RIT15S - A radio frequency ion engine for high specific impulse operation

Since nuclear fusion was first discovered in the 1930s, researchers have shown strong interest on this future clean energy source. The tokamak and rotamak devices are designed to achieve fusion <sup xmlns:mml="http://www.w3.org/1998/Math/MathML" xmlns:xlink="http://www.w3.org/1999/xlink">1</sup> . However, they both require immense efforts to be realized and the rotamak is less likely to achieve fusion than the tokamak. On the other hand, the development of spacecraft thrusters has seen a tremendous growth over just a few decades. Although traditional chemical propulsion is still widely employed in space flights, it cannot meet the more challenging requirements of deep space travel due to its low specific impulse. The advent of electric thrusters has helped humans travel further from Earth and has the potential to be developed for interstellar flights. In addition, electrodeless thrusters do not suffer from electrode corrosion, as with presently operational electric thrusters. This work is attempting to utilize rotamak technology in a plasma engine to accomplish a novel type of thruster with high specific impulse <sup xmlns:mml="http://www.w3.org/1998/Math/MathML" xmlns:xlink="http://www.w3.org/1999/xlink">2</sup> . We have assessed the potential of a rotamak-type system as a small electrodeless thruster. Herein, we present the preliminary works done for the development of a novel rotating magnetic field driven plasma engine for space propulsion. The schematic and simulation of a new type of thruster will be presented. This work will also discuss its main principles and advantages.

The search of ways to decrease the per-unit cost of space vehicles injection into high operational orbits and to increase their power-to-weight ratio at the present time is centered on the promising propulsion systems with high specific impulse and with high specific electric power. Such system makes it possible to decrease significantly the propellant mass, as well as on the promising power systems. While SV injects from LEO to final operational orbit, the SPPS must heat hydrogen to temperatures required by specific impulse and generate auxiliary electric power. This paper deals with a solar power and propulsion system with a thermionic energy conversion. The SPPS performance data are given.

Ion thrusters utilize high‐voltage grids to electrostatically accelerate ions to high velocity to produce thrust. Efficient ionization of the propellant carried onboard the spacecraft is achieved by optimized hollow cathode, RF or microwave discharge chambers, and electrons from an external hollow cathode are injected into the ion beam to provide charge neutralization in order to avoid spacecraft charging. The ion thruster propulsion subsystem also contains a power processor, a propellant control system and a gimbal mechanism. The high exhaust velocities produced in present flight thrusters by acceleration of the ions to energies in the range of 1–2 keV produces a specific impulse that exceeds 3000 s, and energies of over 10 keV have been demonstrated. The low thrust produced by these devices of typically a fraction of a Newton requires that thruster lifetimes measured in years are needed for most deep space prime propulsion and earth‐orbiting station‐keeping applications. Ion thrusters feature the highest efficiency (60 to &lt; 80%) and very high specific impulse (2000 to over 10000 s) compared to other thruster types.

A 2-cm electron cyclotron resonance ion source has the advantages of long life and high specific impulse, which can meet the requirements for space gravitational waves detection. In the experiment on finding the lower limit of thrust, it is found that when the ion source operates under the extreme condition of 0.5-W microwave power and 0.1-sccm gas flow rate, increasing the voltages of grid system excessively may cause flameout. The plasma discharge level is controlled by the gas supply, microwave, and power supply system, and their small disturbances will make experimental results different, thus the flameout of the ion source appears randomly and transiently. Besides, it is difficult to observe the flameout phenomenon experimentally, because the probe diagnosis has big interference to low-density plasma, and the optical diagnosis is blocked by the grid system. Therefore, the integrative simulation with the full particle-in-cell method is used to simulate the operating process of the ion source, whose calculation range includes the discharge chamber, grid system, and plume. Through simulating the processes of plasma discharge and ion beam extraction continuously in space and time, the flameout phenomenon can be reproduced artificially after increasing the voltages of grid system. The simulation results show that the ambipolar diffusion between the antenna and discharge chamber is the fundamental reason for the flameout of the ion source. In the circuit, the antenna does not touch the discharge chamber but for bulk plasma, which makes its surface gradually accumulate charges until it reaches the floating potential. Because the increase of the voltage of antenna lags behind that of grid system, a strong electric field will appear between the antenna and chamber. Then, electrons and ions respectively move toward the chamber and antenna, the ambipolar diffusion helps the antenna reach the floating potential rapidly. When the plasma density inside the chamber is low, the ambipolar diffusion will cause flameout. In order to avoid the flameout of the ion source in such an extreme situation, an improvement measure that the voltage of antenna equals the voltage of chamber is proposed, which is verified by the integrative simulation. The study on the flameout phenomenon will provide a theoretical basis for the design and application of the ion source, which can help the ion source operate safely to meet the requirements for space gravitational wave detection.

High specific impulse ion propulsion is of concern as primary propulsion system for various interplanetary explorers. A high specific impulse microwave discharge ion engine has been developed in Japan Aerospace Exploration Agency (JAXA) since 2003. The aim of this development is achieving a specific impulse of 10,000sec with an ion acceleration voltage of 15kV. New electro-static grid made of carbon-carbon composite material and various high voltage insulating components such as a DC block and a propellant isolator were developed. The authors fabricated a prototype model of the ion engine, and successfully operated it. The paper describes the outline and the experimental result of demonstration test of the high specific impulse microwave discharge ion engine

The electron cyclotron resonance ion engine, “μ10,” has a long life and high reliability because of electrodeless plasma generation in both the ion generator and the neutralizer. Four μ10, each generating a thrust of 8 mN, specific impulse of 3,200 seconds, and consuming 350 W of electric power, propel the “HAYABUSA” asteroid explorer that was launched on May 2003. After vacuum exposure and several runs of baking to reduce residual gas, the ion engine system established continuous acceleration. Delta-V Earth Gravity Assist, a new orbit change scheme that uses electric propulsion with a high specific impulse was applied to change from a terrestrial orbit to an asteroid-based orbit. In 2005, HAYABUSA, using solar electric propulsion, managed to successfully cover the distance between 0.86 AU and 1.7 AU in the solar system, as well as rendezvous with, land on, and lift off from the asteroid Itokawa. During the 3-year flight, the ion engine system generated a delta-V of 1,400 m/s while consuming 22 kg of xenon propellant and operating for 25,900 hours.

Performance measurement and evaluation of an ionic liquid electrospray thruster

Through this work we show the development of small hybrid rocket thrusters intended for attitude control systems of small satellites and space-crafts. These thrusters have the advantage of being controllable and having higher specific impulse than the traditional solid fuel thrusters or hydrazine based thrusters. An internal ballistic model of the small hybrid thrusters is shown together with numerical results for the thrust-curves and other relevant parameters. The internal ballistic model is shown to be influenced by the regression speed and the need for higher regression speed fuel-oxidizer pairs is, hence, shown. In the present study we also show how to couple the hybrid combustion process with the unsteady flow. Moderately high specific impulses can be obtained by using polypropylene as fuel and NO2 as oxidizer. Detailed discussion is done on the construction of the thrusters as well as the engineering solutions for the combustion chamber, injector and nozzle. Several experimental thrust-curves are shown together with the estimation of total impulse and specific impulse.

MEMS thruster with blended propellant was fabricated and experimentally tested in order to improve specific impulse of micro scale monopropellant thruster and to improve stability of thrust to be better. 90 wt. % H2O2 blended with 25 O/F ratio ethanol was used as propellant of thruster and platinum on alumina support was used as catalyst for decomposition of propellant. Thruster was made by five layers of photosensitive glasses. Four layers were integrated by thermal bonding method and catalyst was directly inserted into chamber before UV bonding process for the last layer bonding. Results of experimental tests showed ethanol blended hydrogen peroxide had higher specific impulse than unblended hydrogen peroxide. Expected improvement of thrust stability due to the blended propellant was found only in the transient state of thrust. Also, unlike the thrust instability of vertical type thruster of previous research, improvement of thrust stability was found owe to horizontal type thruster pattern on glass, despite aspect ratio limitation of glass fabrication with wet etching process. During the experimental test, combustion phenomena of ethanol with decomposed hydrogen peroxide were observed through glass layer and it made fracture on structure of thruster.

The Space Exploration Initiative requires the development of nuclear thermal and nuclear electric technologies for space propulsion for future Luna and Mars missions. Sandia National Laboratories has proposed a new nuclear thermal propulsion concept that uses fission fragments to directly heat the propellant up to 1000 K or higher above the material temperatures. The concept offers significant advantages over traditional solid‐core nuclear rocket concepts because of higher propellent exit temperatures, while at the same time providing for more reliable operation due to lower structure temperatures and lower power densities. The reactor can be operated in either a steady‐state or pulsed mode. The steady‐state mode provides a high thrust and relatively high specific impulse, as compared to other nuclear thermal concepts. The pulsed mode requires an auxillary radiator for cooling, but has the possibility of achieving very high specific impulses and thrust scaleable to the radiator size. The propellant temperatures are limited only by thermal radiation and transient heat conduction back to the substrate walls.

The concept of remotely heating a rocket propellant with a high intensity radiant energy flux is especially attractive due to its high specific impulse and large payload mass capabilities. In this paper, a radiation receiver-thruster which is especially suited to the particular thermodynamic and spectral characteristics of highly concentrated solar energy is proposed. In this receiver, radiant energy is volumetrically absorbed within a hydrogen gas seeded with alkali metal vapors. The alkali atoms and molecules absorb the radiant flux and, subsequently, transfer their internal excitation to hydrogen molecules through collisional quenching. It is shown that such a radiation receiver would outperform a blackbody cavity type receiver in both efficiency and maximum operating temperatures. A solar rocket equipped with such a receiver-thruster would deliver thrusts of several hundred newtons at a specific impulse of 1000 seconds.

The laser-assisted pulsed plasma thruster is considered a promising propulsion system to support the tasks of microsatellites because of its high specific impulse and low volume. Different from the traditional pulsed plasma thruster, laser-assisted pulsed plasma thruster uses the laser to replace the spark plug for ignition, which can avoid ignition failure and remove the side effect of carbon deposition. Both the thrust efficiency and impulse bit are expected to increase after the plasma flow produced by laser ablation is further ionized and accelerated. Since there are a few macro-performance prediction models in laser-assisted pulsed plasma thrusters, this paper develops a model based on the laser ablation model and electromagnetic acceleration model to capture macro-performances of laser-assisted pulsed plasma thrusters. In this model, the initial velocity and mass of plasma flow can be obtained from the ablation model, and the acceleration model is utilized to describe the electromagnetic acceleration process of plasma flow. With this combined model, the discharge current, voltage, impulse bit, specific impulse, and thrust efficiency can be estimated. The deviation between the predicted results and experimental results was less than 10%, verifying the correctness of the developed model. The effects of different parameters on the performance are further investigated with this model.

The paper investigates the utility of how nuclear thermal rocket (NTR) upper stage propulsion could substantially reduce the interplanetary transit time while greatly enhancing the economics of performing the manned Mars mission. au ECCV GTO IMLEO Isp kg kg/kw km km/sec kw LEO MEV mw NTR sec !N Nomenclature astronomical unit Earth capture crew vehicle geosynchronous transfer orbit initial mass in low Earth orbit specific impulse kilogram kilograms per kilowatt kilometer kilometers per second kilowatt low Earth orbit Mars excursion vehicle megawatt nuclear thermal rocket seconds delta V 1. Background Initial responses to the launch cost issue have been to suppose the development of heavy lift launch vehicles that are assumed to have fundamental launch cost savings. In general, most studies have isolated launch vehicle systems from the interplanetary vehicle assembled in low Earth orbit. This paper investigates a potential synergism between applying nuclear thermal rockets as both upper stages of the launch vehicle and as departure propulsion for a Mars vehicle departure stage. II. InterPlanetary Vehicle Description The mission to Mars is a propulsion technology dominated mission. The velocity change requirements (llV or delta V) are very large and very sensitive to trip time constraints. Due to this, there has been a heavy emphasis on high specific impulse (Isp) propulsion systems for performing the manned missions to Mars. The fundamental problem of high Isp systems is that they are generally thermally limited by material temperature limits. This issue is eliminated by electric propulsion systems that use electro-magnetic forces for accelerating a working fluid, but are thrust limited due to practical limitations in power supply mass. More exotic propulsion systems tend to use plasma working fluids that can only be contained by such methods as magnetic nozzles. The Space Exploration Initiative (SET) has a long range goal to perform a manned mission to Mars. Preliminary design goals of an interplanetary vehicle for this mission stress extremely high propulsion requirements to minimize transit times while allowing significant stay times on Mars. Related to this key trade between propulsion performance and trip time is the space transportation economics of 20 or more Earth launches required of the current launch vehicle architecture to place adequate vehicle mass in low Earth orbit to perform a single manned mission to Mars. The large initial mass in low Earth orbit (IMLEO) to perform the manned Mars mission is recognized as a dominant mission cost driver using the current launch infrastructure. IMLEO has been used as a preliminary figure of merit to evaluate various design approaches. * Senior Advanced Propulsion Engineer, Member AIAA Most analysis to date has focused on nuclear thermal rockets (Isp ... 1000 sec) and megawatt class electric propulsion systems (Isp ... 5000+ sec). The vehicle strategies have been to split the mission between a one way only cargo mission to Mars and a fast transfer manned vehicle. Mars rendezvous is essential for mission success, and techniques to slow at Earth return have used Earth capture crew vehicles (ECCV) with very high excess velocities (9+ km/sec) that are ballistically braked. Analyzed here is a very simple single vehicle. It uses an NTR driven, high thrust, Earth departure stage that is dropped once the vehicle has entered a fast Mars transfer ellipse. Once the vehicle has exited the deep Earth gravity well, it relies solely on a 5000 second Isp electric propulsion system for capture at Mars, departure from Mars, and final mission capture at Mars with zero excess velocity (a conservative assumption) The payload to Mars is broken into two elements. The ECCV is assumed to weigh 40,000 kg for a crew of four, and is the crew's interplanetary travel habitation module. The Mars excursion vehicle (MEV) is the element used to drop the crew from Mars orbit to the planet surface, be the crew's planetary habitation module, carry all science equipment, and carry the ascent module for returning to Mars orbit. The MEV has a baseline weight of 84,000 kg for a nominal 30 day planetary stay time and is scaled with the squareroot of normalized stay time. The MEV is not returned to Earth. Consumables are estimated at 10 kg per person per mission day. optimized for minimum l1V by adjusting the up leg perihelion and aphelion while maintaining the constraints of Earth rendezvous and travel time. Figure 2 (with the radius in millions of kilometers) shows a transfer ellipse for the up leg for the 320 day travel time, 80 day Mars stay time, 0 degree mean Mars true anomaly case. Figure 3 shows the down leg of that mission. The mission l1V requirement of 29.1 krn/sec is divided into 4.2 krn/sec for Earth departure (in heliocentric space), 6.0 krn/sec for Mars capture, 7.4 krn/sec for Mars departure, and 11.5 krn/sec for Earth capture. Transfer Ellipse Up. III. Interplanetary Mission Modeling The problem of analyzing the interplanetary mission can be driven by very exacting issues such as actual planetary configurations for particular mission opportunities, the requirements for travel time, and the Mars stay time. The issue that needs near term resolution, however, is the general efficacy of various propulsion and/or vehicle configuration options for the Mars mission in general. To answer these type of questions, it was necessary to generalize the Mars transfer problem. To do so, simplifying assumptions were that the Mars and Earth orbits were coplanar, and that the Earth orbit was circular. The nominal planetary configurations for Mars, Earth, and Venus are shown in Figure 1. The radius units are in astronomical units (Earth radius =1 au) and the angular units are the true anomaly of the orbital positions in degrees (Mars perihelion = 0 degrees). tojMJoC:: ~ •• ~: ••• . -

Complete burning of liquid propellants droplets is very important to avoid combustion instability and to get higher specific impulse from liquid rocket engines. Combustion takes place in gaseous phase and reaction is fast. So, the time consumed by a droplet for complete burning is the time taken by a droplet to get evaporated. An analytical mono component fuel droplet evaporation model is developed and proposed for droplet evaporation in liquid rocket engine based on heat and mass transfer. Effect of the fuel droplet curvature on vapor pressure is incorporated in the model. The model is applied on a small-scale combustion chamber of bi-propellant liquid rocket engine where fuel concentration away from droplet is not zero. A code is generated and applied. The results of the model show that fuel droplet diameter, combustion chamber pressure, chamber diameter, mass flow rate and droplet relative velocity with hot gas has significant effects on the life of droplet and characteristic length. Effects of chamber pressure and diameter are more significant for larger droplet size.

Pulsed electromagnetic plasma accelerators and, more specifically, those that are inductively coupled potentially have high thrust density, high specific impulse, and long thruster lifetime since there is limited contact between the plasma and coil(s). One-dimensional electromechanical models of these thrusters, which are important for identifying important relationships among circuit parameters, coil geometry, and plasma, usually make the same underlying assumption about the mutual inductance between the primary coil and induced plasma current, based on empirical evidence. This approach is inconvenient for arbitrary coil geometries and if trying to make comparisons with directly coupled plasma accelerators, such as coaxial plasma guns. A generalized slug model was developed in which <i xmlns:mml="http://www.w3.org/1998/Math/MathML" xmlns:xlink="http://www.w3.org/1999/xlink">RLC</i> circuit equations are coupled to a mechanical force equation. The acceleration term was made to be thruster and geometry dependent, so that the same model could be used to study both coaxial plasma guns and conical theta pinches. The calculated mutual inductance from the coil model exhibits the same exponential falloff with distance as in previous models and experiments. Efficiency (plasma kinetic energy divided by initial capacitor bank energy) and exhaust velocity were calculated for the conical theta pinch coil over a 5-D parameter space. For the cases with highest efficiencies, peak coaxial plasma accelerator efficiencies were calculated under certain operating constraints. It was found that, while coaxial plasma accelerators are generally more efficient, under certain conditions, the conical theta pinch performance is higher.