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SP2020_00056
DEVELOPMENT STATUS OF THE ST-40 HALL THRUSTER
O. Petrenko(1), (2), S. Tolok(1), A. Perepechkin(1), D. Shcherbak(1)
(1 )Space Electric Thruster Systems (SETS), 115, Gagarin Ave., Dnipro, 49050, Ukraine
E-mail: stas.tolok@sets.space
(2)Dniprovsky National University (DNU), 72, Gagarin Ave., Dnipro, 49050, Ukraine
E-mail: aleksandr.petrenko@sets.space
KEYWORDS: Hall thruster, heatless hollow
cathode, magnetic system, laboratory tests, power-
processing unit, discharge voltage stabilization,
discharge power stabilization, parameters of the
Hall thruster
ABSTRACT:
Results of development and experimental tests of
Hall thruster ST-40 are presented. Structurally, the
thruster consists of an anode unit and two heatless
hollow cathodes. The magnetic system is made in
the form of four electromagnets (central and three
external) and magnetic poles for providing the
required configuration of the magnetic field. The
magnetic system is made of soft-magnetic alloy
Permendur 49 (Curie temperature is 940°C). For
uniform distribution of the working gas (Xe) inside
the acceleration channel, a slotted gas distributor
is used. The ceramic insulator of the thruster
acceleration channel is made of boron nitride. In
the process of the anode unit optimizing, a
magnetic field in the thruster acceleration channel
was simulated and optimized.
The heatless hollow cathodes using made it
possible to shorten the thruster start-up time and
simplify the cathode structure by excluding the
additional heater.
1. INTRODUCTION
Hall thrusters are widely used on spacecraft
vehicles. The main mission for them: the problems
of orientation and stabilization, maintaining and
changing the parameters of the orbit, deorbiting the
spacecraft after the end of the mission.
The most known Hall thruster are the M-70 and
SPT-100, developed by “Fakel” Company (Russia)
[1]. These thrusters work at the level of input power
660 - 1200 W and are used on spacecraft that have
a sufficient amount of electrical power on board.
The current stage of space technology
development is characterized by a significant
decrease in the spacecraft mass, as a result of
which the level of electrical energy on board the
spacecraft does not exceed 400 - 500 W. That is
why to needs to design, develop and testing the
Hall thruster with 250 - 400 W of electrical power.
2. FORMULATION OF THE PROBLEM
Design of the electric propulsion thruster with
power consumption in the range 250 – 400 W. This
thruster must work with heatless hollow cathode.
Laboratory testing of Hall thruster must conduct at
two regimes of the discharge power supply
operating:
а) discharge voltage stabilization;
b) discharge power stabilization.
As results of the thruster laboratory testing to
determine the optimal regimes of the electric
propulsion thruster operation.
3. SOLUTION OF THE PROBLEM
For solution of the electric propulsion thruster
design with power consumption 250 - 400 W the
structure the Hall thruster was chosen. This
thruster consists of accelerating channel, magnetic
system and heatless hollow cathodes assemble –
ST-40.
The magnetic system is made in the form of four
electromagnets (one central and three external)
and magnetic poles for providing the required
configuration of the radial magnetic field. The
magnetic system is made of soft-magnetic alloy
Permendur 49 (Curie temperature is 940°C). The
stable current for these electromagnets is used
from separate power supply.
For uniform distribution of the working gas (Xe)
inside the acceleration channel, a slotted gas
distributor is used. The ceramic insulator of the
thruster acceleration channel is made of boron
nitride. In the process of the anode unit optimizing,
a magnetic field in the thruster acceleration
channel was simulated and optimized.
The heatless hollow cathodes using made it
possible to shorten the thruster start-up time and
simplify the cathode structure by excluding the
additional heater as element of hollow cathodes.
General view of the ST-40 thruster with the hollow
cathodes’ assembly is presented on Figure1.
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Figure 1. General view of ST-40 thruster
with two hollow cathodes
Table 1. Specification of ST-40 Hall thruster
Parameters
Value
Working substances
Xe (Ar, Kr)
Input power, W
200 …400
Discharge voltage, V
200 …280
Ignition voltage, V
1200
Electromagnet electric power,
W
< 10
Anode mass flow rate, mg/s
0.90…1.40
Cathode mass flow rate, mg/s
0.10 … 0.15
Thrust, mN
10 …20
Specific impulse, s
> 1200
Thrust efficiency, %
> 38
Mass of the thruster (with two
cathodes), kg
1.30
Dimensions (without cathode),
mm
140x117x122
Lifetime (estimation), hr.
5000
For the ST-40 thruster operation the heatless
hollow cathode was designed and developed. It
insures keeping arc discharge in the thruster
acceleration channel and neutralization of the ion
beam. The hollow cathode working current which
keeps auto regime operation is 1,0…1,5 А, the
value of the hollow cathode mass flow rate is in the
range 0.10…0.15 mg/sec.
4. ST-40 thruster laboratory testing
ST-40 thruster laboratory testing was carried out in
the testing laboratory of SETS (Dnipro, Ukraine)
with using the experimental facility, consists of the
vacuum chamber, laboratory storage and feed
system, flight prototype the power processing unit
and laboratory instrumentation rack. General view
of the experimental facility is presented on Fig. 2.
The vacuum chamber is equipped by
turbomolecular pump, which provides the vacuum
1·10-6 Tor at absence of the working gas mass flow
rate and value 2.4·10-4 Tor at the maximal mass
flow rates into anode and hollow cathode. Inside of
vacuum chamber the devise for measurement of
the thrust level is located. This devise can be used
for measurement of the thrust in the range 0.0 …
30.0 mN. The error of the thrust measure is about
± 5% from maximal value.
Figure 2. General view of the experimental facility
for ST-40 testing
Laboratory Xenon storage and feed system
(XFS), which was used for feeding the working gas
into anode unit and hollow cathode. It consists of
the tank with Xenon, the reducer, manometer and
devises of control and measurement the mass flow
rate (Fig. 3). Referenced values the mass flow
rates into anode unit and hollow cathode are
determined by two devises F-201CV Bronkhorst
company.
Figure 3. Laboratory storage and feed system
Flight prototype the power processing unit for
ST-40 thruster (PPU) includes the power supplies:
discharge, electromagnet, ignition of cathode and
also supplies insured the storage and feed system
operation. Electrical scheme of connection ST-40
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thruster to power processing unit is presented on
Fig. 4.
Figure 4 – Scheme of the ST-40 thruster
connection to the power processing unit
The laboratory facility also includes the
instrumentation rack, presented on fig. 5. Here the
measurement-information system, indicators and
additional laboratory facility are located
Figure 5 – Instrumentation rack general view
During the ST-40 laboratory investigation following
characteristics of the thruster were determined:
dependency the thrust from discharge voltage and
discharge power at the fixed levels of the anode
mass flow rate; the thrust from anode mass flow
rate at fixed levels discharge voltage and also the
value of specific impulse from discharge voltage.
Discharge voltage was changed in the range 200
… 280 V; mass flow rate into anode unit was
changed in the range 0.9 … 1.4 mg/s; mass flow
rate into hollow cathode was kept at level 0.10
mg/s: discharge power in anode unit was changed
in the range 200 … 400 W.
5. Results of ST-40 thruster laboratory testing
Typical volt-ampere characteristics of the ST-40
thruster in the range discharge voltage 160 … 280
V and values of mass flow rates 1.0 … 1.3 mg/s are
presented on fig. 6.
Figure 6. Volt-ampere characteristics of ST-40
Graphs (fig. 6) shows that efficiency of the working
substance ionization is very high because the
values of the current are practically constant in
wide range of the discharge voltage.
Dependencies of the thrust on the anode mass flow
rate and discharge voltage are presented on fig. 7
- 8. Graphs presented here show practically linear
dependence of the thrust on values of anode mass
flow rate.
Figure 7. Dependences of the thrust on anode
mass flow rate
Figure 8 – Dependences of the thrust on discharge
voltage
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Experimental dependences of the specific impulse
on discharge voltage at different values of anode
mass flow rates are presented on fig. 9.
Figure 9. Dependences of the specific impulse on
discharge voltage
It’s known [4], the Hall-effect thruster parameters
and characteristics which are obtained with using
the discharge voltage supply can be strongly
differenced from parameters and characteristics
obtained with using the discharge power supply.
That is why during the second stage of the ST-40
testing flight prototype the discharge power supply
was used. This discharge power supply is as
component of the PPU flight prototype and insures
the discharge power stabilizing.
Results of the ST-40 thruster testing with discharge
power stabilizing are presented on fig. 10.
Figure 10. Dependences of the thrust on
discharge power and anode mass flow rate
at the discharge power stabilizing
In frame of the ST-40 thruster laboratory testing
alongside with the static characteristics
determining the cyclogram of the thruster starting
was developed.
Specific of ST-40 thruster is the heatless hollow
cathode application. That is why the cyclogram of
thrust starting very strong difference from starting
process at classic preheated cathode application.
Typical cyclogram of the ST-40 thruster with
heatless hollow cathode starting is presented on
Fig. 11.
At the heatless hollow cathode application, it’s
possible to get ST-40 thruster starting time less
than 25 s.
Figure 11. Typical cyclogram of the ST-40 starting
6. CONCLUTIONS
1. Hall-effect thruster with heatless hollow cathode
ST-40 was designed, manufactured and tested.
Optimal regimes of the ST-40 operating were
obtained.
2. Experimental characteristics and parameters of
the ST-40 thruster were obtained with using the
laboratory power supplies which have proprieties of
flight prototypes.
3. The cyclogram of the ST-40 thruster with
preheated cathode starting was developed.
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4. As result of ST-40 thruster testing it was
improved the possibility of application such type of
the thruster onboard spacecraft in which primary
power less than 300 … 500 W.
7. REFERENCES:
1. Rossi A. Parametric optimization of a Hall Effect
Thruster magnetic circuit / A. Rossi, F.
Messine, С. Henaux, S. Sanogo. Processing of
34th International Electric Propulsion
Conference. IEPC-2015-40, Hyogo-Kobe,
Japan, 2015.
2. Petrenko O. Results of Research of Steady
Work Models of Stationary Plasma Thrusters.
Processing of the 47th International
Astronautical Congress, IAF-96-S.3.03,
Beijing, China, 1996.
3. Petrenko O. The effect of power supply output
characteristics on the operation of the SPT-100
Thruster / O. Petrenko, Hamley, J.A.,
Sankovic, J.M. Processing of the 24th
International Electric Propulsion Conference,
IEPC-95-241, Moscow, Russia, September 19-
23, 1995.
4. Bugrova A.I., Desiatskov A.V., Kaufman H.R., et
al. Design and experimental investigation of a
small closed drift thruster // Proc. of the 27th
International Electronic Propulsion
Conference. 2001. IEPC-2001-344.
5. Polk J. Electric propulsion in the USA // Proc. of
the 30th International Electronic Propulsion
Conference (Florence, Italy, 2007). IEPC-
2007-368.
6. Biagioni L., Cesari U., Saverdi M., (2005),
Development status of the HT 100 miniaturized
hall effect thruster system, Proc. of the 41th
Join Propulsion Conference, AIAA 2005-3875.
7. Tahara H., Fujioka T., Kitano T., (2003),
Optimization on magnetic field and
acceleration channel for low power hall
thrusters, Proc. of the 28th International
Electronic Propulsion Conference, IEPC 2003
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8. Bugrova A.I., Desiatskov A.V., Kaufman H.R., et
al., (2001), Design and experimental
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