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We present a straightforward formulation predicting the characteristic velocity and specific impulse for bipropellant thrusters as a direct function of injection conditions, propellant combination,...

Conventional expressions and definitions describing performance of plasma thrusters, including the thrust, specific impulse, and the thruster efficiency, assume a steady-state plasma flow with a co...

The performance (specific impulse, Isp) of a modular, 150-lbf-class rotating detonation rocket engine (RDRE) was measured with three gaseous fuels (methane, ethane, and ethylene) and gaseous oxygen...

In this study, an electrostatic thruster was newly developed, in which a diverging magnetic field with a cusp around a cathode was applied by using two solenoid coils. The effects of the magnetic field strength on thrust performance under a similar applied-magnetic-field configuration were investigated. Because of its light weight and lower price, argon was used as a propellant. By increasing the magnetic field strength, the thrust efficiency was improved owing to the suppression of the discharge current while an almost constant thrust was maintained. A specific impulse of 3800 s with thrust efficiency greater than 30% was obtained; the corresponding mass-averaged exhaust velocity exceeded the value by full-potential electrostatic acceleration of singly charged ions. From the thrust performance, ion beam current, and ion energy distribution function that were experimentally measured under the representative operation condition, 32% of the ion beam current and 43% of the total thrust were evaluated as the contribution of doubly charged ions.

Cost is one of the biggest obstacles to sending humans to Mars. However, spacecraft costs are typically not estimated until after the preliminary vehicle and mission concepts have been designed. By...

Recent studies identified the addition of metal particles into conventional liquid fuels (jet fuel, gasoline, diesel) as one possible strategy to increase fuel energy density and specific impulse. ...

Numerical simulations are presented that compare the specific impulse of a generic inlet-fueled uniform-compression and nonuniform-compression scramjet over the flight Mach number range of 7–12 wit...

Work is reported on the use of quantitative structural property relationships to estimate the properties of ionic liquids for which no measured property data are available and to classify liquids a...

A helicon plasma thruster has been studied to develop a completely electrodeless electric thruster using high-density helicon plasmas. The proposed helicon plasma thruster involves two processes: the generation of source dense plasma by using a helicon wave, and the additional acceleration of the generated plasma by using the Lorentz force generated by the product of the induced azimuthal current and external radial magnetic field. This additional acceleration method requires additional electrodes or coils, leading to a longer discharge tube. Therefore, it is necessary to find a good configuration that minimizes wall losses within the discharge tube. Here, thrust characteristics such as thrust, thrust-to-power ratio, specific impulse, and thrust efficiencies of argon and xenon gases were studied, using a radio frequency of 7 MHz and an input power less than 3 kW, to optimize the target plasma without employing an additional acceleration method. A helicon plasma source, with electromagnets and permanent ma...

Recent accomplishments related to the performance, application, and analysis of rotating detonation engine technologies are discussed. The pioneering development of optically accessible rotating detonation engines coupled with the application of established diagnostic techniques is enabling a new research direction. In particular, OH* chemiluminescence images of detonations propagating through the annular channel of a rotating detonation engine are reported and appear remarkably similar to computational fluid dynamic results of rotating detonation engines published in the literature. Specific impulse measurements of rotating detonation engines and pulsed detonation engines are shown to be quantitatively similar for engines operating on hydrogen/air and ethylene/air mixtures. The encouraging results indicate that rotating detonation engines are capable of producing thrust with fuel efficiencies that are similar to those associated with pulsed detonation engines while operating on gaseous hydrocarbon fuels....

Orbital flight of CubeSats at altitudes between 150 and 250 km has the potential to enable a new class of scientific, commercial, and defense-related missions. A study is presented to demonstrate the feasibility of extending the orbital lifetime of a CubeSat in a 210 km orbit. Propulsion consists of an electrospray thruster operating at a 2 W, 0.175 mN thrust, and an specific impulse (Isp) of 500 s. The mission consists of two phases. In phase 1, the CubeSat is deployed from a 414 km orbit and uses the thruster to deorbit to the target altitude of 210 km. In phase 2, the propulsion system is used to extend the mission lifetime until propellant is fully expended. A control algorithm based on maintaining a target orbital energy is presented that uses an extended Kalman filter to generate estimates of the orbital dynamic state, which are periodically updated by Global Positioning System measurements. For phase 1, the spacecraft requires 25.21 days to descend from 414 to 210 km, corresponding to a ΔV=96.25 m...

Nanosized aluminum (16% by weight) was added into JP-10 and surfactant (2% by weight) was used to reduce the agglomeration of nanoparticles. Combustion of metalized fuel, as well as pure JP-10, was carried out in a small-scale combustor. The oxygen-to-fuel ratios were set to be 1.7, 1.8, and 1.9, respectively. An additional trial of water injection during combustion was also tested. The pressures at the combustion chamber and nozzle exit, along with the thrust, were measured during the combustion; and intervals of true values of the specific impulse were presented. The results showed that a relatively higher combustion efficiency was achieved with JP-10-based slurry by 3.0 to 9.0% when compared to pure JP-10. However, the specific impulse could be increased only when the combustion-induced heat release was improved enough to overcome the two-phase loss. Depositions at different positions were collected after combustion for deep analysis; and the results of the x-ray diffraction, thermogravimetry analysis,...

Experimental tests are conducted to evaluate the effect of an annular rocket exhaust pattern on the mixing characteristics of a simple ejector. Rocket exhaust is simulated using pure oxygen while air is entrained from the surroundings at static conditions. Results are compared with an equivalent configuration using traditional circular rocket exhaust nozzles. It is shown that an annular rocket exhaust pattern yields a ratio of entrained air to rocket mass flows over 75% higher than an equivalent circular rocket exhaust pattern over a range of rocket chamber total pressures. It is demonstrated that an annular rocket exhaust pattern within a straight ejector duct with a length-to-diameter ratio of 6 is able to produce entrained air to rocket mass flow ratios up to 10% higher than a straight then expanding ejector configuration twice as long employing a circular rocket exhaust pattern. Exit plane measurements of both total pressure and oxygen concentration are also collected at high and low rocket chamber to...

To evaluate the impact of dual-bell nozzles on the payload mass delivered into geostationary transfer orbit by Ariane 5 Evolution Cryotechnique Type A (ECA), detailed studies were conducted. For this purpose, a multitude of Vulcain 2 extension contours were designed. The two variation parameters were the starting point and the inflection angle of the nozzle extension. As the most upstream starting point, the position of the turbine exhaust gas injection was chosen. Geometrical restrictions were imposed by the launch pad ELA 3. Considering these parameters, an analytical and a numerical method were applied to predict the impact of the dual-bell nozzle on the payload mass. The analytical approach yields a correlation between specific impulse, nozzle mass, and payload mass increment. The numerical approach was conducted applying German Aerospace Research Center’s trajectory simulation code Trajectory Optimization and Simulation of Conventional and Advanced Transport Systems. Both calculation procedures yield...

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 performance metrics of a 2-kW-class thruster operated using magnesium propellant were measured and compared to the performance of the same thruster operated using xenon propellant. When operated with magnesium at a 7 A discharge current, the thruster had thrust ranging from 34±0.8 mN at 200 V using 1.8 mg/s of propellant to 39±1.5 mN at 300 V using 1.8 mg/s of propellant. The thrust-to-power ratio ranged from 24±0.5 mN/kW at 200 V to 18±0.7 mN/kW at 300 V. At a 200 V discharge voltage, the specific impulse was 1930±49 s at 23±5.0% efficiency (at 7 A using 1.8 mg/s). At a 300 V discharge voltage, the specific impulse was 2420±130 s at 21±6.4% efficiency (at 5 A using 1.1 mg/s). The performance of the thruster using magnesium propellant was compared to xenon performance at matched molar propellant flow rates: 5 mg/s for xenon and 1.1 mg/s for magnesium. The xenon-fueled thruster produced 76±1.5 mN of thrust, with a specific impulse of 1550±70 s, at an efficiency of 40±2.0% compared to the ...

A rotating detonation engine is experimentally tested with various nozzle configurations for the purpose of measuring the propulsive performance of these devices in terms of thrust and specific impulse. Particular attention is given to comparing different internal nozzle configurations, which include bluff body, aerospike, and choked aerospike arrangements. The nozzle throat exit choke present in the rotating detonation engine exhaust is analyzed to provide insight into the stagnation pressure gain nature of the device.

Plasmonic space propulsion uses solar light focused onto deep-subwavelength nanostructures to excite strong optical forces that accelerate and expel nanoparticle propellant. Simulations predict that light within the solar spectrum can excite asymmetric nanostructures to create plasmonic forces that will accelerate and expel nanoparticles. A peak force of 55 pN/W is predicted for a 50-nm-wide, 400-nm-long nanostructure that resonates at 500 nm. Results for a conceptual design of a plasmonic thruster that has 35 layers, 86 array columns, a multistage length of 5 mm, a 5-cm-diam light focusing lens, and uses 100 nm polystyrene nanoparticles expelled at a rate of 1×106 per second would have a thrust of 250 nN, specific impulse of 10 s, and minimum impulse bit of 50 pN·s.

A new method has been developed for obtaining approximate near-optimal low-thrust interplanetary transfers using solar electric propulsion spacecraft. The trajectory-calculation method consists of analytic curve-fit functions that have been empirically derived from a database of minimum-propellant coplanar transfers between Earth’s orbit and circular target orbits. It is assumed that the transfer begins with a powered arc, followed by a single coast arc, and ends with a second powered arc. The curve-fit functions depend on the mean thrust acceleration and target radius. A scaling technique transforms the Earth-orbit-based solutions in order to provide outward and inward trajectories between circular orbits of arbitrary radius. The inputs to the trajectory tool are the initial and final circular radii, initial spacecraft mass, input power, specific impulse, and thruster efficiency; and the computed (output) values are trip time, transfer angle, coast arc angle, and ΔV. The curve-fitting method is extremely...

Thrust characterization experiments were carried out on a 200 W quad confinement thruster operating on xenon propellant. The thrust was measured using an inverted pendulum-type thrust stand. The anode power, propellant flow rate, and magnetic field were systematically varied in order to understand the impact of each parameter on the performance, which was characterized in terms of specific impulse, thrust, and thrust efficiency. At 200 W of anode power, the specific impulse of the device reached a maximum value of 860 s, and the thrust reached 3.2 mN. The thrust efficiency was modest, and it reached a maximum value of 4.6%. However, the thrust efficiency was found to depend strongly on the strength of the magnetic field. The quad confinement thruster device investigated in this study was limited to a maximum magnetic field of 250 G. Further improvements to the device can likely be achieved by applying much stronger magnetic fields.

Thrust measurements are essential to qualify thrusters for space missions, especially for new technologies. In the framework of the Gaia mission, the 13 newly developed cold gas micronewton thrusters from Thales Alenia Space Italia (TAS-I) were subjected to acceptance tests on the micronewton thrust balance of ONERA in order to measure their thrust noise and specific impulse curve. The balance, based on the principle of a pendulum, has a high-thrust resolution (less than 0.1 μN), with a thrust noise of 0.1 μN/Hz in the range of 0.02–1 Hz. The calibration of the balance is presented, and the postprocessing corrections leading to this performance are detailed. The specific impulse of the 13 cold gas micronewton thrusters flight models has been measured for thrust levels in the range of 0.1–1000 μN, showing a slight increase of Isp from around 50 to 63 s with thrust for all flight models.

The helicon Hall thruster is a two-stage thruster that was developed to investigate whether a radio frequency ionization stage can improve the overall efficiency of a Hall thruster operating at high thrust and low specific impulse. This paper describes an experiment that measured the single-stage and two-stage performance of the helicon Hall thruster operating at 10–25 mg/s anode mass flow rates of xenon at 100–200 V discharge voltages, and also for 6 mg/s of argon at 300 V, and 2.6 mg/s of nitrogen at 200 V. The helicon Hall thruster performance during operation with argon and nitrogen is characterized by low beam divergence efficiency and low propellant utilization efficiency. During two-stage operation, the thrust of the helicon Hall thruster marginally increased with radio frequency power, but the propulsive efficiency and thrust-to-power both decreased with increasing radio frequency power. Probe diagnostics suggest that gains were realized by a slight increase in propellant efficiency, but that t...

Testing of the Variable Specific Impulse Magnetoplasma Rocket VX-200 engine was performed over a wide throttle range in a 150 m3 vacuum chamber with sufficient pumping to permit exhaust plume measurements at argon background pressures less than 1×10−3 Pa (1×10−5 torr) during firings, ensuring charge-exchange mean free paths longer than the vacuum chamber. Measurements of plasma flux, radio frequency power, propellant flow rate, and ion kinetic energy were used to determine the ionization cost of argon and krypton in the helicon discharge. Experimental data on ionization cost, ion fraction, exhaust plume expansion angle, thruster efficiency, and thrust are presented that characterize the VX-200 engine performance over a throttling range from 15 to 200 kW radio frequency power. A semiempirical model of the thruster efficiency as a function of specific impulse indicates an ion cyclotron heating efficiency of 85±7%. Operation at a total radio frequency coupled power level of 200 kW yields a thruster effici...

This paper describes measurements of performance characteristics of micro-cathode arc thruster. In particular, influence of an applied magnetic field B on the impulse bit, specific impulse, and the mass consumption is investigated. The micro-cathode arc thruster was operated in a quasi-steady-state mode to ensure that a measurable thrust stand deflection was produced. It was found that the impulse bit increases from about 0.1 μN·s in the case without a magnetic field up to about 1.1 μN·s in the case of a magnetic field of about 0.3 T. A magnetic field leads to an increase of the mass consumption from about 10 μg/C at zero magnetic field up to about 50 μg/C at a magnetic field of about 0.3 T. Average ion velocity is measured to be about 18 km/s (at B=0 T) and about 31 km/s (at B=0.3 T). Maximum efficiency of the thruster is estimated to be about 7%.

mixtureexhibitedalinearburningrateof4. 8c m∕satapressureof10.7MPaandapressureexponentof0.79.Three motors of internal diameters in the range of 1.91–7.62 cm were studied. Grain configuration, nozzle throat diameter, and igniter strength were varied. The propellants were successfully ignited and combusted in each laboratory-scale motor, generating thrust levels above 992 N in the 7.62-cm-diam motor with a center-perforated grain configuration (7.62 cm length) and an expansion ratio of 10. For the 7.62 cm motor, combustion efficiency was 69%, whereas the specific impulse efficiency was 64%.Increased combustionefficiency and improvedease of ignitionwere observedat higher chamber pressures (greater than 8 MPa).

Highly reactive nanothermites prepared by mixing bismuth trioxide or cupric oxide nanoparticles with aluminum nanoparticles were evaluated as solid propellants for small-scale propulsion applications. Miniaturized engines were fabricated from steel in three-piece configurations without a converging/diverging nozzle. Bismuth trioxide-aluminum generated 46.1 N average thrusts for 1.7 ms durations with a specific impulse of 41.4 s. Cupric oxide-aluminum generated 4.6 N thrusts for 5.1 ms durations with a specific impulse of 20.2 s. Convective and conductive reaction regimes were identified as functions of bulk packing density and confinement geometry. Average thrusts and burning durations differed by greater than an order of magnitude for equivalent nanothermites dependent on the reaction regime. Adding small amounts of nitrocellulose to the nanothermites increased specific and volumetric impulses to maximum values of 59.4 s and 2.3 mN·s/mm3 while controllably reducing average thrusts and prolonging burning...

This paper proposes the use of micron-sized aluminum (pyral, average particle size 3.66 ?m) with a higher specific surface area as a good candidate to enhance the burn rate of the composite propellant. Experiments were performed in the pressure range of 10 to 70 bar in a window bomb for measuring the burn rate. Comparison of these burn rate results with those obtained using micron- and nanosized aluminum found that the performance of pyral was in between that of micron- and nanosized aluminum. The reason for the high burn rates observed with pyral is due to the flake like appearance of pyral with a large specific surface area. It is argued that, if the specific surface area is large, then the thickness becomes the characteristic length scale. This ensures the heat release from the aluminum combustion to occur closer to the propellant surface as the thickness of pyral is in nanometers. Both the x-ray diffraction and heat of formation analyses indicated that pyral had higher purity than nanoaluminum, which has an implication to the specific impulse of the propellant. In addition to this, it shows that the micron-sized catalyst is effective with pyral, whereas it is ineffective with nanoaluminum. This study also reports the mechanical properties of the propellant containing pyral.

The evaluation of an ethanol blended hydrogen peroxide monopropellant thruster was carried out. The specific impulse of hydrogen peroxide was increased by blending it with ethanol. Ethanol was selected because it presented no significant problem in terms of storability. An oxidizer-to-fuel ratio of 50 was selected considering the thermal characteristics of the material used and the higher specific impulse that was obtained in comparison to the theoretical specific impulse of hydrogen peroxide. For the thruster material, 316L stainless steel was chosen based on compatibility test results and other considerations. Barium hexaaluminate was used as a catalyst support because the chamber temperature increased owing to the combustion of ethanol. The surface area of barium hexaaluminate was approximately six times higher than that of alumina after heating at 1200°C. Platinum and manganese oxide with lead oxide catalysts were evaluated on a 10 N class thruster and the catalyst capacity, specific impulse, and effi...

The aim of this work is to provide an upper estimate of the theoretical maximum efficiency and specific impulse of the external burning scramjet. Contrary to the classical Carnot and Brayton cycles for the heat and gas-turbine engine and the known ideal cycle for the ramjet, in this case there is not a reversible cycle to provide a reference. The reason for this is that the heat release to the high-speed flow is accompanied by inherent total pressure losses, and thus the process results are irreversible. Subsequently, to estimate the maximum efficiency an idealized optimization problem inevitably must be analyzed that also takes into account these inherent losses. This optimization problem was solved 1) in the linear approximation, valid for thin bodies and small amounts of applied heat, and 2) in the nonlinear formulation but in the original framework of heat release in the jet stream with a “zero mass flux” condition. The results obtained in the two cases are comparable for small heat release, but they ...

The performance of a radio-frequency plasma thruster is evaluated using a displacement-type, inverted-pendulum thrust stand in the Large Vacuum Test Facility at the University of Michigan. A radio-frequency generator supplies up to 2000 W to the radio-frequency plasma thruster at a fixed frequency of 13.56 MHz. The matching network is placed inside the vacuum chamber at the thruster, and the radio-frequency power is measured at the matching network input port with a dual-directional coupler. A Faraday probe measures the current density in the far-field exhaust of the thruster, a retarding potential analyzer characterizes the ion voltage distribution, and a Langmuir probe measures the plasma potential. The radio-frequency plasma thruster is operated with argon propellant at mass flow rates of 2.4–7.6 mg/s, at corresponding corrected facility pressures of 1.5×10−6–4.6×10−6 torr. The maximum thrust observed is 10.8 mN, and the maximum specific impulse is 303 s. Thrust efficiency is below 1% at all conditio...

There is interest in the use of a helicon plasma source in propulsive applications as both an ion source and a thruster. Development of a helicon thruster requires a performance baseline as a basis for future optimization and modification. For the first time, the thrust of a helicon plasma source is measured using a null-type inverted pendulum thrust stand at an operating pressure of 2×10−5 torr through the operating range of 215–840 W RF power, 11.9 and 13.56 MHz RF frequency, 150–450 G magnetic field strength, and 1.5–4.5 mg/s propellant flow rate for argon. Maximum thrust is found to be 6.3 mN at a specific impulse of 140 s and a maximum specific impulse of 380 s at 5.6 mN. Thrust efficiency is less than 1.4% and demonstrates very-low-power coupling to ion acceleration.

To study the working characteristics of metal/water reaction fuel ramjets the influences of different parameters on engine performance were obtained by experiments with one and two water injections under different first water/fuel ratios and burning rates. The following conclusions have been made: the magnesium-based metal fuel could combust steadily with water under a suitable water/fuel ratio, and the combustion efficiency and ejection efficiency rose at first and then declined with an increasing first water/fuel ratio, which demonstrates the existence of an optimum water/fuel ratio; the time of the pressure climbing period could be reduced, and the engine performance could be improved by the increasing burning rate; and a specific impulse of the engine could be improved by adding secondary water injection, and a specific impulse of 3500 N·s/kg was reached in the experiment on the ground.

This paper describes a simple semi-analytic model for mass-optimized interplanetary solar electric low-thrust trajectories. A description is given of a model that accurately and quickly determines the performance of circular-coplanar low-thrust transfers with a series of simple empirical and physics-based relationships that can be implemented easily in a spreadsheet. The model takes flight time, departure and arrival velocity, initial power, initial mass, and propulsion-system efficiency as inputs and produces the optimum specific impulse, Δv, final mass, and burn time that correspond to the mass-optimum trajectory as outputs. The development methodology is described, governing variables and fundamental relationships are identified, and a model is presented that efficiently calculates these parameters for a wide range of low-thrust trajectories. Models are presented for Earth-Jupiter/Trojan asteroid, Earth-Mars, Mars-Earth, Earth-Venus, and Earth-Main-Belt asteroid transfers using solar electric propulsio...

The area of application of the Hall effect thrusters is constantly expanding, both in the direction of reduced power and toward an increase of power and specific impulse. The two main problems are studied in this paper. The first is to examine the possibility of providing the similarity of plasma parameters in Hall effect thrusters, and the second is to estimate the best achievable output characteristics when scaling Hall effect thrusters. It is shown that a strict similarity of plasma parameters is possible only for the same type of propellant at the same discharge voltage at invariance of at least two dimensionless scaling parameters. Based on the analysis of the similarity criteria and available experimental data, a speculative assertion is made, that in the optimal mode, all the geometrical dimensions of the Hall thruster should be changed in the same proportion. For this type of scaling, the semi-empirical expression for the mass utilization efficiency is obtained under the invariance of the heat flu...

DOI: 10.2514/1.B34488 This work presents a physical model of the combustion behavior of a liquid oxidizer droplet with adjacent solid fuel. It is the first modeling attempt related to a proposed new potential class of liquid-oxidizer-enhanced solid propellants, where liquid oxidizer units/capsules are embedded in a solid fuel/propellant, revealing a theoretical specific impulse increase of up to 12%. A one-dimensional mathematical model is formulated, which describes the main characteristics of the microscale flamelet formed by a volatile (typically polymeric) fuel component and an adjacent endothermically evaporating oxidizer droplet. The model predicts fuel and oxidizer surface temperatures, fluxes, and flame height as functions of operating pressure and droplet size. The transient nature of the combustion processisemphasized,revealingacombustioncycleattheindividualdropletscale,withpressure-dependentaverage burning rates typical to solid propellants.

The performance of a nominal 200 W Hall effect thruster fueled by iodine vapor was evaluated. The system included a laboratory propellant feed system, a flight-model Hall thruster, and breadboard power processing unit. Operation of the Hall thruster with iodine vapor was stable on both short and long time scales, enabling measurements of thruster performance across a broad range of operation conditions. Performance was found to be comparable with xenon. At 200 W, thrust is 13 mN, anode specific impulse is 1500 s, and anode efficiency is 48%. Plume-current measurements indicate a profile typical of xenon Hall thrusters, while E B probe measurements indicate the presence of ionic dimers.

A numerical investigation of three-dimensional flow in a converging–diverging microelectromechanical-systemsbased supersonic nozzle design is reported. Based on microfabrication techniques, the nozzle geometry is a twodimensional pattern etched into a substrate to a specified depth, thus resulting in a three-dimensional ductlike geometry and flowfield. Owing to the low Reynolds numbers associated with the microscale, viscous effects result in substantial subsonic layers that develop along expander walls and reduce nozzle efficiency. In this study, the threedimensional flowfield is simulated with a continuummodel and decomposed hydrogen peroxide as the working fluid monopropellant. Numerical simulations are performed over a range of nozzle expander half-angles (15–45 ), for several nozzle depths (25–400 m), and for varying throat Reynolds numbers (15–800). Simulation results are analyzed to determine subsonic layer growth rates and to delineate their impact on micronozzle thrust and performance. Specific impulse efficiencies are compared with quasi-one-dimensional theory, previous twodimensional simulations, and data from the micronozzle literature, including direct simulation Monte Carlo and experimental studies. It is found that three-dimensional impulse efficiencies can be much less than corresponding two-dimensional model results owing to the presence of viscous, subsonic layers on the additional walls in a threedimensional device. Themicronozzle depth is a critical parameter that impacts thrust and efficiency. For sufficiently shallow nozzles, it is possible for the subsonic layers to merge and decelerate the flow to subsonic conditions, thus degrading nozzle performance.

Microthrusters with conventional converging–diverging and orifice nozzles experience significant performance degradation when operating at lowReynolds numbers. The direct simulationMonte Carlo method has been applied to design amicronozzle with improved performance through simulation. Thrust performance calculations using this method demonstrated that the coaxial micronozzles can achieve millinewton thrust levels with specific impulses on the order of 45 s using a cold gas expansion of argon. Improved coaxial micronozzle designs are proposed using centerbody geometries to exploit pressure thrust and show a large increase in specific impulse compared with a baseline converging nozzle operating at similar conditions and at low Reynolds numbers. Micronozzle performance for different geometries are also compared for various Reynolds numbers.

H IGH-POWER electric propulsion thrusters can reduce propellant mass for heavy-payload orbit-raising missions and cargo missions to the moon and near-Earth asteroids, and they can reduce the trip time of robotic and piloted planetary missions [1–4]. The Variable Specific Impulse Magnetoplasma Rocket (VASIMR®) VX-200 engine is an electric propulsion system capable of processing power densities on the order of 6 MW=m with a high specific impulse (4000 to 6000 s) and an inherent capability to vary the thrust and specific impulse at a constant power. The potential for a long lifetime is due primarily to the radial magnetic confinement of both ions and electrons in a quasi-neutral flowing plasma stream, which acts to significantly reduce the plasma impingement on the walls of the rocket core. High-temperature ceramic plasma-facing surfaces handle the thermal radiation: the principal heat transfer mechanism from the discharge. The rocket uses a helicon plasma source [5,6] for efficient plasma production in the first stage. This plasma is energized further by an ion cyclotron heating (ICH) RF stage that uses left-hand polarized slow-mode waves launched from the high field side of the ion cyclotron resonance. Useful thrust is produced as the plasma accelerates in an expanding magnetic field: a process described by conservation of the first adiabatic invariant as the magnetic field strength decreases in the exhaust region of the VASIMR [7–9]. End-to-end testing of the VX-200 engine has been undertaken with an optimum magnetic field and in a vacuum facility with sufficient volume and pumping to permit exhaust plumemeasurements at low background pressures. Experimental results are presented with the VX-200 engine installed in a 150 m vacuum chamber with an operating pressure below 1 10 2 Pa (1 10 4 torr), and with an exhaust plume diagnostic measurement range of 5 m in the axial direction and 1 m in the radial directions. Measurements of plasma flux, RF power, and neutral argon gas flow rate, combined with knowledge of the kinetic energy of the ions leaving the VX-200 engine, are used to determine the ionization cost of the argon plasma. A plasmamomentum flux sensor (PMFS)measures the force density as a function of radial and axial positions in the exhaust plume. New experimental data on ionization cost, exhaust plume expansion angle, thruster efficiency, and total force are presented that characterize the VX-200 engine performance above 100 kW. A semiempirical model of the thruster efficiency as a function of specific impulse has been developed to fit the experimental data, and an extrapolation to 200 kWdc input power yields a thruster efficiency of 61% at a specific impulse of 4800 s.

T HE miniaturization of space systems, such as microsatellites, has become an important development trend. Using a cluster of microspacecraft with a constellational architecture to replace a traditional spacecraft can greatly reduce the costs of production and launch, increase flexibility, and disperse the risks of a mission. Miniaturized spacecraft are classified based on mass, power, and dimensions. Spacecraft with amass of less than 20 kg are classified as class I microspacecraft [1] and require millinewton-level thrusts for spacecraft control. For microspacecraft, the onboard thrusters must be extremely small and lightweight; microelectromechanical systems (MEMS) are thus employed in microthruster design and fabrication [2]. A number ofmicropropulsion systems have been proposed.Micro cold-gas systems have been constructed and used in practice [3,4]; however, a rather low specific impulse (60–80 s) limits their usage. Micro electric-type thrusters provide a high specific impulse [5], but the requirement of high power for operation limits them to larger spacecraft. Miniaturized solid-propellant thrusters have a simple structure and a high specific impulse [6], but their relatively high thrust level (10–10 mN) and single use restrict their application. Monopropellant thrusters are appropriate for miniaturization due to their simplicity and acceptable working temperatures [7]. The catalytic reaction of monopropellant systems mitigates the constraints of radical quenching and mixing prohibition found in the microcombustion of bipropellant systems. Although hydrogen peroxide/silver systems have been tested and effectively reacted in microreactors, hydrazine is considered a better monopropellant for actual microthruster design and operation [7]. A millinewton hydrazine (N2H4) monopropellant thruster is presented in this work. MEMS technologies are employed in the design. The design considerations and component fabrication are discussed. The vacuum thrust of the designed thruster was measured and its propulsive performance was analyzed.

Various sizing methodologies are currently available to get a first estimate of the required Hall effect thruster dimensions for a given input power and a corresponding thrust and specific impulse level. In this work, a semiempirical approach to compute the three characteristic thruster dimensions, i.e., the channel length, the channel width, and the channel mean diameter, is introduced. The magnetic field strength is also considered. The determination of the scaling relations is based onanalytical relationships deduced from the physicalmechanisms that govern the properties of a Hall thruster discharge. A set of simplifying assumptions naturally specifies the validity domain of the relationships. The existence of a critical propellant atom density inside the channel, which warrants a high-efficiency thruster operation, is revealed and commented. The proportionality coefficients of the scaling relations are assessed by way of a vast database that comprises 33 single-stage Hall effect thrusters covering a power range from 10W up to 50 kW. The sizing method is employed to access the geometry and the operating parameters for a 20 kW-class Hall thruster operating with xenon. Results obtained with two different series of simplifying assumptions are compared. The first set forms a very restrictive frame. The second set offers a more realistic description of the physics at work as the electron temperature, the energy losses andmultiply charged ion species are taken into account.

Analytical and numerical investigations of the performance of a series of potential dual-mode propulsion systems using ionic liquids are presented. A comparison of the predicted specific impulse of ionic liquids with hydrazine and unsymmetrical dimethylhydrazine shows that ionic liquid fuels have a 3–12% lower specific impulse when paired with a nitrogen tetroxide oxidizer. However, when paired with hydroxylammonium nitrate oxidizer, the specific impulse of the ionic liquids is 1–4% lower than that of hydrazine and unsymmetrical dimethylhydrazine paired with nitrogen tetroxide. Analytical investigation of an electrospray electric propulsion system shows that ion extraction in the pure ion regime provides a very high specific impulse, outside the optimum range for potential missions. Results suggest a deceleration grid, a lower ion fraction, or emission of higher solvated states is required. Analysis of a dualmode ionic-liquid-propelled spacecraft shows that the electric propulsion component determines the overall feasibility comparedwith current technology. Results indicate that the specific power for an ionic liquid electrospray systemmust be at least 15 W=kg in order for a dual-mode ionic liquid system to compete with traditional hydrazine and Hall thruster technology.

A radiation-cooled low-power arcjet thruster is tested in order to study the startup characteristics and performancewith nitrogen and ammoniamixtures as the propellant in a vacuumwith chamber pressures lower than 5 Pa. In the ignition tests, two different starter circuits, i.e., high-frequency and high-voltage pulse ignition techniques, are used to initiate the discharge. The results of the ignition tests show that arc current, electrode temperature, inlet pressure, andmole fraction of ammonia in the propellant have significant effects on arcjet startup characteristics. Experimental studies are also carried out to reveal the effects of different molar mixing ratios of nitrogen/ammoniamixtures on arcjet thruster performance. It is found that the important performance parameters of arcjet such as thrust, specific impulse, and thrust efficiency increase with an increased amount of ammonia in the propellant. The molar mixing ratio of the nitrogen/ammonia mixtures has significant effects on arcjet currentvoltage characteristics.

e = elementary charge, C fi = ionization mass fraction of the propellant Id = discharge current, A Ii = exhausted ion current, A Isp = specific impulse, s j = ion current density, A=m M = spacecraft mass, kg m = mass of a xenon atom/ion, kg _ m = flow rate of propellant as determined by the mass-flow controller, kg=s P = kinetic power delivered to the spacecraft, W Ps = supply output power IdVd, W Q = average charge state of the ionized propellant q = charge number T = thrust, N Ue = ion velocity, m=s 2 Vaccel = acceleration potential for a given ion, V Vd = discharge voltage (anode potential), V = average offaxis ion trajectory angle t = trip time, s V = velocity increment, m=s ion = ion kinetic energy B = beam divergence as an efficiency c = current efficiency E = energy efficiency p = propellant efficiency probe = the product of all probe-measured efficiencies T = thrust efficiency v = voltage utilization efficiency vdf = velocity distribution efficiency = angle from thruster axis ctg = cathode to ground potential, V plasma = plasma to ground potential, V

A transversely excited at atmospheric pressure CO 2 laser operated at 10.6 μm and 5 μs pulse width was employed to ablate atomized water droplets for laser propulsion. The momentum was derived from the force sensor data. Experimental results indicated that coupling coefficient C m , specific impulse I sp , and internal efficiency η were increased remarkably. The maximum value of C m was 52.1 dyne/W. The measured value of I sp was 102 s. Internal efficiency η of 26.1 % was achieved. Total impulse I and coupling coefficient C m were dependent largely on the droplets' Sauter mean diameter and speed and their combined interaction. Specific impulse I sp and internal efficiency η were dependent on the droplets' Sauter mean diameter, speed, flow rate of liquid propellant feeding system, and their combined interaction. The flow rate of the liquid feeding system dominated the influencing factors on specific impulse I sp and internal efficiency η.

Since the mass flow rate through nozzle is primarily a function of throat density and axial Mach number, the reductioninthelatterwithswirlexplainstheobservedreductioninmass flowatmatchedreservoirpressure.Greater pressures, on the other hand, result in higher throat densities, which compensates for the reduced axial Mach number,andthemass flowratecanbekeptconstantatitsnonswirlingvalue.Itwasalsofoundthatthedistributionof subsonicMachnumber(andnotanyofitscomponents)inaswirling flowissolelydependentoncross-sectionalarea, similar to nonswirling flows; i.e., nonswirling and swirling flows have the same subsonic Mach number profile. In terms of thrust and specific impulse, the application of swirl at matched nozzle reservoir pressure results in the expected reductions in discharge coefficient, thrust, and specific impulse. At matched mass flow, however, the application of swirl results in the enhancement of both thrust and specific impulse. This is attributed to the considerable degree of underexpansion associated with the swirling flow as a result of the higher nozzle reservoir pressure with swirl.

The fundamental relationship between entropy and aerospace vehicle and mission performance is analyzed in terms of the general availability rate balance between force-based vehicle performance, available energy associated with expended propellant, and the overall loss rate of availability, including the vehicle wake. The availability relationship for a vehicle is analytically combined with the vehicle equations of motion; this combination yields the balance between on-board energy rate usage and rates of changes in kinetic and potential energies of the vehicle and overall rate of entropy production. This result is then integrated over time for a general aerospace mission; as examples, simplified single-stage-to-orbit rocket-powered and air-breathing missions are analyzed. Examination of rate of availability loss for the general case of an accelerating, climbing aerospace vehicle provides a powerful loss superposition principle in terms of the separate evaluation and combination of loss rates for the same vehicle in cruise, acceleration, and climb. Rate of availability losses is also examined in terms of separable losses associated with the propulsion system and external aerodynamics. These loss terms are cast in terms of conventional parameters such as drag coefficient and engine specific impulse. Finally, rate losses in availability for classes of vehicles are described.

Preliminary designs of low-thrust transfer trajectories are developed in the Earth-moon three-body problem with variable specific impulse engines and fixed engine power. The solution for a complete time history of the thrust magnitude and direction is initially approached as a calculus of variations problem to locally maximize the final spacecraft mass. The problem is then solved directly by sequential quadratic programming, using either single or multiple shooting. The coasting phase along the transfer exploits invariant manifolds and, when possible, considers locations along the entire manifold surface for insertion. Such an approach allows for a nearly propellant-free final coasting phase along an arc selected from a family of known trajectories that contract to the periodic libration point orbit. This investigation includes transfer trajectories from an Earth parking orbit to some sample libration point trajectories, including L 1 halo orbits, L 1 and L 2 vertical orbits, and L 2 butterfly orbits. Given the availability of variable specific impulse engines in the future, this study indicates that fuel-efficient transfer trajectories could be used in future lunar missions, such as south pole communications satellite architectures.

System design of liquid rocket engines must consider engine performance, weight, cost, and reliability requirements. A general design optimization framework has been developed in this paper to select the best system parameters for liquid rocket engines with gas-generator cycles. The object is to maximize the specific impulse and vacuum thrust-to-weight ratio of the engine with given system requirements and design assumptions by changing thrust-chamber pressure and mixture ratio. The system analysis, along with the engine weight estimation, is based on a modular scheme. Multidisciplinary design optimization formulations including multidisciplinary feasible and collaborative optimization are used, evaluated, and compared during the optimization process. Several techniques of multi-objective processing are also used to identify the Pareto frontier and the optimal compromise solutions. A proposed cryogenic-propellant engine using liquid oxygen and hydrogen with a gas-generator cycle is studied as a specific example. Moreover, uncertainties in the engine operation, such as thrust-chamber pressure and mixture ratio, are taken into account as random variables in the reliability-based optimization. Results are presented to illustrate the tradeoff between the engine performance and reliability requirements.

ROCKET-GRADE hydrogen peroxide has been used as a monopropellant and a storable oxidizer. However, because of the demand for a higher specific impulse, hydrazine andN2O4 are being used as the monopropellant and storable oxidizer, respectively. Recently, due to concerns regarding propellant toxicity, there has been a renewed interest [1] in the use ofH2O2 in propulsion systems [2–10]. A monopropellant thruster is operated in either the continuous or pulse mode. The thrust force and pressure instability are important issues in the continuous mode. For generating the desired thrust, the catalytic reactor size required for completely decomposing the propellant must be determined [8]. However, in the pulse mode (the main operationmode for attitude control systems), the response characteristics of the thruster are important. The catalyst activity, thruster component design (including the injector design), manifold volume, ullage volume in the reactor, and operating pressure influence the thruster response time. Tian et al. investigated the response time when using a combination of PbO and MnO2 catalysts [11]; they found that Ir=Al2O3 is unsuitable for use as a catalyst in a H2O2 monopropellant thruster [12]. Xu et al. studied the activities of various catalysts during H2O2 decomposition [13]. El-Aiashy et al. reported that the catalyst activity ofMnO2 increased when ZnO was added [14]. Hasan et al. reported that the activity ofMnO2 increased when promoters such as Ni, Cu, Bi, and Ce were added [15]. None of the aforementioned studies have addressed the effect of thruster design parameters on response times, although a few researchers have measured the thruster response time. Optimization of the thruster design (determination of the appropriate injector and ullage volume in the reaction chamber) can also influence the response characteristics. Therefore, we investigate the response characteristics of H2O2 monopropellant thrusters for three different thruster designs andmeasure the response times by varying the injector type, reactor volume, and catalyst grain size. AMnO2=Al2O3 catalyst is used for the decomposition of concentrated H2O2 (90 wt%).