System Analysis and Test Bed for an Air-Breathing Electric Propulsion System

Abstract

Challenging types of mission scenarios include those in Earth orbit (i.e., LEO, GEO), where the residual atmosphere, especially at low altitudes, creates significant drag to the S/Cs and forces their orbit to decay. For drag compensation propulsion systems can be used requiring on-board propellant and electric power. Enhancing lifetime of Earth-orbiting satellites without any substantial increase in costs is an important objective for governmental as well as commercial operators. An air-breathing electric propulsion system (RAM-EP) ingests the air of the residual atmosphere through a mechanical intake and uses it as propellant for an electric thruster. This system theoretically allows a S/C to orbit for an unlimited time without carrying propellant on board. Moreover a new range of altitudes (120-250 km) can be accessed, filling the gap between ramjet atmospheric propulsion and LEO space propulsion, thereby enabling many new scientific missions. Preliminary studies according to [2] have shown that the propellant flow necessary for electrostatic propulsion exceeds the available mass intake with reasonable limits, and that electrode erosion due to aggressive gases, such as oxygen, highly present in LEO, might limit the thrusters lifetime. The electrode-less design of inductive plasma generators-IPG-solves this issue. Characterisation of such plasma generators using pure O 2 and CO 2 gases exists and shows significant electric-to-thermal coupling eciencies [10]. A system analysis is shown within this work to derive main design drivers for a RAM-EP mission application. Atmospheric modelling, orbit considerations, heat fluxes, drag force, air intake, and available mass flow for a wide altitude range have been investigated. Preliminary results have shown that full drag compensation is possible. The small-scale inductive plasma generator IPG6-S of the University of Stuttgart is continually improved and used as test bed for RAM-EP using IPG source. A set of mass flows has been defined, depending on altitude, inlet area, and intake eciency to simulate relevant mission conditions. IPG6-S has been tested for mass flow rates between 120 mg/s down to 0.25 mg/s with air and O2. Mean mass-specific energies of the plasma plume have been assessed and used to estimate exhaust velocities for the system analysis.

Full text

5th Russian German Conference on Electric Propulsion Dresden, Germany, September 7-12, 2014
System Analysis and Test Bed for an Air-Breathing
Electric Propulsion System
F. Romano⇤,B.Massut´ı
†and G. Herdrich‡
Institute of Space Systems (IRS), Universit¨at Stuttgart, Stuttgart, 70569, Germany
T. Sch¨onherr§
Department of Aeronautics and Astronautics, The University of Tokyo, Japan
Abstract
Challenging types of mission scenarios include those in Earth orbit (i.e., LEO, GEO), where
the residual atmosphere, especially at low altitudes, creates significant drag to the S/Cs and forces
their orbit to decay. For drag compensation propulsion systems can be used requiring on-board
propellant and electric power. Enhancing lifetime of Earth-orbiting satellites without any substan-
tial increase in costs is an important objective for governmental as well as commercial operators.
An air-breathing electric propulsion system (RAM-EP) ingests the air of the residual atmosphere
through a mechanical intake and uses it as propellant for an electric thruster. This system theoret-
ically allows a S/C to orbit for an unlimited time without carrying propellant on board. Moreover
a new range of altitudes (120-250 km) can be accessed, filling the gap between ramjet atmospheric
propulsion and LEO space propulsion, thereby enabling many new scientific missions.
Preliminary studies according to [2] have shown that the propellant flow necessary for electrostatic
propulsion exceeds the available mass intake with reasonable limits, and that electrode erosion due
to aggressive gases, such as oxygen, highly present in LEO, might limit the thrusters lifetime.
The electrode-less design of inductive plasma generators - IPG - solves this issue.
Characterisation of such plasma generators using pure O2and CO2gases exists and shows signifi-
cant electric-to-thermal coupling efficiencies [10].
A system analysis is shown within this work to derive main design drivers for a RAM-EP mission
application. Atmospheric modelling, orbit considerations, heat fluxes, drag force, air intake, and
available mass flow for a wide altitude range have been investigated. Preliminary results have
shown that full drag compensation is possible.
The small-scale inductive plasma generator IPG6-S of the University of Stuttgart is continually im-
proved and used as test bed for RAM-EP using IPG source. A set of mass flows has been defined,
depending on altitude, inlet area, and intake efficiency to simulate relevant mission conditions.
IPG6-S has been tested for mass flow rates between 120 mg/s down to 0.25 mg/s with air and O2.
Mean mass-specific energies of the plasma plume have been assessed and used to estimate exhaust
velocities for the system analysis.
Keywords: RAM-EP - Air-Breathing Electric Propulsion - VLEO - Inductively Coupled Plasma
⇤PhD Student, Institute of Space Systems (IRS), romano@irs.uni-stuttgart.de.
†Associate Researcher, Institute of Space Systems (IRS), massuti@irs.uni-stuttgart.de.
‡Head Plasma Wind Tunnels and Electric Propulsion, Institute of Space Systems (IRS), herdrich@irs.uni-stuttgart.de.
§Associate Professor, Department of Aeronautics and Astronautics, schoenherr@al.t.u-tokyo.ac.jp.
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5th Russian German Conference on Electric Propulsion Dresden, Germany, September 7-12, 2014
Nomenclature
LEO : Low-Earth Orbit
VLEO: Very Low-Earth Orbit
S/C : Spacecraft
GIE : Gridded Ion Engine
EOL : End-of-life
BOL : Begin-of-life
HET : Hall E↵ect Thruster
HHT : Helicon Hall Thruster
ICP: Inductively Coupled Plasma
IPG: Inductively Heated Plasma Generator
SA: Solar Array
SSO: Sun-Synchronous Orbit
FMF: Free Molecular Flow
1 Introduction
Missions in LEO are of great value for activities such as surveillance and Earth monitoring. This is of
great importance for weather forecasting, oceanic currents, polar ice caps and fires monitoring, as well
as for military and civil surveillance services. Recently ESA’s mission GOCE has ended. It provided
detailed information of Earth’s geomagnetic field by orbiting as low as 229 km.
However such missions have limited lifetime due to drag, which is caused by momentum transfer of
the residual atmosphere’s particles impacting the S/C and decreasing its kinetic energy.
The lifetime of a S/C orbiting in LEO can, therefore, be significantly increased using a propulsion
system that is capable to compensate the drag. The lifetime of a S/C in LEO is a mission design that
depends highly on the efficiency of the propulsion system and moreover on the propellant carried on
board. The basic idea of an Air-Breathing Electric Propulsion System, shortened RAM-EP, is to use
the air of the residual atmosphere as propellant and process it through a device for generating thrust.
This will decrease, ideally nullify, the on board propellant requirement and will generate thrust to
partially or fully compensate the drag, increasing mission’s lifetime.
In this paper a system analysis for such a propulsion system, which describes how the mission parame-
ters are taken into account, is proposed. The approach of using an ICP generator as thruster candidate
is introduced. ICP generators are electrode-less and this solves the issue of the limitation in lifetime
due to electrodes erosion. IPG6-S from the University of Stuttgart has been used for experimental
activity with O2and air as working gases and the produced thrust has been evaluated, in order to
determine thrust to drag ratio and therefore show the feasibility of such a technology.
1.1 Literature Review
A literature review has been developed as a start point for this research. Many publications dealing
with RAM-EP, as well as cases dealing with low orbiting small S/C, e.g. GOCE and GRACE, using
electric propulsion.
ESA study [2] is a proposal for a technology demonstration mission featuring RAM-EP. It considers
a 1000 kg S/C equipped with 4 ⇥ASTRIUM RIT-10 GIE operating with the incoming air molecules.
The S/C is to be launched and set into a circular SSO at an altitude of h= 200 km for a 7 years
mission. The front area is of 1 m2and the maximum power available for propulsion is of 1 kW enabling
thrust from 2 to 20 mN. The SA surface will be of 19.74 m2to provide a power in EOL of 2.9 kW, this
will be combined with a 612 W h Li-Ion battery.
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5th Russian German Conference on Electric Propulsion Dresden, Germany, September 7-12, 2014
Diamant’s [3] proposes a mission for drag compensation on small S/C at an altitude of 200 km
with a 2-stage cylindrical Hall thruster and propellant ingested from the atmosphere. The first stage
is an electron cyclotron resonance ionization stage and the second stage is a cylindrical Hall thruster.
The required power is of 1 kW for propulsion, the frontal area is of 0.5m
2with a collection efficiency
of ⌘c= 35%.
The study from Ceccanti et al. [4] proposes a mission into a 296 220 km, 96.52°inclination, orbit
with EP system for drag compensation. The weight is of 450 kg for a 8 years mission. The front
area is of 0.8m
2. The power subsystem provides 200 400 W with a peak of 660 W through 2 SA of
2.5m
2together with Li-Ion battery. The propulsion system is of 3 HET of 650 W operating with Xe
generating a thrust of FT= 40 mN each.
The PhD dissertation from Shabshelowitz [5] investigates RF Plasma applied to a RAM-EP system.
The S/C mass is of 325 kg, to be set into a circular orbit at an altitude of 200 km for a mission duration
of 3 years. The frontal area is of 0.39 m2, a length of 2.1 m, and the S/C is to be covered with solar
cells. The ratio of the frontal area through the inlet area is of Af/Ainlet =0.5 and the collection
efficiency is = 90%. The propulsion system is composed by a single-stage HHT operating with air and
supported by a tank of propellant for ballast. The thruster requires a power of 306W.
The study of Pekker and Keidar [6] considered HET using air of the atmosphere as propellant in
the orbit at an altitude of h= 90 km and h= 95 km. The gas leaving the chamber of the HHT is
considered fully ionized and under this condition the achievable thrust is of FT@90 km = 22 N and of
FT@95 km =9.1 N with a thrust density of 13 mN/kW. The power required at the two altitudes is of
Preq@90 km =1.62 MW and Preq@95 km = 700 800 kW.
The ESA’s GOCE mission successfully ended last year. The S/C had a mass of 1090 kg and orbited
into a 250265 km SSO for a predicted mission lifetime of 2030 months, but it reached finally 4 years
of operation. The frontal area was of 1.1m
2[7]. The S/C was provided with two Ion thrusters derived
from the QinetiQ T5 (one for backup) operating with Xe and providing thrust between T=1.5 and
20 mN. The SA for the power subsystems was providing PEOL =1.6 kW in EOL and was completed
by 78 A h battery.
The BUSEK company [8] developed a study for RAM-EP applied to a small S/C orbiting Mars:
Martian Atmosphere-Breathing Hall E↵ect Thruster - MABHET. An HET has been run with a gas
mixture which reproduces Mars atmosphere, the most present component is CO2.Thethrusttopower
peak ratio of HET has been measured around 30 mN/kW with a low peak of 19 mN/kW. MABHET has
an inlet area of 0.15 m2and a frontal area 0.30 m2. The collector efficiency is of ⌘c= 35%. Compression
of the incoming air flow is required to achieve better performance of the thruster. MABHET may
work better in Mars atmosphere than in Earth’s, because of the lower density and temperature of the
atmosphere and of the accommodation coefficients.
The study from JAXA [9] has shown a concept for an Air-Breathing Ion Engine - ABIE, in which
the low density atmosphere surrounding the satellite is used as propellant for the Electron Cyclotron
Resonance (ECR). A S/C has been proposed orbiting in a circular polar SSO of h= 170 km for at
least 2 years. The frontal area is of 1.5m
2and the inlet area is of 0.48 m2. The propulsion system
should deliver a thrust to power ratio between 10 14 mN/kW. Morover altitudes of 185 and 145 km
have been investigated showing a power required for the thruster of 470 W and 3.3kW.
A summary of the literature review is briefly shown in Tab. 1.
The thrusters used in the di↵erent studies are GIT, HET, HHT and they are all a↵ected by the
issue of limited lifetime due to corrosion of the electrodes.
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5th Russian German Conference on Electric Propulsion Dresden, Germany, September 7-12, 2014
Table 1: Summary of Literature Review Results
S/C Mass <1090 kg
Inlet Area 0.31m
2
Orbit SSO, 90 250 km
Lifetime 2 8 years
Thrust Density 10 59 mN/kW
Power Generated 0.660 2.9kW
Collection Efficiency 0.35,0.9
80 100 120 140 160 180 200 220 240 260 280 300 320 340
1010
1015
1020
Altitude vs. Atmosphere Constituents
Mean Solar and Geomagnetic Activities
NRLMSISE−00 Model (F10.7 = F10.7avg = 140, Ap = 15)
Altitude, km
Numerical Density, m−3
N2
O2
Ar
O
He
H
N
Anomalous O
VLEO LEO
Figure 1: Atmosphere’s components vs. altitude.
2 System Analysis
2.1 Atmospheric Model
Considering an Air-Breathing Electric Propulsion System, an estimation of the mass flow that can be
collected by the system as well as an estimation of the level of generated drag, are needed. Therefore,
an appropriate atmospheric model for the system analysis has to be selected.
The chosen model is NRLMSISE-00, compared to the common MSISE-90 and JR-71, it provides better
estimation of the air density below 350 km of altitude and it is the most accurate model for compo-
sition of residual atmosphere in LEO and VLEO. NRL stays for Naval Research Laboratory, MSIS
stands for Mass Spectrometer and Incoherent Scatter Radar and E indicates that the model extends
from the ground to space [11]. It is an empirical global model for describing the Earth’s atmosphere
under di↵erent conditions of solar and geomagnetic activities.
The model has been generated through the NRLMSISE-00 model website [12] and loaded into MAT-
LAB and has shown that the most dominant elements in VLEO and LEO are O2and N2, with the first
more dominant in higher altitudes, as in Fig. 1. Particular care must be taken concerning the solar
activity which cycles every 11 years. This will result in change of the density vs. altitude profile as it
compress and release the atmosphere by the time. This must be taken into account when designing
the mission as it a↵ects both the mass flow and the drag. In detail, this variation with the solar
activity is more evident, in LEO and VLEO ranges, in higher than in lower altitudes as it is shown in
Fig. 2.
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5th Russian German Conference on Electric Propulsion Dresden, Germany, September 7-12, 2014
100 120 140 160 180 200 220 240 250
10−11
10−10
10−9
10−8
10−7
10−6
Altitude, km
Density, kg/m3
Density vs Altitude
NRLMSISE−00 Model, Solar Activity
Solar Average
F10.7=F1.07avg=140, Ap=15
Solar Maximum
F10.7=F1.07avg=250, Ap=45
Solar Minimum
F10.7=F1.07avg=65, Ap=0
Figure 2: Density vs. atmosphere and solar activity.
2.2 Orbit
LEO extends in the range from 160 to 2000 km, VLEO from 100 to 160 km.
According to ESA [2] the maximum altitude for an Air-Breathing Electric Propulsion mission is to
be set at 250 km to be competitive against conventional electric propulsion. The minimum altitude
has been set according to JPL, [13], at 120 km, due to heating e↵ects, however this last statement
should be investigated by further thermal analysis on a 3D S/C model. Concerning the orbit’s plane,
considering to continuously generate power with solar arrays - SA -, a sun-synchronous orbit - SSO -
should be chosen. In this way the sun vector will be always perpendicular to the orbit plane, therefore
directing SA in the orbit plane will make them operate at maximum power condition for the most of
time. However this depends on the mission requirements, if a particular orbit is required, depending
on the propulsion system requirements in terms of electrical power, a thrust profile related to the
eclipse and sunshine periods has to be investigated, as the power subsystem might not be able to
deliver the required power of the propulsion system all the time.
2.3 Intake
Figure 3: JAXA Air Intake. [15]
The propulsion system needs a device which collects and delivers the
air particles to the thruster. According to [14] a mechanical device
should be used as the ionization degree in LEO and VLEO is too low
to use a magnetic one. JAXA [15] developed an intake, see Fig. 3,
which proofed a collection efficiency, defined as the ratio between
collected and incoming particles ⌘c=˙mc/˙min of 35%. According to
Sch¨onherr [14] a collection efficiency up to 40% and a compression fac-
tor between 100 200 are achievable, a pressure of 1 mPa is achieved
at the thruster head but this is not enough for most electric thrusters
[14]. In the Fig. 4 the mass flow vs. altitude is plotted for average
solar activity considering an intake area of 1 m2and three collection
efficiencies of ⌘c=1.0; 0.9; 0.35.
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5th Russian German Conference on Electric Propulsion Dresden, Germany, September 7-12, 2014
120 140 160 180 200 220 240 250
10−1
100
101
102
Mass Flow vs Altitude
Ainlet=1 m2
NRLMSISE−00 Model (F10.7 = F10.7avg = 140, Ap = 15)
Altitude, km
Mass Flow, mg/s
c = 1
c = 0.90
c = 0.35
Figure 4: Mass flow vs. altitude.
2.4 Drag
An estimation of the drag is needed to design the mission as the S/C orbits in LEO and VLEO where
the presence of residual atmosphere is not negligible. First step is to determine if the flow is to be
considered continuum or free molecular flow - FMF, as the mean free path length of the molecules
might become comparable to the size of the S/C. The Knudsen number has been therefore calculated
and showed that for altitudes above 120 km and mean length of 0.3, 1, 2, and 3 m the flow is FMF. A
sensitivity analysis on the average molecules size and on the solar activity has been done and shown
very small variations. In particular the e↵ect of solar activity is more appreciable on higher than on
lower altitudes.
A numerical model for the calculation of the drag in FMF, see [16], has been implemented and it is
as following:
~
FD=Aˆnp +✓ˆnsin ⇣ˆvrel◆✓ ⌧
cos ⇣◆ (1)
p
q1
=⇢2n
p⇡sin ⇣+n
2srTs
Ta⇢1
ses2sin ⇣2+p⇡[1 + erf (ssin ⇣)] sin ⇣+
+2n
2s2[1 + erf (ssin ⇣)]
(2)
⌧
q1
=t⇢1
sp⇡es2sin ⇣2+erf(ssin ⇣)] sin ⇣(3)
Here ~
FDis the drag force, Ais the area encountered by the flow, pis the total pressure, ⌧the
shearing stress and q1the dynamic pressure given by q1=1
2⇢1v2
rel. The quantity ˆnis the outward-
pointing unit normal vector, nand tare the normal and tangential accommodation coefficients, Ts
is the absolute temperature of the surface, set to Ts= 490 K,as the average value in VLEO orbit
calculated, and Tais the atmospheric temperature, altitude dependent, taken from the model. An
important component of the above equations is sthe air speed, nondimensionalized by the mean
molecular speed of the atmosphere as shown in Eq. 4:
s=sMav2
rel
2RspecTa
(4)
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5th Russian German Conference on Electric Propulsion Dresden, Germany, September 7-12, 2014
80 100 120 140 160 180 200 220 240 250
10−2
100
102
104
106
Drag vs Altitude
Free Molecular Model vs Continuum
S/C: Af=1.00 m2, i=0.00 deg
NRLMSISE−00 Model − F10.7 = F10.7avg = 140, Ap = 15
Altitude, km
Drag, mN
Free Molecular Flow
Continuum
Free Molecular Flow
Continuum
Transition
Figure 5: Drag vs. altitude.
Mais the mean molar mass of the atmosphere calculated from the atmospheric model, and Rspec is
the universal gas constant.
All the parameters are altitude-dependent except for the universal gas constant. In the calculation
t=n=0.9 are typical values, according to [16], and ⇣=ˆnTˆvrel is the pitch angle set to 0°,ˆnand
ˆvrel are the unit vectors of the normal vector and the relative velocity vector.
This means that also temperature and kind of material of the S/C will influence the drag as they will
result in di↵erent accommodation coefficients that define the angle in which particles are deflected
and with which amount of energy. The result of this calculation is shown in Fig. 5. In particular the
transition region is defined as the region where the Knudsen number variates from 0.1 (continuum
flow) and 1 (FMF), this is in the altitude range between 100 and 110 km for a mean length of 1 m.
2.5 Power Supply
The power supply system must provide electrical power for the propulsion system and for all the other
subsystems. A common approach for S/Cs orbiting Earth or Mars is the use of solar arrays - SA -
together with batteries. Batteries compensate the fluctuation of required power and provide electricity
when the SA are not illuminated by the Sun. When the SA are illuminated by the Sun they provide
power to all the subsystems and recharge the batteries.
One physical quantity which becomes important when reaching the lower altitudes in VLEO is heat.
Heat can be converted directly into electricity by the use of a thermionic generator operating on the
principle of the Seebeck e↵ect, which describes the phenomena of voltage generation in a conductor
or semiconductor when subjected to a temperature gradient. Thermionic generators are not yet a
mature technology for an application in space under these conditions, but a recent study, see [17],
calculated a maximum efficiency of ⌘= 42% in optimum condition.
The use of this kind of generator might allow to reduce the requirement of the surface of the S/A, in a
way to decrease the surface generating drag and the mechanical complexity, as a thermionic generator
has no moving parts. It has to be kept in mind that only ⌘= 42% of the heat is converted in electrical
energy and the other 56% of heat must be taken away through the thermal subsystem. SA with a
minimum average BOL efficiency of ⌘= 29.5% have been considered, [18]. The Sun vector has been
considered always perpendicular to the SA surface, as in an SSO. The calculated areas are in the
following Tables for both power and voltage considering the panel degradation over a 7 years long
mission, as in ESA study [2].
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5th Russian German Conference on Electric Propulsion Dresden, Germany, September 7-12, 2014
Table 2: Power vs. SA Area - EOL
Pmax ASA
kW m2
0.5 1.98
1 3.96
1.5 5.93
3 11.87
3.5 13.84
5 19.6
Table 3: Voltage vs. Solar String Area - EOL
Voltage Number of Cells Astring
V-m
2
550 319 0.85
850 493 1.30
1000 579 1.54
2.6 Thrust
A thrust profile must be investigated and chosen for the mission. In this study continuous thrust
compensation has been considered, the start point enthalpy value of a previous characterization of the
Inductively Plasma Generator - IPG6 - of the University of Stuttgart [10], has been taken to evaluate
the feasibility of this study.
A maximum specific plasma enthalpy of hcal =7.5 MJ/kg at a mass flow of ˙m= 60 mg/s operating with
air has been determined experimentally. With the assumption that all the plasma energy, measured
by a calorimeter, is converted into kinetic energy, the exhaust velocity has been estimated as following:
ce=p2htot = 3872.98 m/s (5)
This is an estimation which neglects frozen losses and it must be taken as an upper limit for the
exhaust velocity, with this assumption the thrust is estimated as following:
T=˙m(h)ce=⇢(h)vrel(h)Af⌘cce= 232.38 mN (6)
Considering constant this value over the altitude and comparing it to the value of drag - altitude
dependent - lead to the possibility of achieving full thrust compensation whit the previous assumptions.
IPG6 is not yet optimized as a thruster.
3 Experimental Set-Up
An inductively coupled plasma source - ICP has been selected as a candidate for a RAM-EP thruster.
The main advantage of using IPC sources is their electrode-less operation. No electrodes means no
issues concerning lifetime due to their corrosion. Presence of O and O2is high in LEO and VLEO,
which are the main responsible for electrode corrosion, which is one of the first issues which limits
S/C’s lifetime.
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5th Russian German Conference on Electric Propulsion Dresden, Germany, September 7-12, 2014
3.1 IPG Principle of Operation
In an ICP a coil is wrapped around a quartz tube - the discharge channel - and it is fed by an HF AC
current. It operates in a way similar to a transformer where the primary winding is the coil and the
secondary is the gas inside the discharge channel. The current flowing in the coil induces an oscillating
magnetic field in the discharge channel which accelerates ions and electrons of the gas, plasma is created
and a chain reaction established that increases temperature and electrical conductivity of the plasma
itself.
3.2 IPG6-S
Figure 6: IPG6-S.
IPG6-S, Inductively heated Plasma Generator available at the Uni-
versity of Stuttgart, see Fig. 6, has been used for the tests, it has
been selected because of its size and power levels scalable for an ap-
plication on small S/Cs [19]. IPG6-S facility main parameters are a
maximum input power Pmax = 20 kW, a maximum voltage of 1.7 kV,
and a variable frequency between f=3.54.5 MHz, depending on
the impedance of the IPG, in case of IPG6 in the current configura-
tion is of f⇠4 MHz. It is water cooled, the discharge channel has a
diameter of 40 mm, a length of 80 mm and the coil has 5.5 turns with
an inductance of 0.489 µH. A twin facility, IPG6-B, is installed at the
University of Baylor, Waco, Texas, USA [20].
3.3 Tests
The input required for the test are in terms of kind of gas, mass flow and voltage. As result from
the system analysis N2and O are the elements more present in LEO and VLEO. Air has been used
to simulate N2, as it is composed by ⇠78% of N2, to simulate atomic O, O2has been introduced in
the generator as it is difficult to provide atomic O as it recombines very fast and generates O2.The
mass flow, as results from the system analysis and in relation to the facility capability, has been set
between 0.245 and 120 mg/s for both gases.
Three voltages have been selected for the tests: 0.55, 0.85 and 1.00 kV.
Figure 7: IPG6-S Calorime-
ter [19].
3.4 Thrust Evaluation
In order to evaluate the thrust produced by the IPG6, the following
procedure has been performed. The calorimeter is a device, shown in
Fig. 7,used to evaluate the plasma energy by measuring the temper-
ature di↵erence of the cooling water between the inlet and outlet of
the calorimeter. The water is heated up by the plasma plume of the
generator.
The equation is shown in Eq. 7, where hcal is the enthalpy measured
by the calorimeter, ˙mgas is the gas mass flow, Pcal is the calorimeter
power, ˙mwater is the water flow in the calorimeter, Cpwater is the spe-
cific heat capacity, and Tis the water temperature at the outlet and
inlet of the calorimeter.
hcal =Pcal
˙mgas
=1
˙mgas
[˙mwatercal Cpwater (Tout,cal Tinlet,cal)] (7)
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5th Russian German Conference on Electric Propulsion Dresden, Germany, September 7-12, 2014
Considering that all the plasma energy is converted into kinetic energy, the exhaust velocity, ce,is
estimated through Eq. 8.
ce=p2htot =p2hcal (8)
Hence thrust is given by Eq. 9.
T=˙mce=⇢(h)vrel(h)Af⌘cce(h) (9)
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5th Russian German Conference on Electric Propulsion Dresden, Germany, September 7-12, 2014
4 Results
4.1 Thrust
The evaluated thrust is plotted as a function of the altitude for Air and O2,Af=1m
2, the three
di↵erent set voltages, and for ⌘c= 1 and 0.35. Thrust reaches a maximum of 250 mN at low altitudes
with O2, slightly less with Air and a minimum of 5 mN at high altitudes for both gases.
100 120 140 160 180 200 220 240 250
10
25
50
100
200
250
T vs Altitude
IPG6−S, Af=1m2, c=1, Air
Altitude, km
T, mN
V=0.55kV
V=0.85kV
V=1kV
(a)
100 120 140 160 180 200 220 240 250
10
25
50
100
200
250
T vs Altitude
IPG6−S, Af=1m2, c=0.35, Air
Altitude, km
T, mN
V=0.55kV
V=0.85kV
V=1kV
(b)
Figure 8: Thrust, Ainlet =1m
2, Air.
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5th Russian German Conference on Electric Propulsion Dresden, Germany, September 7-12, 2014
100 120 140 160 180 200 220 240 250
10
25
50
100
200
250
T vs Altitude
IPG6−S, Af=1m2, c=1, Oxygen
Altitude, km
T, mN
V=0.55kV
V=0.85kV
V=1kV
(a)
100 120 140 160 180 200 220 240 250
10
25
50
100
200
250
T vs Altitude
IPG6−S, Af=1m2, c=0.35, Oxygen
Altitude, km
T, mN
V=0.55kV
V=0.85kV
V=1kV
(b)
Figure 9: Thrust, Ainlet =1m
2, Oxygen.
4.2 Thrust to Drag Ratio for Air and O2
In this section the evaluated thrust to drag ratio is plotted as a function of the altitude for Air and
O2, for the three di↵erent voltages, ⌘cand for an Af=1m
2. The thrust value is divided by the drag
value at the corresponding extracted altitude. In particular it is shown that under these conditions
the use of IPG6 as plasma generator for an Air-Breathing Electric Propulsion, full drag compensation
is always possible for the di↵erent collection efficiencies, on the whole selected altitude range and with
all the di↵erent voltages.
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5th Russian German Conference on Electric Propulsion Dresden, Germany, September 7-12, 2014
120 140 160 180 200 220 240 250
0
20
40
60
80
90
T/D ratio vs Altitude − Free Molecular Model
S/C: Af=1.00 m2, i=0.0 deg
IPG6−S, V=0.55kV, Air
Altitude, km
T/D, −
c = 1
c = 0.9
c = 0.35
(a) 0.550 kV.
120 140 160 180 200 220 240 250
0
20
40
60
80
90
T/D ratio vs Altitude − Free Molecular Model
S/C: Af=1.00 m2, i=0.0 deg
IPG6−S, V=0.85kV, Air
Altitude, km
T/D, −
c = 1
c = 0.9
c = 0.35
(b) 0.850 kV.
120 140 160 180 200 220 240 250
0
20
40
60
80
90
T/D ratio vs Altitude − Free Molecular Model
S/C: Af=1.00 m2, i=0.0 deg
IPG6−S, V=1kV, Air
Altitude, km
T/D, −
c = 1
c = 0.9
c = 0.35
(c) 1.000 kV.
Figure 10: Thrust to Drag Ratio Ainlet =1m
2, Air .
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5th Russian German Conference on Electric Propulsion Dresden, Germany, September 7-12, 2014
120 140 160 180 200 220 240 250
0
20
40
60
80
90
T/D ratio vs Altitude − Free Molecular Model
S/C: Af=1.00 m2, i=0.0 deg
IPG6−S, V=0.55kV, Oxygen
Altitude, km
T/D, −
c = 1
c = 0.9
c = 0.35
(a) 0.550 kV.
120 140 160 180 200 220 240 250
0
20
40
60
80
90
T/D ratio vs Altitude −Free Molecular Model
S/C: Af=1.00 m2, i=0.0 deg
IPG6−S, V=0.85kV, Oxygen
Altitude, km
T/D, −
c = 1
c = 0.9
c = 0.35
(b) 0.850 kV.
120 140 160 180 200 220 240 250
0
20
40
60
80
90
T/D ratio vs Altitude − Free Molecular Model
S/C: Af=1.00 m2, i=0.0 deg
IPG6−S, V=1kV, Oxygen
Altitude, km
T/D, −
c = 1
c = 0.9
c = 0.35
(c) 1.000 kV.
Figure 11: Thrust to Drag Ratio Ainlet =1m
2, Oxygen .
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5th Russian German Conference on Electric Propulsion Dresden, Germany, September 7-12, 2014
5 Conclusion and Outlook
System analysis investigation set orbit and altitude ranges, collection efficiencies, input gases, drag to
compensate and power available.
IPG6-S has been selected and investigated as a plasma generator for an Air-Breathing Electric Propul-
sion System application.
Facility has been improved and the generator tested with O2and Air, for mass flows representing
di↵erent altitude ranges for di↵erent intake efficiencies.
Enthalpy of the plasma produced by IPG6 has been evaluated through a calorimeter. Subsequently,
the exhaust velocity has been calculated considering a total conversion of the the plasma energy at
the calorimeter into kinetic energy.
From the exhaust velocity, the thrust has been calculated.
Anode power reached a minimum of 0.5kW and a maximum of 3.5 kW which is an acceptable power
level for a small S/C, however the power absorbed by the plasma is expected to be even lower as the
cooling system absorbs most of the power [10].
Di↵erence of pressure between injection and tank is not enough to achieve supersonic discharge, hence
a pump with greater suction capabilities, as well as a bigger vacuum tank, is required to obtain better
simulation conditions.
Three screen voltages have been applied for the experimental investigation, 0.55, 0.85 and 1.00 kV.
Low voltages yield higher enthalpies for low mass flows. Vice-versa high voltages yields to higher
enthalpies for high mass flows.
Air showed better results for high mass flows, when O2showed better results for low mass flows. Low
mass flows means higher altitudes. At high altitudes the predominant component is O, that means
the performance of the thruster will increase by the altitude as the amount of O will increase.
Thrust to drag ratio has been calculated for the three selected voltages and ⌘c, for both O2and Air,
for the inlet area of Af=1m
2in the RAM-EP altitude range.
Results have shown that the thrust to drag ratio is always greater than one and full drag compensation
might be achieved in all the test conditions.
5.1 Outlook
For further work, the use of a multiple stage vacuum pump and bigger tank are required for achieving
better simulation conditions.
A 3D S/C model is required for better estimation of S/Cs temperature and for DSMC for the drag.
The assumption of all plasma energy converted into kinetic energy is a simplifying assumption, there-
fore an analysis of the acceleration strategies is required to better evaluate the exhaust velocity, hence,
the thrust.
Moreover the realisation of a RAM-EP S/C model for the IRS software REENT is required to evaluate
S/Cs lifetime and orbit decay.
The biggest challenge is, however, a downscaling of the IPG6-S to a suitable thruster size.
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[20] Dropmann, M.; Herdrich, G.; Laufer, R.; Puckert, D.; Fulge, H.; Fasoulas, S.; Schmoke, J.;
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