3C: Plasma Thrusters
3C0102: Extending Ion
Engine Technology to NEXT and Beyond M.T. Domonkos, M.J. Patterson,
J.E. Foster, V.K. Rawlin, G.C. Soulas, J.S. Sovey, S.D. Kovaleski, R.F. Roman,
and G.J. Williams, Jr. NASA John H. Glenn Research Center Mail Stop
301-3 21000 Brookpark Road Cleveland, OH
44135
3C03: Kinetic Models for Plasma Transport in the Hall
Effect Thrusters O. Batishchev, V. Blateau, M. Martinez-Sanchez MIT 77 Mass. Ave, Cambridge MA 02139, USA
3C04: Modeling,
Simulation, and Design of an Electrostatic Colloid Thruster David Kirtley ERC, Inc. Dr. John Michael Fife Air Force Research Laboratory
3C05: On the Efficiency
of Several Thruster Configurations* A. Fruchtman Holon Academic Institute of Technology 52
Golomb St., Holon 58102, Israel
3C06: Ionization Layer at the Edge of the Accelerating
Plasma in a Pulsed Plasma Thruster Michael Keidar1, Iain. D. Boyd1 and Isak I.
Beilis2 1. Department of Aerospace Engineering, University of Michigan Ann
Arbor MI 48109, 2. Electrical Discharge and Plasma Laboratory, Tel Aviv
University, P. O. B. 39040, Tel Aviv 69978, ISRAEL
Extending Ion Engine
Technology to NEXT and Beyond
M.T.
Domonkos, M.J. Patterson, J.E. Foster, V.K. Rawlin, G.C. Soulas, J.S. Sovey
S.D.
Kovaleski, R.F. Roman, and G.J. Williams, Jr.
NASA John H. Glenn Research
Center
Mail Stop 301-3
21000 Brookpark Road
Cleveland, OH 44135
Extending ion engine technology beyond the current state-of-the-art primary interplanetary electric propulsion system, the 2.3-kW NASA Solar Electric Propulsion Technology and Applications Readiness (NSTAR) system, will require thrusters with improved propellant throughput and total impulse capability. Many of the design choices that culminated in the NSTAR thrusters must be revisited, and their application to next generation ion engine technology must be evaluated. The concept of derating, which was successfully employed in NSTAR, has been applied to the 40-cm NASA Evolutionary Xenon Thruster (NEXT) currently under development at NASA Glenn Research Center (GRC). At 5-kW, NEXT operates with the same average beam current density as NSTAR, and at 10-kW, the peak beam current density is only ten percent greater than NSTAR. The result is that similar ion optics technology is expected to yield comparable lifetime. Thick-accelerator-grid ion optics are also being tested to realize additional lifetime benefits. A 40-A discharge cathode is being developed for NEXT based on scaling the NSTAR design. Nevertheless, the experiences of the NSTAR ground tests and the thruster on the Deep Space One spacecraft indicate that the discharge cathode wear must be studied experimentally and theoretically to ensure that it meets the lifetime requirements. Although NEXT is in its infancy, investigations have already begun to examine possible modifications to engine design for even higher-power and higher-specific impulse engines. Ion optics using alternate materials such as titanium, graphite, or carbon-carbon composite are currently being investigated due to their low sputter yields at high voltage. To avoid the difficulties encountered using electrodes at high-currents, the use of a microwave-based ion thruster is under investigation for potential high-power ion thruster systems requiring long lifetimes. Additionally, alternative propellants are being considered for applications requiring high-specific impulse (>> 5000 s) and extremely long-life (>> 15,000 hr). Testing requirements make condensable propellants attractive for high-power engines. Although the NSTAR ion engine demonstrated the flight maturity of ion thruster technology, many challenges remain for the development of thrusters with improved propellant throughput and power handling capabilities.
Kinetic Models for Plasma Transport in the Hall Effect Thrusters
O. Batishchev, V. Blateau, M. Martinez-Sanchez
MIT
77 Mass. Ave, Cambridge MA 02139, USA
Specifics of plasma
transport phenomena in the Hall effect thrusters were a subject of many
experimental and numerical studies [1-3]. In the present paper we present and
discuss results from two different kinetic PIC models.
First one is a model of
plasma flow in the SPT-type thruster in the idealized planar approximation. The
system is periodic in the transverse coordinate to resemble the azimuthal
dimension. We study excitation of collective plasma modes that might be
responsible for high-frequency oscillations at 0.1-10Mhz, and anomalous
transport observed in the experiment [3]. We also anticipate strong resonance
effects for the cases when plasma Debye length is close to the electron
gyroradius in the internal magnetic field of a Hall thruster.
The second model is a further development of the
fully kinetic model for plasma and neutral gas in the 2D3V axisymmetrical
approximation [2]. A previously developed computational method [4] is applied
to the realistic P-5 thruster geometry. We add new elementary plasma-chemistry
reaction and modify boundary conditions to capture self-consistent dynamics of
high ionization states of xenon atoms. We study thruster performance at wide
range of applied voltages and make comparison to the experimental data.
[1] M.Hirakawa,
Electron transport mechanism in a Hall thruster, IEPC-97-021 technical paper,
1997.
[2] J.Szabo,
Fully kinetic numerical modeling of a plasma thruster, PhD thesis, MIT,
2001.
[3] N.B.Meerzan,
W.A.Hargus, M.A.Capelli, Anomalous electron mobility in a coaxial Hall
discharge plasmas, Phys. Rev. E, 63, 026410-1, 2001.
[4] V.Blateau,
M.Martinez-Sanchez, O.Batishchev, J.Szabo, PIC Simulation of High Specific
Impulse Hall Effect Thruster, IEPC-01-037 paper, 27th International
Electric Propulsion Conference, Pasadena CA, 15-19 October, 2001
David Kirtley
ERC, Inc.
Air Force Research Laboratory
Electrostatic colloid thrusters fill a unique niche in space propulsion; small (micro/milli-Newton range), high efficiency (70%+), and a large Isp range (500-5000s). Colloid thrusters electrostatically create charged liquid particle beams to generate thrust, allowing higher charge to mass ratios than ion thrusters. Presented is a simulation and design for a low-voltage (few kV) system using commercial-off-the-shelf 3D modeling and E/M mapping. Techniques and results are explored as well as performance implications of non-ideal acceleration grid effects. Particle tracking analysis is done by building on an unstructured mesh in the acceleration volume and a static electric field generated by COSMOS/EMS design package. Finally, exploration and optimization of the performance characteristics of a colloid thruster is done.
On the Efficiency of Several Thruster Configurations*
A. Fruchtman
Holon Academic Institute of Technology
52 Golomb St., Holon 58102, Israel
Limits on the efficiency of
several thruster configurations are discussed. The efficiency of the Pulsed Plasma
Thruster (PPT) is reduced when part of the magnetic field energy that is
converted into particle energy does not become directed kinetic energy but
rather a thermal energy. This thermal energy can still be used for propellant
ionization. The partitioning of the power when the propellant exhibits slug,
snowplow or specular-reflection accelerations is analyzed. Steady acceleration
is examined in two configurations: the Magneto-Plasma Dynamics (MPD) and the
Hall thrusters. The smooth acceleration to supersonic velocities in the two
configurations1 is compared. A limit on the efficiency of the MPD
thruster in a non-diverging geometry is derived. The efficiency of the Hall
thruster in the limit of intense full ionization is discussed. Two steady-state
flows in the Hall thruster are compared for two different boundary conditions
at the anode. The first flow is of a zero ion current and velocity at the
anode. The second is the recently-analyzed2 case of an ion backflow
at the anode, in which the profile of the electric potential along the thruster
is expected to be nonmonotonic. By employing certain asymptotic relations2
together with considerations of momentum and energy balance3,
analytic expressions for the efficiency of the Hall thruster for the two flows
are derived. The possible existence of these two types of flow for different
values of the applied voltage due to the characteristics of secondary electron
emission2,4,5, is discussed.
*Partially supported by the
US-Israel Binational Science Foundation and by the Israel Space Agency.
1.
A.
Fruchtman, N. J. Fisch, and Y. Raitses, Phys. Plasmas 8, 2000 (2001).
2.
E.
Ahedo, P. Martinez-Cerezo, and M. Martinez-Sanchez, Phys. Plasmas 8, 3058 (2001).
3.
A.
Cohen-Zur, A. Fruchtman, J. Ashkenazy, and A. Gany, 27th International
Electric Propulsion Conference, Pasadena, CA (2001), paper IEPC-01-26.
4.
E. Y. Choueiri, Phys. Plasmas 8, 5025 (2001).
5.
M.
Keidar, I. D. Boyd, and I. I. Beilis, Phys. Plasmas 8, 5315 (2001).
Michael Keidar1,
Iain. D. Boyd1 and Isak I. Beilis2
1. Department of Aerospace Engineering, University of Michigan
Ann Arbor MI 48109, keidar@engin.umich.edu
2. Electrical Discharge and Plasma Laboratory, Tel Aviv University
P. O. B. 39040, Tel Aviv 69978,
ISRAEL
There are different
characteristic sub-regions near the surface namely space-charge sheath, Knudsen
layer, presheath and ionization layer, where transition to ionization
equilibrium occurs. In this work we describe the phenomena associated with a
contact wall surface-plasma under conditions of strong plasma acceleration.
Such a case is realized in pulsed plasma thrusters.
There are two specific
problems that we address in this work, namely plasma generation phenomena
(ablation) and formation of the ionization layer. Considering ablation
phenomena we couple the non-equilibrium, Knudsen layer, with the hydrodynamic
layer, that provides solution for the ablation rate. The ablation rate is
determined by the flow velocity at the edge of the Knudsen layer. Considering
an electromagnetic pulsed plasma thruster we found that, depending on the
current density in the hydrodynamic layer, this velocity varies from very small
(compared to the sound speed) up to the sound speed with current density
increase. It was shown how plasma acceleration under external forces affects
the boundary condition at the edge of the Knudsen layer.
The second problem is
important in the case of small plasma devices, where the spatial extent of the
ionization layer becomes comparable to the size of the device. The model
considers the current distribution in the thruster near field plume and its
effect on the ionization layer. For instance, it was found that the thickness
of the ionization layer is approximately inversely proportional to b, the ratio of Alfven velocity to the sound
velocity. It was concluded that in these devices, significant ionization is
obtained when the ionization and acceleration regions are separated.
A specific example of a calculation,
applied to a micro-pulsed plasma thruster developed at the Air Force Research
Laboratory, is considered.