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Hani Kamhawi

Publications and source records attributed to Hani Kamhawi.

At least 37 records · Page 2
All sourcesNASA NTRSDOE OSTI
NASA NTRSDate not supplied

Risk-Based Methodology for the Determination of Hall Thruster Performance Specifications

This work adapts manufacturing and metrology industry standards to create a risk-based approach for the determination of electric propulsion performance specifications. The developed process is applied to 208 total thrust measurements acquired using three different NASA Hall Effect Rocket with Magnet Shielding (HERMeS) Technology Demonstration Unit (TDU) 12.5-kW Hall thrusters in three different test facilities in order to generate thrust specifications at the 300 V/6.25 kW, 400 V/8.33 kW, 500 V/10.42 kW, and 600 V/12.5 kW operating conditions. These thrust specifications are used to perform a set of sample acceptance tests to illustrate the importance of adding additional conservatism using a parameter known of the Test Accuracy Ratio in order to control the risk to the thruster producer and consumer from false rejection and acceptance of thruster units.

HERMeS↗
NASA NTRSDate not supplied

Extended Performance Characterization of the 12.5-kW Advanced Electric Propulsion System Engineering Test Unit Hall Thruster

This work presents a summary of the first detailed performance assessment of the Advanced Electric Propulsion System (AEPS) Engineering Test Unit 2 (ETU-2) thruster produced by Aerojet Rocketdyne at the throttle conditions most relevant for AEPS application on the Gateway Power and Propulsion Element (PPE). First, an assessment was performed of ETU-2 performance and backpressure sensitivity at the previously-tested throttle points of 300 V/6.25 kW, 400 V/8.33 kW, 500 V/10.42 kW, and 600 V/12.5 kW that confirmed ETU-2 and the test facility were both operating nominally and consistent with historical baselines. ETU-2 performance and stability were then assessed throughout the PPE throttle range of 2.6-12 kW and shown to be in-family with predictions based on the scaling of previous results. Taken together, these results will help inform application of AEPS on PPE as the AEPS design progresses towards its Critical Design Review.

HERMeS↗
NASA NTRSDate not supplied

Risk-based Methodology for the Determination of Hall Thruster Performance Specifications

This work adapts manufacturing and metrology industry standards to create a risk-based approach for the determination of electric propulsion performance specifications. The developed process is applied to 208 total thrust measurements acquired using three different NASA Hall Effect Rocket with Magnet Shielding (HERMeS) Technology Demonstration Unit (TDU) 12.5-kW Hall thrusters in three different test facilities in order to generate thrust specifications at the 300 V/6.25 kW, 400 V/8.33 kW, 500 V/10.42 kW, and 600 V/12.5 kW operating conditions. These thrust specifications are used to perform a set of sample acceptance tests to illustrate the importance of adding additional conservatism using a parameter known of the Test Accuracy Ratio in order to control the risk to the thruster producer and consumer from false rejection and acceptance of thruster units.

HERMeS↗
NASA NTRSDate not supplied

Ion Velocity Characterization of the 12.5-kW Advanced Electric Propulsion System Engineering Hall Thruster

During development testing of the 12.5-kW Advanced Electric Propulsion System engineering unit Hall thruster, which is magnetically shielded, a laser-induced fluorescence test was performed. During this test, a third medium-energy ion population was found near the inner front pole cover in addition to two low-energy counter-streaming ion populations that were found in previous testing. This newly found ion population matched in characteristics with the single population found near the outer front pole cover. The measured characteristics of the medium-energy ions matched the behavior expected of them if they were energized by a plasma wave with magnetized electrons, such as a lower hybrid wave. Comparison of the data from this test to prior tests showed that this engineering thruster had very similar ion characteristics as the precursor laboratory thruster. The acceleration zone was found to move upstream with increasing background pressure, decreasing anode flowrate, and increasing magnetic field strength. For the low-energy ions, the energy of the ions arriving at the inner pole did not vary noticeably with background pressure but did increase with increasing magnetic field strength and decreasing anode flow rate. For the medium-energy ions, the energy of the ions increased with decreasing background pressure, decreasing anode flow rate, and increasing magnetic field strength. Testing at different cathode flow fraction showed that the energy of the low-energy ions from the cathode decreased with increasing cathode flow.

Electric propulsion↗
NASA NTRSDate not supplied

Expanded Performance Characterization of the NASA HERMeS Hall Thruster

This work presents a summary of a detailed performance assessment of the Hall Effect Rocket with Magnetic Shielding (HERMeS) Technology Demonstration Unit 3 (TDU-3) thruster at the throttle conditions most relevant for the Gateway Power and Propulsion Element (PPE). First, an assessment was performed of TDU-3 performance and stability at the previously-tested throttle points of 300 V/6.25 kW, 400 V/8.33 kW, 500 V/10.42 kW, and 600 V/12.5 kW that confirmed TDU-3 was operating nominally and consistent with historical baselines. TDU-3 performance and stability were then assessed throughout the PPE throttle range of 2.6-12 kW and shown to mirror the characteristics observed at the previously tested conditions. Finally, a statistical estimate for the thrust specification at the expanded set of throttle conditions was computed and found to yield similar tolerance ranges to those calculated for the heritage conditions. Taken together, these results provide strong evidence that the demonstrated performance characteristics of the HERMeS design are preserved at the expanded set of throttle conditions relevant to application on PPE.

HERMeS↗
NASA NTRSDate not supplied

Extended Performance Characterization of the NASA HERMeS Hall Thruster

This work presents a summary of a detailed performance assessment of the Hall Effect Rocket with Magnetic Shielding (HERMeS) Technology Demonstration Unit 3 (TDU-3) thruster at the throttle conditions most relevant for the Gateway Power and Propulsion Element (PPE). First, an assessment was performed of TDU-3 performance and stability at the previously-tested throttle points of 300 V/6.25 kW, 400 V/8.33 kW, 500 V/10.42 kW, and 600 V/12.5 kW that confirmed TDU-3 was operating nominally and consistent with historical baselines. TDU-3 performance and stability were then assessed throughout the PPE throttle range of 2.6-12 kW and shown to mirror the characteristics observed at the previously tested conditions. Finally, a statistical estimate for the thrust specification at the expanded set of throttle conditions was computed and found to yield similar tolerance ranges to those calculated for the heritage conditions. Taken together, these results provide strong evidence that the demonstrated performance characteristics of the HERMeS design are preserved at the expanded set of throttle conditions relevant to application on PPE.

HERMeS↗
NASA NTRSDate not supplied

Overview and Performance Characterization of Northrop Grumman’s 1 kW Hall Thruster String

Northrop Grumman (NG) Tactical Space SystemsDivisionhas embarked on the development and qualification of a high throughput, low power Hall Thruster String (HTS) using hardware designed and built in-house.Following the success of Mission Extension Vehicles 1 and 2, NG is currently developing the next generation in its lineupof satellite servicing capabilities, the Mission Robotics Vehicle(MRV)and Mission Extension Pod (MEP).MEP’s mission profile imposes highly demanding requirements upon the electric propulsion system. When surveying the industry for available systems, NGwas unable to identify any mature Hall thrustersystems that could satisfy the performance and lifetime requirements for MEP. Eventually, it was decided to vertically integrate the EP development process, partneringwith NASA Glenn Research Center to leverage ongoing development at GRC of a high throughput, low power Hall thruster. The components of the HTS have successfully passed PDR and are currently in the engineering development and test phase. Based on performancecharacterization testingof the development hardware,NG’s low power Hall thruster, dubbed the NGHT-1X, promises to deliver state-of-the-art performanceand lifetime for a sub-kW Hall thruster, achieving total efficiencies of 50-55% over a wide range of throttle conditions. Results of a seriesof characterization tests including integrated systems testingwith the PPUare presented.Qualification methodology andthruster lifetime verification is also discussed. Environmental qualification of the HTS components is expected to completein mid-2023 with a first flight in mid-2024.

Propulsion↗
NASA NTRSDate not supplied

Emitter Surface Temperature Measurements of a 25-A Class Hollow Cathode

The axial emitter surface temperature profile is experimentally characterized for a 25-A class hollow cathode operating across a range of operating conditions. Profiles were measured using aratio pyrometry diagnostic featuring a fiber optic inserted into the emitter region using a fast motion stage. Emitter temperature is found to increase with increasing discharge current, decreasing flow rate, and increasing magnetic field strength. Across a range of 10 –29 A, the maximum emitter temperature increased 176 ºC, an order of magnitude more than the changes measured for flow rate or magnetic fields strength. Profiles were generally self-similar across all operating conditions, and the difference between maximum and minimum emitter temperature never surpassed 10%, indicating that this cathode design results in satisfactory emitter utilization across its operating envelope

space propulsion↗
NASA NTRSDate not supplied

Extended Wear Testing of the 12-kW Advanced Electric Propulsion System Engineering Test Unit Hall Thruster

This work presents a summary of the first wear test of the 12 kW Advanced Electric Propulsion System (AEPS) intended to isolate the impact of discharge current on component wear rates. Testing was led by Aerojet Rocketdyne and performed with the AEPS Engineering Test Unit 2 (ETU-2) thruster. In total, approximately 860 hours of operation were accumulated split between operating conditions of 600 V/9 kW, 600 V/11 kW, and 600 V/12 kW. Thruster performance and stability were invariant throughout the wear test for all thruster throttle conditions and shown to be equal to the values previously measured with ETU-2. Inner front pole cover erosion rates were found to be invariant to discharge current as the measurements at all three operating conditions were equal to within the empirical uncertainty. Outer front pole cover erosion rates were found to be equal for the 600 V/9 kW and 600 V/12 kW conditions with operation at 600 V/11 kW yielding elevated erosion rates. Overall, the data shows that the AEPS thruster design has a high probability of meeting lifetime and performance requirements as the design proceeds to flight production and qualification.

Jason D Frieman↗
NASA NTRSDate not supplied

PPE Electric Propulsion Advancing to CDR: 12kW String

String level testing is a vital part of electric propulsion development. With the Power and Propulsion Element (PPE) spacecraft flying two new models of Hall Thrusters, this work is even more critical. PPE will have both 6kW and 12kW thruster strings onboard, all powered by Maxar power processing units (PPU). Both strings have recently undergone end-to-end string level hot fire testing to verify the ability of all components to work together. This entails operating the thruster with the Xenon Flow Controller (XFC) and PPUs under both nominal and off-nominal conditions to stress the system and determine weak points. Such testing has long been a standard Maxar practice since our first 1.5kW electric propulsion system in 2004. This paper will focus on Phase 1 of the 12kW testing recently completed at NASA Glenn. The primary conclusion of the test effort was that the 12kW string operated nominally and that the PPU and XFC could operate the thruster within specifications.

Electric propulsion↗
NASA NTRSDate not supplied

The Side Plume of Magnetically Shielded Hall Thrusters

A combination of near-thruster laser-induced fluorescence and far-field plasma probe measurements collected from two magnetically shielded Hall thrusters have shown that there exist an independent ion population of medium-energy (on average, 60-190 eV) exiting out the side (60° to 100° from firing axis) of the thrusters. Studying the trends with background pressure and known discrepancies between the laser-induced fluorescence and far-field plasma probe data suggest that at zero pressure the side plume is likely to exhibit higher energy and lower density than at facility background pressures. The possibility that these ions were produced by the modified two-stream instability is explored and while the agreement is good, there are areas of notable disagreement. Based on the idea that production of the side plume and channel exit striations are due to the same plasma wave actions, a theory was explored that explains why striations form under certain circumstances but not others. Another theory, based on the idea that side plume production only occurs over certain portion of the global discharge oscillation cycle, was also explored and areas of disagreement were identified. While the exact mechanism for side plume acceleration remains unclear, if the theory behind side plume production is correct, then side plume can be expected to be present in a wide range of magnetically-shielded as well as non-shielded Hall thrusters.

Hall thruster↗
NASA NTRSDate not supplied

The Side Plume of Magnetically Shielded Hall Thrusters

A combination of near-thruster laser-induced fluorescence and far-field plasma probe measurements collected from two magnetically shielded Hall thrusters have shown that there exist an independent ion population of medium-energy (on average, 60-190 eV) exiting out the side (60° to 100° from firing axis) of the thrusters. Studying the trends with background pressure and known discrepancies between the laser-induced fluorescence and far-field plasma probe data suggest that at zero pressure the side plume is likely to exhibit higher energy and lower density than at facility background pressures. The possibility that these ions were produced by the modified two-stream instability is explored and while the agreement is good, there are areas of notable disagreement. Based on the idea that production of the side plume and channel exit striations are due to the same plasma wave actions, a theory was explored that explains why striations form under certain circumstances but not others. Another theory, based on the idea that side plume production only occurs over certain portion of the global discharge oscillation cycle, was also explored and areas of disagreement were identified. While the exact mechanism for side plume acceleration remains unclear, if the theory behind side plume production is correct, then side plume can be expected to be present in a wide range of magnetically-shielded as well as non-shielded Hall thrusters.

Electric propulsion↗
NASA NTRSDate not supplied

High-Propellant Throughput Sub-kW Electric Propulsion System for Deep Space Science and Exploration

The National Aeronautics and Space Administration (NASA) is maturing high-propellant throughput sub-kilowatt electric propulsion technologies to enable small spacecraft deep space science and exploration missions with high delta-v requirements. The pathfinder model (PM) propulsion system consists of the H71M-PM Hall-effect thruster, a breadboard 1-kW power processing unit (PPU), and a propellant flow control system. The propulsion system requirements balance the needs of various high delta-v NASA and commercial industry mission concepts to achieve a design that both enables a variety of NASA small spacecraft deep space missions, while remaining viable for select commercial applications. The H71M-PM thruster has completed performance characterization and three 500-h short duration wear tests (SDWT). The propulsion system provides stable thrust generation over a wide range of operating conditions from 200 W to 1 kW, and 200 V to 400 V. The thruster has demonstrated a thrust as high as 68 mN at 300 V and 1 kW. The thruster has similarly demonstrated a specific impulse of 1850 s at 400 V and 1 kW. Key surfaces were machined between each SDWT to simulate accelerated discharge channel and pole cover erosion. Profilometry scans across masked pole cover surfaces were conducted to determine erosion rates. SDWT results support that a target thruster lifetime of 14 kh with 50% margin is achievable. Component testing has demonstrated propellant azimuthal flow uniformity better than ± 2 percent of the nominal value, azimuthal magnetic field uniformity better than ± 0.5 percent of the nominal value, and cathode heater cycle testing to greater than 30,000 cycles. Propulsion system integrated system testing is planned to use the H71M-PM and a breadboard 1-kW PPU. Pathfinder model test results are now supporting the design of the H71M-EM engineering model thruster. A second-generation breadboard PPU has been fabricated and is currently under test. NASA has made these technologies available to U.S. industry through a no cost, nonexclusive license agreement.

Hall↗
NASA NTRSDate not supplied

Emitter Surface Temperature Measurements of a 25-A class Hollow Cathode

Measurements of the emitter surface temperature of a hollow cathode were made using an optical ratio pyrometer. These measurements were performed to provide insight into the effect of various cathode operating parameters on emitter surface peak temperature and temperature distribution

Hollow Cathode↗