Performance measurement and evaluation of an ionic liquid electrospray thruster

An experimental study was performed to evaluate the feasibility of arcjet operation at very low power levels ranging 5 W 35 W. The very low power arcjet was run using nozzles with different material and geometry. Nitrogen gas was used as the propellant. Both propulsive performances and thermal characteristics at the constrictor exit were investigated for conventional nozzles which consist of an assemble of tungsten nozzle parts, and modified nozzles, consisting of an assemble, an insulator and a tungsten anode. In the modified nozzles, a ceramic material or an insulator was used as a part of a constrictor to allow an arc column penetrate further downstream of the constrictor or to maintain the high-voltage mode discharges and to reduce the electrode losses. Stable operations with the specific impulse levels of ~ 270 sec at very low power levels ranging about 5 W 35 W with the constrictor diameter of 0.3 mm or 0.5 mm were confirmed at efficiencies between 30 and 40 percent, except a singular case, glow discharge, in which little effect in propulsive performance was observed with the expense of electrical power. At higher specific powers the specific impulse was relatively independent of mass flow rate. At lower specific powers, the specific impulse for the lower mass flow rate was slightly above that for the higher mass flow rate. The constrictor diameter was found to have significant effect on the thermal characteristics (heavy particle temperature and thermal efficiency) of the internal gas flow and the performance of the device. With partially insulated nozzles the specific impulse and thrust efficiency were significantly increased compared to conventional nozzles.

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 present paper describes performance evaluation of a Dimethyl Ether (DME) propellant arcjet thruster. We have proposed to apply DME to arcjet thrusters owing to its favorable characteristics for space propulsion device such as non toxicity, weak chemical reactivity and storability. Thrust measurement using a vertical-pendulum-type thrust stand showed that thrust was almost monotonously increased with mass flow rate at a given discharge current, and rose with specific power. Specific impulse was also enhanced with specific power whereas thruster efficiency was attenuated. DME propellant presented higher thrust, specific power and thruster efficiency than nitrogen propellant at a given discharge current or specific power. Thrust measurement yielded a thrust of 0.19N, a specific impulse of 330s, thruster efficiency of 0.14 at a DME flow rate of 60-mg/s with the corresponding discharge voltage, power, and specific power of 90V, 2350W, and 39MJ/kg, respectively.

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.

The significant increase of available electric power for the propulsion unit onboard satellites makes high and ultra-high specific impulse thrusters attractive for different applications and missions. Especially thrusters of the radio frequency type show a great potentiality for high Isp applications. Two different rf-thrusters are described: The RIT15LP engine optimized for a moderate input power consumption and the RIT15S designed for a high specific impulse. The main, but only difference between these engines is the extraction grid system. During the development of RIT15LP, principle improvements of rfthrusters were invented. The given description of RIT15LP is focused on these performance improvements whereas the advantages of the rf-principle for (ultra-) high specific impulse applications were pointed out on basis of the RIT15S thruster. The function principle of gridded ion thrusters can be described as a two step process. In the first step the propellant is ionized and after that, in the second step it is accelerated in the electrostatic fields of an ion extraction and acceleration grid system. Both tasks a nearly independent from each other. The acceleration has no effect on the ionisation process and therefore the beam velocity can simply be controlled by the selection of the voltages applied to the grid system. Amoung different types of electric thrusters this is a unique feature of all gridded ion engines. In contrast, the use of other electric propulsion concepts requires often deep changes in the thruster design for an increased beam velocity. For example, it was necessary to invent a two stage concept for SPTs to reach a beam velocity comparable with that of gridded thrusters. Such efforts are not required in case of the later ones. Presently, the beam velocity of gridded ion thrusthers is merely limited by their electric input power consumption and not in a conceptional way. It is important to be aware that the specific input power consumption (the total consumption of electric power Ptot over the thrust F, usually given in>luWatt per Millinewton') of any kind of thruster is a simple function of the beam velocity Vbeam and the electric efficiency of the thrusters rjel:

Nuclear thermal propulsion is an evolving technology that can be utilized for long-distance space travel. This technology yields the advantage of a high thrust and specific impulse, but requires an examination of the potential design adjustments necessary to enhance its feasibility. The development of nuclear thermal propulsion requires a comprehensive understanding of the system-level behavior during transient and steady-state operation. This paper extends our previous research by including the proper handling of turbomachinery with multi-channel thermal hydraulic simulations only for steady-state solutions. The system-level approach presented here enables the treatment of the turbopump components through non-dimensional analysis that eliminates the assumption of constant efficiencies. All the other components within the system (e.g., reflector and core) can be discretized to multiple channels and layers, in which the full thermal hydraulic solution is established. The approach chosen here enables the realistic modeling of the propellant flow within the expander cycle by capturing the pressure losses, mass flow rate splits, and enthalpy gain for various operational conditions. The verification of the package is completed through point comparisons of previous investigations into similar system designs. Furthermore, sensitivity studies are used to benchmark the capabilities of the package and investigate solution variations due to the perturbation of operational conditions and regimes. The sensitivity studies performed here are important to capture variation in flow characteristics (e.g., temperature, pressure, mass flow rates) for different design objectives such as the thrust and specific impulse. This work demonstrates that system-level simulations lacking multi-channel capability and proper turbomachinery treatment may yield higher uncertainties in understanding the engine’s response and characteristics to changing various requirements. This is extremely important when screening the design space of such propulsion systems and when transient simulations are required.

The BPT-4500 and method of tuning the acceleration by focusing a plasma lens is described. Test results show that overall thruster efficiency and plume shape is improved by adjusting the curvature of magnetic field lines in the acceleration zone. Thrust efficiency is improved to 50% from 21% at 150 V and from 58% to 62.5% at 350 V using this technique. At 125 V, the BPT-4500 operates at 45% thrust efficiency and 1030 s specific impulse yielding a thrust/power ratio only 10% less than a 600 s hydrazine arcjet. Introduction Multi-mode Hall thruster capability enables one thruster to serve constant power applications needing the highest possible thrust for orbit transfer and high specific impulse for station keeping. Towards this objective, we have evaluated a GENERAL DYNAMICS (formerly PRIMEX Aerospace Company) approach for tuning the BPT-4500 for the best possible performance over a range of accelerating potentials. This paper describes our initial exploratory efforts to improve thruster efficiency by actively changing magnetic field shape. Thruster Description The BPT-4500 is a 4.5 kW Hall field plasma accelerator based on an earlier 4000 watt device''. These thrusters employ a short discharge to minimize contact with the insulators and pole pieces to shape the curvature of the magnetic field in the ion creation zone. As is well known in Hall type accelerators, the magnetic field lines become nearly equipotential lines defining the acceleration vector of the ions. Also, the technique of adjusting electric potential shape by magnetized electrons in a magnetic field is commonly called a plasma lens. By virtue of the discharge properties in the BPT and shaping these fields, we have attempted to Copyright (c) 2001 by GENERAL DYNAMICS. Published by the American Institute of Aeronautics and Astronautics, Inc., with permission. create a plasma lens with adjustable focus. Plasma Lens Focusing Morozov in the Soviet Union has shown that using two sets of electromagnets to change the gradient of the radial magnetic field can have a significant effect on the behavior of ion acceleration in Hall type discharges. Gavryushin and Kim have shown that the axial gradient of the radial magnetic field regulated by varying the degree of screening of the accelerator channel. These accelerators are characterized by relatively long acceleration lengths in channels that typically have a length to width of 2:1. Insofar as these accelerators had a long discharge, the curvature of the field was not emphasized in the description of this early work. Instead, the researchers discuss primarily the axial gradient of Curvature of ion path changed by Anode focal length Anode potential contour Cathode potential contour Figure 1 Hall Acceleration with Plasma Lens Concept (c)2001 American Institute of Aeronautics & Astronautics or Published with Permission of Author(s) and/or Author(s)' Sponsoring Organization. the magnetic field. The BPT discharge is configured to take place in a very short channel, typically with a length to width of 1:2. As such, physical evidence from a variety of geometry studies suggests that the BPT discharge creates ions in a thin, -1-2 mm layer, occurring approximately 3-4 mm downstream of the anode, schematically illustrated in Fig 1. As such, we hypothesize the BPT discharge embodies Hall acceleration within a plasma lens. Our innovation is to exploit the features of this short discharge by adjusting the focal length of a plasma lens using a trim or focusing coil. Variation of the curvature of the fields in this layer has the effect of changing the initial trajectory of ions. Changes in the shape downstream of the ionization zone are minimized by the fact that magnetic fields are manipulated inside the primary magnetic structure. The BPT-4500 thruster shown in Fig 2 is composed of a primary magnetic structure with four outer coils and an inner coil. The BPT-4500 has a 13 cm mid-diameter and a 2 cm annulus providing slightly more discharge area than the BPT-4000. The schematic cross section of the thruster is shown in Fig. 3. This shows the anode is surrounded by a secondary magnetic structure to add or subtract from the magnetic field radius of curvature in the annulus between insulator rings. Figure 2 BPT-4500 Laboratory Style Thruster Flight Version A flight capable version of the thruster is under development in the Lockheed Martin funded Hall Thruster Propulsion System (HTPS) program. This flight thruster is referred to as the BPT-4000. This thruster has a fixed magnetic field and plasma lens taken from the optimal configuration found here. GENERAL DYNAMICS is on track to deliver a flight qualified thruster, power processor and xenon flow control system in 2003.

Eight magnetoplasmadynamic thruster configurations were tested with argon propellant at power levels between 20 -130 kW to study the effects of geometry and applied magnetic field strength on thruster performance. The discharge voltage, thrust efficiency, and specific impulse increased monotonically with increasing applied-field strength for all geometries. The highest measured performance was 23% thrust efficiency at 2300s specific impulse at a power of 113 kW. Both cathode and anode radii fundamentally influenced the efficiencyspecific impulse relationship, whereas their lengths influenced only the magnitude of the applied magnetic field required to reach a given performance level. At a given specific impulse, large electrode radii resulted in lower efficiencies, although the rate at which efficiency increased with applied-field strength was higher with larger anode radii. Anode power deposition was the largest efficiency loss, and represented between 50-80% of the input power. The fraction of the input power deposited into the anode decreased with increasing applied-field strength and anode radius.

Microplasma thruster of electrothermal type has been investigated with feed or propellant gases of He and H2. The thruster consisted of an azimuthally symmetric microwave-excited microplasma source 1.5 mm in diameter and 10 mm long with a rod antenna on axis, and a converging-diverging micronozzle 1 mm long with a throat 0.2 mm in diameter. Surface wave-excited plasmas were established by 4.0-GHz microwaves at powers of ≤ 6 W, with the source pressure in the range 0.5–12 kPa at flow rates of 2–70 sccm. The microplasma generation, micronozzle flow, and thrust performance with He were numerically analyzed by using a two-dimensional fluid model, coupled with an electromagnetic model for microwaves interacting with plasmas in the source region. In experiments, the plasma electron density and gas temperature in the microplasma source were measured at around the top of the microwave antenna, or just upstream of the micronozzle inlet, by optical emission spectroscopy with a small amount of additive gases of H2 and N2. In the case of He propellant, the Stark broadening of H Balmer-β line and the vibronic spectrum of N2 2nd positive (0, 2) band indicated that the electron density was in the range (2-5)×1019m-3 and the gas or rotational temperature was in the range 600–700 K. The thrust performance was also measured by using a target-type microthrust stand, giving a thrust in the range 0.04–0.51 mN, a specific impulse in the range 150–270 s, and a thrust efficiency in the range 2%–12%. These experimental results were consistent with those of numerical analysis, depending on microwave power and gas flow rate. Similar plasma characteristics and thrust performance were obtained with H2 propellant, where the specific impulse of ≤ 450 s was more than 1.5 times higher than that with He, owing to a difference in mass between He and H2. A comparison with previous studies with Ar propellant [T. Takahashi et al., Phys. Plasmas 16, 083505 (2009)] indicated that in the presence as well as absence of plasma discharge, the specific impulse was enhanced by more than 3–5 times with light-mass propellants He and H2 as has been known for large-scale propulsion systems. Thus, it follows that in the microplasma thruster of electrothermal type, the high diffusivity and thermal conductivity of He and H2 in the microplasma source of high surface-to-volume ratios do not lead to a deterioration of the thrust performance, primarily owing to a more significant thermal energy gain due to elastic collisions between electrons and heavy particles in He and H2.

Nitrous oxide is a monopropellant with potential for use in propulsion systems for orbit correction, positioning and attitude control of satellites. This paper presents experimental results with a 2N thruster prototype employing nitrous oxide gas decomposed by a rhodium oxide catalyst supported on alumina. Initially, the thruster design, the test bench and test procedures are described. Then the main experimental results, including thrust curves, mass flow rates, pressures and temperatures along the thruster are presented for pulsed operation. At the end specific impulses, characteristic velocities and thruster efficiencies are determined, thus indicating a catalytic decomposition efficiency of about 88%.

This paper presents an optical method and system for contactless measurement of the mass flow rate of melts by digital cameras. The proposed method is based on reconstruction of melt stream geometry and flow velocity calculation by cross correlation, and is very cost-effective due its modest hardware requirements. Using a laboratory test rig with a small inductive melting pot and reference mass flow rate measurement by weighing, the proposed method was demonstrated to have an excellent dynamic response (0.1s order of magnitude) while producing deviations from the reference of about 5% in the steady-state flow regime. Similar results were obtained in an industrial stone wool production line for two repeated measurements. Our method was tested in a wide range of melt flow rates (0.05-1.2kg/s) and did not require very fast cameras (120frames per second would be sufficient for most industrial applications).

Hollow cathodes have been proposed as reaction control thrusters for all-electric and small spacecraft. This paper makes an assessment of modified T5 and T6 hollow cathodes for use as millinewton range thrusters. The influence of terminal parameters such as discharge current, mass flow rate, and cathode/anode geometry on thrust production is discussed. The data indicate that the T5 cathode may be able to develop specific impulses in the range of 150–250 s with argon at reasonable thrust efficiencies of up to 14%. As such, hollow-cathode thrusters may meet some limited applications. However, it is unlikely that this type of thruster could be improved significantly or could compete with similar thrusters in the same operating range. High specific impulse operation is also shown to develop large discharge voltage fluctuations, which may significantly limit the lifetime of such a device.

The piloted deep space exploration missions envisioned by the NASA Human Exploration and Development of Space initiative will require the development of advanced electric propulsion systems capable of providing high specific impulse for extended periods of operation. Current electric propulsion thrusters are well suited for orbit maneuvering and robotic exploration, but at present they cannot provide the combination of specific impulse, lifetime, and efficiency required for piloted deep space missions. The theta-pinch thruster concept is a high power plasma rocket that can potentially meet these future deep space propulsion requirements. Efficient, partial preionization of a gas propellant followed by rapid adiabatic magnetic compression is used to generate, heat, and expel a high velocity, high density plasma to provide thrust. The concept is electrodeless, and radial compression of the plasma by the magnetic field of the discharge coil mitigates material erosion to ensure long thruster life. Because the heated plasma is free to flow along axial magnetic field lines during compression, a magnetic mirror located at the entrance to the discharge chamber is used to direct the plasma flow out of the thruster. The thrust and specific impulse of the engine can be tailored for a given mission scenario through the selection of propellant species, mass flow rate, compression coil discharge current, and/or the compression coil repetition rate, making this a unique and versatile electric propulsion system.

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.

A type of pulsed arcjet is examined which operates with gaseous helium propellant at powers from 100 to 1500 W and pulse rats from 360 to 6000 pulses per second. During the pulse, peak power is 50 to 250 kW generating chamber pressures of 10 to 50 atmospheres. The high operating pressure substantially reduces ionization in the 2.5 mm diameter x 12.5 mm long capillary and lowers frozen flow losses in the nozzle. The thruster thermal efficiency, measured calorimetrically, is 43 percent at 440 W and 8 mg/sec helium mass flow rate. Thruster performance trends are predicted by a time-dependent lumped-parameter model which includes heat exchange between the propellant and the wall. The model substantially underpredicts the experimental thermal efficiency values. Alternative thruster configurations yielding significant improvements in thrust efficiency and specific impulse are discussed.