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Plasma stability in plasma propulsion
An MPD thruster was tested on board the Japanese Space Flyer Unit as part of EPEX (Electric Propulsion EXperiment) that was launched March 18, 1995 and retrieved by space shuttle mission STS-72 January 20, 1996. To date, it is the only operational MPD thruster to have flown in space as a propulsion system.
Enthalpy of vaporization: Iron 3134 K 6090J/g, Water 2257J/g, Gallium 258.7 kJ/mol, Li 145.92kJ/mol
The CCU has a capability of tmnsforming the SFU unregulated bus voltage (32 - 52 Vdc) up to 350V dc to charge up in 0.55 sec the CAP of the pulse forming network (PFN) (Fig. 5). The pulse forming network is configured by a CAP of 2,240uF, consisting of 8 pair of 140uF plastic film capacitors, and a CL of 2 uH with 8 bifilar windng power lines each of which is connected to the 8 segmented ancdes anda cathode (Fig. 6). The trigger is applied by a 1 kV of 20 usec pulse duration.
The primary tank of 16 cm diam. employed a surface-tension liner sustaining the 130 g of liquid hydazine propellant with nitrogen pressurant gas. The liquid hyckazine in the primary tank is pressurized at 24kgf/cm2 gauge (2.35 MPa absolute) and supplied to the gas-generator through a propellant valve to be decomposed into nitrogen and hydrogen gases. The decomposed gas is stored in the secondary tank and supplied to the HDS through 2 sets of FAV’s.
On June 2, 1995 ckuing the EPEX firings at the visible pass of Uchinoum-Station, we switched on an S-band communication antenna which is closest to the MPD amjet thruster in a dstance of 0.6 m. At that time the telemetry fmmssynchronization was intermittently unlocked and apparently the communication was disturbed. The automated telemetry gain control at Uchinoura-Station observed 0.5 Hz well-regulated notches &ring the EPEX firings. This wasanalyzed that the plasma plume from the MPD arcjet thruster plume tlarpassed the S-band link at a distant of 7 m ahead of the SFU and the telemetry signal of PCM-PSK (Pulse Code Modulation- Phase Shift Keying) transmission was unintentionally modulated by the existence of plasma. To the contrary, this was also a good measure of plasma plume of the EPEX firing.
This measurement technique resuhedin 3.6 mN.sec thrust impulse per shot and revealed good agreement with the ground test data We can also calculate the propellant consumption rate from the pmssuredecayof secondary tank as shown in Fig. 14. The specific impulse was deduced fromthesedataandit wasptovedtobe 1,l00sec as the peak value which is correspondent with the ground test data. The total accumulated firings amounts to 43,395 pulses on-orbit during the assigned experimental period.
In this section we take a more detailed look at the conclusions of each group with respect to several thruster design considerations; Geometry and Electrode Design, Applied Field Shape, Current Outflow, and Propellant and Propellant Injection. Our current scaling parameter, ξB, while it has some sensitivity to propellant (through ao) and geometry (through R¯), does not take into account these differences between thrusters. We will see in the following paragraphs however, that they can have a significant effect on AF-MPDT performance.
The most extensive study of geometry effects in AFMPDTs has been conducted by Myers [47, 48]. Eight thruster geometries were tested at a constant current of 1000 A and a constant argon flow rate of 100 mg/s. Several groups in Russia have also done detailed empirical and theoretical studies of MPDT geometry. The MAI thruster design is based upon at least a decade of empirical studies, many of which are still inaccessible but are summarized in MAI technical reports from 1994-1998 [59, 58].
7.2 Applied Magnetic Field Shape 7.3 Current Outflow 7.4 Propellant, Mass Flow Rate, and Propellant Injection
In this work, we have reviewed much of the existing experimental performance data for magnetoplasma dynamic thrusters operating with applied magnetic fields (AF-MPDTs). The majority of the data has been obtained at power levels below 30 kW. Much of the performance data obtained with gaseous propellants was obtained at questionable facility background pressures (Pb > 1 mTorr). As a result, most of the performance data reviewed in this work was obtained with low-power lithium-vapor thrusters. Low power hydrogen, helium, and argon data from the University of Tokyo and argon and hydrogen data at 100 kW (AF100kW) taken at background pressures of 10−3 − 10−4 Torr were also included. Only two thruster designs were operated in the 100-250 kW power range for which performance data was obtained, the 100 kW thruster (Myers) and the MAI 150kW and 200kW lithium thrusters. The only AF data obtained at powers above 250 kW was taken with the quasi-steady MW-class (Osaka University) thruster operating above 500 kW. Clearly there is a gap in AF-MPDT knowledge in the 250-500 kW power range. Lithium thrusters have demonstrated the best performance to date; 69% at 5500 s with the 20 kW EOS LAJ-AF-2 thruster in the low pressure NASA-Lewis facility. Similar performance (55% at 4500 s) was obtained at EOS with a similar thruster and at MAI (50% at 4240 s), also with lithium, but at higher powers.
The thruster designs and operating regimes were found to span a large parameter space. A nondimensional scaling parameter (ξB) based upon Tikhonov’s critical current (J) was presented to assist in the interpretation of the data. This parameter showed some promise in scaling thrust at modest powers (30-100 kW) and mass flow rates (10-100 mg/s). Thrust was found to increase linearly with increasing ξB at ξB . 2. The slope of the thrust vs ξB curve increased as ξB decreases. As a performance scaling parameter ξB is lacking in several respects as is seen by the relatively large scatter in the data at different operating regimes and thruster designs. Some additional parameters which are not included in J but have been found to influence AFMPDT behavior are; magnetic field shape, current distribution (outflow), propellant injection location, and electrode design.
The lithium Lorentz-force accelerator (Li-LFA) uses a multichannel hollow cathode and lithium propellant to substantially reduce the cathode erosion problem while significantly raising the thrust efficiency at moderately high power levels. For example, a 200-kW Li-LFA has demonstrated essentially erosion-free operation over 500 hr of steady thrusting at 12.5 N, 4000 sec Isp, and 48% efficiency. Since no other electric thruster has yet shown such a high power processing capability, the Li-LFA is at the forefront of propulsion options for nuclear-powered deep-space exploration and heavy cargo missions to the outer planets.
Lorentz Force Accelerators (LFAs) _ Electric Propulsion and Plasma Dynamics Laboratory.pdf
Overview on IRS electric and advanced electric propulsion activities, uni-stuttgart, EPIC, 2017
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5 years ago Pulsed Inductive Thruster (suggested by ncc21382
Due to this ability, it has been suggested to use PITs for Martian missions: an orbiter could refuel by scooping CO2 from the atmosphere of Mars, compressing the gas and liquefying it into storage tanks for the return journey or another interplanetary mission, whilst orbiting the planet.[3]
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5 years ago Magnetic Nozzle
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A centre-triggered magnesium fuelled cathodic arc thruster uses sublimation to deliver a record high specific impulse
