure 3-24: Historical data for water resistojets.
2.5
2.0
y = 4.1265x + 0.1122
R2 =0.4415
~1.5
~-1.0
0.5
0.0
0
0.05
0.1
0.15
0.2
0.25
Thrust (N)
0.3
0.35
Figure 3-25: Historical data for water resistojets.
42
0.4
one for water. The sparse data for water resistojets raises the question as to whether
they could be further optimized for small satellites, since they don't carry the risks
that monopropellants do. Resistojets also enjoy an abundance of other propellant
options that could open up further possibilities. A team at the University of Arkansas
recently developed a 1U resistojet propulsion module for cubesats that uses R-134a
refrigerant. [69]
3.5
Arcjets
Arcjets are very similar to resistojets but use an electric arc to heat propellant instead
of resistors. This allows for higher input power thresholds and correspondingly higher
thrust and specific impulse, although efficiency is compromised. As with resistojets,
ammonia is a common propellant choice. Despite the storage problems associated
with light gases, some research has focused on creating high power hydrogen arcjets
with the expectation that large manned space systems would have an excess on board
due to boil off from cryogenically stored hydrogen.[7]
As with resistojets, the first arcjets used hydrazine to take the place of hydrazine
monopropellant thrusters. In a monopropellant hydrazine thruster, hydrazine is decomposed into ammonia and nitrogen in an extremely exothermic reaction, while
some fraction of ammonia is further decomposed into nitrogen and hydrogen in an
endothermic reaction. Good design minimizes this fraction in order to keep the exhaust gases as hot as possible. However, the addition of heat in resistojets and arcjets,
while more than compensating for the heat lost in the decomposition of ammonia, can
actually raise the fraction decomposed to the point that there is no real difference between using hydrazine and hydrogen nitrogen mixture as propellant.[66] In hydrazine
arcjets, for example, the heat added through electrical means totally decomposes the
ammonia and drowns out much of the benefit that the first chemical decomposition
of hydrazine might still have to offer. The question designers are faced with is why
to use hydrazine at all given its toxicity and instability. I,,
is still dependent on
propellant choice, especially at high power levels, but it is the molecular mass of the
43
80
70
60
y = 0.7834x2
R = 0.6326
50
v40
30
20
10
0
0
20
40
60
Power (kW)
80
100
120
Figure 3-26: Historical data for arcjets using various propellants.
propellant (or its products after decomposition) that are decisive rather than any
chemical reaction.
Arcjets straddle the middle ground between purely chemical bipropellant thrusters
and high powered Hall thrusters or ion engines in terms of I
and required power,
striking a balance between tank mass and power system mass. As with resistojets,
the modeling process is complicated by the variety of the propellants used, but the
high dependency performance characteristics have on electrical power makes the data
much more organized. Three sets of equations have been developed for ammonia,
hydrogen, and hydrazine propellants. Unfortunately, no mass data could be found
for ammonia or hydrogen arcjets and feed pressures could only be found for hydrazine
arcjets. The mass data for all of the arcjets, regardless of propellant, was propagated
out to the maximum power recorded to estimate thruster mass, as shown in Figures
3-26 and 3-27. It is uncertain how accurate this equation is in the high thrust, high
power regime. It is likely an overestimation of thruster mass, which typically increases
slightly less than linearly with relation to power.
44
2
1.8
1.6
1.4
1.2
vsI
0.8
y = 0.001xO-9677
0.6
0.R=
0.4
0.6326
0.2
0
0
500
1000
1500
Power (W)
2000
2500
3000
Figure 3-27: Historical data for arcjets using various propellants, zooming in on the
low thrust regime.
30
25
20
20
y =5.8358x2 +1.1817x + 0.2925
R2 = 0.9998
15
10
0
0
0.5
1
1.5
2
Thrust (N)
Figure 3-28: Historical data for ammonia arejets.
45
2.5
900
800
700
~600
500
400
300
0
5
10
15
Power (kW)
20
25
30
Figure 3-29: Historical data for ammonia arcjets.
120
100
80
y = 18.886x'.0172
R2 = 0.9592
-60
40
20
0
0
1
2
3
4
Thrust (N)
Figure 3-30: Historical data for hydrogen arcjets.
46
5
1600
1500
1400
1300
1200
.C1100
0 0
y =1270.6x -1 9
1000
R2 = 0.9844
900
800
700
600
0
1
2
3
4
5
Thrust (N)
Figure 3-31: Historical data for hydrogen arcjets.
2.5
2
0.5
0 '0.05
0.1
0.15T
0.2
Thrust (N)
0.25
Figure 3-32: Historical data for hydrazine arcjets.
47
0.3
1.7
1.6
1.5
1.4
11.3
1.2
1.1
1
0.9
0.8
1
1.2
1.4
1.6
Power (kW)
1.8
2
2.2
Figure 3-33: Historical data for hydrazine arcjets.
650
600
550
-
_500*
450
400
350
0
0.5
1
1.5
Power (kW)
2
Figure 3-34: Historical data for hydrazine arcjets.
48
2.5
1.85
1.8
1.75
S1.7
1.65
.6
1.55
1.5
1.45
0.05
0.1
0.15
0.2
0.25
0.3
Thrust (N)
Figure 3-35: Historical data for hydrazine arcjets.
3.6
Ion Engines
Ion engines accelerate ionized gas using an electric field created by a potential difference between two grids in order to create thrust. The gas is usually ionized using
electron bombardment or radio frequency-induced gas discharge in a chamber before
being accelerated through biased grids.
Electric propulsion in general favors propellants with high molecular mass and low
ionization potential. Early ion engines used mercury or cesium, which were largely
abandoned due to their toxicity and tendency to contaminate the spacecraft.[88] Most
ion engines today use xenon, an inert gas which is much easier to handle. The
equations have been developed using data only from xenon ion engines.
In comparison with the thrusters discussed so far, ion engines offer a much higher
specific impulse but limited thrust to power ratio. The high specific impulse translates
to propellant mass savings, but high power requirements in turn drive up the mass
of the required power system. Most models are clustered around the low power end
of the spectrum, but recent research has focused on high power models for use in
49
45
40
35
y = 64.664x2 + 18.216x + 0.0462
30
R2 = 0.9848
25
20
C.15
10
5
0
0
0.1
0.2
0.3
0.4
Thrust (N)
0.5
0.6
0.7
0.8
Figure 3-36: Historical data for xenon ion engines.
50
45
40
y
61.162x+0.6/61
R2= 0,9152
35
30
25
20
15
10
5
0
0
0.1
0.2
0.3
0.4
Thrust(N)
0.5
0.6
0.7
Figure 3-37: Historical data for xenon ion engines.
50
0.8
12000
10000
8000
-C 6000
= -4.3967x2 + 357.29x + 2661.7
FAy
R2 =0.8464
4000
*
2000
0
0
10
30
20
Power (kW)
40
50
Figure 3-38: Historical data for xenon ion engines.
interplanetary travel. [25] [39]
As with many of the electric propulsion options to be discussed, the significantly
reduced propellant mass brought on by higher I.,, and the complexity of the thruster
makes the thruster itself a more significant contributor to the mass of the system as
a whole than in chemical systems. This means that ion engines have been the focus
of quite a bit of optimization, especially in the low thrust regime, and that the data
lends itself to tight, orderly relationships.
3.7
Hall Effect Thrusters
Hall thrusters accelerate ionized gas using a combination of electric and magnetic
fields in order to create thrust. Propellant is fed into a ring shaped chamber with
magnets arranged to create a radial magnetic field and an anode and cathode arranged
to created an axial electric field pointing out of the open end of the thruster. Ions
are accelerated out of the thruster towards the cathode by the axial electric field.
51
80
70
y
0.8441x3 - 0.0386x2 + 18.768x +
0.1972
2
R = 0.9694
60
.. 50
S40
30
20
10
0
0
0.5
1
1.5
2
Thrust (N)
2.5
3
3.5
Figure 3-39: Historical data for xenon Hall thrusters.
The electrons, instead of accelerating in the opposite direction towards the anode,
are trapped in a Hall current by the perpendicular electric and magnetic fields and
exert a force on the magnets which propels the thruster. Some of the electrons make
it to the anode and are in turn propelled out of the cathode to neutralize the exhaust
plume.
Xenon is usually chosen as a propellant for its high molecular mass, low ionization
potential, and because it is inert and easy to handle. The possibility of using bismuth
has been explored in high power hall thrusters for it's even higher molecular mass,
lower ionization potential, and lower cost, often resulting in higher I,.[92] However,
not enough data on bismuth hall thrusters exists to identify any relationships and the
equations were developed using data from only xenon Hall thrusters.
Research on Hall thrusters has been continuing for decades and is now concentrated on the high and low power ends of the spectrum. The consistent, ordered data
reflects the high level of effort directed towards the optimization of Hall thrusters,
and the sudden decrease in performance in the low power regime is indicative of the
52
90
80
70
60
150
40
30
20
10
0
10
0
20
30
50
40
Power (W)
60
70
80
Figure 3-40: Historical data for xenon Hall thrusters.
4500
4000
3500
3000
2500
C.
22000
1500
1000
500
0
0
10
20
30
50
40
Power (kW)
60
70
Figure 3-41: Historical data for xenon Hall thrusters.
53
80
the challenges of miniaturizing this propulsion option. In general, hall thrusters offer
higher thrust but lower I,, than ion engines.
3.8
Pulsed-Plasma Thrusters
Pulsed-plasma thrusters (PPTs) also use a combination of electric and magnetic fields
to accelerate propellant. Arc discharges are pulsed through the propellant, which is
accelerated by the force created from the current of the discharge and its associated
magnetic field.
PPTs are low thrust, low efficiency, and low Ip compared to ion engines. Their
advantage is their simplicity and low power requirements. Most models use solid
Teflon as propellant, which is ablated and vaporized by the discharge. This makes
propellant storage on PPTs extremely simple and reliable. The low minimum impulse
bit resulting from the pulsed firing of PPTs also makes them a good choice for attitude
control.
300
250
0
y = 56.634xo 6e
R2= 0.867
200
150
0
100
50
0
0.00
1.00
2.00
3.00
Thrust (mN)
4.00
5.00
Figure 3-42: Historical data for PPTs.
54
6.00
9
8
7
6
15
y=
4.9919xO.2-57
R2 = 0.2343
3
2
1
0
0.00
1.00
2.00
3.00
Thrust(mN)
4.00
5.00
6.00
Figure 3-43: Historical data for PPTs.
3000
2500
y = 917.28xo.29 1
R2 = 0.4414
2000
-1500
IA
1000
500
0
0.00
1.00
2.00
3.00
Thrust (mN)
4.00
5.00
Figure 3-44: Historical data for PPTs.
55
6.00
Although trends can be observed in the characteristics of PPTs, the data is still
very chaotic and suggests that further optimization is possible. In addition, thrust
to power ratio was found for many high I, models without any accompanying max
thrust or max power data, implying that these relationships are not representative of
PPTs' full potential.
3.9
Vacuum Arc Thrusters
Vacuum Arc Thrusters (VATs) are very similar to PPTs except that the discharge
ablates propellant off of the cathode itself, which replaces the Teflon block and retains
the advantage of solid propellant storage. Performance is very similar to that of PPTs
but with higher Ip, although some models utilize permanent magnets to substantially
increase thrust. Very few VATs have been built, but they have been tested and
scheduled to fly on microsatellite missions.
Propellant
Max Thrust (mN)
Mass (g)
IS, (s)
Min Impulse Bit (pNs)
Chromium
1
100
2000
0.25
Table 3.3: Specifications of Alameda Applied Sciences Corporation pVAT for Illinois
Observing NanoSatellite (ION).
Since data on only four VATs could be gathered, two of which were augmented
with magnets and had very different characteristics, only one was chosen to represent
this propulsion option. The pVAT, built by Alameda Applied Sciences Corporation,
was chosen by the University of Illinois for deployment on their satellite, ION.
3.10
Field-Emission Electric Propulsion
A form of electrostatic propulsion like ion engines, field-emission electric propulsion
(FEEP) extracts and propels ions from liquid metal propellant (usually cesium or
56
120
y = 72.257x'o5
R2 = 0.9161
100
80
~60
0
CL4
0
20
0
0.00
0.20
0.40
0.60
0.80
1.00
Thrust (mN)
1.20
1.40
1.60
Figure 3-45: Historical data for FEEPs.
indium) using a potential difference. The potential difference causes Taylor cones on
the surface of the propellant, which in turn allows the emission of ions from the tips.
FEEPs offer the highest I, and the smallest minimum impulse bit out of any
propulsion option, but thrust is limited to the micronewton and low millinewton
regime. Most applications of FEEPs take advantage of their low minimum impulse bit
to provide high accuracy pointing or control, but their high I,, makes them favorable
for any high AV, low thrust mission.
One difficulty with FEEPs is propellant management, since the indium or cesium
must be melted before emission. On the other hand, solid propellant lends itself to
simple storage, and since the thruster is emitting ions, the propellant can be fed by
capillary forces and requires no pressurization.
Except for I,,, the data for FEEPs is relatively consistent, though sparse.
57
6
5
0
y = 0.2141xo I
R2 = 0.6261
4
3
2
1
0
0
20
40
60
80
100
Power (W)
Figure 3-46: Historical data for FEEPs.
14000
12000
.
10000
8000
--
6000
4000
e
2000
0
0
20
40
60
80
Power (W)
Figure 3-47: Historical data for FEEPs.
58
100
3.11
Colloid Thrusters
Colloid thrusters are very similar to FEEPs except that they extract and accelerate
charged droplets instead of ions and use different propellant. Colloid thrusters were
one of the earliest forms of electric propulsion to be investigated, but the lack of
suitable propellants inhibited their development until the invention of liquid salts at
room temperature.
Today's colloid thrusters offer a higher thrust to power ratio than FEEPs, ion
engines, and Hall thrusters but a lower specific impulse. A major advantage of colloid
thrusters is the ease of propellant management. Liquid salts are easily manufactured
with the right equipment, require little pressurization, and are safe to handle and
store.
Figure 3-48: Historical data for colloid thrusters.
Unfortunately, no relationships can be discerned from the data on colloid thrusters
and flat averages must be taken in order to predict most characteristics. This is probably because most of the data is from old research before liquid salts were invented,
and is hardly indicative of the potential of colloid thrusters. Although a variety
59
0
1
2
3
4
Power (W)
56
7
Figure 3-49: Historical data for colloid thrusters.
1600
1400
1200
1000
C800
600
400
200
0
0
2
4
6
Power (W)
8
10
Figure 3-50: Historical data for colloid thrusters.
60
12
of propellants have been used, the program assumes the liquid salt EMI-BF4 in its
propellant storage calculations.
3.12
Electrospray Thrusters
Electrospray thrusters are conceptually identical to FEEPs except that they use liquid
salt as propellant, and identical to colloid thrusters except that they emit ions instead
of droplets. In fact, it may be possible to develop a single thruster using the same
propellant that can act as both a colloid thruster and an electrospray.
Although electrosprays offer a lot of promise, they have not yet been flown in space.
Current research has achieved thrust and specific impulse levels similar to standard
ion engines. Propellant management is identical to liquid salt colloid thrusters except that the propellant can be passively fed into the thruster using just capillary
forces. Miniaturization of electrosprays using MEMS technology could enable them
to replace ion engines in many applications due to their low mass and easy propellant
management.
Another unique characteristic of MEMS electrosprays is their scalability. Electrosprays consist of hundreds of microscopic needles from which liquid salt is propelled
using a voltage difference. Scaling the thruster is simply a matter of creating more
needles. For low thrusts, this can be achieved by increasing the density of the needles,
and the mass of the thruster remains constant. At some point the a density limit is
reached and larger area is required, at which point mass will increase with thrust.
Efficiency and specific impulse remain constant throughout.[55]
Because of both their scalability and lack of historical data, this research models
electrospray thrusters using constant scaling equations instead of equations derived
from historical specifications. Even if there was extensive flight data on electrosprays,
it would be more logical to use that data to find constants for the known scaling
equations than to work entirely off of historical specifications, as has been done for
the other propulsion types. The constants used in this research were measured from
MIT Space Propulsion Laboratory's MEMS electrosprays, which are due to be tested
61
in orbit by 2014.
Propellant
Specific Mass (kg/N)
I,, (s)
Efficiency (%)
EMI-BF4
6.32
3500
80
Table 3.4: Specifications of MIT Space Propulsion Lab's electrospray arrays.
62
Chapter 4
Modeling Propellant Storage and
Management
The database and the resulting equations provide the tools needed to quickly and
repeatedly estimate the mass of a given thruster from the required thrust. However,
as shown by the focus on developing high I, electric propulsion systems, the mass
of the propellant and the storage and management requirements that come with it
is often a driving force in propulsion design and a deciding factor when choosing a
propulsion system. Fortunately, the physics of how much propellant is needed and
how big its tank needs to be are relatively simple.
The equations developed in the previous chapter can provide I,,. From there, the
mass of the propellant (mp,p) is found from the required AV and the initial mass of
the spacecraft (ms/c) (provided by the user) using the rocket equation, where g is
the acceleration of 9.81 m/s 2 due to gravity.
M,,.4 =
ms/c(l - e6
)
3P9"v1
(4.1)
The mass of the propellant tank (mtank) is then calculated by assuming spherical
tanks with a factor of safety (FoS) of two and some given material properties. The
material properties are read from within the database and can tuned specifically to
each thruster type. Titanium alloy is preferable, but aluminum alloy was used in
63
Yield Strength - Uyield (MPa)
Density - Ptank (kg/m)
Propellants
Titanium
6Al-4V
Aluminum
2014-T6
924
4430
Nitrogen
Helium
Water
414
2790
Hydrazine
Hydrogen Peroxide
Hydrogen
Ammonia
EMI-BF4
Xenon
Table 4.1: Choosing a tank material.
light of the chemical properties of certain reactive propellants.
Under these assumptions, the mass of a propellant tank can be calculated using
the equations for thin walled pressure vessels and the volume of a sphere. All that's
left is to find the volume (V,,.p) and maximum pressure (Pma)
mtank =
ttank =
rtank =
41r[(rtank
- ttank)3 -
FoS ma2 rtank
rtank]ptank
within the tank.
(4.2)
(4.3)
207yield
L
(4.4)
3
rtank and ttank represent the radius and wall thickness of the tank, respectively.
All tank masses calculated by this method are multiplied by two to account for
other factors such as acceleration and dynamic loads, stress concentrations, and weld
efficiencies. [35]
For thrusters that use pressurized gas (nitrogen, helium, hydrogen, and xenon), the
propellant is assumed to be an ideal gas, the minimum pressure (Pmin) is set to the feed
pressure estimated from the database, and a leftover mass of propellant (mleftoverprop)
is accounted for that will fill the propellant tank at the required minimum pressure.
The mass of the leftover propellant is added to the mass of the tank, and the following
64
equations along with equations 4.2, 4.3, and 4.4 are optimized for minimum tank mass.
mltoverprop = minitialpro, - myo,
(4.5)
R
(4.6)
_maminitialproRT
(4.7)
Vr=
mi
minitialpro represents the mass of all of the propellant initially stored in the tank and
m,
represents just the mass of the propellant that's expelled in order to achieve
the required AV. R, T, and M are, respectively, the universal gas constant (8.3145
J/mol/K), temperature of propellant (set to 293.15 K), and molar mass of each the
propellant (varies by propellant).
For thrusters using pressurized liquid propellant (water, ammonia, hydrazine, and
hydrogen peroxide), the propellant tank is assumed to be at constant feed pressure
and mass is estimated using equations 4.2, 4.3, and 4.4 by calculating the volume
from the density of the liquid (pprp).
V,-,
= m'""
(4.8)
Pprop
A titanium nitrogen pressurant tank is then modeled using a process similar to
that for pressurized gas thrusters except that the mass of the gas (mpres,) is constant,
final volume is the combined volume of the pressurant (Vank) and propellant tanks
(Vop), and minimum pressure is the pressure of the propellant (which has already
been set to the feed pressure for the thruster). As with the pressurized gas tank
method, the mass of the pressurant is added to the mass of the pressurant tank and
the equations are optimized with 4.2, 4.3, and 4.4 for minimum pressurant tank mass.
mp.ess
-
PPro(V±op+ Vank)M
mesRT
(4.9)
(.0
(4.10)
Pmax = m,ek
For bipropellant thrusters, two liquid propellant tanks are calculated, one for the
65
fuel and one for the oxidizer, and a single pressurant tank with enough nitrogen to
fill all three tanks at the feed pressure is calculated.
Colloid thrusters are pressurized, but only slightly. Many models use pumps
instead of pressurization, and recent models have used a compressible tank attached
to a spring. This program calculates the mass of colloid thruster tanks the same as
other liquid propellants (equations 4.2, 4.3, 4.4, 4.8) but neglects the pressurant tank.
For thrusters that do not use solid or unpressurized liquid propellants (PPTs,
VATs, FEEPs, and electrosprays) the propellant system mass was estimated to be
5% of propellant mass.
Finally, 10% is added onto the combined mass of the thruster and tanks of all
propulsion types to account for feed lines and support structures.
The mass of the power system is also estimated using an assumed power density
of 15 kg/kW. This mass includes power source and processing, and is representative
of solar panels.[35]
66
Chapter 5
Validation
Several difficulties are encountered when validating the results of the program. The
first is that complete mass breakouts of propulsion systems can be difficult to find.
The second is that when such data is found, it's often from systems that violate
assumptions used in the program. For example, many propulsion systems use multiple
thrusters even when a single thruster with the necessary thrust has been designed,
either for redundancy or for attitude control. Some spacecraft end their missions
without using all their propellant, and the achieved AV actually corresponds to a
much smaller tank. Usually the tanks themselves are not spherical or made out of a
single material, which are assumed in the tank estimation process. All of these are
indicative of the fact that there are many factors in propulsion system design that this
program does not take into account, such as cost, reliability, volume, and schedule.
Quite simply, the program predicts what is possible, but not necessarily what is.
That said, comparing the results with what data can be gathered from real world
systems yields some interesting results. The four systems in Table 5.1 are those for
which the most complete specifications could be found.
5.1
Aerojet Cubesat Propulsion Module
The first system is a 1U hydrazine propulsion module being developed by Aerojet
for cubesats. This is an example of a system for which the technology exists and
67
Propulsion Type
Thrust (N)
dV (m/s)
Mass: S/C (kg)
Propulsion (kg)
Propellant (kg)
Thruster (kg)
Tanks (kg)
Isp (s)
Power (W)
Max Pressure (MPa)
Tank Volume (m3)
Aerojet Cubesat
Propulsion Module
Hydrazine
Monopropellant
P
2.81
4501
2
BepiColombo
Xe Ion EnRine
1
0.1451
57601
4100
Uof Minnesota
Shackleton Crater
MMH/MON-1
Bioronellant
41
102111
445
Cyde Space
Teflon PPT
100.E-051
3.53
Table 5.1: Comparison of results from program with actual propulsion systems.
feasible thrusters are recorded in the database, but which has never been integrated
before into such a tight package. Aerojet uses MR-140A thrusters, for which no mass
data could be found and are therefore not included in the database.
The inputs
in yellow are the capabilities Aerojet advertises for this system on a hypothetical
cubesat. Although Aerojet is still working to make sure that the system meets range
safety requirements, the program confirms that the design is feasible.[91]
There are a few key differences between the system designed by Aerojet and the
one that the program predicts. First of all, Aerojet's system is volume limited and
designed to fit entirely within a 10 cm cube, which changes the shape of the tank
from an optimal sphere. Second, Aerojet's system cannot provide its max thrust of
2.8 N at end of life, whereas the program designs a system so that the max thrust
can still be achieved until all the propellant needed to achieve the required AV has
been expelled.
This accounts for the vast difference in maximum tank pressures,
since both are designed with blow down propellant feed systems. The first difference
should result in Aerojet's system being heavier, and the second should result in it
being lighter than what the program predicts. Another difference is that Aerojet's
system actually includes four thrusters in its mass figures, but the thrust value is for
an individual thruster. The calculated numbers have been adjusted so that the mass
of the predicted thruster is multiplied by four to account for this.
68
This comparison is especially exciting because the program has predicted a system
that is on the edge of what is feasible, and it just so happens that a very similar system
is in the process of being built and tested. It validates the program's ability to predict
systems in the low thrust regime, where much of the data is from experimental systems
and the margin for error is especially tight.
5.2
BepiColombo Ion Engines
The second system is the electric propulsion system on BepiColombo, a European
Space Agency mission to Mercury that will be launched in 2013 and will take 6
years and 5760 m/s AV to complete the journey from Earth. This is an example
of a fairly straightforward electric propulsion system within the range of what has
been accomplished before. It uses T6 ion engines, which are well documented in the
database.[271[111]
The data for BepiColombo was assembled from two sources which were not in
complete agreement. It is possible that the numbers were from different periods of the
design phase, or that they actually referred to different things. Mass of the propulsion
system, for example, could be taken to mean simply the mass set aside for propulsion
or the mass of the propulsion module designed to ferry the spacecraft from Earth
to Mercury, including guidance and control. BepiColombo actually consists of two
spacecraft on a single propulsion module which will separate upon reaching Mercury,
and it is unclear whether some of the numbers refer to combined propulsion on all
three systems or just the propulsion module.
Regardless, this case scenario provides another interesting comparison. The I,
for BepiColombo's engines is higher than that predicted by the program. This is to
be expected since the program estimates Ip from its historical correlation with thrust
and power, but in reality I,, is somewhat independent in the design process. The
mission designers no doubt picked a higher I,, to design for because of the large AV.
The probability that BepiColombo plans to keep some propellant in reserve after it
reaches Mercury can account for the higher propellant and propulsion masses. Data
69
on the tanks was not found, so it is unsure what effect they have on the results.
5.3
University of Minnesota Shackleton Crater Reconnaisance Mission Bipropellant System
The third system is a design of a reconnaissance mission to Shackleton Crater on
the south pole of the moon by students at the University of Minnesota. This is an
example of a chemical propulsion system that is within the range of many systems
that have been built in the past.[23]
Although the spacecraft uses the 420 N EADS Astrium S400-12 as its main
thruster to provide the 1021 m/s of AV, it also uses 4 N EADS Astrium S4 as
guiding thrusters. The mass budget for one 4 N thruster was added to the mass
budget for tanks and propellant for comparison. This assumes that the 4 N thruster
provides the same I,, as the 420 N, although it as actually quite lower.[5] As can be
expected, the mass of propellant is predicted to be higher than the what is budgeted
for the Shackleton mission, along with the tank masses and total mass of propulsion.
The difference between the actual and predicted total propulsion mass is almost entirely accounted for by the discrepancy between propellant and tank masses. The
pressure of the tank reveals that the propellant storage system is different than the
one the program designs, perhaps using pumps. In addition, it is known that the
actual design uses composite tanks.
Still, the programs results are remarkably similar to the actual design. This shows
the validity of the model within a well known regime for chemical systems just as the
BepiColombo example does for electric systems.
5.4
Clyde Space PPT Module
The PPT propulsion module build by Clyde Space is a bit of a different case. First of
all, it should be noted that the only data found for this system was for mass, power,
and I,.[57] The inputs were chosen to predict the same propellant mass at the lowest
70
mass and power.
Glancing at the database (see Table A.13), it is obvious why the program predicts
higher power and mass requirements than Clyde Space's module. No thruster that
requires that little power is recorded. Furthermore, Clyde Space doesn't provide any
thrust data on this model, so it can't be used to extend the equations for PPTs into
this regime. The program is already using the lowest thrust recorded in the database,
and this system is simply outside the range of what has been recorded.
This example illustrates a limitation of the program: it can only predict within
the regime of what has been recorded. Clyde Space's PPT propulsion module provides an I.,, that surpasses most other PPTs at a highly optimized mass and power
requirement. Although the program returns reasonable characteristics, it should always be kept in mind that it is possible to design a system with characteristics much
worse, and sometimes much better, than predicted.
71
72
Chapter 6
Applications
As previously described, one of the most useful applications of this program is within
larger simulations where mass and propellant fractions of hypothetical spacecraft
must be generated for a range of scenarios.
The program was initially written as
part of a parametric model for DARPA's System F6 project that analyzed collision
avoidance in satellite clusters. An example of the results generated by the program
as it was then can be seen in Figure 6-1.
104.
10
10
10
Average Propulsion System Mass Fraction
0
10
Figure 6-1: Results from application of a preliminary version of the program for the
System F6 project.
73
These results were generated for a range of mission profiles in which the mass of
the satellite and the required acceleration were varying. At that point, the program
included equations for cold gas, monopropellant, and Hall thrusters, and did not
differentiate according to propellant.
When the mass of the satellite was brought into the nanosatellite range (1-10 kg),
the program showed that monopropellant thrusters were the most feasible option.
Thrusters that met the requirements were found in the database that were being
sold by Micro Aerospace Solutions.[94] Later it was found that Aerojet was in the
process of designing an integrated system that would fly on satellites down to 2 kg in
mass.[91]
Lowest Mass Propulsion Option for 100 kg Satellite
10,
X
+
+
+
+
0
+
+
+
+
+
+
+
+
+
X
+
x
+
+
+
+
x
x
x
+
+
+
X
x
x
x
x
x
x
+
+
+
+
+
+
+
+
+
+
-
+
102+
+
+
+
+
+
x '
-
+
+
+
+
+
x
+
+
x
+
+
+-
+-
+
+
+-
+
+-
-+
+
+
+
-
+ ++
+
+~
+
+
+K+
+
+
+
+
+
4-
+
+
+
10
x
x
x
x
x
x
x<
X
x
x
x
X
x
x
x x
x.
x
x
xIon\Engine(Xe)
X
Arcjet(NH3)
*
FEEP
x
0
X
x
xa
X
X
a
x
Y
0
C
D
0
K~ a
a
a
a
0
01
1
x
+ +
x
x
a1
[
a
+
x
*
*
*
*
*
*
*
*
*
*
-
*
xx
-
VAT (Cr)
xx
x
+
+
-
+ Monopropellant (H4N2)
+ Monopropellant (H202)
Blpropellant (NTO/MMH)
X
x
x
x
x
+
±-[
x
x
+
+
+
+.
+
X
+
++
+ +
+-
+
4-
10
+
X
x
+
+
+
x
X
x
X
+
+
+
+
+
1
+
+
+
+
+
+
+
+
+
+
+
+
X
O
E
a
a
a
C O
*
-
10
[
1
0
aE
0
C
a
a
O
a
O
a
a
-
0
*
-*
*
-
--
103
102
01
(m/s)
Figure 6-2: Using the program to compare propulsion options for a 100 kg spacecraft.
Electrosprays are not compared. Mass fractions over 90% are not shown.
Much more extensive applications are possible with the most recent version of the
program. For demonstration purposes, the program was iterated for a wide range
74
of thrust, mass, and AV combinations. The first example in Figure 6-2 compares
propulsion options on a 100 kg spacecraft.
It should be noted that some propulsion options cannot provide thrust as low as
shown in these plots. If the thrust was lower than the range in the database for that
type of propulsion, the program reset the thrust to the lowest recorded and calculated
mass accordingly. So even though monopropellant thrusters cannot provide thrust
less than 1 mN, at low AV they can still result in a smaller system than those that
do. These plots assume that having more than the required thrust is acceptable, and
contain no information with regard to minimum impulse bit.
Lowest Mass Propulsion Option for 100 kg Satellite
10
x
x
+
+
100 +
+
10
x
x
x
x
x
x
x
x
x
x
x
-
x
X
+ ++ +
+ + +
+
+
+
+
+
+
+
+
+
+
+
+
x
x
x X
X
+
+
+
+
+
+
+-
A-
A-
A-
A-
+
+-
X
x
+
A-
+
+
+
A-
AA-
+
+
+
xX
+
+
+-
x-
+-
+
+
A-
+
+
+A-
+
+
A-
+
+
+
x
X
x X
1
.
.
X
X
x
X
-
x
X
x
X
x x
X
x x
x
X
x
x
x
X
x
x
x
x
x
x-
x
x
x
x
x
x
x
X
X
X
XX
x
x
X
x
x
x
X
X
M
+
A
onopropellant (H4N2)
+ Monopropellant (H2O2)
Bipropellant (NTO/MMH)
x
Arcjet(NH3)
FEEP
Electrospray(EMI-BF4)
x
X
-
10
X
-
-
-
-
.
-
-
-
1
C10
-+4
IT
+
3 AA-
10,
.
+
10-2
A+
+
A +
+
+
A-A+
+- A-+
A- 44.
A
A -A-
A
4
10..
,
10
.
...
.
.
.
.
.
10
OW (m/s)
..-
.
.
..
l01010p
Figure 6-3: Using the program to compare propulsion options for a 100 kg spacecraft,
including electrosprays. Mass fractions over 90% are not shown.
Figure 6-3 includes electrospray thrusters in the comparison. As can be expected,
they dominate a sizable portion of mission scenarios, replacing ion engines and VATs.
In order to better observe the capabilities of other electric propulsion options, elec-
75
trosprays were excluded from the other plots.
Two other interesting cases are the high and low mass ends of the spectrum.
Lowest Mass Propulsion Option for 10 kg Satellite
10
x
+
4-
+
+
+-
+
x xx
+
+
x
x
x
x
+ + +-
+ + + + + + + + + + +
10
+4
+
+
4-
+
+
+
+-
-+
+
+
+
+
+
+
+
++
+
+
+
+
+
+-
+
4-
+
+
+
4-
1
4
4-
7
4-
10
+
+
4-
10
*
+
+
+-
+ + +
+4- 4
+
+ ±
+-
+ 4+ +4+-
+
+
+.
+ +
+
+
+
+
+
+
+
+
+
+
+
+
+
x
x
x
x
4-
4-
+
+-
x
x
x
x
+
x
x
xx
+
+
X
X-
+
+
x
+
+4
+
+
+
+
++
+
+
+
+*******----
+
+
+
+
4+
4-
+
+
*
+
+
+
+
+
+
+
-
-
-
*
4-
t
4-
4-
4-
,+,,t
4-
4
+
4-,
10 1-
4-
+
4-
4-
4-
+
4-
4 -+
44
0 4 4-4+
4-
+
4-
4
"
x
I on Engine (Xe)
VAT(Cr)
FEEP
x
x
4
4
*
*
+4- 4+
x-
-4-7o
C 4
+
+
+ 4-
x
X
+ + ± + + 4
+
+. x
4-
x
x4
x
4x4
+
+
x
x-4-x
4
Bipropellant (NTO/MMH)
-
x
+
+
Monopropellant (H4N2)
+ Monopropellant (H202)
+
x x
+
+
+
+
+ 4-
+
+-
x
x
+ + + +X X
+
+
+
+
+-+
+f
+ +
10~
10
+ +
+
x
x .
x
+
+ +
++-
x
x
x x x X x x x
10P + + + + + + + + ++ +
+ + +- + + +
Cold Gas (N2)
x
x
x
+ + + +-+ + + + + + x
+ ± +- + +- + +
0
r
*
*
*
-
-
-
,
+
*
,
-
-
*
4-*
4-.
103
ice
4-
4-i
4
-
-
0/ (m/s)
Figure 6-4: Using the program to compare propulsion options for a 10 kg spacecraft.
Electrosprays are not compared. Mass fractions over 90% are not shown.
76
Lowest Mass Propulsion Option for 10000 kg Satellite
x
10
x
e
z
1o
UF
N
NFl
N
CO
Dro
e
2102
IN
10~
F
*
o
o
0 o
0 o
o
a
o
o
10
10~
x
o
A
A
~>
$
U
N
N
N
N
N
o
o
e
o
0
C1
NU~
N
0
71
C
.
.
1Bipropellant (NTO/MMH)
ArcJet (NH3)
ArcJet (H4N2)
Ion Engine (Xe)
Hall Thruster (Xe)
-VAT (Cr)
-FEEP
UN.
o
D
.
-
-
+*--
104
10~~
10
10
10
10
10
10
e/ (m/s)
Figure 6-5: Using the program to compare propulsion options for a 10,000 kg spacecraft. Electrosprays are not compared. Mass fractions over 99% are not shown.
77
78
Chapter 7
Conclusion
This research has resulted in a useful tool for spacecraft design and simulation. Although it has limitations that are important to understand, it also carries the unique
capability of providing quick preliminary design estimates that would otherwise be
much more time consuming. In addition, the database itself can be useful if searching
for a particular thruster model. The validity of the program lies in its use of historical data. Assumptions are made, but a lot of these are constants that can easily be
adjusted.
There are several improvements that could be made to the program that are still
unexplored. The first is the calculation of power system mass. Instead of using a
single constant for power density, different constants could be used for different types
of propulsion to represent the diverse power conditioning requirements for each type
of propulsion. Scaling equations instead of constants could be used as well, reflecting
the difficulty of scaling power systems into the low mass regime. If the data on power
processing units and power system design is readily available, this would be very easy
to add into the program.
A second possibility for improvement would be in the propellant storage calculations. Particularly, storage mass for unpressurized propellant is assumed to be a
constant 5% of propellant mass. This may not be accurate across a wide range of
propellant mass. Perhaps some estimation of surface area of the propellant can be
used, but this would have to take into account the highly integrated nature of un79
pressurized propellant storage at low AV in FEEPs, PPTs, VATs, and electrosprays.
In all practicality, some of these systems may never be designed for high AV because
of their low thrust, and their propellant calculations may only be relevant within a
certain regime anyways.
Also, the calculations for pressurized propellants do not account for cryogenic
storage, composite tanks, or fuel pumping. The reason for this was that the program
was focused on systems providing less than 10 N of thrust, and the calculations
require more precision as mass decreases. Systems in this regime are less likely to
be optimized with these features. However, for high AV, these features open up the
possibility to further optimization.
Two specific sets of equations that could use some more attention are those for
ion engines and for PPTs. The ion engine equations estimate I.
from thrust, and
this may be more accurately correlated to AV. The data for PPTs is extensive but
chaotic. Perhaps more complete data can be found, or perhaps older, less optimized
models can be excluded from the database. There are some PPT models that are
competitive with the single VAT model used, but since they are averaged in with
other models the equations portray VATs in a much better light.
Some thrusters, such as PPTs, FEEPs, and VATs, are commonly integrated with
their propellant and power systems. Although the program calculates thrusters, propellant storage, and power separately, exceptions could be made for these thrusters
if the necessary data was gathered.
Most other possibilities would require a significant addition to the database, which
would be very time consuming. For example, propellant storage calculations could
follow the same method as thruster calculations if a database of tank specifications
was created.
Information on the the cost, reliability, or flight status of thrusters
could be added and filters could be implemented to only return data on, for example,
commercially available flight qualified thrusters. Ultimately, this direction leads away
from simulation and prediction and towards choosing an actual thruster model for the
designer. It is worth noting, however, that data on efficiency and minimum impulse
bit have already been gathered for many thrusters in the database and could easily
80
be incorporated into any modifications.
The real strength in the program, as said, lies in how quickly it can compare so
many different options. As a spacecraft moves from its preliminary design phase,
more extensive and specific calculations are required. For assistance in comparing
propulsion options or estimating mass fractions, however, this research fills a much
needed role.
81
82
Appendix A
Thruster Data
Table A.1: Cold gas thrusters
Tmax
Thruster
(N)
m (g)
ISP
Ibt
Pnax
Pin
(s)
(mNs)
(W)
(MPa)
0.75
1.5
0.14
[pog]
3.5
0.25
[8o][63]
0.10
["u7
0.25
[121
0.70
[117][82][110
ALMASat CGMPS
0.00075
AMPAC SV-14
0.04000
75.00
0.00100
300.00
70
0.00200
175.00
60
DASA CGT1
0.02000
120.00
67
EG&G
0.10000
75
5.0
ESMO-CGS
0.20000
140
10.0
Marotta
0.05000
70.00
2.36000
70.00
65
Marotta MCGT
0.44500
50.00
73
Marquardt
4.50000
5.40
AMPAC-ISP
ref.
VP-03-001
Bradford
Engineering
4.5
PMT
Marotta
Gold
Gas
[52]
1.0
0.69
44
1.0
15.44
[71][62]
44
1.0
3.45
[61][115
8.84
[70]
[117]
Micro-Thruster
Continued on next page
83
Table A.1 - continued from previous page
Tmax
Thruster
(N)
m (g)
Isp
Ibit
Pmax
Pin
(s)
(mNs)
(W)
(MPa)
ref.
Moog 58-102
1.11000
15.00
30.0
0.63
117][82]
Moog 58-103
5.55000
15.00
30.0
0.63
[117][382]
Moog 58-112
1.11000
15.00
30.0
0.49
[1171[82]
Moog 58-113
3.33000
15.00
30.0
0.63
[117]
Moog 58-115
2.89000
13.00
30.0
1.46
[in73[o]
Moog 58-118
3.60000
23.00
65
30.0
1.59
[71][281
Moog 58-125
0.00450
7.34
65
2.4
0.03
[117[70][61]
Moog 58E142
0.12000
16.00
57
35.0
2.07
[71[28][29)
Moog 58E143(-6)
0.04000
40.00
60
10.0
0.25
[711][28
Moog 58E151
0.12000
70.00
65
10.5
2.76
[71[(28]129]
Moog 58X125A
0.00440
9.00
65
10.0
0.34
[71][281
Moog B
0.00500
5.50
73
[61]
RDMT-0.8
0.80000
100.00
70
[6]
RDMT-5
5.00000
350.00
70
[6]
Sterer
1.00000
174.00
68
8.89600
376.00
VACCO
Gas Cold Thruster
84
0.1
6.0
0.35
[82)
20.0
1.72
[37]
Table A.2: Cold gas thrusters with liquid propellant (mass
includes integrated tank)
IN t
Pmax
(s)
(mNs)
(W)
35
100
0.25
Tmax
Thruster
Propellant
(mN)
m (g)
I,
liq SF6
5
VACCO MiPS
liq Butane
55
456
65
VACCO Piezo
liq Butane
25
30
70
CNAPS
85
ref.
[22]
[71][38][36]
1
[71]
Table A.3: Hydrazine monopropellant thrusters
Tmax
Thruster
AMPAC
(N)
m (g)
Isp
Ibut
Pmax
Pin
(s)
(mNs)
(W)
(MPa)
ref.
1.00000
330
232
11.00
[45][811
5.00000
370
230
90.00
[41[81]
DASA CHT 0.5
0.50000
195
222
DASA CHT 1
1.10000
290
220
DASA CHT 10
10.00000
240
225
DASA CHT 2.0
2.00000
210
221
200.00
6.00
DASA CHT 5
6.00000
220
222
100.00
8.37
10.00000
600
229
[6]
DOT-5
5.00000
900
230
[6]
HmNT
0.12900
40
150
0.05
1.42000
330
226
10.00
4.45000
380
230
MONARC 1
AMPAC
MONARC 5
DOK-10
Marquardt
KMHS
10.00
7.50
2.20
[116][11o[5[1
15.90
2.20
[11o][5][61]
7.40
2.20
[5][116][110][1]
[5]
2.20
8.25
[s][116)[11o
[71]
[61][70][16)
10
Marquardt
KMHS
2.97
[70][61][16]
17
MEMS /Yuan
162
0.00144
71]
and Li
pAerospace M005
0.00500
10
150
0.17
0.50
0.35
[94)
pAerospace M010
0.05000
12
160
6.00
0.50
0.35
[941
pLAerospace M050
0.50000
15
170
30.00
1.00
0.35
[94]
pAerospace M100
1.00000
30
200
30.00
2.00
0.35
[94]
Primex
0.10000
220
10.00
[52]
Continued on next page
86
Table A.3 - continued from previous page
Isp
Tmax
Thruster
(N)
m (g)
bit
(s)
(mNs)
Pmax
Pin
(W)
(MPa)
ref.
Primex MR-111C
5.30000
330
228
80.00
13.64
2.76
Primex MR-111E
2.20000
330
219
20.00
13.64
2.55
Rafael LT-1N SP
1.10000
250
210
32.00
9.00
2.20
Rafael LT-5N SP
6.00000
230
200
120.00
9.60
2.20
Redmond MR-103
1.13000
330
215
13.30
14.57
2.83
[701[6][61][103][45][36)
2.04000
1270
217
30.00
29.14
2.76
[6]
SEP CHT 2
3.50000
460
216
11.20
2.20
[116]
TRW MRE-0.1
1.00000
500
216
15.00
2.41
[70][17[110]
TRW MRE-1
5.00000
500
218
2.76
[70][61][17]
Redmond
[116][70][61][45)[61
[70][61][6]
[8i]
[u16[1]
MRM-103
87
17.80
Table A.4: Hydrogen peroxide monopropellant thrusters
Tmax (N)
Thruster
m (g)
I., (s)
Ist
Pmax
Pin
(mNs)
(W)
(MPa)
153
ref.
ARC Seibersdorf
0.800
Fotec
0.900
60.0
153
Kuan et al.
0.221
5.8
125
piAerospace M005HP
0.005
10.0
100
0.12
0.5
0.35
[94]
pAerospace M010HP
0.100
11.0
110
0.90
0.5
0.35
[94]
ptAerospace M050HP
0.500
15.0
120
4.00
1.0
0.35
[8][94)
pAerospace M100HP
1.000
20.0
120
8.50
2.0
0.35
[94]
[71]
1.5
0.60
[po6]
[71]
Table A.5: Bipropellant thrusters
Thruster
Propellant
Tmax
m
IV
Pmax
Mixture
Pin
(N)
(g)
(s)
(W)
Ratio
(MPa)
290
ref.
285
[110][5[116][70
275
[81]
EADS S4
MON/MMH
4.00
AMPAC LTT
NTO/MMH
9.25
EADS S10
MON/MMH
12.50
350
287
Fotec
RP-1/H202
3.00
100
320
LEROS 10
N202/MMH
9.50
460
285
Marquardt
NTO/H4N2
4.45
430
280
MMH/NTO
10.00
1.64
3
2.3
[116][70o[110][5
1.2
[16
[45]
1.65
[116[70
R-2/R-2B
Marquardt R-52
Marquardt R-53
Rocketdyne
NTO/MMH
295
[70]
12.90
410
295
1.65
4.50
730
300
1.6
RS-45
88
[16]
[116]
Table A.6: Water resistojets
Thruster
Tmax
m
ISP
Pmax
(mN)
(kg)
(s)
(W)
Mark IV
50.0
Marquardt
44.0
MightySat II
75.0
1.500
1.240
Pin
7)
(%)
ref.
(MPa)
180
100
0.50
[18
220
200
0.25
1521
152
100
1.00
[53][52][59]
180
904
2.20
[521
140
700
Morren @ NASA Lewis
360.0
Othman, Makled
118.0
Rocketdyne Technion coupled
122.0
138
505
0.20
[52
Rocketdyne Technion decoupled
240.0
144
708
0.20
[52]
SSTL Micro Resistojet
SSTL UoSAT-12
0.400
3.3
0.013
45.0
1.240
89
75
3
152
100
[77
[71][8]
34
[1181
Table A.7: Hydrazine resistojets
Tmax
Iep
Ibit
Pmax
Pin
(s)
(mNs)
(W)
(MPa)
ref.
(mN)
m (g)
DCA PACT
500
360.0
Redmond
177
Redmond MR-501B
369
889.0
300
2.20
487
2.41
[52][82][119][118][45[6][36]
Redmond MR-502A
800
870.0
299
88.96
899
2.65
(52][82][118][45[6[36]
Thruster
515
20.00
306
2.20
350
300
[52][31)
10
Sol Rad-10
TRW HiPEHT
490
TRW
245
226.8
[85][65]
550
295
12
4.45
223
[118][119]
[82][118]
1.55
[72]
Table A.8: Ammonia arcjets
Pmax
Tmax
Thruster
ESEX Arcjet
I,
m (g)
(N)
2.000
IRS
(s)
(kW)
800
26.000
600
2.000
rM
(%)
ref.
11o][8216][36]
[36]
IRS VELARC
0.028
520
0.375
18.5
[1]
IRS VELARC
0.035
350
0.240
25.0
p1
LeRC Mini LPATS
0.051
360
0.250
36.0
[1]
PSU Microwave
0.300
550
1.100
[52)
U of Stuttgart
0.115
480
0.750
[52][120][47]
480
90
Table A.9: Hydrogen arcjets
Thruster
Pmax (kW)
I,, (s)
Tmax (N)
77(%)
ref.
HIPARC
3.000
1500
100.000
22
[r
HIPARC-R
4.000
1400
100.000
28
[7][87]
IRS ESA BMDO
1.364
1300
10.000
40
[36]
IRS VELARC
0.023
865
0.365
26
[
IRS VELARC
0.014
820
0.310
18
1
Table A.10: Hydrazine arcjets
Tmax
Thruster
(mN)
m (kg)
I,
(s)
Pmax
Pin
(kW)
(MPa)
ref.
BPD/Centrospazio
130
440
1.000
DCA HAJ1
100
530
0.250
GE Arcjet
210
502
1.600
po]
ISAS SEGAMI-I
176
600
1.500
p96
OSAKA-IHHI
150
475
1.200
IN]
Redmond MR-507
220
1.50
465
1.400
[118]
Redmond MR-508
230
1.34
502
1.808
1.51
[118][6][?][45]
Redmond MR-509
254
1.48
502
1.808
1.76
[82][6][45]
Redmond MR-510
258
1.58
600
2.000
1.79
[45][6][82][36][1o2]
Redmond MR-512
254
1.03
502
1.788
1.76
[6]
U of Stuttgart
230
1.30
520
1.200
1.00
91
[96]
1.80
[a8][1uo]
p96
Table A.11: Ion engines
Thruster
13cm XIPS
25cm XIPS Hughes
Tmax
m
Isp
Pmax
(mN)
(kg)
(s)
(kW)
17.80
6.50
165.00
r
(%)
ref.
2507
0.500
68.0
[70][21][96][36[45][108][6][611
3500
4.500
65.0
[97[96] [45
25cm XIPS INTELSAT
63.00
9.70
2800
1.300
[21][97][82][36][75][108][511
ARTEMIS
15.00
1.60
3000
0.600
[96]
CIT 3 cm
0.50
3703
0.024
37.7
[18]
DERA T5
25.00
1.70
3110
0.644
60.0
[82][36][118][108][61]
DERA T6
150.00
6.20
3470
3.900
65.0
[36][118][61]
ETS-6 Kaufman
23.30
3.70
2910
0.600
ETS-8 Kaufman
23.20
2665
0.611
EURECA
10.00
3300
0.440
GRC 8 cm
3.60
1760
0.100
31.0
[18][36]
9620
39.300
80.0
[24][25][36)
10000
25.000
3200
0.900
66.0
[61][70)
HiPEP
670.00
HiPER DS3G
450.00
JPL
31.00
1.50
46.50
2.50
[20][96]
49.7
[2o][1m8]
[96]
[39]
Keldysh Res Center
6.00
3650
0.162
66.0
[61]
Keldysh Res Center
19.00
3500
0.494
65.0
[61]
p 10 ECR
8.10
2910
0.340
36.0
[76][73][108]
pNRIT-2.5
0.60
0.21
2861
0.034
25.5
[1][71]
MiXI
1.50
0.20
2850
0.050
40.0
[71]
3500
0.600
NAL 14 cm
25.00
NASA 30cm
178.00
7.00
3610
4.880
NASA LeRC
72.00
7.00
2130
1.500
[96,
67.0
[79]
[96]
Continued on next page
92
Table A.11 - continued from previous page
Thruster
NASA Lewis
Tmax
m
I.,
Pmax
(mN)
(kg)
(s)
(kW)
11.00
28.70
7 (%)
ref.
2650
0.308
47.0
[61]
8500
25.000
78.0
[108][26][24][67][36]
4070
6.860
59.0
[95][36]
NEXIS
470.00
NEXT
238.00
NSTAR
92.70
8.20
2500
2.325
61.8
[6][98][61][51][821[361[45][1081
RIT-10
15.00
1.00
3058
0.459
36.0
[5][61][36111][108][82][118]
RIT-10 EVO
41.00
1.80
3300
1.050
67.0
[5][61][36][70]
RIT-15PL
50.00
3600
1.350
67.0
(611
RIT-22
250.00
7.00
6400
5.000
RIT-25
25.00
1.75
3060
0.800
67.0
161][a8]
RJT-XT
210.00
4448
6.850
75.5
[5][61]
3600
0.480
55.0
[13][101][61][36]
31.6
[18]
RMT
12.00
2.00
RUS 5 cm
1.60
2900
0.072
SCATHA P78-2
0.14
350
0.045
93
[5][36]
[6][97][99]
Table A. 12: Hall Thrusters
Thruster
Tma
m
(mN)
(kg)
BHT-1500
102.0
BHT-200
17.0
BHT-20k
1080.0
I., (s)
'bit
Pmax
(mNs)
(kW)
ref.
1.700
54.6
[10]
0.300
32.5
[112][36][71][18]
2750
20.000
70.0
[10][67]
1820
1.1
77(%)
1390
2
17.0
0.9
1600
0.600
45.0
[61]
512.0
20.0
1900
8.000
60.0
[10[15][36]
BHT-HD-1000
55.5
3.5
1700
1.000
56.0
[61]
BHT-HD-600
36.0
2.2
1700
0.600
50.0
[61][36)
2500
4.500
60.0
[19]
BHT-600-X2B
BHT-8000
D-100
D-100-II
65.0
4250
15.000
65.0
[6][461
D-20
17.0
2000
0.300
40.0
[19]
D-35
82.0
1263
1.230
40.0
[61]
D-38
35.0
2000
0.750
45.0
[19][104][61]
D-50
48.0
1700
4.4
D-55
DS-HET
300.0
12.0
[6]
1600
1.400
50.0
[191[82)
3000
5.000
50.0
[poi]
Fakel SPT-25
6.4
948
0.154
19.0
[18[61][112]
Fakel SPT-35
10.0
1200
0.196
30.0
p
HIVHAC
0.4
2500
8.000
62.0
[6][441
HTX
9.3
1350
0.200
31.0
[112
IT-100
18.0
3100
0.500
50.0
[46]
IT-50
5.0
2900
0.140
55.0
[46]
Jacobson, Jankovsky
1.0
3250
25.000
60.0
[87]
Continued on next page
94
Table A.12 - continued from previous page
Thruster
Tmax
m
(mN)
(kg)
I., (s)
bit
Pmax
(mNs)
(kW)
t;
(%)
ref.
Keldysh K-15
16.0
1718
0.400
36.0
[61]
Keldysh KM-32
10.4
1410
0.156
45.0
[61]
Keldysh KM-37
18.4
1635
0.294
49.0
[61]
Keldysh X-40
35.0
1750
0.525
56.0
[61]
KeRC
5.7
895
0.109
22.9
[18]
KM-20
8.8
1850
0.210
39.0
[1121
MHT-9
16.6
1676
0.481
29.0
[112]
1.8
865
0.126
6.0
1170
0.260
30.0
MIT 50W
Moskow SPT-30
13.0
0.4
[71](112
[61][112]
NASA Glenn
1.0
0.500
[40]
NASA Glenn
2.5
1.000
[40]
NASA Glenn
2.5
1.500
[40
NASA Glenn
3.0
2.000
[40]
NASA Glenn
4.3
2.000
[40]
NASA Glenn
5.5
4.500
40]
NASA Glenn
6.5
5.000
40]
NASA Glenn
7.0
5.000
[40]
NASA Glenn
8.0
5.500
[40]
NASA Glenn
12.0
8.000
[40]
NASA Glenn
12.0
10.000
[40]
NASA Glenn
12.0
11.000
[40]
NASA Glenn
20.0
15.500
[40]
Continued on next page
95
Table A.12 - continued from previous page
Thruster
Tm.
m
(mN)
(kg)
I,
(s)
Pmax
(mNs)
(kW)
r
(%)
16.500
18.0
NASA Glenn
Iuit
ref.
[40]
342.0
3000
7.872
59.6
[33]
NASA-300M
1130.0
2640
20.000
67.0
[67][41]
NASA-457M
2900.0
2900
75.000
56.0
[67][6[87][41]
NASA-457Mv2
1280.0
2520
25.200
63.0
[41]
246.0
2326
5.000
PP Annul
3.5
1086
0.098
19.0
[18]
PP Cylind
3.7
1136
0.103
20.0
[18]
0.300
23.5
[71]
1325
0.185
23.5
[71]
NASA-137Mv2
P5
PPPL CHT 2.6
12.0
PPPL CHT 3.0
6.0
PPS 1350-G
PPS-20k ML
80.0
[6]
90.0
4.5
1650
1.500
45.0
[19][61][o][93]
1050.0
25.0
2500
22.400
60.0
[121][46][67]
PPSX000
340.0
2480
6.000
Princeton
54.4
1550
0.870
46.0
[61]
[82][36]
Redmond BPT-2000
0.1
5.2
1765
2.200
48.0
[6]
Redmond BPT-4000
0.3
7.5
1950
4.500
58.0
[6]
100.0
3.5
1600
1.400
50.0
SPT-100
[36][45][108
SPT-140
300.0
1750
5.000
55.0
[108]
SPT-160
400.0
2500
4.500
60.0
[19]
SPT-200
498.0
2250
11.000
63.0
[40
SPT-290
1500.0
23.0
3300
30.000
70.0
[40][67][15]
13.0
1.0
973
0.258
25.0
[71]
SPT-30
Continued on next page
96
Table A.12 - continued from previous page
Thruster
Tmax
m
(mN)
(kg)
SPT-50
20.0
SPT-60
30.0
SPT-70
40.0
I,
0.9
2.0
(s)
INt
Pmax
(mNs)
(kW)
7 (%)
ref.
1100
0.350
35.0
[19][61][108][4]
1300
0.510
38.0
[61]
1500
0.700
45.0
[108][82](45]
Stanford
1.6
575
0.040
12.5
1112]
Stanford
11.0
544
0.277
10.6
[18]
T-100
82.4
1573
1.400
47.0
[61][821
T-140
300.0
2000
4.500
T-220
1000.0
1950
20.000
T-220HT
1180.0
[g9][74][45]
62.0
22.000
[40][19]36
[74s36]
T-27
9.5
1430
0.201
33.0
[112][61]
T-40
20.0
1300
0.400
60.0
[19][74]
1114.0
2400
25.000
66.0
[40]
3500
3.000
58.0
[50[49][36]
TAL TM-50
6.0
Thales HEMP
152.0
Thales HT-100
12.5
1650
0.400
29.0
[112]100]
U of Hifa
39.0
1656
0.663
49.0
[61]
97
Table A.13: Pulsed plasma thrusters
Thruster
Tmax
m
Isp
Ibit
Pmax
(mN)
(kg)
(s)
(pNs)
(W)
2.800
37 min parabolic
Sinica
Academy
280
445
133.00
'q (%)
ref.
2.30
[68]
4.00
[68]
8.00
[61]
Geom-Varying
advPPT
4.200
AFRL
0.030
APPT
5.200
1700
0.025
ARCS
133.0
515
2.00
20.0
650.00
250.0
2.00
4.0
1616
490.00
Busek jPPT
827
80.00
China Lab
990
448.00
CSSAR 40J
1188
605.00
CU UI PPT-11
1400
CU UI PPT-8
360
533.00
56.10
ASU
Dawgstar
0.112
1.720
483
ETS-IV (Japanese)
0.029
6.600
300
808
Fairchild, NASA
1.700
FalconSat-3
FOTEC pLPPT
lIT7
U
0.010
Singapore
[71]
17.00
[71]
10.50
10.0
[11]
9.00
[68]
15.00
[36]
4.00
[68]
10.2
1.80
[114][91]ts[71][113]
2.2
1.92
[9][68]
6640.00
10.00
8.40
827
10.00
1600
30.00
600
450.00
2620
1200.00
[68]
[83]
100.0
500
[68][61
2.0
[68]
[113][11][68
[106]
6.00
[68]
self-ignited
IL PPT-3 Lab
IRS ADD-Simplex
1.370
[83]
68.0
14.00
[1][68]
Continued on next page
98
Table A.13 - continued from previous page
Thruster
Tmax
m
1
(mN)
(kg)
(s)
Jiaotong nozzled
KIT,
U
Tokyo
,p
INt
Pmax
(pNs)
(W)
rq
(%)
ref.
765
82.00
8.00
[68]
1675
25.00
8.00
[68]
68
1.12
38.00
[68]
liq prop
Lab
Combustao
Propulsao
LES 6
0.026
1.400
300
26.00
2.5
2.90
LES 8/9 (MIT)
0.600
6.700
1000
297.00
25.5
7.00
[9][83][68][61][102]
[14] [9][83][68]
[61][82][102][45]
585
Mars Space pPPT
MDT-2A (Chinese)
0.064
280
[68]
[9]
0.40
0.014
MILIPULT
4.48
29.00
[71]
150.0
18.00
[68][61[
13.00
[83][68]
1210
2230.00
1130
2250.00
600
454.00
NASA 450J
3000
8259.00
27.00
[68]
NASA HEPPT
2770
1300.00
23.00
[68]
5.31
[14]s"][102][45]
30.0
7.80
[9][68][61]
70.0
14.00
[61]
13.50
[68]
7.40
[18]
Millipound
4.450
1.130
MIPD-3
MIT Lab
NOVA
0.378
6.350
543
378.00
NOVA (TIP-Il)
0.375
7.100
850
378.00
OS-1
1.400
4.950
1400
445
Osaka IT
PPT
1.000
PPT
0.090
PPT
2.000
1500
183]
200.00
100.0
20.0
0.500
800
100.0
[61]
8.00
[61]
Continued on next page
99
Table A.13 - continued from previous page
Thruster
Tm.
M
I,,p
it
Pma
(mN)
(kg)
(s)
(iNs)
(W)
77(%)
ref.
PPT
0.140
500
12.5
3.00
[61]
PPT-4
0.450
1250
15.0
18.00
[61]
PPT-5
0.750
1750
50.0
13.00
[61]
Primex
4.500
1500
120.0
PRIMEX EO-1
1.400
4.950
Princeton
1136
100.00
400
650.00
130.8
[5211821
5.50
[14u[113][114][83][68]
6.00
[r,8
UM AZPPT
1.420
PRS-101
RRC
4.740
Kurchatov
100.0
1350
[6[45]
3000
41.21
13.00
[68]
3000
1600.00
22.00
[68]
4200
3883.00
37.00
[68]
720
660.00
10.00
[68]
9-APPT array
RRC
Kurchatov
High Energy
RLRC
Kurchatov
Very High Energy
Shairf U of T
0.133
SMS
4.100
22.0
400
800
STSAT-2
2.900
Teflon PPT
100.0
745
3.00
[9][68][61][1021
1.34
[68]
10.40
[18]
945
17.00
3.30
[68]
1057
35.00
5.00
[68]
TMU Co-III 3-12mm
400
49.00
10.20
[68]
TMU PPT-B20
960
22.00
3.10
[68]
2025
4.00
5.50
[68]
TMIT-5
TMIT-6A
Tokai
U
Laser
Assisted
Continued on next page
100
Table A.13 - continued from previous page
Thruster
Tmax
m
Isp
Ibit
Pmax
(mN)
(kg)
(s)
(pNs)
(W)
7 (%)
ref.
423
469.00
3.00
[8s][68]
UCV XPPT-8
1210
320.00
9.00
[68]
UI PPT 3,4,5
2200
150.00
19.50
[68]
U Tokyo External B
Zond 2
2.000
5.000
8.04
50.0
410
[9][68][61]
Table A.14: Vacuum arc thrusters
Tmax
Thruster
(mN)
m (kg)
I-, (s)
Ibit
Pmax
(pNs)
(W)
0.25
(%)
100
AASC VAT (ION)
1.000
100
2000
AASC MVAT
0.125
150
2000
100
10.00
Lun
56.000
500
160
10
0.05
MPNL CT
10.000
140
3500
2
14.00
101
0.10
ref.
[36][32][71][89]
[36][8][84][107]
[w]o
[43][42]
Table A.15: Field emission electric propulsion
Tmax
m
I.,p
Ibi
Pmax
'r
Propellant
(mN)
(kg)
(s)
(nNs)
(W)
(%)
Alta FEEP-100
Cs
0.120
Alta FEEP-150
Cs
0.300
Alta FEEP-1000
Cs
1.200
ARCS GOCE MTA
In
ARCS InFEEP-100
Thruster
ref.
10000
7.2
82
[61]
1.41
9000
20.0
90
[3[781
5.00
10000
72.0
82
[61]
0.650
12000
80.0
In
0.100
12000
10.0
61][711
In
1.000
12000
80.0
[36][71]
ARCS inFEEP-25
In
0.025
7.10
10000
Centrospazio
Cs
0.800
3.50
8000
60.0
[70
Centrospazio
Cs
0.100
0.45
8000
9.0
[70]
Centrospazio FEEP
Cs
1.000
6000
60.0
Centrospazio
Cs
0.040
0.60
9000
2.7
Cs
1.400
1.20
9000
93.0
ARCS
[71]
InFEEP-1000
161136]
(521
65
[61][71]
FEEP-5
Centrospazio
[61][71]
FEEP-50
[86]
34
In
0.100
2.50
10000
10
6.7
In or Ga
0.300
0.30
4000
5
7.0
50
[106]
In
0.015
0.08
5000
5
0.5
50
[106]
InFEEP
In
0.120
95
[4]
In-FEEP
In
0.100
2.50
8000
5
95
[105]
Slit Emitter FEEP
Cs
1.200
2.20
9000
1
FEEP
FOTEC mN-FEEP
FOTEC
Single FEEP
12000
102
13.0
[4]
Table A.16: Colloid thrusters
Thruster
Busek
Busek ST-7
Tma
m
(pN)
(kg)
189.0
Pin
Pmax
Ip (s)
390
2.50
(W)
r
(%)
6.00
[61][34][451
103.4
240
35.8
ref.
(kPa)
[71][36][45
8.0
7.33
700
0.03
78
[61]
GRC
200.0
2.50
390
0.50
70
[18]
Hruby et al.
190.0
400
10.00
[2]
30.00
[2]
Electro Optical System
MAI
1000.0
Stanford
500.0
0.50
200
6.00
[45][48]
Stanford EMERALD
315.0
0.25
600
6.00
[48]
1.0
1450
0.01
69
[61]
TRW/ Edwards AFB
159.0
1382
1.40
77
[61]
TRW/ Edwards AFB
129.0
1405
1.15
77
[61]
TRW/ Edwards AFB
335.0
1029
2.41
70
[61]
Velasquez
450.0
TRW
4.95
300
103
[2
104
Appendix B
Source Code
i
function
propulsion-mass-f raction,
[propellant-mass-fraction,
inert-mass,
propellant.mass,
...
. ..
max-power ]=thruster-mas s (deltaV, total-init
ial-mas s, max-accelerat ion,
thruster-type)
2
3
%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%
4
%
s
% Filename:
thruster-mass.m
6
% Filetype:
Function
7
% Project:
F6
8
% Author:
Thomas Chiasson
9 % Mod Date:
19/8/2011
10
%
ii
% PURPOSE: Estimates the mass of a propulsion system.
12
%
13
% USAGE:
This is the main program called for propulsion system
14
%
estimation.
is
%
another program. The equations formed from the database are
16
%
hardcoded and propellant and tank constants are read from
17
%
the database,
18
%
and equations are in this program, with programs for tank
19
%
optimization called from within this code.
It can be used stand-alone
or called within
'thrustertradespace.xlx'. All other constants
105
20
21
% INPUT:
22
% 'deltaV'
23
%-
(m/s)
(kg)
(10-10,0 00)
25
(10^-6-1 0^-3)
27
including propulsion and propellant.
acceleration expected
from the
thruster.
Row number of desired thruster
'thruster-type'
28
The total initial mass of the satellite,
(m/s^2) Maximum
'max-acceleration'
26
in velocity that the thruster
should provide over its entire lifetime.
'total-initialamass'
24
Change
in
'thrustertradespace.xlx.'
29
30
31
% OUTPUT:
32
% 'propellantamass-fraction'
fraction of total initial mass that is
33
%
propellant
34
%
35
%
36
%
37
%
38
% 'propellant-mass'
(kg)
Mass of propellant from rocket equation.
39
% 'max-power'
(W)
Estimated maximum required power.
'propulsionmass-fraction'
fraction of total initial mass that is
propellant or inert propulsion mass
'inertamass'
(kg)
Mass of thruster, tank(s),
power system,
lines, supports
40
41
%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%
42
43
%clc
44
global g R FS;
45
46
g = 9.81;
47
R = 8.3145;
48
FS = 2;
49
specificpower = .015;
5o
tank-factor
si
ullage-volume = 5;
%m/s2
%J/(mol K)
%kg/W
= 2.5;
%%
52
53
54
55
% Import the file
[numbers, strings]
if
= xlsread('thrustertradespace.xlsx',
~isempty(numbers)
106
'Input Calcs');
thrusterData.data
56
57
end
58
if
numbers;
~isempty(strings)
thrusterData.textdata =
59
6o
=
strings;
end
61
62
% Create new variables in the base workspace from those fields.
63
vars = fieldnames(thrusterData);
6
for i
assignin('base', vars{i}, thrusterData.(vars{i}));
65
66
1:length(vars)
end
67
68
% extract values from file
69
tank-density=thrusterData.data(thruster-type,
7o
tank-yield-strength=thrusterData.data(thruster-type,
14);
13);
71
72
%Calculate thrust
73
max-thrust
74
thrusters=1;
*max-acceleration;
= total-initial-mass
75
76
if
max-thrust>10
|| max-thrust<.000009999999999
fprintf('\nERROR:
77
input
10 <
Thrust out of range
78
< 10,0000
total-initialamass
max-acceleration
for this model. Please
and 10e^-6 <
.
...
< 10e^-3\n')
end
79
8o
if
max-thrust>thrusterData.data (thruster-type,
fprintf('\nWARNING:
81
6)
The required thrust is higher than that of
any thruster in the database. Multiple thrusters will be ...
simulated.\n')
while max-thrust/thrusters>thrusterData.data(thruster-type,
82
thrusters=thrusters+1;
83
84
end
85
max-thrust=max-thrust/thrusters;
86
end
87
107
6);
s
if
max-thrust<thrusterData. data (thruster-type,
fprintf('\nWARNING:
89
4)
The required thrust is lower than that of
...
any thruster in the database. Thrust is being raised to that
%i N\n',
of the lowest thrust model in the database:
4) )
thrusterData. data (thruster.type,
max-thrust = thrusterData.data(thruster-type,4);
90
91
...
end
92
93
9
%Calculate specs and masses
95
switch thruster..type
(see thrustertradespace.xls
for equations)
96
97
case 1 %Nitrogen Cold gas thruster
98
exhaust-velocity = thrusterData. data (thruster-type,
99
max-power = 11.607*max-thrust^0.2294;
100
7) *g;
thruster.mass = thrusterData. data (thruster-type,
10);
101
102
%Calculate propellant mass
103
propellant-mass=total.initial.mass.*(1-exp(-deltaV./exhaust-velocity));
104
105
%Calculate tank mass
106
if
max-thrust<=.12
propellant._pres sure=88030*log (max..thrust) +741342;
107
ios
else
propellant.pres sure=3165825;
1o9
110
mll
end
112
propellant-molar.mass=thrusterData.data (thruster-type,15);
113
initial.propellant-mass
=
...
...
fminsearch(@ (initial.propellant..mass)
cold-tank.nass (initial-propellant-mass,
propellant..temperature,
propellant.pres sure,
propellant..mass,
propellant-molar..mas s,
tank..yield.strength,
...
tank-density) ,propellant..mass);
114
16);
propellant-temperature=thrusterDat a. data (thruster.type,
tank..mass=tank..factor*cold.tank..mass
propellant-mass,
(initial.propellant.mass,
propellant.temperature,
108
...
...
propellant-pressure,
propellant.molar-mass,
tank-yield-strength,
tank-density);
115
116
if
initial-propellant..mass
117
fprintf('\nERROR:
118
thruster-type
119
deltaV
120
total.initial.mass
121
max-acceleration
122
< propellant-mass
tank optimization incorrect\n')
end
123
124
%syms mO real;
125
%s impIify(estimate-tank-mass (mO,
propellant _temperature,
propellant-pressure,
126
propellant-mass,
propellant-molar-mass,
propellant.temperature,
propellant-pressure,
tank-density),
tank-density))
tank-yield-strength,
%ezplot(cold.tank.mass (mO, propellant.mass,
...
.. .
propellant..molar-mass,
tank-yield-strength,
...
[propellant-mass, propellant-mass*1.3] )
127
128
case 2 %Helium Cold gas thruster
129
exhaust.velocity = thrusterData. data (thruster-type,
130
max-power = 11.607*max-thrust^0.2294;
131
thrustermass
= thrusterData.data (thruster.type,
7) *g;
10);
132
133
%Calculate propellant mass
134
propellant-mass=total.initial-mass
.*(1-exp(-deltaV./exhaust-velocity));
135
136
%Calculate tank mass
137
if
max-thrust<=.12
prope llantpressure=88030*log (max-thrust) +741342;
138
139
else
propellant.pre s sure=3165825;
140
141
end
142
propellant-temperature=thrusterData. data (thruster-type,
143
propellant-molar-mas s=thrusterData.data (thruster.type,15)
109
16);
;
144
=
initial..propellant-mass
...
...
fminsearch (@ (initial-propellant-mass)
propellant.mass,
cold..tank-mass (initial.propellant-mass,
propellant.temperature,
propellant-pressure,
propellant-molar..mass,
tank.yield-strength,
...
...
..-
tank.density) ,propellant-mass) ;
145
tank-mass=tank.factor*cold-tank.mass (initial-propellant.mass,
propellant-mass,
propellant-temperature,
propellant-pressure,
propellant.molar-mass,
tank-yield-strength,
...
...
tank-density);
146
147
if
< propellant..mass
initial.propellant..mass
148
fprintf('\nERROR:
149
thruster...type
iso
deltaV
151
total-initial..mass
152
max.acceleration
153
tank optimization incorrect\n')
end
154
155
%syms mO real;
156
%simplify(estimate-tank-mass (mO, propellant-mass,
propellant-temperature,
propellant.pres sure,
157
propellant-temperature,
tank-density),
propellant-mass,
...
tank-density))
tank-yield-strength,
%ezpiot(cold-tank_mass(mO,
propellant-pressure,
propellant-molar..mass,
...
propellantmolar-mas s,
...
...
tank-yield-strength,
[propellant-mass*.7,propellant-mass*1.3])
158
159
case 3 %Butane Cold gas thruster
160
exhaust.velocity = thrusterData. data (thruster-type,
161
max-power = thrusterData.data(thruster-type,
162
thruster-mass = thrusterData.data(thruster-type,
7) *g;
9);
10);
163
164
%Calculate propellant mass
165
propellant..mass=total-initial-mass.*(1-exp(-deltaV./exhaust-velocity));
166
if
propellant.mass/thrusters
>
110
.053
fprintf('\nWARNING:
167
This thruster is integrated with its
propellant tank and the required deltaV is higher ...
than what can be provided. Multiple thrusters will
...
be simulated.')
168
while propellant-mass/thrusters
169
thrusters=thrusters+1;
.053
end
170
171
>
end
172
173
%tank mass included in thruster mass
174
tank-mass=O;
175
176
case 4 %Hydrazine Monopropellant thruster
en
exhaust-velocity
178
thrustermass
179
max-power
= 206.21*max-thrust^0.048*g;
= 0.161*max-thrust^0.5778;
= 28.684*thrusternass^0.7943;
180
181
%Calculate propellant mass
182
propellant.mass=total-initial-mass.*(1-exp(-deltaV./exhaust-velocity));
183
184
%Calculate propellant tank mass
185
propellant-pressure=2558.7*exp(0.0311*exhaust-velocity/g);
186
propellant-tank-volume=(1+ullagevolume/100)
187
*propellant-mass/...
thrusterData.data(thruster-type,12);
(4/3*pi))
i8
tank-radius=(propellant-tank-volume./
189
tank-thickness=propellant-pressure.
190
propellant-tankamass=tank-factor*4/3*pi* ( (tank-radius+...
191
.(1/3);
*tank-radius*FS/tank-yield-strength/2;
tank-thickness) .^3-tank-radius.^3) .*tankdensity;
192
193
%Calculate pressurant tank mass
194
pressurant-temperature=thrusterData.data(1,
195
pressurant-molar-mass=thrusterData.data(1,15);
16);
% use N2
gas
as pressurant
i9
pressurant-tank-density=thrusterData.data(1,14);
197
pressurant-tank-yield-strength=thrusterData.data(1,13);
use Ti alloy as tank material
111
%
...
...
198
199
%syms VO real
200
%pretty
(simplify (mono.tank-mass (mO,
pressurant..molar.mass,
pres surant..temperature,
propellant-pres sure,
propellant-tank-volume,
...
pressurant-tank-yield-strength,
...
pressurant-tank-density)))
201
pre s surant _temperature,
%ezplot(mono.tank..mass (VO,
pressairant_molar-mass,
propellant-pressure,
propellant-tank-volume,
...
pressurant-tank-yield.strength,
pressurant.tank-density),
...
[0, propellant..tank.volume*.2])
202
203
pressurant.tank.volume =
fminsearch (@ (pressurant_tank-volume)
...
mono..tank.mass (pressurant-tank-volume,
.. .
pressurant.temperature, pressurant-molar-mass,
propellant.tank.volume,
propellant-pressure,
...
...
pre s surant-t ank-yield-s trength, pressurant.tank-density) , O);
204
pressurant-tank-mass = ...
t ank-factor*mono-t ankimass (pressurant-tankvolume,
pressurant-temperature, pressurantmolar-mass,
propellant.pressure,
propellant-tank..volume,
pressurant..tank-yield-strength,
...
...
pressurant-tank-density);
205
206
%initial-pressurant.volume =
fminsearch(@ (initialpressurant-volume)
...
mono.tank-mass2 (initial-pressurant..volume,
pressurant-temperature,
pressurant-molar..mass,
propellan t.pres sure,
propellant-tank-volume,
tank-yield-strength,
...
...
tank-density) ,propellant-tank-volume*.1)
207
%tank-mass = mono.tank-mass2(initialpressurantvolume,
pressurant-temperature,
pressurant-molar-mass,
propellant_pres sure,
propellant...tank-volume,
tank-yieldstrength,
tank-density)
208
209
if pressurant.tank-volume <
0
112
...
...
...
210
fprintf('\nERROR:
211
thruster-type
212
deltaV
213
total-initial-mass
214
max-acceleration
215
tank optimization
incorrect\n')
end
216
217
%propellant-mass
218
%propeIlant-tank-volume
219
%propellant-tankanass
220
%pressuranttank-volume
221
%pressurant-tank-mass
222
223
tank-mass=propellant-tank-ass+pressuranttbank-mass;
224
225
case 5 %Hydrogen peroxide Monopropellant thruster
= 136.76*max-thrust^0.0649*g;
226
exhaust-velocity
22T
thruster-nass
228
max-power = 0.4715*exp(1.3858*max-thrust);
= 0.0077*exp(1.4637*max-thrust);
229
230
%Calculate propellant mass
231
propellant-mass=total-initialamass.*(1-exp(-deltaV./exhaust-velocity));
232
233
%Calculate propellant tank mass
234
propellant-pressure=103421*exp(0.011*exhaust-velocity/g);
235
propeilant-tank-volume= (1+ullage-volume/100) *propellant-mass/...
236
thrusterData.data(thruster-type,12);
(4/3*pi))
237
tank-radius=(propellant-tank-volume./
238
tank-thickness=propellant-pressure.
239
propellant-tank-mass=tank-factor*4/3*pi*(
240
(1/3);
*tank-radius*FS/tank-yieldstrength/2;
(tank-radius+...
tankathickness) .^3-tank-radius.^3) .*tank-density;
241
242
%Calculate pressurant tank mass
243
pressurant-temperature=thrusterData.data(1,
244
pressurant-molar.aass=thrusterData.data(1,15);
as pressurant
113
16);
% use N2 gas
...
245
pressurant-tank-density=thrusterData. data (1, 14) ;
246
3
pressurant-tank-yield-strength=thrusterData.data(1,1 );
% ...
use Ti alloy as tank material
247
248
%syms VO real
249
%pretty
(simplify (mono-tank-mass (mO,
pressurant-temperature,
propellant.pres sure,
pressurant-molar-mass,
ank-yield-strength,
pressurantt
propellant-tank-volume,
...
...
pressurant.tank-density)))
250
pres suranttemperature,
%ezplot(mono-tank-mass (VO,
propellant-pressure,.
pressurant-molar-mass,
propellant-tank-volume,
.
..
pressurant-tank-yield-strength,
...
[0, propellant-tank-volume*.2])
pressurant-tank-density),
251
252
pressurant-tank-volume
fminsearch(@(pressurant-tank-volume)
mono-tank-mass (pressurant..tank-volume,
...
pressurant-temperature, pressurant-nolar-mas s,
propellant.tank-volume,
propellant-pressure,
pressurant-tank.yield-strength,
253
pressurant.tank-mass =
...
...
pressurant..tank-density) , O);
...
tank-factor*mono-tank..mass (pressurant-tank-volume,
pressurant-temperature, pressurant-molar-mass,
propellant-pressure,
propellant-tank-volume,
pressurant-tank-yield.strength,
...
.
pressurant-tank..density);
254
255
%initial-pressurant-volume =
fmlnsearch(@ (initial-pressurant-volume)
...
mono..tank-mass2 (initial-pressurant.volume,
pressurant-temperature,
pres surant.molar.mas s,
propellant.pressure,
propellant.tank-volume,
tank-yield-strength,
...
...
tank-density) , propellant..tank-volume*. l)
256
%tank-mass = mono.tank-mass2 (initialpressurant-volume,
pressurant-temperature,
propellant-pressure,
pressurant.molar-mass,
propellant.tank-volume,
114
...
...
.
tank-yield-strength,
tank-density)
257
258
if
pressurant-tank-volume
< 0
259
fprintf('\nERROR: tank optimization incorrect\n')
260
thruster-type
261
deltaV
262
total..initial-mass
263
max-acceleration
264
end
265
266
%propellant-mass
267
%propellant-tank-volume
268
%propellant..tank-mass
269
%pressurant-tank-volume
270
%pressurant-tank-mass
271
272
tank-mass=propellant.tank.mass+pressurant..tank-mass;
273
274
case 6 %Nitrogen tetroxide monomethylhydrazine bipropellant thruster
= thrusterData.data(thruster-type,7)
275
exhaust.velocity
276
max...power = thrusterData.data (thruster-type,
277
thruster-mass
*g;
9) ;
= thrusterData.data (thruster..type, 10);
278
279
%Calculate propellant masses
280
oxidizer-fuel-ratio
281
propellant-mass=total-initial-mass.*(1-exp(-deltaV./exhaust.velocity));
282
fuel-mass=propellant.mass/
283
oxidizer-mass=propellant-mass-fuel.mass;
284
propellantpres
= 1.65;
(1+oxidizer..fuel..ratio);
sure=thrusterData. data (thruster..type, 11);
285
286
%Calculate fuel tank mass
287
fuel.density = 878; %kg/m3
288
fuel.tank-volume=(+ullage.volume/100) *fuel-mass/fuel-density;
289
fuel.tank-radius=(fuel.tank.volume.
29W
fuel-tank-thickness=propellant-pressure.
291
(MMH)
tank.yield..strength/2;
115
/(4/3*pi) ) .^ (1/3) ;
*fuel-tank.radius*FS/...
292
293
fuel-tankmass=tank-ffactor*4/3*pi*(
(fuel-tank-radius+...
^3) .*tank-density;
fuel..tank-thickness) .^3-fuel-tank-radius.
294
295
%Calculate oxidizer tank mass
29
oxidizer-density
297
oxidizer.tank-volume=(1+ullage..volume/100)
298
oxidizer..tank-radius=(oxidizer-tank-volume.
2
oxidizer-tank-thickness=propellant-pressure.
300
301
= 1440;
%kg/m3 (NTO)
*oxidizer.mass/oxidizer-density;
/ (4/3*pi)) .^ (1/3) ;
*oxidizer-tank-radius...
*FS/tank-yield..strength/2;
oxidizer.tank-mass=tank-f actor*tank-factor*4/3*pi*
302
oxidizer-tank-radius+oxidizer-tank-thickness)
303
oxidizer-tank.radius. ^3) .*tank-density;
((.*.
.^3-.
304
305
%Calculate pressurant tank mass
306
pressurant.temperature=thrusterData. data (1, 16);
307
pressurant-molar.mass=thrusterData.data (1, 15) ;
% use N\12 gas
as pressurant
3So
pressurant-tank-density=thrusterData.data(1,14);
309
pressurant..tank.yield.st rength=thrusterData.data (1, 13) ;
% ...
use Ti alloy as tank materiai
310
311
%syms VO real
312
%pretty
(simplify (mono-t ank-mas s (mO,
pressurant-molar-mass,
pressurant-temperature,
propellant-pressure,
...
pressurant-tank-yield-strength,
propellant-tank-volume,
...
pressurant-tank-density)))
313
pre s surant-temperature,
%ezplot(mono-tank-mass(VO,
pressurant-molarmass,
propellant-pressure,
fuel-tank-volume+oxidizer-tank-voiume,
pressurant-tank-yieLd-strength,
pressurant-tank-density),
...
...
...
[0, (fuel-tank-volume+oxidizer-tank-voiume)*.2])
314
315
pressurant-tank-volume
=
fminsearch(@(pressurant-tank-volume)
mono..tank-mass (pressurant..tank-volume,
pressurant-temperature,
...
...
pressurant..molar.mass,
116
...
. . .
propellant-pressure,
fuel-tankvolume+oxidizer-tank-volume,
pressurant..tank.density) , O);
pressurant-tank-yield-strength,
316
pressurant-tank...mass
...
=
...
t ank-f actor*mono..t ankmas s (pressurant..t ankvolume,
pressurant-molar-mass,
pressurant-temperature,
propellant-pressure,
...
.. .
fuel-tank-volume+oxidizer-tank-volume,
pressurant-tank..yield.strength,
...
pressurant..tank.density);
317
318
%initial pressurant-volume
=
fminsearch(@ (initial-pressurantvolume)
...
mono-tank-mass2 (initial-pressurant.volume,
pressurant-temperature,
pressurant-molar-mas s,
propellant-pres sure,
propellant-tank-volume,
tank-yield-strength,
...
...
tank-density) ,propellant-tank-volume*.1)
319
%tank-mass = mono-tank-mass2 (initial-pressurant-volume,
pressurant-temperature,
pressurant..molar-mass,
propellant-pressure,
propellant-tank-volume,
tank-yield-strength,
tank-density)
320
321
if
pressurant.tank-volume
< 0
322
fprintf('\nERROR: tank optimization incorrect\n')
323
thruster-type
324
deltaV
325
total-initial..mass
326
max.acceleration
327
end
328
329
%propellant.mass
330
%fuel.mass
331
%fueltank-volume
332
%f ue l..t ank..na s s
333
%oxidizer-mas s
334
%oxidizer-tank.voIume
...
117
335
%oxidizer-tank-mass
336
%pressurant-tank-volume
337
%pressurant-tank-mass
338
339
tank-mass=fue1-tank-mass+oxidizer.tank.mass+pressurant-tank-mass;
340
341
case 7 %Water Resistojet
342
exhaust-velocity
343
max-power
344
thruster.mass = 0.0149*max-power^0.7736;
7) *g;
= thrusterData. data (thruster-type,
= 5156.5*max..thrust^1.2627;
345
346
%Calculate propellant mass
347
propellant-mass=total.initial.mass.*(l-exp(-deltaV./exhaust-velocity));
348
349
%Calculate propellant tank mass
350
propellant-pressure=4E6*max.thrust
351
prope llant-_tank-volume= (1+ullage-volume/100) *propellant.mass/...
352
+ 112214;
thrusterData.data(thruster-type,12);
(1/3);
ank-volume. / (4/3*pi))
353
tank-radius= (propellant-t
354
tank...thickness=propellant.pressure.
355
propellant-tank.mass=tank-factor*4/3*pi* ( (tank..radius+. . .
356
*tank-radius*FS/tank..yield-strength/2;
tank-thickness) .^3-tank-radius.^3) .*tank.density;
357
358
%Calculate pressurant tank mass
359
pressurant-temperature=thrusterData.
360
pressurant..molar-mas s=thrusterData.data (1, 15) ;
data (1, 16);
% use N2 gas
...
as pressurant
361
pressurant-tank-density=thrusterData. data (1, 14);
362
pres surant-tank..yie ld.strength=thrusterData.data
(1, 13) ;
%
use Ti alloy as tank material
363
364
%syms VO real
365
%pretty
(simplify (mono-tank.mass (mO,
pressurant-molar-imass,
propellant.tank-volume,
pressurant.temperature,
propellant...pres sure,
pressurant-tank-yield.strength,
pressurant-tank-density)))
118
...
...
366
pres surant.temperature,
%ezplot (mono-tank-mass (VO,
pressurantmolarmass,
propellantpres
propellant-tank-volume,
.
sure,
..
pressurant-tank-yield-strength,
pressurant-tank-density),
...
[0, propellant-tank...volume*.2])
367
368
pressurant..tank-volume
=
fminsearch(@(pressurant-tank-volume)
...
mono-tank.mass (pre ssurant.tank-volume,
pressurant.temperature,
propellant-pres sure,
pressurant-molar-mass,
...
...
propellant.tank-volume,
pressurant..tank-yield-strength,
369
...
pressurant..tank-density),0);
pressurant-tank..mass = ...
t ank.factor*mono.t ank..mas s (pres surant..t ank.volume,
pressurant-molar-mass,
pressurant-temperature,
propellantpres
sure,
...
...
propellant.tank-volume,
pressurant..tank..yield.strength,
pressurant-tank-density);
370
371
%init ial .pressurant-volume
=
fminsearch (@ (initial-pressurant-volume)
...
mono-tank-mass2 (initial-pres surant-volume,
pressurant-temperature,
...
pressurant.molar.mass,
propellant_pres sure,
propellant-tank-volume,
tank-yieldcstrength,
...
...
tank-density) , propellant-tank-volume *.1)
372
%tank-mass = mono-tank-mass2 (initial.pressurant-volume,
pressurantmolar-mass,
pre ssurant.temperature,
sure,
propellantpres
tank-yield-strength,
propellant-tank-volume,
tank-density)
373
374
if
pressurant.tank-volume
375
fprintf(
376
thruster..type
377
deltaV
378
total.initial-mass
379
max-acceleration
380
'\nERROR:
<
0
tank optimization
end
119
...
incorrect\n')
...
381
382
%propellant-mass
383
%propellant-tank-volume
3M4
%propellant..tank-mass
385
%pressurant.tank-volume
386
%pressurant.tank-mass
387
388
tank-mass=propellant-tank..mass+pressurant-tank-mass;
389
390
case 8 %Hydrazine Resistojet
+ 770.3*max..thrust
391
max-power = 373.54*max..thrust^2
392
exhaust-velocity = 186.93*max-power^0.075*g;
393
thruster..mass
+ 52.949;
= 0.1119*max-power^0.2762;
394
395
%Calculate propellant mass
396
propellant-mass=total-initial..mass.*(1-exp(-deltaV./exhaust.velocity));
397
398
%Calculate propellant tank mass
399
propellant..pressure=818 966*log (max..thrust)
400
prope llant..t ank-volume= (1+ullage..volume/100) *propellant..mass/. . .
401
+ 3E6;
thrusterData.data(thruster-type,12);
(1/3);
402
tank-radius= (propellant.tank-volume. / (4/3*pi))
403
tank..thickne s s=prope llant.pres
404
propellant..tank.mass=tank-f actor*4/3*pi* ( (tank-radius+. . .
405
tank..thickness) .^3-tank-radius.^3) .*tank-density;
.^
sure. *tank-radius*FS/tank..yield-strength/2;
406
407
%Calculate pressurant tank mass
408
pressurant..temperature=thrusterDat a. data (1, 16);
409
pressurant-molar.mass=thrusterData.data (1, 15) ;
% use N2 gas
as pressurant
410
pres surant-t ank-dens ity=thrusterData. data (1, 14);
411
pres surant-tank-yie1d-strength=thrusterData.data
(1, 13);
%
use Ti alloy as tank material
412
413
%syms VO real
414
%pretty
(simplify (mono.tank.mass (mO,
120
pressurant.temperature,
...
pressurant.molar.mass,
propellant-pressure,
rength,
...
ank-yield-strength,
...
ank-yieldst
pressurant-t
prope llant-tank-volume,
. . .
pressurant-tankdensity)))
415
pre s surant-temperature,
%ezplot(mono-tank-mass(VO,
propellant..pressure,
pressurant-molar-mass,
prope llant-t
ank-volume,
pressurant..
pressurant.tank...density),
...
[O,propellant-tank-volume*.2])
416
417
pressurant.tank-volume =
fminsearch(@ (pressurant-tank-volume)
...
mono-tank.mass (pre ssurant-tank-volume,
. ..
pressurant..temperature, pressurant-molar.mass,
propellant.pressure,
418
pressurant..tank-mass =
...
propellant..tank-volume,
pressurant..tank.yield-strength,
...
pressurant..tank.density) , 0);
.. .
tank-factor*mono-t ank-.mass (pressurant.tank.volume,
pressurant-temperature, pressurant.molar-mass,
propellant-pres sure,
...
...
propellant-tank-volume,
pressurant-tank.yield-strength,
pressurant-tank-density);
419
420
%initial.pressurant-volume =
fminsearch(@ (initial-pressurant-volume)
mono-tankrmass2 (init
ialipres
pressurant.temperature,
...
surant-volume,
...
pressurant-nolar-mass,
propellant-pressure,
propellant-tank-volume,
tank-yield-strength,
...
...
tank-density) ,propellanttank-volume*.1)
421
%tank-mass = mono-tank-mass2 (initial-pressurant..volume,
pressurant.temperature,
pressurant..molar-mass,
propellant-pres sure,
propellant.tank-volume,
tank-yield-strength,
tank-density)
...
422
423
if
pressurant-tank..volume < 0
424
fprintf('\nERROR: tank optimization incorrect\n')
425
thruster.type
426
deltaV
121
...
427
total-initial-mass
428
max.acceleration
429
end
430
431
%propellant-mass
432
%propellant-tank-volume
433
%propellant-tankrnass
434
%pressurant-tank-volume
435
%pressurant-tank-mass
436
437
t ank-mass=propellant-tank-mas s +pres surant-tankmas s;
438
439
case 9 %Ammonia arcjet
+ 1181.7*max-thrust
440
max-power = 5835.8*max-thrust^2
441
exhaust-velocity = (89.808*log(max-power)
442
thruster-mass = 0.001*max-power^0.9677;
-
+ 292.45;
97.058)*g;
443
4
445
%Calculate propellant mass
propellant-mass=total-initial-mass .* (1-exp(-deltaV./exhaust-velocity));
446
447
%Calculate propellant tank mass
448
propellant-pressure=thrusterData.data (thruster-type,11);
449
propellant-tank-volume= (1+ullage-volume/100) *propellant-mass/...
450
thrusterData.data(thruster-type,12);
(4/3*pi))
(1/3);
451
tank-radius=(propellant-tank-volume./
452
tank-thickness=propellant-pressure.
453
propellant-tank-mass=tank-f actor*4/3*pi* ( (tank-radius+...
454
.^
*tank-radius*FS/tank-yieldstrength/2;
tank-thickness) .^3-tank-radius. ^3) .*tank-density;
455
456
%Calculate pressurant tank mass
457
pressurant-temperature=thrusterData.data(1,
458
pressurant-mo laramass=thrusterData.data(1,15);
16);
% use N2 gas
as pressurant
459
pressurant-tank-density=thrusterData.data (1,14);
460
pressurant-tank-yield-strength=thrusterData.data(1,13);
use Ti alloy as tank material
122
...
461
462
%syms VO real
463
%pretty
(simplify (mono.tankmass (mO,
pressurant..molar-mass,
pressurant..temperature,
propellant-pressure,
...
propellant..tank-volume, pressurant-tank-yield-strength,
pressurant-tank-density)))
464
%ezplot(monotankmass(VO, pressurant..temperature,
pressurant-molar..mass,
propellant.pressure,
propellant-tank-volume,
...
pressurant-tank-yield-strength,
pressurant-tank.density),
[0, propellant-tank-volume*. 21)
465
466
pressurant.tank.volume
fminsearch(@ (pressurant-tank-volume)
...
mono.tank-mass (pre ssurant-t ank-volume,
pressurant...temperature,
pressurantmolar.mass,
pressurant.tank-yield-strength,
...
.
propellant-tank-volume,
propellant.pressure,
467
.. .
pressurant.tank.density) , O);
pressurant-tank-mass = ...
t ank..factor*mono-tank.mas s (pres surant-tank-volume,
pressurant-temperature, pressurant-molar-mass,
propellant-pressure,
...
...
propellant-tank-volume,
pressurant-tank-yield-strength,
...
pressurant-tank-density);
468
469
%initial.pressurant-volume
fminsearch (@ (init
=
ial...pres surant _volume)
...
mono-tank..nass2(initial-pressurant-volume,
pressurant-temperature,
...
pressurantmolar-mass,
propellant-pressure,
propellant-tank-volume,
tank-yield-strength,
...
tank-density) ,propellant..tank-volume*.l)
470
%tank-mass = mono.tank-mass2 (initial.pressurant-volume,
pressurant-temperature,
pressurant..molarrmass,
propellant -pressure,
propellant.tank-volume,
tank.yield-strength,
tank-density)
471
472
if
pressurant-tank-volume <
0
123
...
...
tank optimization incorrect\n')
473
fprintf('\nERROR:
474
thruster-type
475
deltaV
476
total-initial-mass
477
max-acceleration
478
end
479
480
%propellant-mass
481
%propellant-tank..volume
482
%propellant-tank-mass
483
%pressurant..tank-volume
484
%pressurant-tank-mass
485
486
tank-mass=propellant-tank-mass+pressurant.tank-mas
s;
487
488
case 10 %Hydrogen arcjet
489
max.power = 18886*max-thrust^1.0172;
490
exhaust-velocity
491
thruster-mass
=
(1270.6*max-thrust^0.1019) *g;
= 0.001*max-power^0.9677;
492
493
%Calculate propellant mass
494
propellant-mass=total-initial-mass.*(1-exp(-deltaV./exhaust-velcity));
495
496
%Calculate propellant tank mass
497
propellant-pres sure=thrusterData. data (thruster-type,
498
propellant-temperature=thrusterData.data(t'hruster.type,
499
propellant.molar.mass=thrusterData.data(thruster-type,15);
soo
initial-propellant.mass
=
...
cold-tank-mass (initial.propellant-mass,
propellant-pressure,
16);
...
fminsearch(@ (initial-propellant-mass)
propellant-temperature,
11);
propellant..mass,
propellant-molar-mass,
tank.yield.strength,
...
...
tank.density) ,propellant.mass) ;
501
t ank.mass=t ank.f actor*c old.tank-mas s (initial-propellant.mass,
propellant.mass,
propellant.temperature,
propellant-molar..mass,
...
propellant-pressure,.
124
..
...
tank-yield-strength,
tank-density);
502
503
if
< propellant-mass
initial-propellant.mass
504
fprintf('\nERROR:
505
thruster-type
506
deltaV
507
total-initial-mass
508
max-acceleration
509
tank optimization incorrect\n')
end
510
511
%syms mO real;
512
%simplify(estimate-tank-mass (mO,
propellant _temperature,
propellant-pressure,
513
propellant-temperature,
tank-density),
propellant..molar-mass,
tank-density))
tank-yieldstrength,
%ezplot(cold.tank-mass(mO,
propellant.pressure,
propellant-mass,
propellant-mass,
...
propellant-molar-mas s,
...
...
tank-yield.strength,
[propellantmass,propellant-mass*l.3]
)
514
515
%propellant-mass
516
%tank-mass
517
518
case 11 %Hydrazine arcjet
519
max-power = 1428.1*log(max-thrust)
520
exhaust-velocity
521
thruster-mass
+ 3762.3;
= thrusterData. data (thruster-type,
= thrusterData.data (thruster-type,
7) *g;
10);
522
523
%Calculate propellant mass
524
propellant-mass=totalinitial..mass.*(-exp(-deltaV./exhaust-velocity));
525
526
%Calculate propellant tank mass
527
propellant..pressure=thrusterData.data (thruster-type,
528
propellant.tank.volume= (1+ullage.volume/100) *propellant.mass/ ...
529
11);
thrusterData.data(thruster-type,12);
530
tank-radius= (propel1ant-tank-volume. / (4 /3*pi))
531
tank-thickness=propellant.pressure.
125
(1/3);
*tank-radius*FS/tank..yield.strength/2;
532
propellant.tank-mass=tank-f actor*4/3*pi* ( (tank-radius+...
533
tank..thickness) .^3-tank-radius.^3) .*tank-density;
534
535
%Calculate pressurant tank mass
536
pres surant.temperature=thrusterData. data (1, 16);
537
pressurant-molar-maass=thrusterData.data (1, 15) ;
% use N2 gas
as pressurant
ank-dens ity=thrusterData. data (1, 14);
538
pres surant-t
539
pres surant-tank-yield-strength=thrusterData.data
(1, 13) ;
% ...
use Ti alloy as tank material
540
541
%syms VO real
542
%pretty
(simplify (mono-tank..mass (mO,
pressurant-temperature,
propeIlant-pressure,
pressurant-molar.mass,
...
pressurant-tank-yielidstrength,
propellant.tank-volume,
pressurant-tank-density)))
543
%ezplot(mono-tank-mass (VO,
pres surant _temperature,
sure,
propellantpres
pressurant.molar-mass,
ank-yield-st rength,
pressurant-t
prope llant.t ank-volume,
..
...
[O,propellant.tank-volume*.2])
pressurant-tank-density),
544
545
pres s urant-tank-volume
=
fminsearch (@ (pres surant-tank-volume)
...
mono..tank-mass (pressurant-tank-volume,
pressurant-temperature,
propellant-pres sure,
pressurant
546
pressurant-molar-mass,
propellant..tank-volume,
-tank-yield-strength,
pressurant..tank..mass
.. .
...
pressurant..tank-density) , 0);
= ...
t ank-factor*mono-t ank.mass (pressurant-t
pressurant..temperature,
propellant-pressure,
ank-volume,
pressurant..molar.mass,
propellant..tank..volume,
pressurant..tank-yield-strength,
%initial-pressurant-volume =
fminsearch(@ (initial-pressurant-volume)
...
mono-tank-mass2 (initial-pressurant-volume,
126
...
...
...
pressurant-tank-density);
547
548
...
pressurant-molar-mass,
pressurant-temperature,
sure,
propellantpres
propellant.tank-volume,
tank-yield-strength,
tank.density)
549
...
...
...
propellant-tank-volume*.1)
%tank-mass = mono.tank.nass2(initial-pressurant-volume,
pressurant..molar..mass,
pressurant-temperature,
propellant-pres sure,
propellant..t ank-volume,
tank-yield-strength,
tank-density)
...
...
550
551
if
pres surant.tank.volume
552
fprintf('\nERROR:
553
thruster.type
554
deltaV
555
total..initial-mass
556
max-acceleration
557
<
0
tank optimization incorrect\n')
end
558
559
%propellant.mass
560
%propellant-tank-volume
561
%propellant-tank-mass
562
%pressurant-tank-volume
563
%pressurant-tank-mass
M64
565
tank...mass=propel lant.tank-mass+pressurant-t ank.mas s;
566
567
case 12 %Xenon Ion Engine
5"
max..power=64724*max.thrust^2
569
exhaustvelocity
=
+ 18165*max-thrust
(-4E-6*max-power^2
+ 0.357*max-power
2665) *g;
570
thruster-mass=61.162*max.thrust
+ 0.6761;
571
572
%Calculate propellant mass
573
propellant.mass=total-initialmass.*(1-exp(-deltaV./exhaust-velocity));
574
575
%Calculate tank mass
576
propellant-pressure=thrusterData. data (thruster-type,11);
127
+ 54.78;
+
577
propellant-temperature=thrusterData.data (thruster.type,
578
propellant-molar.mass=thrusterData.data
579
initial-propellant..mass = ...
...
propellant-mass,
cold-tank-mass (initial-propellant.mass,
propellant-pressure,
15);
(thruster-type,
fminsearch(@ (initial-propellant..mass)
propellant.temperature,
16);
propellant_molar.mass,
tank-yield-strength,
.
...
...
tank-density) ,propellant.mass);
s8o
t ank..mas s=t ank-f actor*cold-tank.mas s (init
propellant-mass,
ial.propellant..mas s,
propellant-temperature,
propellant.pressure,
propellant..molar-mass,
tank-yield-strength,
...
...
tank-density);
581
582
if
initial-propellant..mass
583
fprintf('\nERROR:
584
thruster-type
585
deltaV
586
total-initialmass
587
max-acceleration
58s
< propellant.mass
tank optimization incorrect\n')
end
589
590
%syms mO real;
591
%simplify(estimate.tank..mass (mO,
propellant-molar-mass,
propellant-temperature,
propellant..pressure,
592
tank-yied.strength,
%ezplot(cold-tank..mass(mO,
propellant-mass,
tank-density),
...
tank.density))
...
propellant-molar-mas s,
propellant.temperature,
propellant-pressure,
propellant-mass,
tank-yield.strength,
...
[propellant-mass, propellant-mass*1.l])
593
594
%propellant.mass
595
%tank...mass
596
597
598
599
case 13 %Hall thruster
max.power=837.41*max-thrust^
3
18713*max..thrust + 222.35;
128
-
3.0262*max.thrust^2
+
...
=
(518.65*max..power^0.1702) *g;
600
exhaust-velocity
601
thruster..nass=0.0107*max.power^0.7786;
602
60s
%Calculate propellant mass
604
propellant..mass=total-initial-mass.*(l-exp(-deltaV./exhaust-velocity));
605
606
%Calculate tank mass
607
propellant-pressure=thrusterData.data (thruster.type,11);
608
propellant..temperature=thrusterData.data
609
propellant-molar.mass=thrusterData.data (thruster.type,
610
initial-propellant-mass
=
...
propellant..mass,
cold.tank.mass (initial.propellant.mass,
propellant-pressure,
15);
...
fminsearch(@ (initial-propellant-mass)
propellant-temperature,
16);
(thruster-type,
propellant..molar.mas s,
...
...
tank...yield.strength,
tank.density) ,propellant-mass) ;
611
ank-ma s s (init
t ank-ma ass=t ank.f actor*cold.t
propellant..mass,
ial-propellant..mas s,
propellant-temperature,
propellant-pressure,
propellant..molar-mass,
tank-yield-strength,
...
tank-density);
612
613
if
< propellant.mass
initial-propellant.mass
614
fprintf('\nERROR: tank optimization
615
thruster.type
616
deltaV
6i7
total..initial-mass
618
max-acceleration
619
incorrect\n')
end
620
621
%syms mO real;
622
%simplify(estimate-t ank-mass (mO,
propellant-temperature,
propellant-pressure,
623
propellant.molar.mass,
tank.density))
tank-yield.strength,
%ezplot(cold-tank-mass(mO,
propellant.temperature,
propellant-pres sure,
propellant-mass,
propellant-mass,
...
propellant-molar-mas s,
tank-yield-strength,
129
...
...
tank-density),
[propellant-mass,propellant-mass*1.1])
624
625
%propellant-mass
626
%tank-mass
627
628
case 14 %teflon pulsed plasma thruster
= 4784.7*max-thrust^0.2391*g;
629
exhaust-velocity
630
max-power = 6506.7*max-thrust^0.6868;
631
thruster-mass = 29.394*max-thrust^0.2567;
632
633
%Calculate propellant mass
634
propellant-mass=total-initialamass.*(1-exp(-deltaV./exhaust-velocity));
635
636
%Unpressurized solid propellant
637
tank-nass=.05*propellantamass;
638
639
case 15 %Chromium vaccum arc thruster
*g;
= thrusterData.data(thruster-type,7)
64o
exhaust-velocity
64u
max-power = thrusterData.data(thruster-type,
642
thruster-nass
9);
= thrusterData.data (thruster-type,
10);
643
61
%Calculate propellant mass
645
propellant-mass=total-initialinass
.*(1-exp(-deltaV./exhaust-velocity));
646
647
%Unpressurized solid propellant
648
tank-mass=. 05*propellantamass;
649
650
case 16 %field emission electric propulsion
651
max-power = 99014*max-thrust^1.0456;
652
thruster-mass
653
exhaust-velocity
= 0.2141*max-power^0.6287;
= thrusterData.data(thruster-type,7)
*g;
654
655
%Calculate propellant mass
656
propellantmnass=total-initial-mass.*(1-exp(-deltaV./exhaust-velocity));
657
658
%Unpressurized solid propellant
130
659
tankrnass=. 05*propellantamass;
660
661
case 17 %colloid thruster
exhaust-velocity
663
max-power = 47501*max-thrust^1.1542;
64
thruster-rnass
*g;
= thrusterData.data(thruster-type,7)
662
10);
= thrusterData.data(thruster-type,
665
666
%Calculate propellant mass
667
propellant.mass=total-initial-nass.*(1-exp(-deltaV./exhaust-velocity));
668
669
%Calculate tank mass
670
propellant-pressure=thrusterData.data(thruster-type,11);
671
propellant-tank-volume= (1+ullage-volume/100) *propellantxnass/. ..
672
thrusterData.data(thruster-type,12);
673
tank-radius=(propellant-tank-volume./
674
tank-thickness=propellant-pressure.
675
tank-nass=tank-factor*4/3*pi*(
676
-tank-radius. ^3) .*tank-density;
(4/3*pi))
(1/3);
*tank-radius*FS/tank-yield-strength/2;
(tank-radius+tank-thickness) .^3...
677
678
case 18 % Electrospray -historical
equations are used instead of
...
data
= thrusterData.data(thruster-type,7)
*g;
679
exhaust-velocity
680
efficiency = thrusterData.data(thruster-type,8)/100;
681
682
propellant-nass=total-initial-mass.*(1-exp(-deltaV./exhaust-velocity));
683
684
specific-mass=6.32;%kg/N
with 300 micrometer
685
(representative
of a Ni ILIS array
spacing)
thrusteranass=specific-nass*max-thrust;
686
687
% unpressurized liquid propellant
688
tankamass=.05*propellantamass;
689
690
max-power=max-thrust*exhaust-velocity/2/efficiency;
691
692
otherwise
131
fprintf('\nERROR: The thruster type you have requested is
693
unavailable. Please see thrustertradespace.xls')
694
end
695
696
697
%if
(thruster-type,
thrusterData.data
698
%Calculate power system mass
699
power-mass
700
%else
701
%
702
%end
2)==O
= max-power*specific-power*thrusters;
power-mass=O;
703
704
%Calculate inert mass, add %10 for feed lines and supports
705
margin =
706
inert.mass=thruster-mass *thrusters+tank..mass+power-mass+margin;
.1*(thruster-mass*thrusters+tank.mass);
707
708
propulsion-mass...fraction
...
= (inert..mass+propellant-mass) /total-initial.mass;
709
propellant-mass.fraction
=propellant..mass/total-initial-nass;
710
711
%thruster-mass
712
%thrusters
713
%propellant-mass
714
%tank.mass
715
%power-mass
716
%margin
717
%inert-mass
718
%propulsion-mass-fraction
719
%max.power
132
i
2
function tankamass
= mono-tank-ass(pressurant-tank-volume,
pressurant-temperature,
pressurantamolaramass,
propellant-tank-volume,
tank-yield-strength,
...
final-pressure,
...
tank.density)
%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%
3
4
% Filename:
mono-tankamass.m
5
% Filetype:
Function
6
% Project:
F6
7
% Author:
Thomas Chiasson
8
% Mod Date:
19/8/2011
9
10
% PURPOSE:
Calculates the mass of a pressurant tank.
11
12
This program calculates the mass of a pressurant tank
% USAGE:
(including pressurant) for a blowdown system of a liquid
13
It is called iteratively in an optimization
14
propellant.
is
routine in
16
propulsion systems using pressurized liquid propellant.
17
thrustermass.m while estimating the mass of
%
INPUT:
'pressurant-tank-volume
(m3)
Volume of pressurant tank
'pressurant-temperature'
(K)
Temperature of pressurant
(assumed constant)
'pressurant-molar-mass'
(kg/mol) Molar mass of pressurant
(Pa)
Minimum required pressure for
'propellant-tank-volume'
(m3)
Volume of propellant
'tank-yield-strength'
(Pa)
Stress at which tank material yields
'final-pressure'
...
thruster
'tank-density'
(kg/m3)
Density of tank material
% OUTPUT:
%
'tank-mass'
(kg)
Mass of pressurant tank
(including pressurant)
%
%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%
global R FS;
133
34
35
pres surant..ma s s=final.pres sure* (pressurant.t
...
propellant...tank-volume) *pressurant.molar.mass/ (R*pressurant..temperature);
36
37
ank-volume+
initial-pressure=pressurant.mass*R*pressurant-temperature/
pres surant-t
38
...
ank-volume*pres surant..molarma s s);
/ (4/3*pi))
(1/3);
39
tank-radius=(pressurant-tank-volume.
40
tank-thickness=initial.pressure.*tankradius*FS/tank-yield..strength/2;
41
tank.mass=4/3*pi* ( (tank...radius+tank-thickness) .^3-tank..radius.
tank.density+pressurant..mass;
42
43
44
%init
ial.pres
sure
134
^3)
.*.
i
function tankamass = cold-tankamass(initial-propellantamass,
propellant-mass,
propellant-temperature,
propellant-pressure,
tank-yield-strength,
...
propellantamolarmass,
tank-density)
2
%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%
3
%
4
% Filename:
cold-tankmass.m
5
% Filetype:
Function
6
% Project:
F6
7
% Author:
Thomas Chiasson
8
% Mod Date: 19/8/2011
9
%
10
% PURPOSE: Calculates the mass of a spherical cold gas propellant tank.
11
% USAGE:
This program calculates
%
tank. It is called iteratively in an optimization
%
routine
%
propulsion systems using pressurized gas propellant.
in
the mass of a gas propellant
thrusteramass.m while
estimating the mass of
%
% INPUT:
% 'initial-propellant-mass
'
(kg)
Mass
of propellant before thruster
use
% 'propellant-mass'
(kg)
Mass of propellant used over thruster
lifetime
(K)
% 'propellant-temperature'
Temperature of propellant assumed
constant
% 'propellant-molaramass'
(kg/mol) Molar mass of propellant
(Pa)
% 'propellant-pressure'
Minimum required pressure for
thruster
(Pa)
% 'tank-yield-strength'
% 'tank.density'
(kg/m3)
Stress at which tank material yields
Density of tank material
29
30
% OUTPUT:
31
% 'tank-mass'
32
%
(kg)
Mass of tank, including leftover
propellant
33
34
%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%
135
35
global
R FS;
36
37
leftover-propellant.mass=initial..propellant..mass-propellant-mass;
38
tank..volume=lef tover.propellant..mas s*R*propellant.temperature/
39
40
41
(...
propellant-pressure*propellant..molar-mass);
max.pressure=initial-propellant-mass*R*propellant-temperature/
(...
t ank-volume *propellant-molar.mass) ;
(4/3*pi)) . ^ (1/3);
42
tank-r adius=(tankvolume./
43
tank-thickness=max-pressure. *tank-radius*FS/tank-yield-strength/2;
44
tank..mass=4/3*pi* ( (tank-radius+tank.thickness) .^3-tank.radius.^3)
45
tank..density+leftover..propellant-mass;
136
.*
2
%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%
3
%
4
% USAGE:
This is the script used to generate the graphs in Chapter
5
%
6.
6
%
visibility purposes.
7
%
8
%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%
I recommend toying with marker-type
on line
39 for
9
10
clear all
11
clf
12
clc
13
= logspace(0,5,26);
14
deltaV-space
15
max-thrust-space
16
total-initialamass
= logspace(-5,1,31);
100;
=
17
18
margins =
.2;
19
marker-size
20
mass-fractionlimit
= 60;
=
.9;
21
22
last-thruster=17;
23
last-deltaV=length (deltaV-space);
24
last-max-thrust=length (max-thrust-space);
25
26
figure(1)
27
hold on;
28
set(gca,'FontSize',14);
29
set(gca, 'XScale'
3o
set(gca, 'XLim',
31
set(gca,'YScale','log');
32
set(gca, 'YLim',
33
,'Ilog')
;
[deltaV-space(1)*(1-margins) ,deltaV-space (last-deltaV)
[max-thrust-space (1) * (1-margins) ,max-thrust-space(...
lastmax-thrust)
* (1+margins)]);
34
xlabel(' \DeltaV
35
ylabel('Max Thrust
36
title(['Lowest Mass Propulsion Option for
(m/s) ') ;
(N)');
137
*(1+margins)]);
'num2str (total-initialamass),.
'kg
37
Satellite']);
38
39
marker-type={'bo','gp','r+','ys','g.','kx','g+','r*','g*', 'bs','yp','Iydl,
'k+', 'bp','b.','go','rd', 'mh'};
40
41
marker-label={'Cold Gas
(N2)','Cold Gas
(He)','Cold Gas
(Butane)',...
(H202)',...
42
'Monopropellant (H4N2)','Monopropellant
43
'Bipropellant (NTO/MMH)','Resistojet (H20)','Resistojet (H4N2)',...
44
'Arcjet (NH3)','Arcjet (H2)','Arcjet (H4N2)','Ion Engine
45
'Hall Thruster
46
'Colloid (EMI-BF4)','Electrospray(EMI-BF4)'};
(Xe)',...
(Xe)','PPT (teflon)','VAT (Cr)','FEEP',...
47
48
for n = 1:last-thruster
hgg (n) =hggroup;
49
5o
end
51
52
deltaV-point=1
53
54 while deltaV-point <= last-deltaV
55
56
max-thrust -point=last
-naxJthrust
57
58
while max-thrust-point >
0
59
60
thruster=l;
61
62
while thruster
<= last-thruster
63
64
[propellantamass-fraction,
propulsionxmass-fraction (thruster),
propellant-mass,
inert-mass,
...
max-power]=thruster-mass(deltaV-space(deltaV-point),
total-initial-mass,
...
...
max-thrust-space (max-thrust-point) /total-initial-mass,
thruster);
65
138
..
if propellant-mass-fraction<=0
66
||
.
propulsion..mass-fract ion (thruster)<=O
67
.
|
inert-mass<=O
I|
I
68
max.power<=O
||
69
max-power+inert-mas s+propuls ion-mas s-f ract
.. .
ion+...
imag (propellant-mass-f ract
ion (thruster)
~=0
70
fprintf('\nERROR: thruster characteristics
71
thruster
72
deltaV-space (deltaV-point)
73
max-thrust-space (max-thrust-point)
74
pause
incorrect\n')
end
75
76
78
= thruster+1;
thruster
77
end
79
so
[ lowest-mass.fract ion (deltaV-point, max-thrust.point) ,
min (...
(deltaV-point,max-thrust-point)
81
best..thruster
82
propulsion-mass.fraction (1: last-thruster))
83
84
if
max-thrust..point
< last-max.thrust
&& ..-
lowest-mass.fraction (deltaV.point, max.thrust..point)
>
...
lowest.mass-fraction (deltaV..point, max-thrust..point+1)
(deltaV-point,max-thrust-point ) =
lowest..mass..f raction
85
lowest.mass-fraction (deltaV-point,max.thrust-point+l);
(deltaV-point , max.thrust.point)
best-thruster
86
best-thruster
87
=
. . .
(deltaV..point, max-thrust-point+1)
end
88
89
if
lowest -mass..fraction (deltaV-point, max-thrust.point)
<
mass-fraction-limit
90
91
scatter
(deltaV.space (deltaV-point),
max-thrust-point),
marker-size,
max-thrust.space
'filled',
marker-type{. ..
92
93
best-thruster
(deltaV-point, max-thrust-point)},
'displayname',marker-label{best-thruster(delta-point,...
139
(...
)
},
'parent',
94
max.thrust.point)
95
deltaV.point, maxthrust.point)
set (get (get (hgg (best..thruster
96
hgg (best.thruster
(...
)) ;
(deltaV-point, max.thrust-point)),...
'Annotation'),'LegendInformation'),'IconDisplayStyle','on');
97
end
98
99
max...thrust-point
100
= max..thrust-point-1
end
101
102
deltaV-point
103
104
= deltaV-point+1
end
105
106
%for n = l:last.thruster
107
%
set(get(get(hgg(n),'Annotation'),'LegendInformation'),'IconDisplayStyle','on');
108
%end
109
110
leg=legend(hgg,marker-label,
'location',
'northeastoutside',
140
'FontSize',10);
i
function
[imp,
min-bit,
range,
thrust]=getThrust(thruster-type,
level)
2
3
%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%
4
%
s
% Filename: thrust.m
6
% Filetype:
Function
7
% Project:
F6
8
% Author:
Thomas Chiasson
9
% Mod Date:
18/8/2011
10
%
11
% PURPOSE:
Returns thrust information on given thruster types.
12
13
% USAGE:
This is a very useful program if you want to call
14
%
thrusteramass.m
15
%
when you don't have any information on the propulsion
16
%
types. It will find the thrust range of a propulsion
17
%
type from the database and create a user defined
18
%
logarithmic scale so that the user gets a complete picture
19
%
of a thruster's capabilities.
from within a
iteratively
larger
program
20
21
INPUT:
22
'thruster-type'
Row number of desired thruster in
'thrustertradespace.xlx.'
23
24
'range'
Number of logarithmic steps to divide thrust range
25
'level'
The desired step of thrust out of
'range'
26
27
OUTPUT:
28
'imp'
29
'thrust'
(N)
the thurst at
30
'min-bit'
(Ns)
average min impulse for
1 for chemicai thrusters,
0 for electric
'level' out of the entire thrust range
'thrust'
31
32
%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%
33
3
35
36
% Import the file
[numbers,
if
strings] = xlsread('thrustertradespace.xlsx',
~isempty(numbers)
141
'Input Calcs');
thrusterData.data =
37
38
end
39
if
~isempty(strings)
thrusterData.textdata =
40
41
numbers;
strings;
end
42
43
% Create new variables in the base workspace from those fields.
4
vars = fieldnames(thrusterData);
45
for i
1:length(vars)
vars{i},
assignin('base',
46
47
=
thrusterData.(vars{i}));
end
48
49
% extract values
5o
imp=thrusterData.data(thruster-type,2);
% will
be 1
0 for electric
thrusters,
si
min-thrust=thrusterData.data(thruster-type,4);
52
max-thrust=thrusterData.data(thruster-type,
5
% create thrust range
55
thrusts
6);
= logspace(log10(min-thrust) ,loglO (max-thrust) ,range);
56
57
thrust=thrusts(level);
58
59
switch thruster-type
60
case
62
otherwise
min-bit=thrusterData.data (thruster-type,
63
6
3
min.bit= 0.0097*thrust^0.8823;
61
for chemical
end
142
3);
Bibliography
[1] Common Arcjet Advantages and Mission Examples. A Very Low Power Arcjet
(VELARC) for Small Satellite Missions, pages 1-11. 2011. 83, 86, 89, 90, 91,
92, 93, 94, 98
[2] Tareq Bin Ali. Potential propulsionsystem for microsatellites. Master's thesis,
University of London, 2010. 103
[3] M Andrenucci, L Biagioni, F Ceccanti, M Saviozzi, and D Nicolini. Endurance
tests of 150 n feep microthrusters. October, pages 1-5, 2005. 102
[4] N. Antropov, A. Pokryshkin, G. Popov, A. Rudikov, and F. Scortecci. Comparative Analysis of PPT and SPT Propulsion Systems for a Given Thrust Total
Impulse. In Second European Spacecraft Propulsion Conference, May 1997. 96,
97
[5] EADS Astrium. Astrium Space Propulsion- Lampoldshausen, Germany. http:
//cs. astrium. eads.net/sp/home-lam.html, 2003-2012. 31, 36, 70, 86, 88, 93
[6] Encyclopedia Astronautica. Index: Engines. http: //www. astronautix. com/
engines/index.htm, May 2012. 84, 86, 87, 90, 91, 92, 93, 94, 96, 100
[7] Monika Auweter-kurtz, Thomas Go, Harald Habiger, Frank Hammer, Helmut
Kurtz, Martin Riehle, and Christian Sleziona. High-power hydrogen arcjet
thrusters. Journal of Propulsion and Power, 14(5):764-773, 1998. 43, 91
[8] Olivia Billet. Micropropulsion for nanosatellites. http://www.scribd.com/
jkleynhans/d/52715165-power-point-example, 2006. 88, 89, 101
[9] Travis A. Burton. Exhaust plume characterization of a mini-ppt using a timeof-flight/gridded energy analyzer. Master's thesis, University of Washington,
2002. 98, 99, 100, 101
[10] Inc Busek Co. Busek high power hall thrusters.
indexhtm_files/70008511.pdf, 2011. 94
http://www.busek.com/
Pulsed plasma thrusters.
[11] Inc Busek Co.
technologies__ppt.htm, May 2012. 98
http://www.busek.com/
143
http://
Proportional Micro Thruster.
[12] Bradford Engineering B.V.
bradford-space. com/userfiles/Proportional2OMT.pdf, May 2012. 83
[13] M. Capacci, G. Matticari, G. E. Noci, A. Severi, A. Ricciardi, and F. Svelto.
Radiofrequency with Magnetic field ion Thruster (RMT): review of the Engineering Phase accomplished under ASI Contract. Technical report, LABEN
Proel Tecnologie Division, Viale Machiavelli, 31
2012. 93
50125 Firenze (Italy), May
[14] R J Cassady, W A Hoskins, M Campbell, and C Rayburn. A micro pulsed
plasma thruster (ppt) for the dawgstar spacecraft. 2000 IEEE Aerospace Con-
ference Proceedings Cat No00TH8484, pages 7-13, 2000. 99, 100
[15] Vincent Cate. Replacing Ballast Momentum.
thrust.html, 2003. 94, 96
http: //spacetethers. com/
[16] Aerospace Research Information Center. Launch Vehicle Propulsion. http:
//www.aric.or.kr/trend/accessory/list.asp?classify=11&page=2,2005.
86, 88
[17] Northrop Grumman Corporation. Monopropellant Thrusters. http: //www. as.
northropgrumman. com/products/monoprpellant thrust/index.html, 2012.
87
[18] Jorge Delgado, Michael M. Micci, Sven G. Bilen, and Benjamin Welander. A
Study of a Low-Power Microwave Arcjet Thruster Using Ammonia Propellant.
Technical report, Pennsylvania State University, 2003. 89, 92, 93, 94, 95, 96,
97, 99, 100, 103
[19] New Developments and Second Edition.
Rocket and spacecraft propulsion.
Engineering, 2008. 94, 96, 97
[20] International Electric, Propulsion Conference, and Exploration Agency. Development history and current status of dc-type ion. Group, pages 1-8, 2007.
92
[21] International Electric, Propulsion Conference, William G Tighe, Kuei-ru Chien,
Rafael Spears, and Dan M Goebel. Ion thruster development at 1-3 eti for small
satellite applications. InternationalElectric Propulsion Conference, pages 1-11,
2007. 92
[22] Jesse K. Eyer, Christopher J. Damaren, Robert E. Zee, and Elizabeth Cannon.
A Formation Flying Control Algorithm for the CanX-4&5 Low Earth Orbit
Nanosatellite Mission. Technical Report IAC-07-B4.6.04, University of Toronto
Institute for Aerospace Studies, 2500 University Drive NW, Calgary, Alberta,
Canada, July 2004. 85
144
[23] Trevor Fedie, Jason Breeggemann, Brian Evans, Mike Gavanda, Matt Gildner, Jeromie Hamann, Brian Nackerud, Andrew Smude, and Jordan Stewart.
Shackleton crater reconnaissance mission fdr. http: //www .docst oc .com/docs/
74196627/Shackleton-Crater-Reconnai ssance-Satellite-PDR, May 2012.
36, 70
[24] Douglas Fiehler, Ryan Dougherty, and David Manzella. Electric propulsion
system modeling for the proposed prometheus 1 mission, volume 3891, page
1013. 2005. 92, 93
[25] John E Foster, Tom Haag, Michael Patterson, George J Williams, James S
Sovey, Christian Carpenter, Hani Kamhawi, Shane Malone, and Fred Elliot.
The high power electric propulsion (hipep) ion thruster. AIAA, (September),
2004. 51, 92
[26] Dan M Goebel, James Polk, and Anita Sengupta. Discharge chamber performance of the nexis ion thruster. Jet Propulsion,(July):1-11, 2004. 93
http://nssdc.gsfc.nasa.gov/nc/
Bepicolombo.
[27] Ed Grayzeck.
masterCatalog.do?sc=BEPICLMBO, May 2012. 36, 69
[28] Moog Systems Group. Cold Gas Thrusters. https://wiki.umn.edu/pub/
FormationFlying/RelativeOrbit/spdcat-cold-gas.pdf, May 2012. 84
Spacecraft Components: Cold Gas Thrusters.
[29] Moog Systems Group.
http://www.moog.com/products/propulsion-controls/spacecraft/
components/cold-gas-thrusters/,2012. 84
[30] Ilan Hadar and Alon Gany. Augmentation of low power hydrazine thrusters.
Acta Astronautica, 32(4):275 - 281, 1994. 31, 38
[31] David Michael Harland and Ralph Lorenz. Space Systems Failures: Disasters
and Rescues of Satellites, Rockets and Space Probes. Praxis Publishing Ltd. 90
[32] Jw Hartmann, J Schein, and R Binder. Micro Vacuum arc thruster design for a
Cubesat class satellite, pages 1-7. Number 217. 2002. 101
[33] R R Hofer, P Y Peterson, D T Jacobson, and D M Manzella. Factors affecting
the efficiency of krypton hall thrusters. Performer QSS Group Inc Sponsor
NASA Glenn Research Center, 2004. 96
[34] Vlad Hruby, Manuel Gamero-Castao, Paul Falkos, and Suren Shenoy. Micro
Newton Colloid Thruster System Development, page 01281. 2001. 103
[35] Ronald W. Humble, Gary N. Henry, and Wiley J. Larson. Space Propulsion
Analysis and Design, 1995. 15, 31, 38, 64, 66
145
[36] Frank Gulczinksi III, Ronald A. Spores, and Johanes Stuhlberger. In-Space
Propulsion.Technical Report AFRIL-PR-ED-TP-2003-105, Air Force Research
Lab, 5 Pollux Drive, Edwards AFB, CA 93524, April 2003. 15, 25, 85, 87, 90,
91, 92, 93, 94, 96, 97, 98, 101, 102, 103
[37] VACCO Industries. 2 LBF COLD GAS THRUSTER. http://www.vacco.
com/vacco/pdfs/48003040.PDF, July 2004. 84
[38] VACCO Industries. ChEMS Micro-Propulsion System. http://www.vacco.
com/vacco/pdfs/mips2112.pdf, May 2012. 13, 30, 85
[39] I Introduction. The European HiPER Programme : High Power Electric, pages
1-9. 2011. 51, 92
[40] Robert Jankovsky, Sergery Tverdokhlebov, and David Manzella. High power
hall thrusters. 35th joint propulsion conference, (December):AIAA 992949,
1999. 95, 96, 97
[41] Hani Kamhawi, George Soulas, Luis Pinero, Daniel Herman, Jonathan VanNoord, Wensheng Huang, Rohit Shastry, Thomas Haag, John Yim, and George
Williams. Overview of hall thruster activities at nasa glenn research center.
In 32nd International Electric Propulsion Conference, Wiesbaden, Germany,
September 2011. 96
[42] Michael Keidar. Micro- Vacuum Arc Thruster for Nanosatellites. http://
gsf cir .gsf c .nasa.gov/colloquia/3545, November 2010. 101
[43] Michael Keidar, Taisen Zhuang, Alex Shashurin, Lubos Brieda, Tom Denza,
and Samudra Haque. Micro- Vacuum Arc Thruster for Nanosatellites. http:
//geoscan. jhuap1. edu/wsf iles/day2/4/Zhuang_GW-MPNL .pdf, March 2011.
101
Nasa's new hall thruster passes performance test[44] Sallie A. Keith.
ing. http://www.nasa. gov/centers/glenn/news/pressrel/2005/05-033
Thruster
.html,
August 2005. 94
http://mech371.engr.scu.edu/
Space propulsion.
[45] Vadim Khayms.
Lectures/Wk9/2009X20PropulsionKhayms.pdf, May 2009. 86, 87, 88, 90,
91, 92, 93, 96, 97, 99, 100, 103
[46] Vladimir Kim, Garri Popov, B Arkhipov, V Murashko, 0 Gorshkov, A Koroteyev, Valery Garkusha, Alexander Semenkin, and Sergei Tverdokhlebov.
Electric propulsion activity in russia. IEPC, 5:15-19, 2001. 94, 96
[47] D. Kirtley and J.M. Fife. A Colloid Engine Accelerator Concept Update. Technical Report AFRL-PR-ED-TP-2003-013, Air Force Research Lab, 5 Pollux
Drive, Edwards AFB, CA 93524, January 2003. 90
146
[48] D. Kirtley and J.M. Fife. A Colloid Engine Accelerator Concept Update. Tech-
nical Report AFRL-PR-ED-TP-2003-013, Air Force Research Lab, 5 Pollux
Drive, Edwards AFB, CA 93524, January 2003. 103
[49] Norbert Koch, Hans-Peter Harmann, and Gnter Kornfeld.
First Test Re-
sults of the 1 to 15 kW Coaxial HEMP 30250 Thruster. In 41st AIAA/ASME/SA E/ASEE Joint Propulsion Conference & Exhibit. 97
[50] Gunter Kornfeld, Norbert Koch, and Hans-Peter Harmann. New performance
and reliability results of the thales hemp thruster, pages 267-277. European
Space Agency, Noordwijk, 2200 AG, Netherlands, 2004. 97
[51] Inc. L-3 Electron Technologies. Electric Propulsion. http: //www2.1-3com.
com/eti/product-lineselectricpropulsion.htm, May 2012. 92, 93
[52] Timothy J Lawrence. Research into resistojet rockets for small satellite appli-
cations. October, 1998. 83, 86, 89, 90, 100, 102
[53] Timothy J Lawrence. The european office of aerospace research and developments small satellite propulsion system research program.
System,
(October):16-18, 2000. 89
[54] Dr. Rachel Leach and Kerry L. Neal. Discussion of Micro-Newton Thruster
Requirements for a Drag-FreeControl System. In 16th Annual/USU Conference
on Small Satellites. 102
[55] Robert S Legge and Paulo C Lozano. Electrospray propulsion based on emitters
microfabricated in porous metals. Journalof Propulsion and Power, 27(2):485-
495, 2011. 61
[56] A P London, A H Epstein, and J L Kerrebrock. High-pressure bipropellant
microrocket engine. Journal of Propulsion and Power, 17(4):780-787, 2001. 36
CubeSat
pPulse
Plasma
[57] Clyde
Space
Ltd.
http://www.clyde-space.com/cubesat-shop/propulsion/303_
cubesat-pulse-plasma-thruster, August 2011. 70
Thruster.
[58] Rafael Armament Development Authority Ltd. Lt-1n sp hydrazine monopropellant thruster. http: //www . raf ael. co. il/marketing/SIPSTORAGE/FILES/2/
742.pdf, May 2012. 87
[59] Surrey
lited.
Water Resistojet for Small SatelSatellite Technology Ltd.
http://microsat.sm.bmstu.ru/e-library/SSTL/Subsys-h20rjet_
HQ .pdf, May 2012. 89
[60] Jonathan Lun. Development of a vacuum arc thruster for nanosatellite propulsion. Master's thesis, Stellenbosch University, 2008. 101
147
[61] M. Manente, M. L. R. Walker, J. Carlsson, C. Bramanti, S. Rocca, D. Curreli,
Y. G1, and D. Pavarin. Feasibility study of low power helicon thruster. In 5th
Internationalspacecraft Propulsion Conference, 2008. 83, 84, 86, 87, 92, 93, 94,
95, 96, 97, 98, 99, 100, 101, 102, 103
[62] Ltd. Marotta U.K. Cold Gas Micro-Thruster. https://wiki.umn.edu/pub/
AEMAir_LaunchTeam/ComponentList/ColdGasMicro-Thruster .pdf,
May 2012. 83
[63] Ltd. Marotta U.K. Cold Gas Thruster Valve Model No. SV14. http: //www.
ampacispcheltenham.eu/docs/SV14_marotta.pdf, May 2012. 83
[64] Colleen Marrese-reading, Jay Polk, Juergen Mueller, A Owens, Martin Tajmar,
and Richard Fink. In-feep thruster ion beam neutralization with thermionic
and field emission cathodes. Jet Propulsion,pages 15-19, 2001. 102
[65] M Martinez-Sanchez and J E Pollard.
Spacecraft electric propulsion
an
overview. Journal of Propulsion and Power, 14(5):688-699, 1998. 90
http:
Lozano.
Paulo
and
Martinez-Sanchez
[66] Manuel
//ocw.mit.edu/courses/aeronautics-and-astronautics/
16-522-space-propulsion-spring-2004/lecture-notes/lecturela.pdf,
2002-2012. 19, 43
[67] S. Mazouffrel and A. Lejeune. Hight power electric propulsion for robotic
exploration of our solar system. In Space Access International Conference,
2011. 93, 94, 96
[68] P Molina-Cabrera.
Pulsed Plasma Thrusters : a worldwide review and long
yearned classification. 2011. 98, 99, 100, 101
[69] Gilbert Moore, Walter Holemans, Adam Huang, John Lee, Matthew McMullen,
Jim White, Robert Twiggs, Benjamin Malphrus, Nathan Fite, David Klumpar,
Ehson Mosleh, Keith Mashburn, David Wilt, James Lyke, Stewart Davis, Wes
Bradley, Thomas Chiasson, Jay Heberle, and Pat Patterson. 3D Printing
and MEMS Propulsion for the RAMPART 2U CUBESAT.
AIAA/USU Conference on Small Satellites. 43
In 24th Annual
[70] Juergen Mueller. Thruster options for microspacecraft: A review and evaluation
of existing hardware and emerging technologies. Integration The VIsi Journal,
97:1-29, 1997. 16, 27, 35, 83, 84, 86, 87, 88, 92, 93, 102
[71] Juergen Mueller, Richard Hofer, and John Ziemer. Survey of propulsion tech-
nologies applicable to cubesats. May 2010. 16, 30, 36, 83, 84, 85, 86, 88, 89,
92, 94, 95, 96, 98, 99, 101, 102, 103
[72] Charles K. Murch. Monopropellant Hydrazine Resistojet: Data CorrelationTask
Summary Report. Technical Report 20266-6025-RO-00, TRW Systems Group,
1 Space Park, Redondo Beach, CA 90278, March 1973. 90
148
[73] Kazutaka Nishiyama, Hitoshi Kuninaka, Yukio Shimizu, and Kyoichiro Toki.
30mn-class microwave discharge ion thruster. Methods, pages 1-10, 2003. 92
[74] Georgia Institute of Technology. High-power electric propulsion laboratory:
Thrusters. http://mwalker.gatech. edu/hpepl/thrusters/t-40-het/, May
2012. 97
[75] David Y Oh and Dan M Goebel. Performance evaluation of an expanded
range xips ion thruster system for nasa science missions. Systems Engineering, (July):1-12, 2006. 92
[76] Using Optical and Fiber Probe. Number Density Measurement of Xe I in the
ECR Ion Thruster 10 Using Optical Fiber Probe, pages 1-9. 2011. 92
[77] M. A. Othman and A. E. Makled. Evaluation of resisto-jet thruster engineering
model for space application. In 13th International Conference on Aerospace
Sciences & Aviation Technology, 2004. 89
[78] L. Paita, F. Ceccanti, M. Spurio, U. Cesari, L. Priami, F. Nania, A. Rossodivita,
and Mariano Andrenucci. Altas FT-150 FEEP Microthruster:Development and
Qualification Status. In The 31st InternationalElectric Propulsion Conference,
September 2009. 102
[79] Michael J. Patterson, Thomas W. Haag, and Scot A. Hovan. Performanceof the
NASA 30 cm Ion Thruster. Technical Report IEPC-93-108, NASA, September
1993. 92
[80] AMPAC In-Space Propulsion.
Cold Gas Thruster Valve. Model No. SV14.
http://www.ampacispcheltenham.eu/pages/prsolenoid2.htm,
May 2012.
83
[81] AMPAC In-Space Propulsion. Liquid Rocket Engines, Tanks & PropulsionSystems. http: //www.ampacisp. com/products.php, 2012. 86, 88
[82] F. Rdenauer. Space Propulsion.http: //www. iap.tuwien. ac. at/www/_media/
ivainf o/134176/scprop_06_v2.pdf, April 2010. 83, 84, 90, 91, 92, 93, 94, 96,
97, 99, 100
[83] Jeffrey G Reichbach, Raymond J Sedwick, and Manuel Martinez-sanchez. Micropropulsion system selection for precision formation flying satellites. Design,
(January):1-211, 2001. 98, 99, 100, 101
[84] Small Business Innovation Research.
Magnetically Enhanced Vacuum Arc
Thruster. http://www.sbir.gov/sbirsearch/detail/67396, May 2012. 101
[85] K. J. A. Rycek and B. T. C. Zandbergen. Initial design of a In nulti-propellant
resistojet. In European Conference for Aerospace Science, 2005. 90
149
[86] Leonardo Priami Mariano Andrenucci Salvo Marcuccio, Marco Saviozzi. Experimental Performance of 1 mN-class FEEP Thrusters. In 28th International
Electric Propulsion Conference, March 2003. 102
[87] K Sankaran, L Cassady, A D Kodys, and E Y Choueiri. A survey of propulsion
options for cargo and piloted missions to mars. Annals Of The New York
Academy Of Sciences, 1017:450-467, 2004. 91, 94, 96
[88] Carsten A Scharlemann. Theoretical and experimental investigation of c60propellant for ion propulsion. Acta Astronautica,51(12):865 - 872, 2002. 49
[89] Jochen Schein, Andrew Gerhan, Filip Rysanek, and Mahadevan Krishnan. Vacuum arc thruster for cubesat propulsion. Proceedings of the 28th International
Electric Propulsion Conference, 100:1-8, 2003. 101
[90] H. D. Schmitz and M. Steenborg. Augmented electrothermal hydrazine thruster
development. Journal of Spacecraft and Rockets, 20(2):178-181, 1983. 31, 38
[91] Derek T Schmuland, Robert K Masse, and Charles G Sota. SSC11-X-4 Hydrazine PropulsionModule for CubeSats. 2011. 35, 68, 74
[92] Anita Sengupta, Colleen Marrese-Reading, M Capelli, David Scharfe, S Tverdokhlebov, Sasha Semenkin, Oleg Tverdokhlebov, Iain D Boyd, Michael Keidar,
Azer Yalin, and Et Al. An overview of the vhital program: a two-stage bismuth
fed very high specific impulse thruster with anode layer. 29th International
Electric Propulsion conference Princeton New Jersey October 21Nov 4 2005,
IEPC-2005-:1-14, 2005. 52
http://www.snecma.
Pps 1350-g stationary plasma thruster.
[93] Snecma.
com/IMG/files/fiche-ppsl350g-ang_2011_modulvoir_f ile_fr.pdf, March
2011. 96
[94] Micro Aerospace Solutions. Thrusters. http: //www.micro-a. net/thrusters.
php, May 2012. 74, 86, 88
[95] George C Soulas, Matthew T Domonkos, and Michael J Patterson. Performance
evaluation of the next ion engine. Materials Engineering, (July):1-13, 2003. 93
[96] James S. Sovey, Francis M. Curran, Thomas W. Hug, Michael J. Patterson,
Eric J. Pencil, Vincent K. Rawlin, and John M. Sankovic. Development of
arcjet and ion propulsion for spacecraft stationkeeping. Acta Astronautica,
30:151-164, 1993. 91, 92
[97] James S Sovey, Vincent K Rawlin, and Michael J Patterson. A synopsis of ion
propulsion development projects in the united states: Sert 1 to deep space i.
Test, (c), 1999. 92, 93
150
[98] James S Sovey, Vincent K Rawlin, Michael J Patterson, Nasa John, and Lewis
Field. Ion propulsion development projects in u . s .: Space electric rocket test
i to deep space 1. Journal of Propulsion and Power, 17(3):517-526, 2001. 93
[99] James S. Sovey, Lynnette M. Zana, and Steven C. Ktlowles. Electromagnetic
Emission Experiences Using Electric Propulsion- A Survey. Technical Report
AIAA-87-2028, NASA, July 1997. 93
[100] Alta SpA. Hall Effect Thrusters. http: //www. alta-space. com/index. php?
page=hall-effect-thrusters, May 2012. 97
[1011 Thales Alenia Space. Electric Propulsion Devices. http://www.google.
com/url?sa=t&rct=j&q=&esrc=s&source=web&cd=1&ved=OCEOQFjAA&url=
http%3AX2F%2Fwww.thalesgroup.com%2FPortfolio%2FDocuments%
2Felectric-propulsion-devices-pdf%2F%3FLangType%3D2057&ei=D8a3T_
-UA8eBgwet6dXXCg&usg=AFQjCNHMONp8SFivOZGJ7Iqc7adicYeBAw&sig2=
1M2Fu8EjU31UNndAz140YQ, December 2006. 93, 94
[102] Rr Stephenson. Electric propulsion development and application in the united
states. InternationalElectric Propulsion Conference, 1995. 91, 99, 100
[103] James Stratton. The use of the aerojet mr-103j thruster on the new horizons
mission to pluto. In 55th InternationalAstronautical Congress. 87
[104] A A Sukhanov and N A Eismont. Low thrust transfer to sun-earth 11 and 12
points with a constraint on the thrust direction. Libration Point Orbits and
Applications Proceedings of the Conference, pages 439-454, 2003. 94
[105] M Tajmar, A Genovese, and W Steiger. Indium field emission electric propulsion
microthruster experimental characterization. Journal of Propulsion and Power,
20(2):211-218, 2004. 102
[106] M Tajmar and C A Scharlemann. Development of electric and chemical microthrusters. International Journal of Aerospace Engineering, 2011:1-10. 88,
98, 102
[107] Benjamin Tang, Luke Idzkowski, Michael Au, Don Parks, Mahadevan Krishnan,
and John Ziemer. Thrust improvement of the magnetically enhanced vacuum
arc thruster ( mvat ). October, pages 1-15, 2005. 101
[108] Ion Thrusters. Flight ion and hall thrusters. Fundamentals of Electric Propulsion Ion and Hall Thrusters, pages 429-446, 1994. 92, 93, 96, 97
[109] Paolo Tortora,
Fabio Antonini,
and Roberto Cocomazzi.
Cold Gas
Micropropulsion Satellite.
http://www.almasat.unibo.it/02_projects/
micropropulsion/CGMPS.htm, 2006-2010. 83
151
[110] Purdue University. Satellite Propulsion. https://engineering.purdue.edu/
-propulsi/propulsion/rockets/satellites.html, 1998. 83, 86, 87, 88, 90,
91
[111] Jan van Casteren and Mauro Novara. Bepicolombo mission to mercury. http :/
June
www.brera. inaf .it/schiaparelli20l0/documents/MNovara_1.pdf,
2010. 69
[112] Noah Zachary Warner and Manuel Martnez-snchez. Theoretical and experimental investigation of hall thruster miniaturization by. Measurement, (May),
2007. 94, 95, 97
[113] Bong Wie, David Murphy, Michael Paluszek, and Stephanie Thomas. Robust
attitude control systems design for solar sails , part 1 : Propellantless primary
acs. Space Power, (August):1-28, 2004. 98, 100
[114] Charles Zakrzwski and Scott Benson. Pulsed plasma thruster (ppt). http:/
www. eoc. csiro . au/hswww/ozpi/techf orum/Present ations/25-PPT. pdf,
May 2012. 98, 100
[115] Cm Zakrzwski and with Rhee Thomas, Ma. Highlights of Nanosatellite Propulsion Development Program at NASA-Goddard Space Flight Center. 2000. 15,
25, 83
Modern liquid propellant rocket engines.
C. Zandbergen.
[116] B. T.
Delft University of TechnolInternal Publication,
2000 Outlook,
ogy,
Faculty of Aerospace
Engineering,
Delft,
The Netherlands,
http://www.lr.tudelft.nl/en/organisation/departments-and-chairs/
space-engineering/space-systems-engineering/expertise-areas/
space-propulsion/propulsion-options/chemical-rockets/liquid/,
March 2000. 31, 86, 87, 88
Modern liquid propellant rocket engines.
[117] B. T. C. Zandbergen.
Delft University of TechnolPublication,
Internal
2000 Outlook,
ogy,
Faculty of Aerospace Engineering,
Delft,
The Netherlands,
http://www.lr.tudelft.nl/en/organisation/departments-and-chairs/
space-engineering/space-systems-engineering/expertise-areas/
space-propulsion/propulsion-options/chemical-rockets/cold-gas/,
March 2000. 83, 84
[118] B. T. C. Zandbergen. Performance and operatingdata for typical rocket engines.
http: //www. lr.tudelf t.nl/en/organisat ion/departments-and-chairs/
space-engineering/space-systems-engineering/expertise-areas/
space-propulsion/design-data/performance-and-operating-data/, May
2012. 89, 90, 91, 92, 93
[119] B. T. C. Zandbergen. Resistojet configurations. http://www.lr.tudelft.
nl/en/organisation/departments-and-chairs/space-engineering/
152
space-systems-engineering/expertise-areas/space-propulsion/
propulsion-options/thermal-rockets/resistojets/configurations/,
2012. 90
[120] Ralf Zimmermann. First application of an arcjet thrusterfor the orbit control
of a non-geostationary satellite. http: //www. jamsat . or . jp/p3d/ATOS.html,
November 1996. 90
[121] S Zurbach, N Cornu, and P Lasgorceix. Performance evaluation of a 20kw hall
effect thruster. pages 1-10, 2011. 96
153