Propellant Mass Flow Research Articles

Ionization and acceleration processes in a small, variable channel width, permanent-magnet Hall thruster

A 200 W-class permanent-magnet Hall thruster has been operated with different channel widths. A series of experiments have been carried out for each geometry over a broad range of propellant mass flow rates and applied voltages to investigate the impact of a channel cross-section area variation on discharge and plume properties. Measurement outcomes allow calculation of several quantities, namely the propellant utilization, the beam voltage, the equilibrium wall temperature and the energy flux towards the walls. This study reveals that widening the channel of a low-power Hall thruster enhances ionization and acceleration processes and broadens the operating envelope. Two mechanisms are proposed to explain experimental observations. The surface-to-volume ratio decreases with h, which reduces wall-losses. The magnetic field strength near walls increases with h, which leads to a better plasma confinement.

Performance Characteristics of a DME Propellant Arcjet Thruster

This paper describes the influence of cathode configuration on performance of an arcjet thruster using dimethyl ether (DME) propellant. DME, an ether compound, has suitable characteristics for a space propulsion system; DME is storable in a liquid state without being kept under a high pressure, and requires no sophisticated temperature management such as a cryogenic device. DME can be gasified and liquefied simply by adjusting temperature whereas hydrazine, a conventional propellant, requires an iridium-based particulate catalyst for its gasification. In this study, thrust of a 1-kW class DME arcjet thruster is measured at a discharge current of 13 A, DME mass flow rates ranging 15 to 60 mg/s under three cathode configurations: flat-tip rods of 2 and 4 mm in diam. and 4-mm-diam. rod having a cavity of 2 mm in diameter. Thrust measurements show that thrust is increased with propellant mass flow rate. Among the tested cathodes, the flat-tip rod of 4 mm in diam. with 55 mg/s DME flow rate yielded the highest performance: specific impulse of 330 s, thrust of 0.18 N, discharge power of 1400 W and specific power of 25 MJ/kg.

Semi-Analytical Method for the Analysis of Solar Sail Heliocentric Orbit Raising

L OW-THRUST trajectories are usually investigated within an optimal framework by minimizing a suitable scalar performance index, like the total mission time or the propellant mass consumption. Very often a preliminary mission analysis requires a considerable amount of simulation time to investigate the effect of different parameters on mission performance; therefore the availability of simple analytical and/or graphical means to describe the orbit evolution is very useful. These simplified tools are of great importance formission-design concepts and tradeoff studies because they can provide good estimates of fundamental mission parameters, like the total velocity increment, with a reduced computational effort. In this regard, a famous method was developed by Edelbaum [1] for the study of a circle-to-circle transfer under the effect of a constant and freely steerable propulsive acceleration. This method, especially in the variant proposed by Kechichian [2–4], has been succeedingly extended to deal with more complex mission scenarios [5–7]. An interesting alternative to Edelbaum’s approach is provided by the so-called Alfano transfer [8]. Assuming that the spacecraft is subjected to a constant propulsive thrust, the Alfano transfer basically uses the calculus of variations to minimize the flight time t required for a circle-to-circle transfer as a function of the initial and final orbit radii (r0 and r1, respectively) and of the propulsion system characteristics (in terms of initial acceleration a0 and constant propellant mass flow rate _ mp). Because, by assumption, the thrust acts continuously without any coasting phase, the minimum-time transfers coincide with minimum propellant mass consumption trajectories. The results of the optimization process are then collected into charts [9] that simplify the flight time evaluation as a function of themission characteristics for a number of parameters combinations, like r1=r0, a0, and _ mp=m0, where m0 is the initial spacecraft mass. The aim of these charts is twofold. On one side, they allow the designer to quickly obtain, in a graphical form, the optimal transfer performance, to estimate the sensitivity to a mission parameter variation, and to automatize his/her analysis with the aid of simple interpolations [10]. The second and more intriguing aim is to detect possible relationships to describe the variation of the performance index as a function of the mission parameters. Such information is useful both for validating already available analytical models [5] and to suggest new approximations. According toAlfano [10], in hismethod theflight time is not given explicitly, but it is represented in terms of total accumulated velocity change Va, that is

Operation Characteristics of a Steady-State, Two-Dimensional MPD Thruster Using a Hollow Cathode

This paper reports operation characteristics of a steady-state, two-dimensional magnetoplasmadynamics (MPD) thruster using a hollow cathode. In order to improve cathode performance, a tungsten rod was replaced with a hollow cathode which had better electron emission performance and a longer lifetime. A rectangular acceleration section was employed between electrodes. The discharge voltage and thrust were measured with varying an applied magnetic field, discharge current and propellant mass flow rate.

Thrust Evaluation of an Arcjet Thruster Using Dimethyl Ether as a Propellant

This paper describes the performance of an arcjet thruster using dimethyl ether (DME) as a propellant. DME, an ether compound, has adequate characteristics for space propulsion systems; DME is storable in a liquid state without a high pressure or cryogenic device and requires no sophisticated temperature management. DME is gasified and liquefied simply by adjusting temperature, whereas hydrazine, a conventional propellant, requires an iridium-based particulate catalyst for its gasification. In this study, thrust of the designed kW-class DME arcjet thruster is measured with a torsional thrust stand. Thrust measurements show that thrust is increased with propellant mass flow rate, and that thrust using DME propellant is higher than when using nitrogen. The prototype DME arcjet thruster yields a specific impulse of 330 s, a thruster efficiency of 0.14, and a thrust of 0.19 N at 60-mg/s DME mass flow rate at 25-A discharge current. The corresponding discharge power and specific power are 2.3 kW and 39 MJ/kg.

Current filamentation and onset in magnetoplasmadynamic thrusters

The possible role of current filamentation in the operation of magnetoplasmadynamic thrusters is investigated here by means of a stability analysis of a current-carrying plasma in a simplified coaxial configuration. Magnetoplasmadynamic thrusters are known to enter a strongly unstable regime, named onset in the literature, when operated above a threshold current level, given the propellant mass flow rate. During onset, a transition from diffuse to spotty current pattern occurs, leading to intense fluctuations of thruster terminal voltage and to severe anode damage with commonly employed anode materials. Despite several experimental and theoretical efforts in the last few decades, no complete and definitive understanding of the physical nature of this phenomenon is yet available. In this work it is shown that conditions suitable for azimuthal symmetry breaking and the subsequent development of this instability can actually exist in magnetoplasmadynamic thrusters. A physically coherent explanation of the complex onset phenomenology is then proposed, showing that both the plasma dynamics and the voltage fluctuations can be ultimately explained in terms of the filamentation instability and its effects.

Finite-thrust optimization of interplanetary transfers of space vehicle with bimodal nuclear thermal propulsion

The nuclear thermal rocket (NTR) propulsion is one of the leading promising technologies for primary space propulsion for manned exploration of the solar system due to its high specific impulse capability and sufficiently high thrust-to-weight ratio. Another benefit of NTR is its possible bimodal design, when nuclear reactor is used for generation of a jet thrust in a high-thrust mode and (with an appropriate power conversion system) as a source of electric power to supply the payload and the electric engines in a low-thrust mode. The model of the NTR thrust control was developed considering high-thrust NTR as a propulsion system of limited power and exhaust velocity. For the proposed model the control of the thrust value is accomplished by the regulation of reactor thermal power and propellant mass flow rate. The problem of joint optimization of the combination of high- and low-thrust arcs and the parameters of bimodal NTR (BNTR) propulsion system is considered for the interplanetary transfers. The interplanetary trajectory of the space vehicle is formed by the high-thrust NTR burns, which define planet-centric maneuvers and by the low-thrust heliocentric arcs where the nuclear electric propulsion (NEP) is used. The high-thrust arcs are analyzed using finite-thrust approach. The motion of the corresponding dynamical system is realized in three phase spaces concerning the departure planet-centric maneuver by means of high-thrust NTR propulsion, the low-thrust NEP heliocentric maneuver and the approach high-thrust NTR planet-centric maneuver. The phase coordinates are related at the time instants of the change of the phase spaces due to the relations between the space vehicle masses. The optimal control analysis is performed using Pontryagin’s maximum principle. The numerical results are analyzed for Earth–Mars “sprint” transfer. The optimal values of the parameters that define the masses of NTR and NEP subsystems have been evaluated. It is shown that the low-thrust NEP subsystem with Brayton cycle power conversion system is preferable in comparison with NEP subsystem with thermoemission power conversion system.

Spectral measurements of inductively coupled and helicon discharge modes of a laboratory argon plasma source

An experimental study was conducted to investigate the effects of several operational parameters in the emission spectra, in the 400–850 nm wavelength region, of a laboratory Argon plasma source. In particular, the emission spectra of the inductively coupled plasma and the Helicon plasma modes of operation were compared. Comparisons of spectra point to a significant increase in the ionization fraction of the plasma for the Helicon mode of operation. The spectral measurements allow one to determine the major trends in the plasma electron density for various parameters such as power delivered to the helical antenna, propellant mass flow rate, and applied external magnetic field intensity. Analysis of a prominent Argon single ion line, at 434.8 nm, points out that the plasma electron density increases linearly with the power delivered to the helical antenna, and that there is an optimum propellant mass flow rate for maximum ionization fraction. Additional analysis of the same line shows that above a minimum applied axial magnetic field intensity, the variation in the magnetic field strength has little effect on the plasma electron density.

우주추진용 홀방식의 전기추력기를 위한 제논연료공급장치 개발

A Xenon Feed System (XFS) has been developed for hall-effect thruster to small satellite space-propulsion system applications. The XFS delivers low pressure gas to the Anode and Cathode of thruster head unit from a xenon storage tank. Accurate throttling of the propellant mass flow rate is independently required for each channel of the thruster head unit. The mass flow rate to each channel is controlled using the accumulator tank pressure regulation through a micron orifice and isolation valve. This paper discusses the Xenon Feed System design including the component selections, performance estimation and functional test.

爆轟管同軸回転バルブを用いたパルスデトネーションエンジンの性能評価

We conducted multi-cycle combustion experiments of a coaxial-rotary-valve pulse detonation engine (PDE) system. This PDE system showed a stable operation at the valve rotating frequencies of 5.0, 10, 15, 20, and 33 Hz. We successfully measured propellant mass flow rates, thrusts, and specific impulses in the 2-sec operation duration. The maximum thrust was 32 N at the operation frequency of 33 Hz and at the supply pressure of 0.98 MPa. The maximum specific impulse was 250 sec at the operation frequency of 33 Hz and at the supply pressure of 0.69 MPa. The partial fill ratio was varied from 0.07 to 1.14. The partial fill effect was almost identical to the previous model calculations.

Minimum-time space missions with solar electric propulsion

This paper presents an optimal control strategy suitable for investigating the mission performance of a spacecraft equipped with a NASAʼs Evolutionary Xenon Thruster. Using an accurate mathematical model of the propulsion system, based on both experimental data and a semi-empirical schematization of the power generation system, the minimum-time rendezvous problem with various celestial bodies is investigated through an indirect approach. Unlike a conventional approach, in which the thruster power is varied with continuity within a given range, here the solar electric propulsion system is described using a finite number of operation points, characterized by different pairs of thruster input power and propellant mass flow rate. The model includes the possibility of obtaining coasting phases along the spacecraft trajectory. The optimal control strategy is found for a three-dimensional problem in which the actual shape and relative orientation of the initial and final orbits are taken into account. The paper presents the simulation results for a number of minimum-time trajectories towards inner planets (Mercury, Venus, and Mars) and three Near Earth Asteroids (NEAs) of particular interest for the scientific community: 1998 KY-26, 99942 Apophis, and 2007 VK-184.

Development of a Magnesium and Zinc Hall-Effect Thruster

§This paper describes what are believed to be the first demonstrations of Hall-effect thrusters operating on magnesium and zinc propellant. Pathfinding experiments were performed using consumable anodes that were machined from solid magnesium and zinc, which sublimated under the heat load from the discharge plasma and delivered propellant gas to the thruster. Therefore the magnesium and zinc anodes served as the acceleration electrode and also served as the propellant supply. A retarding potential analyzer was used to obtain plume diagnostics during early operation of the experiments, showing reasonable acceleration of the propellant ions. Two main issues were expected and encountered with the solid magnesium and zinc anodes – 1) the zinc anode displayed localized melting causing liquid zinc to accumulate in the discharge channel and 2) the crude scheme did not feature any means to actively control the sublimation rate of the metal propellant. A new porous anode with internal propellant reservoir was designed and built that could be refilled with either propellant, eliminating liquid intrusion into the discharge channel. A scheme developed earlier for bismuth thrusters was employed wherein shim anodes were implemented to shift discharge current to and from the main anode to control the main anode temperature and hence the metal propellant sublimation rate. Results are reported showing stable operation of a thruster using a porous anode with magnesium propellant for more than 100 minutes. Also demonstrated was the ability of the shim anode scheme to actively control the propellant mass flow rate.

Erratum to: Combustion Characteristics of a Hydroxylammonium Nitrate Based Liquid Propellant. Combustion Mechanism and Application to Thrusters

A new composition of hydroxylammonium nitrate based solution containing ammonium nitrate, methanol, and water was developed for monopropellant in a reaction control system (RCS) as an alternative to conventional hydrazine. In comparison with hydrazine, this solution has a 20% higher specific impulse, 1.4 times higher density, and lower freezing point and toxicity. The linear burning rate of the solution is moderate at the operating pressures of RCS thrusters. It was found that the linear burning rate had some characteristics whose mechanisms had not been understood. The combustion mechanism of this solution was investigated. The burning behavior was observed using a medium speed camera, and a temperature profile for the combustion wave was measured with a 2.5 µm diameter thermocouple. From these results, the instability of the liquid-gas interface may trigger a sudden increase in the burning rate, and methanol was found to be effective in reducing the bubble growth rate in the solution. The reactivity of several catalysts was evaluated in an open-cup test, and the S405 catalyst for hydrazine showed the best performance among them. Thruster tests were conducted using the S405 catalyst with variation in the propellant mass flow rate, catalyst bed configuration, and length-to-diameter ratio of the combustor. As a result, parameters were determined that ensured long operating time. The model thruster operated stably for up to 100 sec with a specific impulse I sp = 240 sec, which corresponds to a 90% efficiency.

Ion diagnostics of a discharge in crossed electric and magnetic fields for electric propulsion

The velocity distribution function (VDF) of metastable Xe+ ions was measured along the channel centerline of the high-power PPS®X000 Hall effect thruster by means of laser induced fluorescence (LIF) spectroscopy at 834.72 nm for various discharge voltages (300–700 V) and propellant mass flow rates (6–15 mg s−1). The development of the on-axis profile of the velocity dispersion reveals the interrelation between ionization and acceleration layers. The ion velocity profiles are in accordance with outcomes of a hybrid numerical model in which the electron mobility is assessed from particle-in-cell simulations. The axial distribution of the effective electric field is inferred from the mean ion velocity profile, despite the parasitic effect due to ions created in the acceleration region. Most of the acceleration process takes place outside the thruster channel. The electric field augments and it moves upstream when the applied voltage is ramped up. The impact of the xenon mass flow rates is found to depend upon the voltage. A novel approach based on the moments of the experimental VDFs in combination with the Boltzmann's equation is introduced in order to determine the real electric field distribution. The method also provides the ionization frequency profile. The LIF diagnostics reveals the existence at the end of the acceleration region of fast ions of which the kinetic energy is above the supplied energy. The fraction of these supra-sped up ions grows when the voltage increases. The ion VDFs were also recorded in the plasma plume far field by way of a retarding potential analyzer (RPA). The shape of the RPA traces as well as their evolution with operating conditions are in agreement with trends observed by means of LIF spectroscopy. Finally, physical mechanisms at the origin of supra-sped up ions are discussed in light of numerical simulation outcomes and a set of new experimental results.

Thrust Stand Micromass Balance for the Direct Measurement of Specific Impulse

A technique has been developed to directly measure the specific impulse from pulsed thruster systems. The technique is especially useful for propulsion devices that use solid propellants, for which a direct measurement of the propellant mass flow is extremely difficult. A torsion balance is used with a horizontal axis of rotation. A thruster is placed on the balance such that the impulse of the thruster firing and the change in mass due to the expelled propellant act in the same direction. A combined impulse and steady-state force measurement (due to propellant mass loss) can then be decoupled to assess the ratio of the impulse to the weight of propellant expended, or the specific impulse. A model has been developed to show the utility of the technique for pulsed systems with a firing time less than the natural period of the balance. An experimental proof of principle study was also undertaken using the laser ablation of engineering-grade Buna, Viton, and Teflon propellants. Specific-impulse measurements on the order of 200 s have been demonstrated with this laser ablation thruster.

Invited Article: Time accurate mass flow measurements of solid-fueled systems

A novel diagnostic method is described that utilizes a thrust stand mass balance (TSMB) to directly measure time-accurate mass flow from a solid-fuel thruster. The accuracy of the TSMB mass flow measurement technique was demonstrated in three ways including the use of an idealized numerical simulation, verifying a fluid mass calibration with high-speed digital photography, and by measuring mass loss in more than 30 hybrid rocket motor firings. Dynamic response of the mass balance was assessed through weight calibration and used to derive spring, damping, and mass moment of inertia coefficients for the TSMB. These dynamic coefficients were used to determine the mass flow rate and total mass loss within an acrylic and gaseous oxygen hybrid rocket motor firing. Intentional variations in the oxygen flow rate resulted in corresponding variations in the total propellant mass flow as expected. The TSMB was optimized to determine mass losses of up to 2.5 g and measured total mass loss to within 2.5% of that calculated by a NIST-calibrated digital scale. Using this method, a mass flow resolution of 0.0011 g/s or 2% of the average mass flow in this study has been achieved.

Optimal trade studies of interplanetary electric propulsion missions

The aim of this paper is to investigate various aspects of an interplanetary electric propulsion mission. To this end, the problem of minimum-time and minimum-fuel-time-fixed interplanetary rendezvous are first discussed. A circle-to-circle two-dimensional transfer is assumed and the problem is solved using an indirect approach. Key features of the analysis are the use of a realistic thruster model. In particular, a second-order polynomial approximation is used for modelling both the thrust and the propellant mass flow rate as a function of the thruster input power. As the mission time and the propellant mass required to reach the target orbit are two classical conflicting requirements, a number of simulations have been performed with different values of mission times and maximum input power supplied by the power processing unit. The numerical examples show that trade-off studies can be conducted with relative ease and with moderate computational effort. The simulation data have then been collected in plots and graphs that are useful for preliminary mission analysis. They are of particular importance for allowing the designer to know exactly what technological level is required before a mission is feasible.

A calibrated infrared imaging study on the steady state thermal behaviour of Hall effect thrusters

The thermal behaviour of Hall effect thrusters was investigated by means of calibrated infrared thermal imaging performed in the 8–9 µm spectral domain. Study on the variation of the steady state temperature of Hall thruster elements like discharge chamber (channel) walls and anodes along with discharge voltage and propellant (xenon) mass flow rate confirms that energy loss mechanisms, which are responsible for the heating of the thrusters, are a direct consequence of interactions between charged particles and surfaces. In order to obtain new insights into plasma surface interactions inside a thruster, the channel wall temperature was monitored over a broad range of electrical power stretching from 400 W to 5.5 kW for three types of thrusters with different designs, dimensions and operation domains, namely SPT100-ML, PPS ® 1350-G and PPSX000-ML. Note that over the range of thruster operating conditions the facility backpressure varies from 10−5 to 6 × 10−5 mbar. In addition, the effect of discharge chamber wall material on temperature field was also investigated using dielectric BN–SiO2 and AlN walls as well as conducting graphite walls. For a given thruster geometry and material, a simple relationship between the mean wall temperature and the input electrical power can be established, in contradiction to the complex dynamics of such a magnetized plasma medium. Besides, thruster thermal history and degree of wear do not have a strong impact on power losses inside the channel.

Effect of solute mixing in the liquid propellant of a pulsed plasma thruster

Sodium chloride or ammonia was dissolved in the water propellant of pulsed plasma thrusters to improve the performance. Pulsed plasma thrusters using liquid propellant utilize water as attractive alternative instead of Teflon. Water propellant enables in controlling propellant mass flow and leads to high specific impulse. However, liquid propellant pulsed plasma thrusters have larger plasma resistance and lower thrust power ratio than the common Teflon propellant thruster. Here, sodium chloride and ammonia solution of water were examined to decrease that plasma resistance. As a result, emission lines attributed from the solute were observed using sodium chloride aqueous solution propellant, and a 5% reduction of the plasma resistance was shown, and the thrust to power was increased. However, ammonia aqueous solution decreased the thruster performance.

宇宙推進用アノードレイヤー型ホール効果プラズマ加速器の性能特性

The Hall-effect plasma accelerators are one of the most advantageous electric propulsion device in space. The Hall-effect plasma accelerators are generally categorized into magnetic-layer type and anode-layer type. The anode-layer type Hall-effect plasma accelerator is expected to be widely accepted as space thrusters with longer-lifetime and higher thrust performance. In this study, the fundamental performance characteristics of an anode-layer type Hall-effect plasma accelerator have been investigated. A 1kW-class anode-layer type Hall-effect plasma accelerator, TALT(Toyonaka Anode-Layer Thruster)-1, was operated under various magnetic fields, discharge voltages, propellant mass flow rates and anode positions and structures. The thrust performance was enhanced with positioning the anode front in the upstream and widening the anode propellant path width. The thrust and the specific impulse ranged from 12mN to 45mN and from 800s to 1739s, respectively, at the discharge voltages of 200-400V under the xenon mass flow rates of 1-3mg/s. The thrust efficiency reached 45%.