62 results found
Staab D, Longhi E, Garbayo A, et al., 2021, ICE: A MODULAR WATER ELECTROLYSIS PROPULSION SYSTEM, SPACE PROPULSION CONFERENCE 2020+1
Rosati Azevedo E, Tirila V, Schwertheim A, et al., 2021, MAGNETIC FIELD ENHANCEMENT OF THE QUAD CONFINEMENT THRUSTER (QCT): DESIGN AND EARLY DEVELOPMENT OF THE QCT PHOENIX, SPACE PROPULSION 2020+1
Masillo S, Lucca Fabris A, Karadag B, et al., 2021, EXPERIMENTAL CHARACTERISATION OF THE NOVEL HALO PLASMA THRUSTER FOR SMALL SATELLITE APPLICATIONS, SPACE PROPULSION 2020+1
Schwertheim A, Knoll A, 2021, PERFORMANCE CHARACTERISATION OF THE WATER ELECTROLYSIS HALL EFFECT THRUSTER (WET-HET) USING DIRECT THRUST MEASUREMENTS, SPACE PROPULSION 2020+1
Muir C, Ma C, Knoll A, 2021, The design, fabrication and test progress summary of the iridium catalysed electrolysis thruster, Space Propulsion 2020+1
Schwertheim A, Rosati Azevedo E, Liu G, et al., 2021, Interlaboratory validation of a hanging pendulum thrust balance for electric propulsion testing, Review of Scientific Instruments, Vol: 92, Pages: 1-11, ISSN: 0034-6748
A hanging pendulum thrust balance has been developed by Imperial College London in collaboration with the European Space Agency (ESA) to characterize a wide range of static fire electric propulsion and chemical micro-propulsion devices with thrust in the range of 1 mN to 1 N. The thrusters under investigation are mounted on a pendulum platform, which is suspended from the support structure using stainless steel flexures. The displacement of the platform is measured using an optical laser triangulation sensor. Thermal stability is ensured by a closed loop self-compensating heating system. The traceability and stability of the calibration are ensured using two separate calibration subsystems: a voice coil actuator and a servomotor pulley system. Two nearly identical thrust balances have been constructed, with one being tested in the Imperial Plasma Propulsion Laboratory and the other in the ESA Propulsion Laboratory. Both balances show a high degree of linearity in the range of 0.5 mN–100 mN. Both instruments have demonstrated a stable calibration over several days, with an estimated standard deviation on thrust measurements better than 0.27 mN for low thrust measurements. The same electric propulsion test article was used during both tests: a Quad Confinement Thruster (QCT) variant called QCT Phoenix. This thruster differed from previous QCT designs by having a newly optimized magnetic topology. The device produced thrust up to 2.21 ± 0.22 mN with a maximum specific impulse of 274 ± 41 s for an anode power range of 50 W–115 W.
Schwertheim A, Knoll A, 2021, Low power thrust measurements of the water electrolysis Hall effect thruster, CEAS Space Journal, ISSN: 1868-2502
We propose that a Hall effect thruster could be modified to operate on the products of water electrolysis. Such a thruster would exploit the low cost and high storability of water while producing gaseous hydrogen and oxygen in-situ as they are required. By supplying the anode with oxygen and the cathode with hydrogen, the poisoning of the cathode is mitigated. The water electrolysis Hall effect thruster (WET-HET) has been designed to demonstrate this concept. The dimensions of the WET-HET have been optimized for oxygen operation using PlasmaSim, a zero-dimensional particle in cell code. We present the first direct thrust measurements of the WET-HET. A hanging pendulum style thrust balance is used to measure the thrust of the WET-HET while operating in the Boltzmann vacuum facility within the Imperial Plasma Propulsion Laboratory. For this test the beam was neutralized using a filament plasma bridge neutralizer operating on krypton. We find thrust, specific impulse, and thrust efficiency all increase linearly with power for values between 400 and 1050 W. Increasing the mass flow rate from 0.96 to 1.85 mg/s increases thrust at the expense of specific impulse. Changing mass flow rate was found to have little impact on the thrust efficiency over this range. An optimal radial magnetic flux density of 403 G at the exit plane is found. Further increases to the magnetic field beyond this point were found to decrease the thrust, specific impulse and thrust efficiency, whereas the discharge voltage increased monotonically with increasing magnetic field for a given input power. It was found that the experimental thruster performance was lower than the simulation results from PlasmaSim. However, the general trends in performance as a function of power and propellant mass flow rate were preserved. We attribute a portion of this discrepancy to the inability of the simulation to model the energy absorbed by the covalent bond of the oxygen molecule. For the powers and mass flow rat
Williams RD, Fabris AL, Knoll A, 2020, Insight into the plasma structure of the Quad Confinement Thruster using electron kinetic modelling, Acta Astronautica, Vol: 173, Pages: 111-118, ISSN: 0094-5765
The behaviour of plasma within the discharge channel of the Quad Confinement Thruster is studied on the basis of electron kinetics. Here we propose that E × B drift of electrons drives the formation of unusual quadrant dependent light emitting structures observed experimentally in the discharge channel of the Quad Confinement Thruster. This assertion is made on the basis of a theory-based analysis and a computational model of the Quad Confinement Thruster. A particle orbit model of electron motion under the influence of applied electric and magnetic fields was used to assess electron transport. Structures strongly resembling that of the observed visible emission regions were found in the electron density distribution within the channel. While the motion of electrons cannot be decoupled from the motion of ions, as in this simple electron kinetic approximation, the results of this analysis strongly indicate the physical mechanism governing the formation of the non-uniform density distributions within the Quad Confinement Thruster channel.
Knoll A, Harle T, Peter S, et al., 2020, Plasma generation, US10595391B2
A plasma torch having an open end from which a plasma plume is emitted in use is disclosed. The plasma torch includes a central cathode rod, a grounded conductive tube having an open end and being arranged around the cathode and spaced therefrom to form a first cylindrical cavity open at one end; and a high voltage electrode having a dielectric barrier material at a radially inward-facing surface thereof and being arranged around the grounded conductive tube and spaced apart therefrom to form a second annular cylindrical cavity open at one end. A constant direct current (DC) electrical power plus a high voltage pulsed electrical power is provided to the cathode producing an arc discharge in the first cavity between the cathode and grounded tube to generate a central thermal plasma emitted at an open end of the first cylindrical cavity. A high voltage alternating current electrical power or pulsed electrical power is provided to the high voltage electrode producing a dielectric barrier discharge in the second annular cylindrical cavity to generate a non-thermal plasma emitted from an open end of the second cavity as a halo around the central thermal plasma.
Karadag B, Masillo S, Moloney R, et al., 2020, Experimental Investigation and Performance Optimization of the Halo thruster, 36th International Electric Propulsion Conference
Schwertheim A, Knoll A, 2020, The Water Electrolysis Hall Effect Thruster (WET-HET): Paving the Way to Dual Mode Chemical-Electric Water Propulsion, 36th International Electric Propulsion Conference
Mörtl M, Knoll A, Williams V, et al., 2020, Enhancing Hall Effect Thruster Simulations with Deep Recurrent Networks, 36th International Electric Propulsion Conference
Williams V, Argyriou V, Shaw P, et al., 2019, Development of PPTNet a Neural Network for the Rapid Prototyping of Pulsed Plasma Thrusters, 36th International Electric Propulsion Conference
Muir C, Knoll A, 2019, Catalytic Combustion of Hydrogen and Oxygen for an Electrolysis Micro-Propulsion System, Journal of the British Interplanetary Society, ISSN: 0007-084X
Schwertheim A, Knoll A, 2019, In situ Utilization of Water as a Propellant for a Next Generation Plasma Propulsion System, Journal of the British Interplanetary Society, ISSN: 0007-084X
Muir C, Knoll A, 2018, Catalytic Combustion of Hydrogen and Oxygen for an Electrolysis Micro-Propulsion System, 16th Reinventing Space Conference
Schwertheim A, Knoll A, 2018, In situ Utilization of Water as a Propellant for a Next Generation Plasma Propulsion System, 16th Reinventing Space Conference
Knoll AK, 2018, Magnetic Holograms
Liu J, Knoll AK, Fabris AL, et al., 2018, Numerical investigation of magnetic neutral points in the Halo thruster, Space Propulsion Conference 2018
Fabris AL, Wantock T, Gurciullo A, et al., 2018, OVERVIEW OF HALO THRUSTER RESEARCH AND DEVELOPMENT ACTIVITIES, Space Propulsion Conference 2018
Fabris AL, Knoll AK, Young C, et al., 2017, Ion Acceleration in a Quad Confinement Thruster, The 35th International Electric Propulsion Conference
Fabris AL, Knoll AK, Dannenmayer K, et al., 2017, Vacuum Facility Effects on Quad Confinement Thruster Testing, The 35th International Electric Propulsion Conference
Gurciullo A, Fabris AL, Knoll AK, 2017, Alternative Neutralization Technologies Enabling the Use of Exotic Propellants in Electric Propulsion, International Electric Propulsion Conference 2018
Gurciullo A, Lucca Fabris A, Knoll AK, 2017, Direct current plasma electron source for electric propulsion applications using atomic and molecular propellants, IEEE Transactions on Plasma Science, Vol: 45, Pages: 2472-2480, ISSN: 0093-3813
The design and performance of a novel direct current (dc) neutralizer for electric propulsion applications are presented. The neutralizer exploits an E×B discharge to enhance ionization via electron-neutral collisions. Tests are performed with helium, argon, xenon, air, and water vapor as working gases. The I – V characteristics and extraction parameters are measured for both atomic and molecular gases. The maximum partial power efficiency is 4.2 mA/W in argon, 2.7 mA/W in air, and 2 mA/W in water vapor. The typical utilization factor is below 1 and the power consumption is less than 120 W. A semiempirical model is derived to predict the performance of dc plasma cathodes using atomic gas. A comparison with existing plasma cathodes and conventional LaB6 cathodes is presented, and design optimizations aimed at improving the performance are proposed.
David AO, Knoll AK, 2017, Experimental Demonstration of an Aluminum-Fueled Propulsion System for CubeSat Applications, Journal of Propulsion and Power, Vol: 33, Pages: 1320-1324, ISSN: 0748-4658
Knoll AK, Bianco P, 2016, Charge separation mechanism
Ryan C, Wantock T, Harle T, et al., 2016, Performance Characterization of the Low-Power Halo Electric Propulsion System, Journal of Propulsion and Power, Vol: 32, Pages: 1544-1549, ISSN: 0748-4658
Performance measurements have been obtained of a novel propulsion concept called the Halo thruster under development within the University of Surrey. The Halo thruster, a type of cusped-field thruster with close similarity to the cylindrical Hall thruster, is motivated by the need for low-power and low-cost electric propulsion for the small satellite sector. Two versions of the device are investigated in this study: a design using permanent magnets at high magnetic-field strength and a design using electromagnets with moderate field strength. While operating at 200 W discharge power, which is of particular interest to power-limited small satellite platforms, the permanent-magnet design achieved a maximum thrust efficiency of 8% at a specific impulse of approximately 900 s using a krypton propellant. By comparison, the electromagnet design achieved a maximum thrust efficiency of 28% at a specific impulse of approximately 1500 s at 200 W using a xenon propellant. For higher levels of power (tested up to 800 W), the performance of the electromagnetic design saturated at approximately 25% thrust efficiency using krypton and 30% using xenon. The thrust efficiency of the permanent-magnet design appeared to increase monotonically up to 600 W reaching a maximum value of 14%.
Gurciullo A, Knoll AK, 2016, Experimental performance characterization of a novel direct current cold cathode neutralizer for electric thruster applications, 52nd AIAA/SAE/ASEE Joint Propulsion Conference, Publisher: AIAA
The performance of a novel neutralizer for space applications based on a E×B dischargeis presented. Preliminary tests were carried out with argon gas and flow rates in therange of 5-10 SCCM. Electrons were extracted through an orifice of diameter 1.8 mm. Themaximum extracted current versus input power reported was 2.4 mA/W. The total powerinput, given by the sum of discharge power plus the extraction power, was in the range of40-90 W. During extraction tests, the discharge current was limited at 0.2 A due to limitin the cooling system. Future work will be focused on tests at various extraction orificediameters and cathode materials. Ultimately, xenon and non-conventional gases would betested as working gases.
Knoll AK, Harle T, Shaw P, et al., 2016, PLASMA GENERATION, GB2532195A
Ahmed O, Knoll AK, 2016, Effect of Fuel-to-Oxidiser Ratio on Thrust Generation of a Hybrid Al + NaOH + H2O Propulsion System for CubeSat Applications, Space Propulsion Conference 2016
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