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.
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) . 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.
Figure 1. General view of ST-40 thruster with two hollow cathodes
Table 1. Specification of ST-40 Hall thruster
|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 3 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
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 , 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
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.
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.
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