William A. Hargus

Edwards Air Force Base

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Mark Cappelli

Stanford University

Michael Nakles

Edwards Air Force Base

John W. Bennewitz

University of Alabama in Huntsville

Stephen Danczyk

United States Air Force Research Laboratory

Blaine R. Bigler

Jacobs (United States)

Brian E. Beal

Edwards Air Force Base

N. B. Meezan

Lawrence Livermore National Laboratory

Christopher Lietz

Edwards Air Force Base

Natalia A. MacDonald

United States Air Force Research Laboratory

Alec D. Gallimore

University of Michigan–Ann Arbor

Cooperative Institutions

Edwards Air Force Base

United States Air Force Research Laboratory

Stanford University

University of Michigan–Ann Arbor

Electric Propulsion Laboratory (United States)

United States Air Force

University of California, Los Angeles

Glenn Research Center

Environmental Remediation Consultants (United States)

U.S. Air Force Institute of Technology

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Ion Velocity Measurements Within the Acceleration Channel of a Low-Power Hall Thruster

IEEE Transactions on Plasma Science (2008)

William A. Hargus Michael Nakles

This paper presents axial ion velocity measurements within the acceleration channel of the 200-W Busek Company Inc. BHT-200 laboratory Hall thruster derived from laser-induced fluorescence measurements of the 5d[4] _7/2 - 6p[3] _5/2 xenon-ion excited-state transition. Acceleration-channel-centerline ion velocities were measured for one nominal and six related cases. These six cases were chosen to be representative of small variations of the applied propellant flow, magnetic field, and discharge potential from the nominal condition. These deviations in operating parameters translate into changes in the plasma density, electron transport, and applied electric field, respectively. The effect of varying the magnetic field, hence influencing the electron transport, is to adjust the location of the internal ion acceleration. Increasing the anode propellant flow, which proportionally increases the plasma density and also influences the electron transport, appears to shift the acceleration upstream. Increasing the discharge potential increases ion acceleration proportionally. Examinations of the fluorescence traces, which have been previously shown to be representative of the ion velocity distributions, are also undertaken. From these data, it is possible to estimate internal axial electric fields and identify regions of ion acceleration and creation.

Ion thruster

Pulsed Power

10.1109/tps.2008.2003967

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Citations (41)

MODAL TRANSITIONS IN ROTATING DETONATION ROCKET ENGINES

International Journal of Energetic Materials and Chemical Propulsion (2019)

John W. Bennewitz Blaine R. Bigler J. J. Pilgram William A. Hargus

Various modal transitions within a rotating detonation rocket engine (RDRE) are investigated in detail within this experimental study. Using direct high-speed visible imaging along with a recently developed processing technique, detonation mode transition events are captured and analyzed with a focus on quantifying the unsteady wave propagation behavior during a modal shift. Specifically, three transition types are discussed including a rotational direction reversal where the number of waves is held constant, an increase in the number of detonation waves, as well as a decrease in the number of waves. The rotational direction reversal exhibits intermittent counter-propagating wave behavior, which eventually results in the opposing set of waves overtaking the dominant set. The ascending and descending modal transitions are attributed to significant galloping-type detonation propagation that grows in severity for the descending shifts and decays during the ascending transitions. In general, the exponential growth/decay rate αavg,pk of the angular separation associated with the wave pairs δθ is larger for the ascending transitions than descending, where the higher wave mode is more stable (i.e., δθ' averages 8° for the three-wave mode and 39° for the two-wave mode). Also, the maximum peak-to-peak δθ' amplitude for the wave pairs separating the threshold between the two- and three-wave modes is on average 135–138° and U'wv/Uwv equals 24% for these two transition types. In total, this work provides notable insight into modal behavior that can affect the stability of RDRE's.

Mode (computer interface)

Rocket (weapon)

10.1615/intjenergeticmaterialschemprop.2019027880

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Citations (56)

Preliminary Study of Arcjet Neutralization of Hall Thruster Clusters

Quentin Walker Mark Cappelli William A. Hargus John H. Schilling

Clustered Hall thrusters have emerged as a favored choice for extending Hall thruster options to very high powers (50 kW – 150 kW). This paper examines the possible use of an arcjet to neutralize clustered Hall thrusters, as the hybrid arcjet-Hall thruster concept can fill a performance niche amongst available propulsion options. We examine missions on which this hybrid concept would be a competitive or favored thruster option, report on fundamental experiments to understand how much electron current can be drawn to a surrogate anode from the plume of low power arcjets operating on hydrogen and helium, and then demonstrate the first successful operation of a low power Hall thruster-arcjet neutralizer package. In the surrogate anode studies, we find that the drawing of current from the arcjet plume has only a weak effect on overall arcjet performance (thrust), with a slight decrease in arc voltage with increased extracted current. A single Hall thruster – arcjet neutralizer package was constructed for the hybrid concept demonstration. The arcjet operated at very low powers (~ 70-120W) on helium, at a mass flow rate of 4.5 mg/s, and was able to effectively neutralize the ~ 200 – 900W xenon Hall thruster causing little measurable departure from the hollow-cathode neutralized Hall thruster VI characteristics up to 250V. At higher helium mass flow rates, the Hall discharge current is slightly perturbed from its expected values, due most likely to the ingestion of helium. Further developments of the hybrid concept to clustered configurations and higher powers will require a vacuum facility that can pump tens of milligrams of helium while maintaining the low pressures needed for normal xenon Hall thruster operation.

Arcjet rocket

Ion thruster

Mass flow rate

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Citations (1)

Thermographic Characterization and Comparison of 200W and 600W Hall Thruster (Postprint)

Taylor S. Matlock William A. Hargus C. W. Larson

Abstract : Thermal analysis of Hall thrusters is necessary for both performance optimization and spacecraft integration. The thermal characteristics of the thruster influence the lifetime, energy losses, and spacecraft integration. The lifetime of most Hall thrusters is limited by the erosion of the insulating materials within the discharge chamber, which may vary with temperature. Thruster thermal data are also indicative of thruster energy losses as well as the heating characteristics of spacecraft interface surfaces. The results of the thermographic imaging of two laboratory Hall thrusters, a 200W (BusekBHT-200-X3) and 600W (BusekBHT-600), are presented. Surface temperature profiles were obtained using an infrared camera (7-13 micrometres), independently verified by thermocouples. Infrared imagery of thruster start-up, steady-state, and shut-down was recorded and used to approximate the transient heating behavior of each thruster. Variation of the nominal mass flow rate (resulting in proportional variation of the anode current and power level) between 85% to 115% resulted in proportional changes to the thruster surface temperatures.

Thermocouple

Transient (computer programming)

Ion thruster

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Ideal Thermodynamic Performance Results for Rotating Detonation Rocket Engine Thrust Chambers Using CEA

AIAA SCITECH 2022 Forum (2020)

Eric J. Paulson Randon C. Kimura William A. Hargus Venke Sankaran

Detonation engine cycles have been extensively studied in the past, using ideal thermodynamic cycles such as constant volume combustion (Humphrey cycle) or modeled as a shock with Rayleigh heat addition (Zeldovich-vonNeumann-Döring cycle). Constant properties along with the assumption of complete combustion, overpredicts the efficiency and specific impulse gains that ideal detonation engines can achieve. In this paper, a modified ZND rotating detonation rocket engine thermodynamic model, for arbitrary gaseous reactants is developed to predict ideal rocket engine performance for parametric concept studies. Next, a tool is developed incorporating NASA's CEA code to calculate chamber specific impulse using equilibrium combustion. The tool is compared with existing 2-D premixed Large Eddy Simulations and existing experimental data. Next the tool is then used to evaluate propellant combinations of methane-oxygen, hydrogen-oxygen, propene-oxygen, and JP-10-oxygen for varying equivalence ratios and temperatures for state-of-the-art combustion chamber pressures. The results show that an ideal Rotating Detonation Rocket Engine can significantly reduce the combustor inlet pressure while matching Isp performance of an equivalent ideal Brayton rocket cycle.

Specific impulse

Rocket engine

Brayton cycle

Thermodynamic cycle

Liquid-propellant rocket

Liquid oxygen

Ideal gas

10.2514/6.2020-0193

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Citations (4)

Interior and exterior laser-induced fluorescence and plasma potential measurements on a laboratory Hall thruster

28th Joint Propulsion Conference and Exhibit (1999)

William A. Hargus Mark Cappelli

Abstract : In this paper, we describe the results of a study of laser induced fluorescence velocimetry of ionic xenon in the plume and interior acceleration channel of a laboratory Hall type thruster operating at powers ranging from 250 to 725 W. Optical access to the interior of the Hall thruster is provided by a 1 mm axial slot in the insulator outer wall. Axial ion velocity profiles for four discharge voltages (100 V, 160 V, 200 V, 250 V) are measured as are radial velocity profiles in the near field plume. Internal neutral xenon axial velocity profiles are also measured at these conditions. For comparison, the plume plasma potential profile is measured with an emissive probe. These probe based potential measurements extend from 50 mm outside the plume to the near anode region for all but the highest discharge voltage condition. For each condition, the axial electric field is calculated from the plasma potential. In addition, an estimate of the local electron temperature is calculated from the Bohm criterion at the location of each plasma potential measurement.

10.2514/6.1999-2721

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Citations (10)

Ignition and Plume Characteristics of Low-Current Cerium and Lanthanum Hexaboride Hollow Cathodes

Journal of Propulsion and Power (2009)

Dustin J. Warner Richard Branam William A. Hargus

This research investigated cerium hexaboride (CeB 6 ) and lanthanum hexaboride (LaB 6 ) as emitters in 6.4-mm-diam hollow cathodes used for low-power electric propulsion applications. Each cathode's ignition performance and plume behavior is presented and discussed. Two identical cathodes, with the exception of the insert material, were assembled and integrated into a cylindrical anode configuration without an applied magnetic field. The experiments tested two orifice geometries and a wide range of xenon flow rates (1.5-4.5 sccm) and anode currents (2-6 A) for each cathode. Also investigated was cathode performance for these conditions with the keeper on and off. A langmuir probe collected plume data to include plasma densities, electron temperature, plasma potential, and floating potential. An oscilloscope monitored the voltage behavior of the keeper electrode and anode. Initially, both cathodes successfully started without excessive keeper voltages (<650 V), heating times (<15 min), or flow rates (<4.5 sccm). The LaB 6 cathode consistently started at these same conditions. However, the CeB 6 cathode demonstrated some degradation after time by requiring higher fiow rates and a longer heating time to start. Both cathodes demonstrated ideal (spot mode) operation for flow rates as low as 1.5 sccm at anode currents as high as 5 A with no heater or keeper power.

Langmuir Probe

Hot cathode

Cold cathode

10.2514/1.41386

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Citations (23)

Detonation Wave Collision Study During Rotating Detonation Rocket Engine Counter-Propagating Modes

AIAA Propulsion and Energy 2020 Forum (2021)

Armani Batista Blaine R. Bigler Mathias Ross Christopher Lietz William A. Hargus

Rocket engine

Rocket (weapon)

10.2514/6.2021-3659

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Citations (4)

Synchronized Measurement of Plasma Characteristics In a Hall Effect Thruster

54th AIAA Aerospace Sciences Meeting (2016)

David A. Cunningham David Liu Carl Hartsfield Carl Mullins Casey C. Farnell

Hall eect thrusters are a key propulsion technology for many satellite missions in low Earth orbit and beyond. Amethod of synchronized data collection was developed by combining high speed imaging data, emissive probe measurements of plasma potential, and magnetic measurement of the Hall current with eective temporal resolution of ± 0.75μs. is new method allows the comparison of features in the visible plasma emission with variations in the plasma potential and Hall current. Observation of plasma characteristics during the breathing mode and passage of azimuthal oscillations in the thruster was demonstrated.

Ion thruster

10.2514/6.2016-1943

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Citations (0)

Spacecraft interaction test results of the high performance Hall system SPT-140

36th AIAA/ASME/SAE/ASEE Joint Propulsion Conference and Exhibit (2000)

John Fife William A. Hargus Donald A. Jaworske C. Sarmiento Lee S. Mason

10.2514/6.2000-3521

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Citations (23)