Air-Breathing Electric Propulsion

Air-BreathingElectricPropulsion
TonySchönherr
DepartmentofAeronau=csandAstronau=cs
2015/11/16
PropulsionandEnergySystems
1
Classifica=onofair-breathingpropulsion
Beamed-energypropulsion
Ramjet/Scramjet
In-spacepropulsion
2015/11/16
PropulsionandEnergySystems
2
Classifica=onofair-breathingsystems
•  Ramjet/Scramjet
–  (Supersonic)air-breathingjetengine
–  Fuel-efficientascenttomesospherical=tudesandhalf-orbitalvelocity
•  Air-breathingbeamed-energypropulsion
(seelectures10/26and11/02)
–  Laserormicrowavepropulsion
–  Fuel-efficientascenttoorbitalal=tudes
•  Atmosphere-breathingin-spacepropulsion
–  UseofatmosphericresidualgasesinspacecraQpropulsion
–  Reduc=onofatmosphericdraginloweral=tudes
–  Extensionofin-orbitlife=me
2015/11/16
PropulsionandEnergySystems
3
Orbitdecaywithoutpropulsion
•  Orbitdecayofsatellitesisdrivingfactorforlife=meinLEO
Ballis=ccoefficient
mSC
=
CD · S
βishigh:
highmomentum
lowdrag
βislow:
lowmomentum
highdrag
Foral=tudeslowerthan
400km,orbitdecay
overtakescommercial
benefit
2015/11/16
PropulsionandEnergySystems
4
ExampleforLEOspacecraQ
•  GOCE(ESA)
www.esa.int
–  260kmorbit
–  Designedlife=me:
•  20months
–  Actuallife=me:
•  56months
•  Lowersolarac=vitythanexpected
–  Useoffuel-efficientionthrusters
–  Veryhighballis=ccoefficientto
extendpossiblelife=me
1077 kg
kg
=
= 323 2
3.7 · 0.9 m2
m
–  Reasonforendoflife:
Actual
Designed
•  Fueldeple=on,decayed
2015/11/16
PropulsionandEnergySystems
5
Purposeofair-breathingpropulsion
•  Extensionofsatellitelife=meinLEOfororbital=tudesless
than400km
–  IncreasedataoutcomeofEarthobserva=onsatellites,scien=fic
pla\orms,reconnaissance/espionagesatellites,etc.
–  Increaseprofitmarginforcommercialsatellites
•  Enablescien=ficmissionsinverylowEarthorbits(VLEO)of
lessthan160kminal=tude
Air-breathingspacepropulsion
2015/11/16
PropulsionandEnergySystems
6
Atmosphere
•  Thermosphereandespecially
theVLEOarescien=fically
interes=ngduetostrong
gradientsinphysicalprofile
160km
•  Satellitesinorbitscloserto
Earthcouldenablenewmissions
– 
– 
– 
– 
Earthobserva=on
Scien=ficapplica=ons
Militaryuse
…
LEO
VLEO
2015/11/16
PropulsionandEnergySystems
7
Atmosphericproper=es
•  Necessaryproper=esfordesignofspacecraQandbreathing
propulsionsystem
–  Massdensity(translatestoeventualusablemassflowrateanddrag
force)
–  Temperature,conduc=vity(influenceoncollectorandsatellitedesign)
–  Atomic,ionic,andmolecularspecies;molarcomposi=on(usageas
propellant)
–  …
•  Defactostandardmodelfordescrip=onoftheatmospheric
proper=esisthe1975Interna=onalStandardAtmosphere
(ISA)withamendmentsin1985and1997
2015/11/16
PropulsionandEnergySystems
8
Atmosphericmodels(abriefreview)
1960
1970
MSIS:MassSpectrometerandIncoherentScager
Jacchiamodel
(1965,1971)
- Totaldensityasf(h)
- Temperatureasf(h)
MSIS(1977)
1980
1990
Currentmostaccurate
modelfortotaldensity
2000
MSIS-86
- Atomicnitrogenadded
- Revisionfor120kmandbelow
(downtomesosphere)
- Anomalousoxygen
- Improvedmodeling
Currentmostaccuratemodelforcomposi=on
JB2008
- Evenbegertemperaturefunc=on
2015/11/16
- 80-220km
NRLMSISE-00
- Improveddensitydescrip=on
- Newtemperatureasf(h)
now
MSIS-83
MSISE-90
JB2006
- Speciescomposi=onincluded
(N2,O2,O,He,Ar,andH)
- 120-150km
Bothmodelsadaptedtothe2008ECSS
standardECSS-E-ST-10-04C
PropulsionandEnergySystems
9
Atmosphericproper=es
•  Atmosphericdensityasafunc=onofal=tudeandsolarac=vity
ECSS E ST
15 Novemb
•  Severalparametersaffectdensity
–  Loca=on
–  Timeofyear
–  Solarac=vity
1.00E-06
High (short term)
1.00E-07
Solarac=vity
1.00E-08
High (long term)
Moderate
Mass density (kg/m3)
Low
1.00E-09
1.00E-10
1.00E-11
1.00E-12
1.00E-13
ECSS-E-ST-10-04C
1.00E-14
1.00E-15
0
100
200
300
400
500
600
700
800
900
Altitude (km)
2015/11/16
PropulsionandEnergySystems
10
Low, moderate,
high long term and high short term solar and geomagenetic activities have the following
me
Composi=onofspecies
Number density n, m−3
•  Stronggradientsoftemperatureandvaria=onofsolar
radia=onyielddistribu=onofspeciesoveral=tude
1023
•  Formostorbit
Meandistribu=on
1021
al=tudes
residualair
1019
consistsof
1017
molecular
1015
LEO
VLEO
nitrogenand
1013
atomicoxygen
1011
•  Negligible
109
degreeof
ioniza=on
107
100
2015/11/16
150
200
250
Altitude h, km
He
O
N2
O2
Ar
H
N
O*
300
350
Figure 3: Atmospheric composition
for mean solar and geomagnetic activities (NRLMSISE-00
m
PropulsionandEnergySystems
11
Atomicdensity
ECSS E ST 10 04C
15 November 2008
•  Influenceofsolarac=vityondensi=esofindividualspecies
1.00E+18
High Activity
1.00E+17
Moderate Activity
Low Activity
1.00E+16
O density (m-3)
1.00E+15
1.00E+14
1.00E+13
1.00E+12
1.00E+11
1.00E+10
1.00E+09
ECSS-E-ST-10-04C
1.00E+08
0
100
200
300
400
500
600
700
800
900
Altitude (km)
2015/11/16
PropulsionandEnergySystems
Low, moderate, and high long term solar
and geomagenetic activities have the following meanings:
12
Consequencesforthesatellite
•  Atmosphericdragforceimposedonsatellite
FD
1
v 2 ·S · CD
=
⇢ · |{z}
|{z}
2 |{z}
JB2006 7.8 km/s
2.1 2.7
Orbitalspeed
v⇡
r
GM
r
Dragcoefficient
G …Gravita=onalconstant
M
…Earth’smass
r …Orbitradius
2015/11/16
PropulsionandEnergySystems
13
Dragcoefficient
•  Dragcoefficientisinfluencedbyseveralparameters
–  Atomicoxygencanadheretosurfacesalteringtheirphysicalproper=es
–  Temperaturechangescanaffectmaterialcharacteris=csofoutersurfaces
–  Long,slenderspacecraQdesignincreasesdragcoefficientduetofric=on
onthesidewalls(e.g.,GOCEhadadragcoefficientof3.7)
–  Solararraysandotherhardwarefeaturesaffectdrag
2015/11/16
PropulsionandEnergySystems
14
Capability to collect the particles flow with properties a
ly the supply for the necessary propellant for Choice
air-breathing
propulsion
systems,thruster that guarantee
of an Electric
Propulsion
2
drag force on the entire satellite. For a 1 m -surface satellite, this drag force was
Capability of the thruster to work with the encountered
Hall thruster study.14 The drag coefficient varies with altitude and shape of the
Capability
of the
work with variable inlet co
2.7.15 For the altitudes of concern, a mean value
of 2.2 can
be thruster
assumed.toUsing
n density from the JB2006 model for different time
sizes of satellite surfaces at mean
Capability
of the
thruster to operate
ows the total drag imposed on the satellite. For
a full drag
compensation,
the within the power an
Dragforce
•  Resul=ngdragforce
•  Orbitdecayasaresultof
Drag force FD , N
o be at least equal to the mean drag to justify the application of air-breathing
3
(atmeansolarac=vity;
A number
ofpar=aldragcompensa=on
trade-offs
havesize
been
during the st
d satellite sizes
include the standard-sized cubesat
of (10
cm)
, a typical
of considered
a
satellites 0.3 m2 , and a standard size of 1 m2 .
forCD=2.2)
10+2
Cubesat (0.01 m2 )
+1
10
Small sat. (0.3 m2 )
Standard (1 m2 )
100
V
L
E
O
10−1
10−2
−3
10
10−4
10−5
10−6
100
150
LEO
200 250 300
Altitude h, km
350
ESAstudy(IEPC-2007-162)
rag force as a function of altitude for mean atmospheric
conditions
Fig. 1 Decay
per orbit for different altitudes
2015/11/16
depending upon percentage of the orbit where
the drag is compensated by an equal thrust. 15
PropulsionandEnergySystems
Conclusionsfromatmosphereanalysis
•  Obviously,assizeofthesatellitedecreases,dragforce
necessarytobecompensateddecreasesaswell
•  Dependingonal=tudeandtargetedlife=me,propellant
necessarytoovercomethetotaldragforcemightbesmall
enoughtobecarriedonboardofthesatellite
•  Resul=ngfromanESAstudy,applica=onofair-breathing
technologytosatellitesinorbitsabove250kmisnot
compe==vewith„conven=onal“electricpropulsion
(propellantbreak-evenpoint)
2015/11/16
PropulsionandEnergySystems
16
V
mission GOCE ended, it provided detailed information of the Earth’s geomagnetic field by
as low as 229 km using ion thrusters to compensate the drag. The amount of propellant on bo
thatperfectly
works
perfectly
in continuum
might
worsen
even counterproductive,
become counterprodu
Air-Intake
thatAir-Intake
works
perfectly
in continuum
might
worsenworsen
and
even
become
counterproductive,
providin
Air-Intake
that works
in continuum
might
and
evenand
become
pr
toless
theparticle
thruster
lessthan
particle
than
without
an
Air-Intake.
The
incoming
flowi
to thetothruster
flow
an
Air-Intake.
The
incoming
particle
flow toparticle
thetointake
the thruster
less particle
flow without
thanflow
without
an Air-Intake.
The incoming
particle
flow
the
defined
by the
front
areaintake
Athe
and
free conditions
stream
(e.g. num
byfactor
theby
open
area
ofarea
the
intake
Aofinthe
and
free
conditions
(e.g.conditions
number
density
n
defined
the
open
front
of
the
Aintake
freeisthe
stream
(e.g.
number
den
instream
inthe
limitingdefined
lifetime
forfront
such
a open
mission,
in
particular
ifand
the
S/C
orbiting
very
low
around
a pla
and
velocity
and velocity
vin
) as:
and velocity
vin ) as: vin ) as:
atmosphere. The atmosphere is indeed responsible
for
thenAdrag,
which
down the S/C and red
Ṅ
nin vin Aslows
Ṅin =Ṅ
nin
(2
Ain
in
in
in v=
in in
inv
in=
Systemdesignforair-breathingpropulsion
total mission lifetime. It is also a limiting factor in terms of costs, as more drag to be compensat
Multiplying
theṄflow
particle
flow
Ṅ with
the particle
average m
particle
mpin
results
in the correspo
Multiplying
the
particle
flow
with
average
particle
in the
corresponding
mass flo
Multiplying
the particle
Ṅ the
the
average
mass
results
the corresponding
m
p results
pmass
longer time
means
an increased
amount
ofwith
propellant
to be mass
carried
onmboard,
which
again increases
ṁ. =
massumption
. With
the
assumption
the
entire
flow
ṁthr by
is accelerated
byth
ṁ = mṁp Ṅ
With
the
that
the
entire
collected
mass collected
flow
isṁaccelerated
theby
thruster,
=. m
With
assumption
that
the that
entire
collected
massṁflow
is accelerated
the thrus
p Ṅthe
thrmass
p Ṅ
thr
mass. The
lifetime
ofproduced
acan
S/C
orbiting
can
be significantly increased by the application of an
thrust
canin
beLEO
produced
thrust
becan
calculated
as
incalculated
Eq.
3: Eq.as3:in Eq. 3:
produced
thrust
be calculated
as in
propulsion system capable to compensate the drag.
Fthr
=vPropulsion
m
vpout
=in
Ṅ
=v)(⌘
(mvcpout
nin
vin
) use th
F = mFp Ṅ
⌘=
Ṅm
v⌘out
=m
nin
vvshortened
vout
A
)vin
(3
=Electric
mvpout
Ṅ
v(m
(m
)(⌘
))(⌘c A
p Ṅ
p ⌘p
c=
in
out
in
out
thr
cthr
in
in
out
c Ṅ
out
pn
in
in
cA
The basic idea of an Air-Breathing
System,
RAM-EP,
isinto
the residual
as
propellant
toof
process
through
aAdescription
device
forof
generating
thrust.
T
Here,
vout
is
the exhaust
out
ofAit
the
thruster.
deeper
description
of the
influencin
Here, atmosphere
vHere,
exhaust
velocity
outand
of velocity
the
thruster.
deeper
description
of the
influencing
factors
for th
is
the
exhaust
velocity
out
the
thruster.
A deeper
the influencing
factors
out isvthe
out
decrease,flow
ideally
nullify,
the the
on
board
requirement
and
will
generate
to inpartially
flow
propulsion
device
will
be described
further
described
as
partbalancing
of model
thethrust
balancing
model
in
into
propulsion
device
will propellant
bewill
further
described
as part
ofpart
the
balancing
in Sec.
IV.
th
flowthe
into
the into
propulsion
device
be further
as
of
the
model
Sec.InIV
compensate
the
A conceptual
scheme
offor
the
inPropulsion
Fig.Propulsion
1. System
context
the
of for
an an
Air-Intake
for is
anshown
Air-Breathing
Electric
Propulsion
System
two
context
ofdrag.
the of
design
ofofan
Air-Intake
Air-Breathing
Electric
two important
poini
context
the
design
ofdesign
an Air-Intake
anS/C
Air-Breathing
Electric
System
two
importan
have
to be
considered.
On hand,
the
one
hand,
for
the
thrust,
not
only
thethe
efficiency
have to
be considered.
On
the
one
for maximizing
themaximizing
thrust,
not
only
the
efficiency
of
thrust
have
tofocus
be
considered.
On
thehand,
one
maximizing
the thrust,
not
only
theatmosphere
efficiency
of
the
This paper
will
on
the
device
needed
tofor
efficiently
collect
and
drive
the
par
•  Air-breathingpropulsionsystemrequiresmorethanjusta
thruster
toof,
be
taken
care
of, but
also
a sufficient
amount
ofto
mass
flow
toEq.
be 3provided.
Eq
has
to has
be itself
taken
care
but
also
a also
sufficient
amount
of mass
flow
be provided.
suggests
tha
itself
to behas
taken
care
of,
but
a sufficient
amount
of
masshas
flow
has
to behas
provided.
Eq. 3 sugge
–  Intaketocollect itself
areaintake
should
be as large
as possible
tothe
collect
most
amount
of mas
the area
thethe
intake
Aof the
should
be A
as large
possible
to collect
the most
amount
of
mass
flow,
theof
area
of
the
Aintake
should
be
as as
large
possible
to collect
mostthe
amount
of
mass howeve
flow,
h
Flight
Direction
the
S/C
front
also
the
drag.
On
the
other
hand,
when
thesystem
feedi
Solar
Array
the S/C
also
determines
thedetermines
drag.
the
other
when
considering
theconsidering
feeding
system
to th
thefront
S/C area
front
area
alsoarea
determines
the On
drag.
On
the hand,
other
hand,
when
considering
the feeding
necessarymassflow
chamber
asthe
part
of the
thehas
collected
gas
has
also
toatbe
fed pressure
at a sufficient
pressur
chamber
as part
the of
Air-Intake,
theAir-Intake,
collected
gas
also
be fed
at
a sufficient
and, inside
a
chamber
asofpart
Air-Intake,
the collected
gas
hastoalso
to be
fed
a sufficient
pressure
and, in
ionization
chamber,
neutral
gasremain
should
remain
aspossible
long
for
an efficient
ionizat
ionization
chamber,
the neutral
gasthe
should
remain
as longas
aslong
possible
for as
an possible
efficient
ionization
process.
Fo
ionization
chamber,
the neutral
gas
should
as
for
an efficient
ionization
proc
andpre-compression
the to
particles
have
to bewhile
slowed
down
while
increasing
pressure.
Therefore,
direct
flo
this, the
particles
have
be slowed
down
increasing
pressure.
Therefore,
a direct
flow ofaflow
free
this,
the this,
particles
have
to be
slowed
down
while
increasing
pressure.
Therefore,
a direct
ofstrea
free
particles
the chamber
would
not be desired.
particles
into the
chamber
would
not benot
desired.
particles
into
the into
chamber
would
be desired.
Incoming
flow
Exhaust
–  Accumula=onand
S/C Core
Air-Intake
III. Air-Intake
Review
III. III.
Air-Intake
Review
Air-Intake
Review
furthercompression
A. Concepts
Basic Concepts
A. Basic
Concepts
A. Basic
intheS/Ccore
The
basic
for design
an Air-Intake
design
aredescribed
hereby
and shown.
The basic
for anconcepts
Air-Intake
are hereby
described
and described
shown.
The concepts
basic
concepts
for
an Air-Intake
design
are
hereby
and shown.
Solarcylinder
Array
A
firstconfiguration
logical can
configuration
be cylinder
a short
section
of the
entire
Sb
A firstAlogical
configuration
becan
a short
cylinder
with the
section
ofcross
the of
entire
S/C,
followed
first logical
be acan
short
withcross
thewith
crossthe
section
the entire
S/C,
follo
–  Thrustersuitabletoa simple
aentrance
simple
cone
asinshown
in cross
Fig.
The
crosscould
section
cone
could
converge
entrance
cone entrance
as
shown
in Fig.
2a.
section
cone
converge
directly
to thedirect
a simple
cone
as shown
Fig.The
2a.
The 2a.
cross
section
cone
could
converge
directly
tosize
the
the
propulsion
system
or,to
alternatively,
to a system
feeding
system
allowing
other
S/C
the propulsion
system
or, alternatively,
a feeding
system
allowing
the other
subsystems
to subsyste
be place
the propulsion
system
alternatively,
to a feeding
allowing
the S/C
otherthe
S/C
subsystems
to b
Figure
1:or,
Air-Breathing
Electric
Propulsion
S/C
Concept.
handleatmosphericgases
behind
the Air-Intake.
behindbehind
the Air-Intake.
the
Air-Intake.
in
2015/11/16
in
in
the thruster, called the Air-Intake. The investigation on an Air-Intake is part of a Ph.D. program
focused on the use of a small inductively heated plasma thruster based on IPG6-S for an Air-B
Electric Propulsion application,14 .13 Emphasizing its crucial design, recent studies involved ESA,4
Inc.8 and JAXA,5 ,7 that proposed and studied di↵erent possible design configurations. The out
these studies are sustained by DSMC simulations and experimental activity on ground. In the fo
an overview of these designs is provided, together with results of DSMC simulations, performed w
in-house code PICLas.10 Additionally,
a Simple
simpleCone
analytical
model based on transmittances
and the ba
(a) Air-Intake
Concept.
(b) Air-Intake
By-Pass Con
(a) Air-Intake
Simple
Cone Concept.
(b) Air-Intake
By-Pass
Concept.
(a) Air-Intake
Simple
Cone Concept.
(b) Air-Intake
By-Pass
Concept.
particle flows is derived, applicable for the analysis and further possible optimization of a generic A
design. The model is compared to the results
of 2:
theAir-Intake
DSMC
simulations
and a sensitivity analysis of t
2: Air-Intake
2:Figure
Air-Intake
ConceptsConcepts
FigureFigure
Concepts
parameters is performed.
PropulsionandEnergySystems
17
However,
theusing
approach
of
using
simple
cone
isone
notasthe
one
as the
flowincontinuu
isthe
notcon
in
However,
the approach
of using
a simple
not
the
best
one
asflow
the
isinnot
However,
the approach
of
a simple
cone aiscone
not isthe
best
thebest
isflow
not
the
Collectorsystem
Figure 5: Air-Intake from Fujita’s 2004 paper,.5
•  Collectorisnecessarytogatherpropellantfromresidual
IV. Balancing Model
atmospheretoyieldamassflowrateforthethruster
•  Duetolowdegreeofioniza=on,electromagne=ccollector
systemscanbediscarded
•  Mechanicalsystemhasmanyparameters
A.
Introduction
In this Section a simple, analytical model for the evaluation of a generic RAM-EP Air-Intake configuration
is presented. The generic design becomes obvious when comparing the introduced designs, where an intake
section collects the particles with free stream conditions and guides them into the propulsion system. In the
context of this model, the intake section is followed by a chamber section in which it is assumed that all
particles have already gone through wall collisions and, thus, have only a thermal movement with regard to
the wall temperature left. By this, the only particle flows directed out of the chamber are due to thermal
diffusion. One flow back through the intake with its desirably low transmittance probability, and another flow
through the outlet. The representation of the outlet flow strongly depends on the specific configuration. For
the JAXA’s design, it is the flow passing through the thruster grids, increased by the acceleration provided
by them. In general, a feeding system and the thruster itself follows. By balancing these particle flows, the
conditions in the separate sections can be estimated.
–  Width,length,material,walltemperature,…
B. Assumptions
–  Numericalsimula=on(e.g.,DSMCPIC)essen=alindesignprocess
The basic assumptions for the analytical model are following the nomenclature in Fig. 6:
•  Performancevalues:
–  Compressionra=o
pprop /pin
–  Collec=onefficiency
Free Stream Condition Intake
pin , nin , Tin , vin
Control Volume, Chamber
pch , nch , Tch , vch
Θintake1 , Ṅintake1
Θout , Ṅout
Ṅin
Θintake2 , Ṅintake2
ṁprop /ṁin
(Ṅaccel )
Aout
Ain
Twall
Figure 6: Balancing Model Scheme.
2015/11/16
PropulsionandEnergySystems
18
of the
Ain and Aout are the respective cross sections for the inflow and the outflow representing those
chamber section. The parameters of the incoming flow are known: number density n (or pressure p ),
Airinletexamples
•  1stcomprehensivestudyconductedatJAXA(Nishiyama/Fujita)
–  ABIE:air-breathingionengine
–  Varia=onofparametersto
analyzepossibleperformance
–  Compressionra=oderivedto
bearound100-200
–  Op=mumcollectordesign
dependsonal=tude
2015/11/16
PropulsionandEnergySystems
19
aerodynamic
or of
a is
hybrid
device.
Aby studies
andintake
each
them
then strikes
otherofmolecules.
The effecttubes
is towith
pusha Direct
a considera
This
supported
simplified collection
Simu
collection system based on electromagnetic effect
[5]
Pinlet
molecules
in
the
desired
direction,
toward
the
passage
to
the
thruster.
Our
study
Pstag = Patm + ½ V2
(DSMC)
would requireCarlo
a sufficient
amount code.
of ions in The results of one such run are shown in Figure 10. Thato
.7 Schematic diagram showing the collector concept
orbit. As most
ofconvinced
the atmosphere
is neutral
the
has
us
thatatsuch
“collisional
pushing”
essential
collection
simple
inlet cylinder,
60 cm
in diameter,
3.7 m islong,
with to
an good
exit aperture
14effi
cm
altitudes of interest no sufficient MFR can be
Initial
focus
was
on larger
sized vehicle,
thus the
larger dimensions.
haveSim
sim
ed by this concept. For the same reason also the hybrid device was discarded.
Therefore
the
aerodynamic
This
is supported
by studies
of simplified
collection
tubes with aWe
Direct
2
was selected as collecting concept. According with the calculation performed,
an areadimensions
of 0.15m was judged
[5]
smaller
at appropriately increased density and found identical results.
Carlo
(DSMC)
code.
cient to collect the required MFR for all the altitude ranges considered
in the study.
Nevertheless,
a higherThe results of one such run are shown in Figure 10. Tha
thethecollection
efficiency,
a totally
absorbing
was
placed
rightaperture
after the14exit
grid
system
used
to stop 60
the cm
area of 0.6m2 was allocated to the collector intake, to take into account
simple
inlet
cylinder,
in diameter,
3.7surface
m long,
with
an exit
c
s at the entrance of the collector, and an extra design margin. The selected
area
was
consistent
with
the
output
figure,
the
curved
equidensity
contours
are
evidence
that
the
incoming
gas
is
2
Initial
larger
sized vehicle, thus the larger dimensions. We have si
tion of 1m front area taken for the calculation of the drag. The collector
lengthfocus
is fixed was
in orderon
to get
the
pushing
theperformed
trapped
d steady level of pressure at the end of the collection chamber. Various
simulations
with
the toward the back of the tube, increasing the collection efficie
smaller
dimensions
atgas
appropriately
increased density and found identical results
-software SMILE were executed varying the collector length from 0.2 m to 1.0 m. The stagnation pressure at
k of the collector, computed for 3 cases, is shown in Fig.8 and in Fig.9.
As showed,
the maximum
is reached a totally absorbing surface was placed right after the ex
the
collection
efficiency,
ngth of 1.0m. In this case the collector can provide at the inlet of the thruster a pressure of 10-3Pa. Further
output
figure,
the atcurved
ions have been performed with a collector 1.3 m long and with varying
collector
configuration
the end: equidensity contours are evidence that the incoming gas
e, straight
and
divergent.
The
first
conclusion
is
that
from
a
certain
point
the
length
has
no
positive
effect
pushing the trapped gasontoward the back of the
tube, increasing the collection effic
E. Collector Design
ESA,2007
Busek,2012
ssure. Changing the collector configuration at the end of the aerodynamic intake with a concave or divergent
The collector system is positioned in the
ems not to result in an improvement either.
Grant No. NNX11AR29G, Final Report
Busek #2
the study it was noted that from the spacecraft operations point of
RAM
direction of the S/C and is meant to
GIE
Collector
V
=
7.7
km/s
tank, located between the collector and the thruster, would help to
provide
the requested
Mass
Rate
(MFR)
In
these simulations,
the entrance
to theFlow
tube had
an area
almost twenty times the ex
ith fluctuations in atmosphere and non-nominal situations. The tank
aperture
to the thruster,
yet thethruster’s
device collectsinlet
almost forty
of the incident gas, and th
be filled when excess mass flow is available and emptied when
directly
to the
at apercent
certain
ow is scarcer. However, to store a quantity of air significant for the
fraction is still going up as the density–diameter product increases.
pressure. Various concepts of collection systems
es above, a very high pressure would be required in the tank, nonThe result makes sense. For an efficient cascade, one needs a collisional mean free pa
P
electromagnetic
scoop,
ible with the current spacecraft configuration.
A spherical tank
T much were
stag
less thanconsidered:
the diameter of an
the tube.
For a head-on collision
that an
drives the struck molecu
x
would be limited by the spacecraft cross section. Moreover the
-19
aerodynamic
intake
or
a
hybrid
device.
10
P
toward the back of the tube, one expects a cross-section of 10 m2 orA
less. To make the mea
uld be filled only during the 11minutes of non-thrusting periods of
free path
a tenth of system
the tube diameter
thatelectromagnetic
means one needs a density-diameter
product at th
collection
based
on
effect
g eclipses. To store all the particles collected in such period of time
20
-2
P
back
of
the
tube
of
10
m
.
Allowing
for
a
100
compression
of
the
gas
in
the
tube, one the
2
inlet
Pstagequation
= Patm + ½
V perfect gas that the
be easily estimated from the
of the
18
-2
would
require
a
sufficient
amount
of
ions
in
needs
at
least
10
m
based
on
the
density
of
the
incoming
gas,
which
is
what
the
simulation
d pressure in the tank would be prohibitively high. The proposed
Fig.7 Schematic diagram showing the collector
concept
Fig.8 Stagnation
pressure
levels
as a As most of the atmosphere is neutral at the
orbit.
show.
tion was to accept the variation in the atmosphere conditions, and
function of the collector length
hat the engine could cope with it.
altitudes of interest no sufficient MFR can be
P
10
Airinletexamples
•  SubsequentstudiesatESAandBusek(UScompany)
Collection Efficiency
Pressurefordifferentaspectra=os
Figure 10 Density distribution predicted by the DSMC code for a collection t
DSMC
Code Predictions
provided by this concept. For the same reason also the hybrid
device was discarded.
Therefore
the aerodynamic
diameter
intake was selected as collecting concept. According with the calculation performed, an area of 0.15m2 was judged
matters
most
is the ratio
of the
collisional
mean-free-path
0.5
as sufficient to collect the required MFR for all theWhat
altitude
ranges
considered
in the
study.
Nevertheless,
a higher to the tube di
2
the product
of the
incident
gas density
the tube
diameter.
Runs were made f
0.4 the times
collector
to take
into account
grid system
to stop
the
surface area of 0.6m was allocated to the Figure
10intake,
Density
distribution
predicted
by
theused
DSMC
code
for a collection
densities
to
determine
this
dependence.
The
results
are
shown
in
Figure
11. We f
particles at the entrance of the collector, and an extra design margin. The selected
area was consistent with the
0.3
diameter
efficiencies
if the density
enough
andtothe
assumption of 1m2 front area taken for the calculation
of are
the achievable
drag. The collector
lengthisishigh
fixed
in order
gettube
the is large enough
0.2
What
matters chamber.
most is the
ratio
of
the collisional
mean-free-path
required steady level of pressure at the end of the
collection
Various
simulations
performed
with the to the tube d
0.1
DSMC-software SMILE were executed varying
the collector
length
from gas
0.2 m
to 1.0 times
m. Thethe
stagnation
pressure Runs
at
the product
of the
incident
density
tube diameter.
were made
0
Fig.9
Pressure
in
the
collector
the back of the collector, computed for 3 cases,
is shown
in Fig.8 and
in Fig.9.
As showed,
is reached
densities
to determine
this
dependence.
Thethe
results
are1.E+18
shown
in 1.E+19
Figure 11. We
9maximum
1.E+16
1.E+17
-3
àabout1mPaachievable
for a length of 1.0m.
In this case the collector
can
provide
at
the
inlet
of
the
thruster
a
pressure
of
10
Pa.
Further
efficiencies are achievable if the density is high enough and the
-2 tube is large enoug
5
Density
x Diameter
simulationsthhave been performed
with a collector 1.3 m long and with varying collector
configuration
at(m
the) end:
The 30 International Electric Propulsion Conference, Florence, Italy
concave, straight and divergent.
The2007
first conclusion is that from a certain point the length has no positive effect on
September 17-20,
Collection efficiency predicted by the DSMC code. The density
2015/11/16
20used here
the pressure. Changing the collector configurationPropulsionandEnergySystems
at the end ofFigure
the 11
aerodynamic
intake with a concave or divergent
the density of the incoming gas
160
140
with collisions
without collisions
straws with collisions
straws without collisions
Airinletexamples
v z ,m/s
120
100
•  StudyatUStuggart(Germany)
– 
– 
– 
– 
– 
80
60
40
Rarefiedflowcomputa=onwithcoupled3Dfull-PICDSMC
Numericalanalysisnecessaryasincidentflowisnotacon=nuumflow
Figure A4: Velocity Along Center Line, BUSEK Design
Geometricalvaria=oncanyieldop=mumcollec=onefficiency
Addi=onofstrawsintheintakeflowincreasesefficiency
Trackingofindividualspeciesfeasible
20
0
0
0.5
1
1.5
2
z,m
Collection Efficiency, JAXA Air-Intake,
70
2.5
3
3.5
4
χ=10.0 over L/R straw ratio
O
N2
65
O2
He
Ar
60
55
ηc ,-
50
45
40
35
30
25
20
6
8
10
12
14
16
18
20
L/R,-
Figure 10: JAXA intake with MSISE-90 model, χ = 10.0 and Θout = 0.2 at 140 km.
2015/11/16
Figure A5: Balancing model applied to the JAXA Design, Collection Efficiency over L/
PropulsionandEnergySystems
21
2.0E-03
Air Intake
Pressure of air intake [Pa]
Reflector
Low aspect ratio
Airinlet–whattoexpect
•  Compressionra=o
1.5E-03
1.0E-03
5.0E-04
–  Achievableresultsaround100-2000.0E+000.0E+00
–  Typicalpproparound1mPa
Exchangeable collimator
•  Collec=onefficiency
High aspect ratio
High aspect ratio
2.0E+14
4.0E+14
6.0E+14
8.0E+14
1.0E+15
Atomic oxygen flux [cm-2sec-1]
Fig. 10: Inside pressure of air intake as a function of
aspect ratio
Low aspect ratio
temperature as the chamber walls and it has a velocity in
– 
Typicalvaluesintheorderof40%
a random direction. The reflector dams up the flow at its
Fig. 8: Schematic of the air intake and the exchangeable
surface and reflects partly in mirror direction with a
little slower velocity and partly in a random direction
with a much slower velocity.
collimator
•  Experimentalverifica=onbyupperatmospheresimulators
III.II Performance of Air Intake
Gas →
Vacuum chamber
Ionization gauge
Collimator
Microwave →
Upper Atmosphere simulator
JAXA
Reflector
The experimental setup for the air intake
performance is shown in Fig. The air intake is set at
10cm downstream from the upper atmosphere simulator.
The inside pressure of the air intake is measured by an
ionization gauge.
Fig. 10 shows the inside pressure of air intake. The
pressure increase more than 1x10-3Pa using the
collimator of high aspect ratio. This experimental result
suggested that the atomic oxygen coming through
collimator are bounced at the reflector and the atomic
oxygen density
increase inside the intake. However, the
ATOXatESA-ESTEC
pressure is affected by the oxygen molecules leaked from
the simulator.
Fig. 9: Experimental setup for the air intake
IV. SUMMARY
2015/11/16performance measurement
PropulsionandEnergySystems
In this paper, we simulated the AO environment of
22
Consequencesforthesatellite
System analysis
RAM EP
System
Analysis
Exhaust velocity required vs
•  Ascollec=onefficiencydecreases,thrustand,hence,exhaust
Francesco
altitude
Romano velocityhavetoincreasetocompensatethedragforce
Introduction
State of the
art
ESA 2007
Diamant
2009
Ceccanti,
Marcuccio
GOCE ESA
System
analysis
Orbit
Intake
Drag
coefficient
Mass capture
efficiency
Mass flow
Thrust
Power
system
Orbit decay
ce
v · CD
2 · ⌘c
Experiment
set up
IPG6-S
Bottom
flange design
2015/11/16
Francesco Romano (Uni Stuttgart)
PropulsionandEnergySystems
RAM EP System Analysis
22 November 2013
19 / 38
23
Possiblepropulsionop=ons
•  Chemicalpropulsion
–  Maximumexhaustvelocityof4.5km/sfarlessthanrequirement
–  Nofavorablereac=onwithmolecularnitrogen
•  Electricpropulsion(seealsolecture10/19)
–  Electrothermal(resistojet,arcjet)
–  Electrosta=c(ionthruster,Hallthruster)
–  Electromagne=c(MPD,PPT)
•  Otherconcepts
IonEngine
ES
HallThruster
EM
PPT
ET
Arcjet
MPDThruster
2015/11/16
PropulsionandEnergySystems
24
Translational Temperature, K
2500
Pre-mixed injection
2000
Electrothermalpropulsion
1500
1000
•  Resistojetsandarcjetsarebasicallycapableofusing
Fig. 4-11 JUTEM Erosion Testing Machine plume and measurement plane.
atmosphericpropellants,buterosionduetooxygenwill
500
Fig. 7. Photograph of a hollow injection plume.
decreasethelife=meofthrusterdrama=cally
0
-15
-10
-5
0
5
10
15
Cathode
Radial Position, mm
Fig. 6. Translational temperature distributions.
around 1260 K, which is much lower than that in the
conventional injection.
4. Discussion
4.1. Cathode erosion
In both hollow
and pre-mixed injection, the number
Arc-heatedplasmagenerator
density of meta-stable oxygen was successfully enhanced.
However, in the usingargon/oxygen
hollow injection, operating time was
Fig. 8. Erosion of thoriated-tungsten hollow cathode before and after
operation.
Fig. 4-12 University of Tokyo arc-heater plume and measurement plane.
86
2015/11/16
Anode(nozzlethroat)
PropulsionandEnergySystems
25
Electrosta=cpropulsion-Ionthruster
•  ABIE(JAXA,2003)
–  Aimedfor150-200kmofal=tude
–  Microwavedischargeionthruster
–  Opera=onatlowinjec=onpressures
of5to500mPa
(àaddi=onalcompressionrequired)
–  T/Paround10mN/kW
–  Sizewouldrequireacompensa=on
of50to100mNofdrag
àsignificantpowerrequirement
2015/11/16
PropulsionandEnergySystems
26
Electrosta=cpropulsion-Ionthruster
he schematically electrical setup
•  Feasibilityevalua=onofatmosphere-breathingionthruster
–  RIT-10testedonatmosphericgases(N2,O2,Xe)
ging the RF – 
power,
voltage and the gas flow. The maximum
450Wnominalpower
total power consumption of 1.24kW and a gas flow of 0.3
–  àThrustlevelofaround7mN(~140km)
pping of the thruster
using Nitrogen and Oxygen respectively,
–  Griderosionnotacrucialissue(aQer500h)
and
IEPC-2013-354
RIT-10
Power(W)
B.
Th
tog
hig
N2
nit
po
nit
the
Xe
Figure 13.
with N2
RIT-10
75%O2
25%O
thruster running
574
540
0.489
0.194
0.170
78.13
115.1
86.93
Thrust(mN)
14.71
6.83
6.79
Isp(s)
3100
3636
4328
MFR(mg/s)
Pprop(mPa)
2015/11/16
Figure
11: RAM-EP-10 thrusterPropulsionandEnergySystems
performance
467
N2
am
com
lev
ach
sam
6sc
scc
bea
hours continuously. During this test, the negative hig
automatic “Turn Off” of the Negative High Voltage. Ho
beam out in this test. This shows that the RF-Generator c
Based on measurements taken on the two grids bef
diameter; any small deviation is below the precision of m
27
Conclusionsionthruster
•  Opera=onbasicallyfeasible
–  NitrogenandoxygensuccessfullyyieldthrustatarangeofMFR
–  Thrustlevelsufficientlyhightoovercomedragforce
–  Erosionnotseenthreateningtothrusterlife=me
•  However:
–  PowernecessarytouselowMFRexceedson-boardcapabili=es
–  Thrustdensity(thrust/surfacearea)toosmall;testedMFRareonly
feasibleifpropellantisstoredduringoff-=mes
–  Propellantstorageandcompressiontoachieveusablepressurewill
increasesystemweightonsatellite
2015/11/16
PropulsionandEnergySystems
28
Electrosta=cpropulsion-Hallthruster
•  ProposedfirstlybyBusekCompanyintheearly2000s
(USpatent6,834,492B2)
•  ProposalofMars-basedatmosphere-breathingsystem
– 
– 
– 
– 
– 
Grant No. NNX11AR29G, Final Report
Thrust/powerof30mN/kWat120sccmofCO
2mixture
20%ofthrustefficiency
Inletpressureoffew100mPa
Mar=anatmospherelessstudied
Plannedmaidenflightby2025
Allimages:CourtesyofBusek
2015/11/16
Air
Busek #288
Grant No. NNX11AR29G, Final Report
Bus
2.0 Phase I Results
To further the goals of the MABHET project, we have run a Hall thruster on
mixture that simulates the Mars atmosphere; analyzed thruster operation with carbon d
(ease of physics versus mixture) and looked at inlet design in the free molecular flow r
The gas mixture simulating Mars was composed of 95.7% CO2, 2.7% N2, and 1.6% Ar.
Figure 3 shows a 1500 Watt Hall thruster running on simulated Mars mixtur
vacuum tank at Busek. The thruster is a standard xenon Hall thruster which has no
modified in any way. The characteristic “jet” can be seen, though somewhat more fai
with xenon, showing that the thruster is in typical high performance operation.
Figure 1
Conceptual MABHET powered vehicle
CO2
3
Hall thruster
running onIn
CO2the
in Busek facility
The Wright brothers flew the world’s first successful poweredFigure
airplane
in 1903.
Asin
will 2003
be shown later,
the
important number for the this
thruster is the thrust to powe
late 1990s, NASA had planned to fly a powered airplane on Mars
to
commemorate
This is because the thruster must overcome drag with some margin
PropulsionandEnergySystems
29for maneuvering/rese
for a givendue
power itto
is only
the thrust that matters.
Since it is and
using in situ propellant, the s
achievement.
The Mars Airplane Package fell short of fruition
restricted
funded
1
impulse is effectively infinite. Figure 4 and Figure 5 display the thrust to power ratio
Electrosta=cpropulsion-Hallthruster
•  PPS®1350(SPT)testedwithatmosphericgases
–  N2,N2/O2+addi=veofxenon
–  Thrustof20mNatpowerinputof1kWand
MFRof2.5mg/s
–  AddingXeincreasesthrustby4-40%
–  Wallerosionverysevere
àNewmaterialsrequired
àReconfigura=onofmagne=cfieldfor
shieldingmightreduceeffects
N2/O2+10%Xe
Fig. 3 Thruster firing during characteriza
350 V
test (N2-O2 mixture with 10% Xe
additio
305
mass flow rate)
30
263
Thrust, mN
25
220
20
Fig. 1 Comparison between phase I N2-O2
mixture characterization () and phase II N2O2 mixture + 10% Xe characterization
()
15
IEPC-2011-224
10
2
2.2
2.4
2.6
2.8
3
3.2
3.4
3.6
3.8
Flow rate, mg/s
Figure 9. Thrust vs. flow rate with discharge voltage as parameter
2015/11/16
PropulsionandEnergySystems
100
30
4
Downloaded by Tokyo University (Daigaku) on March 22, 2013 | http://arc.aiaa.org | DOI: 10.2514/1.6033
Downloaded by Tokyo University on October 11, 2012 | htt
suchtoathe
diffuser
at orbitsofof about 100 km, where the gas mean free
to the
the magnetic layer thickness, LB , that contributes positivelyThe
thrust is calculated as the force opposite
acceleration
annel. A reduction
paths are tens
centimeters
and comparable with the dimensions of
efficiency.
Nevertheless,
to
For constant Lpropellant
we see that
a decrease
of B resultsifinwe
an come back
ions (including
the force due to ion impingement
on of
walls
and
B and Ud ,utilization
ncrease of electron
the[74].
thruster. Furthermore, the use of a diffuser can significantly
for
fixed
U
leads
to
a
Eq.
(19),
we
see
that
an
increase
of
L
increase
of
the
electron
energy
through
the
heating
of
the
electrons
by
neglecting
the
force
exerted
by
the
neutrals)
B
d
y the electric field.
drag
because gas is reflected diffusely in space [3].
reduction
of the
field
the the
acceleration
a
the [see
electric field.
The higher
theelectric
energy is,
theinlower
cost is to region. As
does not the
make
it possible
A magnetic
field below Bmax ! 0:3 $ B0 increase
ergy equation
In the for
present
work
conclusion,
the
gain
of
propellant
utilization
efficiency
results
from
a
create
an
electron-ion
pair
and
the
higher
the
propellant
utilization
to
achieve
the
level
of
thrust
necessary
except
very
largewe propose a simple configuration for an
in the acceleration
airbreathing
Hall-effect
balance
therates,
increase
of the ionization
layer and a decrease
of
efficiency is. For
fixedbetween
mass flow
we observe
some noticeable
propellant
mass flow rates, which is totally
impractical
for suchthruster (shown in Fig. 1; the struts holding
the center piece are not shown), in which the incoming airflow is
the electron
differences between
gases. heating.
As an example, for a propellant mass flow
ionized and accelerated directly in the Hall thruster chamber without
rate(18)
of 4 mg # s"1 and for a magnetic field strength of Bmax !
12
E2
preliminary compression as in [1,2]. Such a design leads to a decrease
0:3 $ B0 , " reaches 0.4 for O2 and remains below 0.2 for N.
E. Calculations of the HET Performance at an Altitude of 250 km
O2
of the possible drag, however, its feasibility warrants careful
10
For Bmax ! 0:3 $ B0, the discharge current and thruster efficiency
oncentrated in the
O
consideration, which is the subject of this paper. The description of
haveconsidered
consideredinthe
of aAssatellite
at an altitude of
are shown in Fig.Finally,
8 for thewe
gases
thiscase
study.
we have
N2
the model and numerical results are provided in Secs. II and III,
250
km.
According
to
the
calculations
of
Sec.
II.A,
a
thrust
in
the
8
already mentioned, a reduction of the magnetic field strength leads to
N
respectively, and the conclusions are given in Sec. IV.
of 20 mNcurrent
to maintain
satellite
in an discharge
LEO is required. At this
an increase ofrange
the electron
and, the
hence,
a large
6
"1 of O and N2 . We show, in
altitude,
the
atmosphere
consists
mainly
(19) (%12 A for O), as we see at m
_ ! 4 mg # s in Fig. 8a.
current
Fig. 10,
T, for varying
mass flowairrates
for these
Another difficulty
forthe
a thrust,
HET working
with ambient
gases
is two gases.
II. Description of the Model
4
The thrust
is calculated
opposite
thetoacceleration of
related to the cathode
working.
Indeed,asasthe
we force
see, the
cathodetohas
of B results in an
ions (including the force due to ion impingement on walls and
The following assumptions are made in the model: 1) no drag due
2
of the electrons by
neglecting the force exerted by the neutrals) [74].
to ambient gas passing through the Hall thruster chamber, illustrated
ower the cost is to
A magnetic field below Bmax ! 0:3 $ BO02does not make it possible
in Fig. 1, corresponding to a condition in which the ambient gas
0
opellant utilization
to achieve the level of thrust necessary
O except for very large
1
2
3
4
freely
flows through
the thruster chamber without a significant
= B0which is totally
ve some noticeable
N2
propellant mass flow Brates,
impractical for such a)
interaction with the wall; 2) the gas jet leaving the thruster chamber is
max
opellant mass flow
N
fully ionized providing the maximum thrust; 3) the plasma in the Hall
trength of Bmax !
12
0.15
thruster is quasi-neutral, ni ! ne ! npl , corresponding to a condition
w 0.2 for N.
in which rD is smaller than the characteristic dimensions of the
O2
10
thruster efficiency
thruster; in our case, Fig. 1, rD must be smaller than rb " ra ; 4) the
0.12
O
study. As we have
electron cyclotron radius is much smaller than the Hall thruster gap,
N2
B8max= 0.5B0
ld strength leads to
rL < rb " ra , as is required for confinement of electrons in the Hall
N
0.09
a large discharge
thruster.
6
g # s"1 in Fig. 8a.
The first assumption is well satisfied when the characteristic time
0.06
mbient air gases is
required for a gas molecule to reach the wall and transfer its
4
the cathode has to 0.4
momentum, !wall , is larger than the time required for the molecule to
0.3
0.03
Bmax= 0.3B02
freely fly through the thruster chamber, !0 ! L=V0 , i.e.
•  Computa=onalstudies
Discharge current Id ,A
Electrosta=cpropulsion-Hallthruster
–  UniversityofToulouse,2012
Thruster efficiency (η)
Discharge current Id ,A
0.2
0
0.1
0.0
1
1
a)
2
2
3
3
4
0.00
1
2
3
fdrag !
4
!wall # V0
>1
L
(1)
b)
poles Hollow cathode
Fig. 8 Discharge current, Id , and thruster efficiency, ", as Anode
a functionMagnetic
of
propellant mass flow rate for different ambient propellant gases for
Bmax ! 0:3 " B0 and for L ! 2:5 cm.
0.15
Fig. 7 Propellant utilization efficiency, !, as a function of propellant
mass flow rate and magnetic field strength for different ambient
0.12
propellant gases and for L !
2:5 cm. Same scale is used.
p
a
i
d
f
d
i
t
d
r
d
c
p
r
a
n
Mass flow, mg.s-1
4
Mass flow, mg.s-1
–  GeorgeWashingtonU/USAirForce,2012
Thruster efficiency (η)
O2
O
N2
N
Propellant efficiency (α)
•  Targetvalues:20mN,1kW,10%@250kmà3mg/s
•  Propellantstoragenecessary
w
r
a
q
i
f
i
fl
c
g
b
c
r
V
m
•  0.09
Highpower(700-800kW),highthrust(9N)foran
0.06
applica=onbelow100km(VVLEO)
S
c
rb-ra
Incoming
ambient 2rb
air
Plume
0.03
w
0.00
1
2
3
4
Mass flow, mg.s-1
b)
nction of propellant
Fig. 8 Discharge current, Id , and thruster efficiency, ", as a function of
2015/11/16
PropulsionandEnergySystems
different ambient
propellant mass flow rate for different ambient propellant
gases for
used.
Bmax ! 0:3 " B0 and for L ! 2:5 cm.
L
Fig. 1
Schematic of airbreathing Hall thruster (not to scale).
31
!
n
b
ConclusionsHallthruster
•  Opera=onbasicallyfeasible
–  Usageofnitrogen,oxygen,airmixture,CO2-basedmixture
–  Higherthrustdensityyieldssmallerdragtobecompensated
–  Measurestoreducewallerosionexist
•  However:
–  MinimumMFRforopera=onhighcomparedtocollectableinflow
–  On-boardpropellantstoragenecessary
–  “TheHall-effectthrusterapplica4ontocompensatetheatmospheric
dragforceforaspacecra;inaloworbital4tudeisnotpossible
becausealargeamountofpropellantmustbestoredtocompensate
forthecon4nuousforceac4ngonthespacecra;foralong4me,which
increasestheweightofthespacecra;andmissionrequirements.“(U
Toulouse,2012)
2015/11/16
PropulsionandEnergySystems
32
Electromagne=c-Pulsedplasmathruster(PPT)
•  Short-=mepulsedelectricalarcdischarge
–  Acrossasolidornon-vola=leliquidpropellantsurface(abla=vePPT)
–  Or:throughaninjectedliquidorgaseouspropellant
PPT
•  Usableatverylowpowerinputs(fewwags)and
verylowpropellantmassespershot(fewµg)
Solid
3-9
LP-PPT
Water
2015/11/16
PropulsionandEnergySystems
3-10
LP-PPT
33
Electromagne=c-Pulsedplasmathruster(PPT)
•  PPTneverexplicitlyevaluatedonair-likepropellants
•  Gas-fedPPTresearchstartedasearlyasthe1960s,
e.g.atGeneralDynamics,GeneralElectric
– 
– 
– 
– 
UsageofN2,Xe,H2
Highpropellantu=liza=onefficiencyof60%forN2
Thrust/Powerra=oaround10-20mN/kW
Highdischargeenergiesof50-100J
OUN FLANOE -
OUN FLANOE
OUTER ELECTROOE
FIRST STAGE
TRIGGER
ELECTRODE^
INNER
I
"ELECTROOE
I—I.27CM0IA.
1—2.S4CM DIA.
•  540J-GPPTatRepublicAvia=on(later:FairchildIndustries),1960s
I— 5 CM OIA.
20 CM
1
Figure 1.5: Schematic drawing of General Electrics A-7D GFPPT. Da
from Ref. [22].
–  Successfuligni=onwith1020m-3(fromsimula=onatinlet:1018m-3)
2015/11/16
PropulsionandEnergySystems
changes were to make the injection ports larger for better propellant
tance and to make them face in more of an axial direction to prevent p
self-triggering. Another major contributor to better propellant utiliza
the addition of a fast acting solenoid valve. The the o-ring sealed v
driven by a 4.4 kV supply permitting a 1 ms cycle time. The disch
initiated by six 18 kV trigger pins 0.7 ms after the valve cycle was
With the relatively slow thermal velocity of xenon, over 90% of ehe gas
before each pulse remained within the electrode volume for tha A-7D
Fast ion-gauge neutral density measurements were used for the mass u
efficiency calculation and showed that most of the mass was near the
(center electrode) and well away from the backplate. The A-8D and A
similar results, however, not as high of performance with slightly mor
lant loaded closer to the cathode for the A-8D and too much propellan
too quickly (conductance was too large) in the A-9D. 34
After 1965, the only reports of GE performance came from confer
Electromagne=c-Pulsedplasmathruster(PPT)
•  1990s/2000sPrincetonUniversity
–  Lowenergyathighfrequency(uptoseveralkHz)
106
–  6mN/kWwithargon
- 25 5. GFPPT PERFORMANCE MEASUREMENTS
CHAPTER
10Impulse Bit
-*
2
a.
8-
()
i i
r 6
^2
a
^
4—
Efficiency
i o
3
n
h 10 ^
J
PT8,2J/Pulse,
0.33 mg/s H20
i
0
1
m
h 15 5'
n
2H
0-
- 20
1
r~
2000
4000
6000
Pulse Frequency (Hz)
- 5
- 0
Princeton,2001
Figure 5.12: Unfiltered Imacon Fast-Framing Camera photos of a typica
charge looking straight into the discharge chamber using 0.5 µg argon
8000
energy per pulse. The photos seen here are spaced by 1 µs going from le
and then top-right to bottom-left. A white outline is provided to indicate
tion of the cathode and spark plugs.
Figure 3.2: Impulse bit and efficiency as a function of pulse frequency for PT8
rently under development at EPPDyL. In all the performance measureme
using water vapor for propellant at 2 J per pulse.
2015/11/16
Large Molecular Weight Propellant.
in the next section, however, the current configuration where all four sp
fire simultaneously will be used as an effective discharge initiator.
PropulsionandEnergySystems
35
At a fixed plenum temperature,
Electromagne=c-Air-breathingPPT
•  Performancees=ma=on–whattoexpect
Specific impulse Isp , s
–  Foratypicalra=oof5Jof
dischargeenergyperµgof
injectedmass
Ø  Isp=5000s
Ø  T/P=15-20mN/kW
(unop=mized;1960s)
à≈30mN/kW(op=mized)
Ø  Dischargefrequencyvariable
T/P
h
2015/11/16
Drag=f(h)
14000
12000
10000
AIAA-2012-4278
"
Worldwidedata
"
8000
△
•
△
•
!!
•
•
6000
!
!!
•
× "
!
PTFE ×
×
••
"
!
×
!!
△
!
Water △
!
△•
••
•
!
4000 ×
×
•
!
DME ◦
!"
×
•
× △• • •
×
!
×
×
•
Argon
× ••
×
"
• •
!
×
×
•
"
•
•
2000 ×
•
×
×
Xenon !
"
×
×△
△
•
!
×
• •• ••• • • •
×
•••
×
×
"
△
!
Nitrogen
•
△
×
×
"
×
•◦
×
!
•◦×
×
◦
△
◦
×
×
◦
◦
!
×
◦
◦
0
0
5
10
15
20
25
30
Specific bank energy E0 /mbit , J/µg
f
35
Figure 1. Specific impulse versus mass-specific energy for solid (PTF
gaseous (argon, xenon, nitrogen11 ) propellants.
Necessary
power
Necessary
energy
PropulsionandEnergySystems
5J/µg
Necessary
mbit
36
Air-breathingPPT–exemplarymission
to-power ratio of about 30 mN/kW (currently: 10-15 mN/kW).
Assuming a mission where a full drag compensation is required for a small satellite of 0.3 m2 surface area,
the drag derived from Figure 2 immediately defines the required power by the thrust-to-power ratio. PPT
2 level without interfering
have the advantage that the discharge frequency can be regulated to vary the power
much with the performance per pulse. That is, for a given power, it is possible to ignite the thruster with
a high frequency, thereby requiring low energy and small mass per pulse, or to ignite at low frequency with
scaled quantities. As an example, Figure 5 shows the results of discharge energy and mass shot per pulse
for 100 Hz and 4 kHz.
•  Exemplarymissionscenario1
–  FULLdragcompensa=onforasmallsatellite(0.3m )
–  Then:Op=mumdischargefrequencydependsonal=tude
–  S=llrequiresstorageofpropellant,thusfullcompensa=onmightbe
106
103
difficult
E4000
m4000
105
102
E100
m100
101
103
100
102
10−1
101
10−2
100
10−3
10−1
10−4
10−2
10−3
100
10−5
Meansolarac=vity;CD=2.2
120
140
160 180 200
Altitude h, km
Mass bit mbit , µg
Pulse energy E, J
104
220
240
10−6
2015/11/16Figure 5: Estimated performance for
PropulsionandEnergySystems
an air-breathing PPT on a 0.3 m2 -sized satellite at full drag compen-37
Air-breathingPPT–exemplarymission
e thruster head is shifted to a challenge on the injection control and the propellant handling s
•  Exemplarymissionscenario2
doing so, the lifetime of the satellite can be extended in a way to enable further scientific ana
–  PARTIALdragcompensa=onresul=nginaslowed-downorbitdecay
nology demonstration.
With the data given in Section 1, and assuming a constant discharge ener
e discharge– frequency
needs to be adjusted to produce the corresponding mean thrust. The resu
Asal=tudedecreases,PPTdischargefrequencycanbeadjustedto
ed in Figure 6.
Discharge frequency f , Hz
keepenergyandmassbitconstant(thus:constantsingle-pulse
performance)105
2015/11/16
104
103
102
101
Atleast4kHzopera=on
shownfeasible
Powerand
requiredmassflux
increaserapidly
Star=ngpoint
100
100 120 140 160 180 200 220 240
Altitude h, km
PropulsionandEnergySystems
re 6: Discharge frequency for a constant discharge energy of 5 J (and m
38
= 1 µg) and fu
Conclusionselectromagne=cpropulsion
•  Opera=ontheore=callyfeasible
–  Preliminarytestswithgaseouspropellantsshowhighpoten=al,buttests
withatmosphericgasess=llmissing
–  Thrust/powerra=oinsimilarmagnitudeaselectrosta=cpropulsion
–  Opera=onatlowenergyandmassinputsreducesneedforon-board
propellantstorage
•  However:
–  Electrodeerosionphenomenaunclear
2015/11/16
PropulsionandEnergySystems
39
Comparisonofbreathingelectricpropulsionsystems
Parameter
Electrothermal
Ionthruster
Hallthruster
PPT
Opera=onwith
atmospheric
gases
Par=allytested
Feasible
Feasible
Par=allytested
Opera=onatlow
MFR
Notfeasible
Athighpower
Notfeasible
Feasible
Necessary
exhaustvelocity
Notfeasible
Feasible
Feasible
Likelyfeasible
Thrust/power
Sufficientlyhigh
Sufficientlyhigh
Sufficientlyhigh
Sufficientlyhigh
Thrustdensity
High
Low
High
Medium
Severe
Low
S=llsevere
Unknown
Verynecessary
Verynecessary
Verynecessary
Necessary
Erosion
Propellant
storage
2015/11/16
PropulsionandEnergySystems
40
Advancedconcept-Induc=velycoupledplasma
•  IPG6thgenera=onatUStuggart
–  Massflowratesof20to400mg/s(air,CO2)
–  CurrentPGpower:15-20kWwith22%
–  Electrode-lessdesignimpliesnoerosion,thus,suitableforall
composi=onsofatmosphericgases
2015/11/16
PropulsionandEnergySystems
41
Advancedconcept-ICPthruster
•  ICPpropulsionsystemdesigncriteria
–  4.4mNper1mg/sfor10MJ/kg(es=ma=onfromPGdata)
–  AsPG,sameoverallefficiencyforcon=nuousopera=onandpulsed
opera=onof1-10ms
–  Noop=miza=onyetforpropulsionpurposes,buthighpoten=al
–  Scalingforlowerpowerandreducedmassflowratesnecessary
•  Extensioninfuturetopropulsionsystemincludesapplica=on
ofnozzletechnology(mechanical,magne=c)
2015/11/16
PropulsionandEnergySystems
42