Mpd Thrusters

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High-power MHD thrusters are less developed than ion engines, and to a large extent are still laboratory items. Exploiting the Lorentz force, MHD acceleration occurs when a flow of charges, e.g., electrons and ions, in all respects equivalent to a current, J, moves in a magnetic field B. The Lorentz force is F = J x B: it accelerates charged species moving in the magnetic field B and according to Newton's Third Principle, creates thrust. The state of a gas containing charged species, that is, an ionized gas, is called ''plasma state''. So, a plasma can be accelerated by the Lorentz force and produce thrust. Accordingly, this type of rocket engine is called a Magneto-Plasma-Dynamic thruster (MPD thruster).

The regime of an MPD thruster can be steady in the strict sense, or quasi-steady. The thrust of a quasi-steady MPD may occur in pulses or bursts; when these last long enough, or when the burst repetition rate is high enough, the averaged thrust is said to be quasi-steady. Quasi-steady MPD thrusters have been tested far more than steady MPD, one of the reasons being their lower power demand, and another their relative simplicity. For high-power applications steady MPD are better, but without a nuclear generator there is no way they can become effective space engines.

The simplest nuclear-powered MPD concept consists of a nuclear reactor generating electricity and driving an MPD accelerator. Laboratory MPD engines are of course powered by photovoltaic (solar) cells, have Isp in the order of 103 to 104 s, but their weight and size are much larger than those of ion engines. A laboratory MPD thruster may have a mass/power ratio of order 1-103 kg/kW, depending on scale. Most of this mass is that of the electric conductors (wiring), especially those of the magnetic coils. If superconducting wires replaced copper, coils and windings mass could be reduced by 1-2 orders of magnitude [Bruno and Giucci, 1999; Casali and Bruno, 2008].

In fact, recent advances in MPD technology have brightened the prospective of this type of electric propulsion. MPD propulsion has been dormant because the power required to reach acceptable efficiency was too large for commercial satellites and space vehicles (it takes hundreds of kilowatts to achieve efficiencies greater than 30%), and also because such power is unattainable with solar cells. Historically, MPD propellant acceleration suffers from many losses, for instance, (a) propellant composition ''frozen'' during expansion, preventing conversion of internal energy, as in arcjets; (b) plasma instabilities, the bane of all plasma applications, increasing plasma resistivity, driving unstable currents and wasting power; (c) excess heating of, and near, the anode; and especially (d) cathode erosion/evaporation, reducing cathode life.

A drawback of MPD engines is also their low thrust density, by a factor 5-10 lower than other electric thrusters [Auweter-Kurtz and Kurtz, 2003]. The reason is that plasma pressure must be low enough so that collisions do not prevent charges from following the magnetic force-lines. The consequence is large internal and nozzle volumes for given thrust or power.

Together with that of power, the major issue of MPD was and still is cathode life. Because of the low thrust, missions using NEP may last 5 to 6 years [Oleson and Katz, 2003], depending on Isp, and mass per unit power. Over months or years of operation even tungsten cathodes erode at the rate of approximately 0.2 mg/coulomb [Choueiri, 2000]. This figure may look insignificant, but a 20-kW MPD thruster, such as those considered for the JIMO mission, will need 20 A when operated at 1,000 V, that is 20 coulomb/s. In a day alone about a third of a gram of tungsten will have been eroded. When Russian technology and know-how on steady plasma thrusters became available after the end of the Cold War, the pace of progress in this area quickened. Interest by USAF in a particular type of MPD propulsion (Hall thrusters) is contributing to advance this field.

In fact, the most important recent development in MPD is probably the replacement of hydrogen propellant (with ionization energy, of order 13.8 eV) with propel-lants with much lower ionization potential, in particular lithium (its ionization potential is 5.37 eV). Lithium can extend cathode life by orders of magnitude [Choueiri, 1998]. Some MPD laboratories (MAI/RIAME in Moscow, CalTech Jet Propulsion Laboratory and Princeton University's Plasma and Electric Propulsion Laboratory) are now collaborating in this specific area. The Russian company NPO Energia has tested a RIAME-designed 130-kW, 43% efficiency Lorentz force MPD thruster using lithium propellant and found very low cathode erosion. Cathode lifetimes of more than 1,000 h are now within reach. Measured Isp was 3,460 s, with a thrust of order 3.2 N. Thrust of order 25N/MW looks feasible. Future plans (in the 2010 timeframe) include a 100-kW and a 120-kW steady MPD thruster.

Before Project Prometheus was started, NASA was planning improbable 20 MW, solar-powered MPD experiments in 2012, and 100 MW in 2024, clearly for interplanetary missions such as a Mars mission. After then NASA Administrator, S. O'Keefe, put emphasis on nuclear power, the future of these plans was uncertain for some time, but still indicated that MPD propulsion was considered viable for long interplanetary missions. The major questions in this context center on the power and type of thruster, that is, below or above 100 kW and whether ion or MPD; until SEI, mission analysis by NASA is focused on a 25-kW ion engine for future unmanned JIMO mission to Europa, Callisto and Ganymede [Bordi and Taylor, 2003].

What power and which type of electric thruster to choose are issues that could have benefited from the NASA decision to fund electric thruster research under Project Prometheus [Iannotta, 2004]. An Advanced Electric-Propulsion Technologies Program would have compared MPD and pulsed inductive thrusters, developed at Princeton University and Northrop Grumman, respectively. The first used lithium, while the second thruster used liquid ammonia, a much cheaper propellant. The power was to be about 10 times that for the JIMO mission, that is, of order 200 kW. Thrust conversion efficiencies predicted were about 70% for the Northrop thruster, vs. 60% for the lithium MPD thruster of Princeton University.

Mpd Thruster

Effective exhaust velocity (m/s)

Figure 7.30. The thrust vs. Isp dilemma at fixed power (thrust conversion efficiency assumed to be 0.8) [Andrenucci, 2004].

Effective exhaust velocity (m/s)

Figure 7.30. The thrust vs. Isp dilemma at fixed power (thrust conversion efficiency assumed to be 0.8) [Andrenucci, 2004].

Once cathode life and propellant issues are solved, to be competitive with ion engines in fast interplanetary missions MPD thrusters must show they can handle much more than the 20 kW power of a JIMO mission: for comparison, the maximum laboratory-tested ion engine power known is 1 MW [Fearn, 2003]. Power is a key element of any NEP trajectory, because it determines the thrust and thus mission length. Figure 7.30 [Andrenucci, 2004] is indicative of the trade-off between 7sp and thrust typical of fixed power propulsion (a thrust conversion efficiency = 0.8 has been assumed in this figure). Because power P ~ Isp x F, the curves are hyperbolas, showing the main limitation of electric propulsion (in fact, of any propulsion system) is power available.

In this context it is probably useful to dispel the myth of solar power as a viable energy source for future interplanetary missions. To collect 1 MW by solar cells in LEO one would need 5,330 m2 of cells, the area of a football field, assuming an average 15% cell efficiency over the entire mission, or 3,320 m2 at an optimistic 25%. Furthermore, the solar constant decreases with the square of the distance from the Sun: near Mars the solar constant is 2.2 times lower than near Earth. This means that Mars missions using solar power should be either very long, or use two or three football fields of solar cell arrays. For missions to the outer planets, such as Jupiter, the solar constant decreases so much that a practical 1-MW power source for an MPD thruster cannot be solar. A 100-MW thruster, e.g., for a manned mission, would need half a million square meters of cells. The sheer weight and cost of orbiting such array would be staggering [Koppel et al., 2003].

Although lagging behind ion engines, marrying MPD technology to nuclear power seems ideal for faster interplanetary missions, the more so because lithium is a very good coolant for advanced nuclear reactors [Buffone and Bruno, 2002]. The reactor could generate all the thermal power needed by the MPD thruster. However a 100-MW nuclear reactor is not a significant challenge as is the electric generator: there is hardly any known experience of generating 100 kW of electric power in space, let alone 100 MW. Probably this is the single most critical technology area in all NEP.

In whatever form, NEP, and in particular nuclear MPD propulsion, is a multi-technology field. For maximum performance MPD-based NEP should integrate superconductivity, electric thruster and nuclear reactor in a single electrically and thermodynamically efficient package. Assuming an MPD core mass reduction by two orders of magnitude, made perhaps possible by future superconducting wiring, the MPD accelerator could weigh 10-50 g/kW, resulting in an engine mass of order 1 ton for a 100-MW engine. An important implication is that scaling laws for MPD thrusters should be derived prior to actual engine sizing. Such laws have been derived to miniaturize much smaller self-field MPD, see [Choueiri, 1998; Casali and Bruno, 2008], but have never been tested in extrapolating to higher power (scaling for ion engines can be found in [Fearn, 2004, 2008]).

A final remark on nuclear electric propulsion is that the thermal power rejected by the thermodynamic cycle to produce electricity is of the order of 50% or more. It could be put to good use, for instance to vaporize and perhaps ionize a propellant with low ionization potential such as lithium. This would result in an additional thrust, with a lower Isp, of order O(103) s, simultaneously reducing mass and surface of space radiators. The negative face of this proposal is a much more complex engine. Nevertheless, given their potential higher thrust, mixed ("hybrid") thermal-NEP systems warrant further study, and appear a possible interim solution for interplanetary missions. In fact, still at the conceptual stage, they are the subject of several recent investigations, and for this reason are briefly discussed below.

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  • MARTA
    How close is mpd twave propulsion to working?
    4 months ago

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