Propellant Mass Flow Research Articles

Analytical model of a Hall thruster

Hall thrusters are one of the most successful and prevalent electric propulsion systems for spacecraft in use today. However, they are also complex devices and their unique E×B configuration makes modeling of the underlying plasma discharge challenging. In this work, a steady-state model of a Hall thruster is developed and a complete analytical solution presented that is shown to be in reasonable agreement with experimental measurements. A characterization of the discharge shows that the peak plasma density and ionization rate nearly coincide and both occur upstream of the peak electric field. The peak locations also shift as the thruster operating conditions are varied. Three key similarity parameters emerge that govern the plasma discharge and which are connected via a thruster current–voltage relation: a normalized discharge current, a normalized discharge voltage, and an amalgamated parameter, α¯, that contains all system geometric and magnetic field information. For a given normalized discharge voltage, the similarity parameter α¯ must lie within a certain range to enable high thruster performance. When applied to a krypton thruster, the model shows that both the propellant mass flow rate and the magnetic field strength must be simultaneously adjusted to achieve similar efficiency to a xenon thruster (for the same thruster geometry, discharge voltage, and power level).

Supervised machine learning-based Hall thruster scaling

The scaling methodology described in this paper to find the geometry and working parameters of Hall Thrusters is based on algorithms of supervised Machine Learning. The approach considers the determination of the geometrical sizes, propellant mass flow rate and discharge voltage taking thrust and specific impulse as requirements. The magnetic field is also considered. The Gradient Boosting Regression is found as the most suitable algorithm for our purpose. Scaling relies on a specific database of 54 thrusters for the determination of all parameters. The database includes measurements carried out with xenon, krypton and argon as propellant. A unique analytical approach based on the GBR algorithm has been developed and validated to determine the suitable design for a Hall thruster according to space mission requirements.

Effect of swirl injection parameters on cooling performance of vortex-cooled bipropellant thruster

The cooling performance of a 10N-class gaseous Propane/GOx bi-propellant thruster with vortex cooling was investigated in this experimental study. Injecting oxidizer with a vortex was proposed as a potential cooling method for a semi-thermoplastic constructed thruster and demonstrated as a promising integrated cooling method that can be fabricated using 3D printing technology. Compared to other cooling types, vortex cooling has low system complexity and can be 3D printed as a single solid piece along with a thruster, reducing design complexity, weight, and manufacturing time. The experimental testing examined various injection configurations and their associated parameters to determine the most optimal injection combination. The test collected a comprehensive amount of data to evaluate the cooling performance with respect to varying operational parameters: oxygen-to-fuel ratio, propellant mass flow rate, and injection configuration. The study found that the acrylic chamber operated safely with a proper injection configuration identified maintaining the temperature of the thruster chamber walls below 60 °C, lower than the melting point of the acrylic material, with no damage observed. It was found that the downstream swirl injection had more favourable vortex cooling performance than the upstream injection for the scale of the thruster considered in this study.

Far-field plume characterization of a low-power cylindrical Hall thruster

A fully cylindrical Hall thruster prototype was tested in the power range of 30–300 W with the objective of understanding the behavior of the discharge as a function of input parameters. Various operating conditions were compared, including two magnetic field configurations, a set of propellant mass flow rates, and a range of discharge voltages. Plasma properties were measured in the plume, with a Langmuir probe, a retarding potential analyzer, and a Faraday cup. The experimental results showed that the mass flow rate strongly affects the ionization and, consequently, other related properties such as the plasma density, currents, and propellant utilization. The discharge voltage also appeared to influence the ion energy and propellant utilization. The performance accessible from the measured magnitudes is assessed, resulting in a maximum thrust efficiency of about 18% at 0.35 mg s−1 and 168 W.

Effect of Using Fractals as Grain Configuration on the Performance of Solid Rocket Motors

Grain design is a very crucial consideration for determining the performance of solid propellant rocket motors. A rocket’s performance and combustion characteristics (thrust profile, chamber pressure variation, propellant mass flow rate, and sliver among other important performance parameters) depend very strongly on the grain configuration. We propose fractal shapes (Koch star, Koch snowflake, Gosper, and Cesaro) as new port geometries and theoretically predict their performance using simulations based on the fast-marching method implemented in the open-source software Open-Motor. The initial thrust for fractal geometry grains was found to be significantly higher than the commonly used industry standard BATES and Finosyl ports. The fractal grains also showed a significant decrease in sliver as compared to the Finocyl port that is generally used to increase performance as compared to the circular BATES ports. Especially, The Koch snowflake port had a drastically lower sliver. These new grain geometries proposed in this work can be used for applications in which low initiation and ignition times along with high initial thrust are crucial. Moreover, these geometries can improve the performance of the solid rocket motor (SRM) owing to their better regression characteristics which promise complete utilization of the entire grain web.

Throttling simulation and control of liquid oxygen/liquid methane rocket engine pressurized by the electric pump

This article presented simulation and control of a liquid oxygen/liquid methane variable thrust rocket engine pressurized by the electric pump in the throttling process. To get a proper injection pressure drop, the variation of the injection area was investigated and determined. In the control scheme, the combustion pressure and mixture ratio were the targets. To simplify the control scheme, the pressure and mixture ratio were correlated to the oxygen mass flow rate and methane mass flow rate, respectively. By controlling the propellant mass flow rate, the target engine thrust was reached. In addition, the influence of mass flow rate measurement points on the control effect was analyzed. The simulation results showed that the designed controller can reduce the overshoot of the critical engine parameters. Therefore, the maximum temperature of the thrust chamber wall after the regenerative cooling was decreased. Consequently, the thrust control errors were within 3% except for the points where the thrust changed suddenly, which can meet the mission requirements.

Characteristics of spray and combustion in gas-centered swirl coaxial injector with varying gas nozzle diameter

Gas-centered swirl coaxial (GCSC) injectors, a type used in the combustion chamber of liquid rocket engines, have been widely investigated because their characteristics are key parameters for the performance of the engine. This study investigated the spray and combustion characteristics of GCSC injectors used in the combustion chamber of oxidizer-rich staged-combustion cycle engines. When the propellant mass flow rate and oxidizer-to-fuel mass flow ratio (O/F ratio) were identical, the diameter of the gas nozzle was changed such that the ratio of the gas-to-liquid momentum could be changed; thus, the swirl strength of the liquid also changed because of the increase in the injector diameter. In the spray experiments, the spray characteristics were identified in accordance with changes in the diameter of the gas nozzle through the spray angle, droplet mean velocity, and droplet size by using backlight photography and the ParticleMaster imaging system. For combustion tests, four types of reduced combustion chambers were designed to spray propellants through the same two injectors used in the spray experiments. To generate an oxidizer-rich combustion gas, combustion tests were performed using a pre-burner. The changes in combustion characteristics according to the change in the injector shape under the same combustion chamber shape, propellant flow rate, and O/F ratio were identified through the characteristic velocity and heat flux of the combustion chamber. As the gas–liquid momentum flux ratio increased in accordance with the change in diameter of the gas nozzle, the characteristic velocity increased from at least 1700 m/s to a maximum of 1721 m/s, which was related to the decrease in the Sauter mean diameter observed in the spray experiments. Few studies have investigated GCSC shape variables by simultaneously conducting atmospheric spray experiments and high-pressure combustion tests on GCSC injectors using kerosene and oxidizer-rich combustion gases as propellants. This study observed the changes in spray characteristics according to the change in the diameter of the injector gas nozzle through atmospheric spray experiments of the GCSC injector and compared and analyzed the combustion characteristics through combustion tests using the same injector. The results of this study will potentially aid the design optimization of GCSC injectors and facilitate future developments.

Development of a modular hollow cathode for ground testing of plasma thrusters

A hollow cathode with a modular design has been developed to assist with laboratory testing of plasma-based thrusters for satellite applications. This novel modular design includes interchangeable components for varying the geometry and tailoring the configuration to specific applications, as well as easing the replacement of individual components in the case of damage. The modular hollow cathode also presents unconventional design features aimed at improving the heating efficiency: the heater is in direct contact with the emitter and the keeper is not in physical contact with the cathode base. The modular hollow cathode development has been based on a combination of theoretical modelling and experimental testing. The influence of the novel mechanical assembly has been investigated by characterising the operational envelope at different propellant mass flow rates for xenon and krypton. The modular hollow cathode has demonstrated stable operation by sustaining discharge currents between 0.5 and 4 A at different conditions. Finally, the cathode has been coupled with a Hall-type plasma thruster operating in the 0.3–2.5 A anode current range. This paper outlines development, experimental validation of this peculiar mechanical cathode configuration, covering plasma and thermal modelling, standalone testing, and coupled Hall-type thruster operation.

Experimental and numerical investigation of transverse combustion instability in a rectangle multi-injector rocket combustor

Rocket engines are prone to suffer from transverse combustion instabilities, manifesting as periodic oscillations in pressure and heat release rate. Self-excited transverse instabilities in a rectangle multi-injector rocket combustor powered by n-C12H26 and O2, are investigated experimentally and numerically in this paper. Transverse combustion instabilities are captured during repeatable experiments, while dynamic characteristics in the combustion chamber and driving mechanisms of instabilities are analyzed via numerical simulation based on stress-blended eddy simulation. Dynamic mode decomposition analysis reveals coupling behaviors between the transverse mode in the combustion chamber and longitudinal mode in the oxidizer post, resulting in propellant mass flow rate oscillations. Substantial evidence demonstrates that the intensities of heat release rate near the end walls are higher than that detected around the chamber center due to its intense coupling to the local first transverse (1W) acoustic mode. Flow field animations indicate strong deflection of oxygen posts and liquid particles by a transverse moving gas mixture leading to flame interactions and collision between adjacent share injectors. The propellant injected into the chamber does not burn immediately and is transported to the combustor end wall. Propellant mixing performance is significantly enhanced during interactions between acoustic pressure and end wall, causing a sudden heat release near 1W pressure anti-node, thereby strengthening pressure oscillations. Finally, a combustion instability driving mechanism associated with propellant mass flow rate oscillations, and interactions between unsteady acoustic pressure and heat release rate is proposed. The insights obtained from this work can aid comprehension of transverse combustion instabilities encountered by rocket engines.

Driving Mechanisms in Low-Order Modeling of Longitudinal Combustion Instability

Several test cases in the literature have shown that both transverse and longitudinal high-frequency combustion instability can be driven by the injector dynamics. In these cases, pressure oscillations result in fluctuations in propellant mass flow rate, which yields pulsing heat release. This fundamental mechanism is the focus of the present work, with the aim of including this effect in a quasi-1D nonlinear model of Euler equations suited to studies of longitudinal combustion instability. In particular, the injection dynamics is represented through a simplified formulation, which is the core of the proposed response function. The analysis also addresses the influence of combustion efficiency on the main characteristics of the resulting limit cycle (frequency and amplitude). The obtained model is tested comparing the quasi-1D simulations against the experimental data of the continuously variable resonance combustor available in the literature, considering three different geometrical configurations, with different lengths of the oxidizer post. The proposed formulation is capable of reasonably reproducing the unstable behavior, as well as providing a simple model that explains the mechanism that leads to a low average combustion efficiency during unstable operation.

The Experimental Investigation of a 98% Hydrogen Peroxide Monopropellant Thruster Comprising the Metal-Foam-Supported Manganese Oxide Catalyst

The article presents alternative metal-supported catalysts for decomposition of the highest-class hydrogen peroxide: 98%+ (Type 98 HP, according to MIL-PRF-16005F). The aim of this study was the experimental investigation of an alternative solution for decomposition of 98%+ hydrogen peroxide, strictly for chemical propulsion. High-porosity open cell metal foams have been identified as structures with great potential. Low density, good mechanical and thermal properties, availability of various materials and alloys as well as new technologies of manufacturing, make metal foam a potential solution for many different propellants, not only hydrogen peroxide. Open cell NiCrAl foam has been processed to prepare several catalysts, with different content and dispersion of the active phase. Cleaning and drying were performed to prepare carriers for further processing: wet impregnation, slow drying, and calcination. Simple drop tests with 98% hydrogen peroxide have been conducted to estimate activity level of catalysts, in a simplified scale: low, medium, high. Then, real-environment tests have been performed in a catalyst bed. Temperature, pressure along the bed and propellant mass flow rate were measured while testing. The analysis of the test results provided a general conclusion that metal foam supported manganese oxide catalyst is a promising solution for hydrogen peroxide propulsion.

Design and performance evaluation of a novel H2/O2 electro-chemical hybrid thruster

A H2/O2 electro-chemical hybrid thruster with a novel design for efficient coupling of hydrogen–oxygen detonation combustion and plasma acceleration is constructed and tested. The comprehensive effects of various parameters, such as the delay time, the background pressure, the capacitance, the external magnetic field, and the propellant mass flow rate, on the performance of the thruster have been explored in single-pulse mode. Several strategies for enhancing the efficiency have been proposed. It is found that, by adding an external magnetic field and decreasing the storage capacitor under the low pressure of 30 Pa and higher single-pulse gas intake at appropriate delay time, the average thrust and efficiency can be enhanced up to 73.2 mN and 15.58%, respectively. In addition, the maximal specific impulse of 676 s is also achieved. The results strongly indicate the great potential of the proposed electro-chemical hybrid thruster in achieving large thrust and high-level specific impulse.

Hot Firing Tests of a Novel Green Hypergolic Propellant in a Thruster

This study investigates the ignition and combustion behavior of the recently developed green hypergolic combination named HIP_11 in a small thruster. The combination HIP_11 consists of highly concentrated hydrogen peroxide as oxidizer and a fuel based on the ionic liquid 1-ethyl-3-methylimidazolium thiocyanate (EMIM SCN) with addition of copper thiocyanate (Cu SCN) to facilitate the hypergolic ignition. A dedicated thruster, featuring a 2-on-1 unlike impinging jet injector, was designed and manufactured. The initial tests were conducted with the injector only, without the combustion chamber mounted. These tests were conducted with two different amounts of catalyst: 1% and 5% Cu SCN. Subsequently, three hot firing series (around 35 tests) were conducted with the complete thruster, using the propellant combination 97% high-test peroxide and EMIM Cu SCN. The propellant mass flow rates were varied between 7 and and the oxidizer-to-fuel ratio from 3 to 7.

Design and Optimization of vacuum Intake for Atmosphere-Breathing electric propulsion (ABEP) system

Atmosphere-breathing electric propulsion system is a concept for space mission for ultra-low Earth orbit, which can collect the rarefied atmospheric particles as propellant to compensate the aerodynamic drag. In the best case, it can extend the lifetime of spacecraft without carrying any propellant. As the key component of Atmosphere-breathing electric propulsion system, intake device can capture the ambient particles and drive propellant to electric thruster. Considering the requirement of propellant mass flow and ionization efficiency, the performance of intake needs to be improved, including capture efficiency and compression ratio. The previous studies about intake device are summarized and analyzed firstly in this paper, and a new intake device is proposed that consists of tapered chamber, grid ducts and tube. Then, the inlet boundary conditions are set based on the NRLMSISE-00 atm model, such as particle species, natural temperature, number density and so on. Lastly, the intake performance is analyzed under different gas-surface interaction models through Direct Simulation Monte Carlo method. Simulation results show that the combination of parabola chamber, honeycomb grid ducts and a certain length of tube is benefit to intake device under the mixture reflection model of specular and diffuse. The capture efficiency is 65.79% with a high compression ratio of 210.2.

Characterization of a radio-frequency inductively coupled electrothermal plasma thruster

A radio-frequency (RF) inductively coupled electrothermal plasma thruster operating with argon is experimentally characterized for different mass flow rates, RF powers, and propellant injection configurations. Depending on the propellant mass flow rate, significant neutral gas heating is observed with effective stagnation temperatures around 2000 K (giving a maximum estimated thrust and specific impulse of about 100 mN and 125 s, respectively) for absorbed powers between 300 and 500 W. A self-consistent theoretical discharge model is developed and used to study the basic physics and operation of RF electrothermal thrusters, and predictions of the gas temperature are in good agreement with experimental measurements. The model identifies primary power inefficiencies as electron-neutral excitation losses and neutral gas heat losses to the thruster walls. Both experimental and theoretical results indicate that a relatively high stagnation pressure (of the order of 100 Torr or higher) is critical for high performance. For pressures significantly below this the electron-neutral collisional power transfer is too low to effectively heat the neutral gas.

Combustion Characteristics of Multi-Element Swirl Coaxial Jet Injectors under Varying Momentum Ratios

The combustion characteristics of a staged combustion cycle engine with an oxidizer-rich preburner were experimentally studied at different momentum ratios of multi-element injectors. Propellants were simultaneously supplied as a liquid–liquid–liquid system, and an injector was designed in which a swirl coaxial jet is sprayed. The injector burned the propellants in the inner chamber which had a temperature greater than 2000 K. To cool the combustion gas, a liquid oxidizer was supplied to the cooling channel outside the injector. To prevent the turbine blades from melting, the temperature of the combustion gas was maintained below 700 K. To confirm the combustion characteristics at different momentum ratios of the high-temperature combustion gas inside the injector and the low-temperature liquid oxidizer outside the injector, three types of injectors were designed and manufactured with different momentum ratios: MR 3.0, MR 3.3, and MR 3.7. In this study, the results of the combustion test for each type were compared for 30 s. For ORPB-A, a combustion pressure of 18.5 MPaA, fuel mass flow rate of 0.26 kg/s, oxidizer mass flow rate of 15.3 kg/s, and turbine inlet temperature of 686 K were obtained in the combustion stability period of 29.0–29.5 s. The combustion efficiency was 98% for MR 3.0 (ORPB-A), which was superior to that for other momentum ratios. In addition, during the combustion test for MR 3.0, the fluctuations in the characteristic velocity, combustion pressure, and propellant mass flow rate were low, indicating that combustion was stable. The three types of combustion instability were all less than 0.8%, thus confirming that the combustion stability was excellent.

Numerical investigation of the plasma processes for propellant heating in electrothermal plasma thruster for nanosatellites

AbstractNumerical investigation of the plasma processes in a cylindrical chamber with small dimensions of a novel microwave electrothermal plasma thruster for nanosatellites has been conducted. The absorbed microwave power from the electrons in the plasma column of the surface wave discharge is included in the computational model as a heat source with Gaussian distribution. The computational model takes into account the elastic and inelastic collisions of the electrons with the atoms in the ground state and two excited states (−s, −p) and the processes of recombination and deactivation of the plasma species in the volume and on the walls of the chamber. The computational model includes the flow of neutral gas and the processes in the plasma for effective heating of neutral particles by collisions not only with electrons but also with ions. Selected combinations of input power and propellant mass flow rates are used as initial parameters for the numerical investigation. The results show that at higher mass flow rates the heating of the neutral gas is more effective and at power levels of 4 W and propellant mass flow rate of 3 mg/s the electrothermal plasma thruster demonstrates effective performance and thrust levels in the order of 1 mN.

Cathode erosion site distributions in an applied-field magnetoplasmadynamic thruster

Erosion can influence cathode life, and is thus considered to be one of the main factors limiting the application of applied-field magnetoplasmadynamic thrusters. In this paper, erosion sites on graphite cathodes are studied so as to identify the influence of applied magnetic field and the ratio of propellant mass flow rate supplied from cathode and anode. The experiment results show that the application of applied magnetic field can significantly reduce the erosion rate of the cathode compared to that without magnetic field. The erosion sites on the cathode vary with the relative position of the convergent-divergent magnetic field, and are mainly distributed in the divergent part of the field. The erosion sites on the cathodes are found to be related to the propellant supply. The decreasing anode mass flow rate enlarges the range of erosion. These results are much helpful for the analysis of cathode erosion site location since they provide evidences of erosion mechanisms and point out the directions for further research.

Performance Parameter Analysis of a Hall Effect Thruster With Modified Bohm Parameter Model

The Hall effect thruster (HET) is an electric propulsion system in which electrons confined in a certain region collide with neutral particles due to the Hall effect. The neutral particles are ionized, and the ions are accelerated by the electric field to produce thrust. Thus, a Hall thruster has a simpler structure than an ion thruster. To improve the accuracy of the HET performance analysis, the energy change of electrons is predicted by considering thermal conduction. The number of ions produced by ionization and the change in electron energy via multivalent ionization are investigated. To minimize the independent variables, the boundary condition is modified after considering the physical correlation of each variable, and a semi-experimental model equation for the Bohm parameter is proposed. The performance results have greater similarity to the experimental results than those of the previous analytical model. Based on this, the performance changes according to the thruster operating conditions are predicted and compared. It is confirmed that the thrust increases by about 10–20% according to the change of propellant mass flow rate due to the increase of ion exit velocity.

Experimental Investigation of Decomposition with Cylindrical Pellets Catalyst

An experimental study on the catalytic decomposition of the nitrous oxide with catalysts based on palladium as active phase on alumina pellets support has been carried out and is presented in this paper, with the aim of inquiring the potentialities of nitrous oxide for space propulsion applications, especially as monopropellant for low-thrust engines. Different tests were carried out, with different system configurations and varying the propellant mass flow rate, to investigate the main dependences of the decomposition efficiency with the characteristic parameters of the system. The results showed that an optimal operating condition exists in terms of the catalytic bed load and the gas hourly space velocity. Finally, a procedure was designed for increasing the catalytic system initial temperature, and a new experimental campaign was carried out, highlighting the effect of the latter parameter on the nitrous oxide decomposition behavior.