PlaS-40 Development Status: New Results

Transcription

1 PlaS-40 Development Status: New Results IEPC /ISTS-2015-b-9 Presented at Joint Conference of 30 th International Symposium on Space Technology and Science 34 th International Electric Propulsion Conference and 6 th Nano-satellite Symposium, Hyogo-Kobe, Japan M.Yu. Potapenko 1, V.V. Gopanchuk 2, S.V. Olotin 3 FSUE EDB Fakel, Kaliningrad, , Russia Abstract: This paper presents test results of the PlaS-40 updated engineering model. Possibility of reaching the total thrust efficiency up to 50 % in a small-sized thruster is demonstrated herein. Results of life time test, mechanical test and plume characteristics measurement are given in this paper. The PlaS-40 constructive development performed corresponds to prequalification development level. B r,max Z B r I d U d G a I i I e U cg K i Nomenclature = maximum value of the magnet induction radial component in a gap between the magnet poles; = axial component of gradient of the magnet induction radial component; = discharge current; = discharge voltage; = anode gas flow rate; = ion current; = electron current; = "cathode-ground" voltage; = gas ionization rate. I. Introduction n the epoch of high technologies and information flow increase it is necessary to develop and use global Itelecommunication systems. Such tasks can be solved using a heavy spacecraft (S/C) placed on GEO or deploying a low-orbit S/C constellation consisting of several tens of small satellites. For these purposes, it is reasonable to use low-orbit S/C systems because it is time and cost effective. Many leading IT-companies in US and Europe are eager to create their own global telecommunication systems. The today s tendencies in spacecraft development are oriented to increase some part of payload reducing mass-and-dimension parameters of on-board electric propulsion system (EPS). Besides such EPS should be high-efficient. The main operating element of EPS is a thruster. Electric propulsions (EP), in particular, Stationary Plasma Thrusters (SPT) are applied more often today. To reach high efficiency in small-sized thrusters is known to be a great challenge because of near location of all elements of a thruster and their interference. The search for effective solution results in appearance of more different constructive modifications of the SPT classical scheme and fundamental new scheme of Hall thrusters are appeared nowadays [1]. 1 Design Engineer, Ph.D, engineering department 2 Leading Designer, engineering department 3 Leading authority, test department 1

2 II. PlaS-40 design development and optimization The EP of the most challenging scheme is a thruster of PlaS-type parametric family [2]. EDB Fakel is developing the low-power thruster PlaS-40 (fig. 1) of a new constructive scheme. Named thruster operates in the power range from 100 to 650 W. The middle diameter of Accelerating channel is equal to 40 mm, mass of the thruster with one cathode is 1.2 kg. To improve PlaS-40 thrust performances its design optimization was performed. The fulfilled design updating consisted in the thruster magnet system improvement in order to enhance thrust efficiency and life performances. The central task was minimization of the DCh walls erosion by decrease of the magnet "lens" length and its displacement from the discharge chamber (DCh) depth to the Figure 1. PlaS-40 Engineering Model (EM) DCh cut.. Parametric test III. PlaS-40 experimental research PlaS-40 test was performed in the EDB Fakel vacuum chamber at discharge voltage from 100 to 500 V in the anode flow rate range of xenon from 1.25 to 2.5 mg/s and the cathode flow rate of 0.18 mg/s in the horizontal vacuum chamber at dynamic pressure not more than mm Hg (in terms of air). The range of the updated thruster PlaS-40 performances at a discharge power of up to 650 W with account of discharge power and power consumption for magnetic field generation, and also with account of propellant flow through the cathode and adjustment for the level of the vacuum chamber pressure is given in table 1. Results of the parametric tests of the updated thruster including current-voltage characteristics measurement at different anode flow rate level are given in figure Current-voltage characteristics (fig. 2) were measured at the specified anode flow rate, which corresponds to the ensured discharge current of 1.00 to 2.25 with an increment of Magnetic field has been optimized by minimum discharge current at each operating point. Table 1. The range of the main performances of the PlaS-40 thruster Performances Value Discharge voltage, V Discharge current, Discharge power, W Thrust, mn up to 40 Specific impulse, s up to 1880 Efficiency, % up to 50 Power-to-thrust ratio, W/mN Mass, kg 1.2 Overall dimension, mm When PlaS-40 is ON, the current-voltage characteristics are of conventional type (fig. 2). Area of the higher discharge current becomes apparent at discharge xenon anode flow rate increase in the range of discharge voltage from 100 to 150 V, after that discharge current keeps constant at the discharge voltage increase up to ( ) V. Insignificant growth of the discharge current is observed at discharge voltage increase more than 350 V, what, as will be showed below, caused by increase of the ion current part. Dependencies of the thrust, power-to-thrust ratio, specific impulse and efficiency of the PlaS-40, shown in figures below, are of traditional type. Thrust and Specific impulse increase at anode flow rate through accelerating channel increase and at discharge voltage increase. Stable and regular growth of the specific impulse indicates sufficiency of plasma concentration in the whole tested range of the discharge voltage to reach high degree of the atom flow ionization in accelerating channel at high discharge voltages. Figure 2. PlaS-40 C-V characteristics 2

3 Maximum thrust determined at the anode flow rate of 2.58 mg/s and at the discharge voltage of 300 V is 44.0 mn. Minimum power-to-thrust is 12.3 W/mN and corresponds to discharge current of 2.0 A and discharge voltage of 150 V. The total specific impulse calculated with an allowance for the discharge power and power consumption for magnet field generation, and also with an allowance for cathode flow rate and for the level of vacuum chamber pressure is 1880 s at maximum discharge voltage of 500 V and power of 500 W. Figure 3. Thrust dependencies on discharge voltage Figure 4. Specific impulse dependencies on discharge voltage Figure 5. Efficiency dependencies on discharge voltage Figure 6. Power-to-thrust ratio dependencies on discharge voltage The thruster PlaS-40 operating modes tested are shown in the table below. Table 2. PlaS-40 operating range I d, A (G a, mg/s) U d, (1.35) 1.25 (0.63) 1.50 (1.85) 1.75 (2.07) 2.00 (2.31) 2.25 (2.56) Notes: - stable operating mode; - flame out; - unstable operating mode; - untested operating mode. 3

4 The cumulative diagram also demonstrates the operating range of PlaS-40 thrust performances (fig. 7). Figure 7. PlaS-40 performances envelope RMS discharge current amplitude dependencies on discharge voltage are shown in figure 8. As shown, discharge current oscillations decrease sharply by the discharge voltage increase, which minimum 20 was detected at the discharge voltage U d =150 V. Minimum of the Power-to-thrust ratio was defined at this discharge voltage also. Analysis of gas utilization was made with flow rate-to-discharge current ratio G a /I d (fig. 9). For PlaS-40, this ratio is more than 1 (G a /I d > 1), and it is more preferable, because the value of spurious electron current is low given gas ionization is fully ensured. Figure 8. RMS discharge current amplitude dependencies on discharge voltage Figure 9. G a /I d ratio dependences on discharge voltage for thruster PlaS-40 Electron value I e, which dependence on discharge voltage is shown in figure 10, has been estimated by test whereby measured ion current at the operating point with discharge current of I d =1,5 A. Such function usually decreases monotonously. Results of the ion current measurement indicate that ion current increases sufficiently at the discharge voltage growth for all operating points, what points to the gas ionization efficiency improvement. However, as shown also, operating point with discharge voltage of 100 V is absolutely ineffective for this thruster, because ion current part and electron current part are comparable. This operating mode is also followed by high amplitude of the discharge current oscillations, as figure 11 shows. 4

5 Figure 10. Ion current, electron current and its part (at I d =1,5 A s) dependences on discharge voltage Figure 11. Ion current dependences on discharge voltage Electron Current-to-Discharge Current relation (I e /I d ) is also shown in figure above, which relation doesn't exceed 20 % of the discharge current value, what is indicative of good gas ionization and acceleration processes (fig. 10). According to the tests results, gas ionization rate K i has been determined, and its value varies in the range of 0.65 to 0.90, what is indicative of partial gas ionization (fig. 12). It is necessary to note also that at such values of the ionization rate (which values are lower than those typical for SPTs) much higher Efficiency values are reached in PlaS-40 EM (fig. 5). Figure 12. Gas ionization rate K i dependencies on discharge voltage b. Comparison of the PlaS-40 thruster models characteristics After thruster design updating dependence type of thrust performances at discharge voltage increase (fig ) does not changed. In comparison with basic design [2] sufficiently increase of thrust and also specific impulse level is observed at discharge voltage more than 250 V Figure 13. PlaS-40 thrust dependence at I d =1.25 before and after updating Figure 14. PlaS-40 specific impulse dependence at I d =1.25 before and after updating Significant growth of the thrust efficiency up to 4-6 % at discharge voltage increase followed by power-to-thrust ratio decrease is observed during test. 5

6 Figure 15. PlaS-40 thrust efficiency dependence at I d =1.25 before and after updating Figure 16. PlaS-40 power-to-thrust ration dependence at I d =1.25 before and after updating Thus, PlaS-40 thruster updating relative to magnet system improvement and to change of outlet ceramic rings material allowed to improve thrust performances up to 5-8% without change overall-mass and energy characteristics in comparison with to base design of the thruster. c. Mechanical test Thruster withstands mechanical test. By the test results the basic eigen frequency of the PlaS-40 thruster amounts to 353 Hz and is caused by cathode unit oscillations. Shock test was performed also. Peak shock acceleration during test was 80g at impulse duration of 4 ms. Shock quantity in the one direction of each axis is 12. PlaS-40 #2 shock test results proved thruster structural strength to mechanical stress of level called for small satellites. d. PlaS-40 life test For life test at the discharge power of 200 W operating mode with low power-to-thrust ratio value is more interesting for research, because test was performed at operating mode with U d =160 V and I d =1.25 during 200 hours. Besides at this mode thruster ensures thrust about 15 mn, total specific impulse about 800 s and thrust efficiency about 31 %. Thruster long operation was performed with duration of running in of 8 hours in the scope of 200 hours and 25 running in. External view of the thruster after 200 hours operation time is given in figure below. The cathode -1, which had been used in PlaS-40 thruster, operated stably at the cathode flow rate of 0.15 mg/s during the test. There is dynamics of the "Cathode-ground" voltage U cg in figure 18, which defines cathode operation efficiency at that operating point. Figure 17. PlaS-40 thruster after 200 hours operation time Figure 18. "Cathode-ground" voltage dependence on discharge voltage Variation of the thrust, power-to-thrust, total specific impulse and total efficiency during life test is shown in figures 19 and 20. Average value of the thrust during life test amounts to 15.0 mn, average value of the specific 6

7 impulse calculated with allowance for the discharge power and power consumption for magnet field generation, and also with an allowance for cathode flow rate and adjusted for level of a vacuum chamber pressure is 820 s. Average value of the total efficiency is 31 %, power-to-thrust ratio is 13.2 W/mN. Total impulse of the PlaS-40 thruster after 200 hours operation amounts to 10.8 KN s. Figure 19. Performances of the PlaS-40 engineering model during life test: ) thrust and b) power-to-thrust ratio Figure 20. Performances of the PlaS-40 engineering model during life test: ) specific impulse and b) total efficiency During operation low level of the discharge current oscillations was noted (fig. 21). Level of the discharge voltage oscillations did not change during test. Figure 21. RMS ) discharge current and b) discharge voltage amplitude dependence on discharge voltage of the PlaS-40 during life test 7

8 Temperature of the PlaS-40 mounting surface and its external magnet pole was monitored also during the life test (fig. 22). The mounting plate temperature when the thruster ON at the discharge power of 200 W did not exceeded 105. Temperature of the external magnet pole was less than 205. Figure 22. Temperature variation of ) the mounting plate and b) the external magnet pole during life test Erosion rate of the DCh walls was defined by measurement of the DCh wall depth by microscope -2 in 4 sections. Measurement was performed before firing test, and also after 200 hours of the operating time at specified mode. Average value of erosion after 200 hours operating time (fig. 23) is: - for the internal wall 0.23 mm; - for the external wall 0.08 mm. Figure 23. Erosion profile of ) internal and b) external insulator By test results, erosion rate of the internal wall within initial operating time 200 hours 2.5 to 2.7 times more than the erosion rate of the DCh external wall. Most probably this fact is caused by plasma plume overfocusing to the thruster longitudinal axis and by distinction in 1.5 times between magnet fields near DCh walls (fig. 24). Depth of the erosion rate was also measured after test. Average value of the erosion rate depth for external ring made of the ceramic BN is 1.2 mm, for internal ring it is 2.0 mm (fig. 24). Figure 24. External view of the DCh erosion rate of the PlaS-40 Thruster with a measured standard sample 2 ) internal and b) external wall of the DCh 8

9 Prediction of radial erosion of the DCh ceramic rings during life test was performed by known dependence [3]: r = A ln( 1+ ω t) (1) where r is linear value of DCh insulator end erosion; t is operation time; A, ω are normalization factors depending on ceramic material, design parameters and thruster performances. Results of calculations of the insulator end walls radial erosion are given in figure 25. Insulators walls thickness of the DCh outlet ceramic rings made of BN is 2.65 mm, that is enough to ensure life time more 4000 h at Figure 25. Calculations of the insulator end walls radial erosion the operating point with discharge power of 200 W (U d =160 V and I d =1.25 ). Thus, PlaS-40 DCh has adequate supplies of insulators walls thickness. e. Plasma plume divergence measurement Divergence of the PlaS-40 plasma plume was checked for Xe at the discharge current of 1.25 A and at the discharge voltage of 160 and 200 V, and calculated with an allowance of 95 % ion flow coverage. When plume divergence measuring probe moved in horizontal plane containing thruster axis along a circle with radius of 0.5 m and with centre placed in the intersect point of thruster axis with its outlet plane. As test showed, plasma plume focusing improves (plume divergence decreases) when discharge voltage increases Results of plume divergence estimation by calculation of the ion current density angular distribution agrees with common tendency of ion flow divergence decrease at discharge voltage increase. Discharge voltage growth leads to narrow of flow core and to some relative increase ion current density in the peripheral plume area [4]. Plume divergence semi-angle with an Figure 26. The angular distribution of the ion current density allowance of peripheral ion part is 60 for of the PlaS-40 operating at the Xe anode flow rate of 1,5 mg/s operating point with U d =160 V, and 43 for operating point with U d =200 V (fig. 26). Some ion flow axes displacement relative to thruster axes, which decreases at discharge voltage growth, is observed during test. This fact is most probably caused by some azimuthally thruster magnet field nonuniformity of 1.4 % - local magnet field level increase in area of the cathode-compensator placement, which has parts made of magnetic material influencing on magnet field distribution. IV. Conclusion Outcomes of the PlaS-40 thruster new model test confirmed effectiveness of the performed design updating, as a result of which thruster is able to ensure more than 5 % higher thrust performances in comparison with the thruster base configuration. Mechanical scheme of the Plas-40 design has an acceptable endurance to mechanical stress. Compared to the known analogues the thruster with a hollow magnet anode ensures: lower values of the thrust-to-power ratio; reduced level of discharge current and discharge voltage oscillations; higher thrust efficiency. 9

10 Performed long operation of the PlaS-40 #2 with duration of 200 hours at the discharge power of 200 W (I d =1.25, U d =160 V) demonstrated stability of the thruster performances during long operation. It is determined that initially erosion rate of the DCh internal wall is two times more than erosion rate of the external wall insulator. In consideration of initially erosion rate of the ceramic rings predicted life time of the updating design thruster at the discharge power of 200 W will be more 4000 hours. References 1 M.Yu. Potapenko, V.V. Gopanchuk. Characteristic Relationship between Dimensions and Parameters of a hybrid Plasma Thruster // IEPC , 32 nd International Electric Propulsion Conference, Wiesbaden, Germany, September 11 15, M.Yu. Potapenko, V.V. Gopanchuk. Development and Research of the Plasma Thruster with a hollow magnet Anode PlaS-40 // IEPC-2013, 33 rd International Electric Propulsion Conference, The George Washington University, Washington, D.C., USA, October 6 10, N.V. Belan, V.P. Kim, A.I. Oransky, V.B. Tichonov. Stationary Plasma Thrusters, Kharkov, A.S. Arkhipov. Research of the stationary plasma thrusters (SPT) plume characteristics at high discharge voltages. PhD dissertation, RIAME,