Engineering topics

Soulas, George C.

Publications and source records attributed to Soulas, George C..

At least 91 records · Page 5

NASA NTRS2001

Design and Performance of 40 cm Ion Optics

A 40 cm ion thruster is being developed at the NASA Glenn Research Center to obtain input power and propellant throughput capabilities of 10 kW and 550 kg. respectively. The technical approach here is a continuation of the "derating" technique used for the NSTAR ion thruster. The 40 cm ion thruster presently utilizes the NSTAR ion optics aperture geometry to take advantage of the large database of lifetime and performance data already available. Dome-shaped grids were chosen for the design of the 40 cm ion optics because this design is naturally suited for large-area ion optics. Ion extraction capabilities and electron backstreaming limits for the 40 cm ion optics were estimated by utilizing NSTAR 30 cm ion optics data. A preliminary service life assessment showed that the propellant throughput goal of 550 kg of xenon may be possible with molybdenum 40 cm ion optics. One 40 cm ion optics' set has been successfully fabricated to date. Additional ion optics' sets are presently being fabricated. Preliminary performance tests were conducted on a laboratory model 40 cm ion thruster.

Soulas, George C.↗

NASA NTRS2001

Processes For Cleaning a Cathode Tube and Assemblies In A Hollow Cathode Assembly

The present invention is a process for cleaning a cathode tube and other subassemblies in a hollow cathode assembly. In the disclosed process, hand covering elastomer gloves are used for handling all cathode assembly parts. The cathode tube and other subassemblies are cleaned with a lint-free cloth damped with acetone, then wiped with alcohol, immersed in ethyl alcohol or acetone, and ultrasonic agitation is applied, heating to 60 C. for ethyl alcohol or 56 C. for acetone. The cathode tube and other subassemblies are dried by blowing with nitrogen gas.

Patterson, Michael J.↗

NASA NTRS2001

Next-Generation Ion Propulsion Being Developed

The NASA Glenn Research Center ion-propulsion program addresses the need for high specific-impulse systems and technology across a broad range of mission applications and power levels. One activity is the development of the next-generation ion-propulsion system as a follow-on to the successful Deep Space 1 system. The system is envisioned to incorporate a lightweight ion engine that can operate over 1 to 10 kW, with a 550-kg propellant throughput capacity. The engine concept under development has a 40-cm beam diameter, twice the effective area of the Deep Space 1 engine. It incorporates mechanical features and operating conditions to maximize the design heritage established by the Deep Space 1 engine, while incorporating new technology where warranted to extend the power and throughput capability. Prototype versions of the engine have been fabricated and are under test at NASA, with an engineering model version in manufacturing. Preliminary performance data for the prototype engine have been documented over 1.1- to 7.3-kW input power. At 7.3 kW, the engine efficiency is 0.68, at 3615-sec specific impulse. Critical component temperatures, including those of the discharge cathode assembly and magnets, have been documented and are within established limits, with significant margins relative to the Deep Space 1 engine. The 1- to 10-kW ion thruster approach described here was found to provide the needed power and performance improvement to enable important NASA missions. The Integrated In-Space Transportation Planning (IISTP) studies compared many potential technologies for various NASA, Government, and commercial missions. These studies indicated that a high-power ion propulsion system is the most important technology for development because of its outstanding performance versus perceived development and recurring costs for interplanetary solar electric propulsion missions. One of the best applications of a highpower electric propulsion system was as an integral part of a solar electric propulsion (SEP) stage to send a payload to outer planet targets. The IISTP studies showed that either trip time or launch vehicle class could be significantly reduced when compared with state-of-the-art systems.

Patterson, Michael J.↗

NASA NTRS2000

Performance of Titanium Optics on a NASA 30 Cm Ion Thruster

The results of performance tests with two titanium optics sets are presented and compared to those of molybdenum optics. All tests were conducted on a 30 cm ion thruster that was nearly identical to the NASA Solar Electric Propulsion Technology Application Readiness (NSTAR) thruster design. Optics performance tests were conducted over a thruster input power range of 0.5 to 4.6 kW. Optics performance including impingement-limited total voltages, electron backstreaming limits, screen grid ion transparencies, near-field beam current density profiles, beam divergence angles, and beam divergence thrust correction factors were determined throughout this power range. The impingement-limited total voltages for titanium optics were within 10 to 55 V of those for molybdenum optics. Electron backstreaming limit magnitude as a function of peak beam current density for both molybdenum and titanium optics were within a few volts of each other, indicating similar hot grid gaps for these two grid materials during steady-state operation. Beam divergence half-angles at 90 percent of the total beam current and thrust correction factors for both titanium optics sets were within one degree and one percent, respectively, of those for molybdenum optics. When thruster power was increased to 2.3 kW immediately following discharge ignition, the titanium screen grid came into contact with the accelerator grid within five minutes of ignition. Relative to molybdenum, titanium's larger thermal expansion and smaller thermal conductivity likely caused the screen grid to thermally expand more relative to the accelerator grid during startup.

Soulas, George C.↗

NASA NTRS2000

Plume and Discharge Plasma Measurements of an NSTAR-type Ion Thruster

The success of the NASA Deep Space I spacecraft has demonstrated that ion propulsion is a viable option for deep space science missions. More aggressive missions such as Comet Nuclear Sample Return and Europa lander will require higher power, higher propellant throughput and longer thruster lifetime than the NASA Solar Electric Propulsion Technology Application Readiness (NSTAR) engine. Presented here are thruster plume and discharge plasma measurements of an NSTAR-type thruster operated from 0.5 kW to 5 kW. From Faraday plume sweeps, beam divergence was determined. From Langmuir probe plume measurements on centerline, low energy ion production on axis due to charge-exchange and direct ionization was assessed. Additionally, plume plasma potential measurements made on axis were used to determine the upper energy limits at which ions created on centerline could be radially accelerated. Wall probes flush-mounted to the thruster discharge chamber anode were used to assess plasma conditions. Langmuir probe measurements at the wall indicated significant differences in the electron temperature in the cylindrical and conical sections of the discharge chamber.

Foster, John E↗

NASA NTRS2000

Plume and Discharge Plasma Measurements of an NSTAR-type Ion Thruster

The success of the NASA Deep Space 1 spacecraft has demonstrated that ion propulsion is a viable option for deep space science missions. More aggressive missions such as Comet Nuclear Sample Return and Europa lander will require higher power, higher propellant throughput and longer thruster lifetime than the NASA Solar Electric Propulsion Technology Application Readiness (NSTAR) engine. Presented here are thruster plume and discharge plasma measurements of an NSTAR-type thruster operated from 0.5 kW to 5 kW. From Faraday plume sweeps, beam divergence was determined. From Langmuir probe plume measurements on centerline, low energy ion production on axis due to charge-exchange and direct ionization was assessed. Additionally, plume plasma potential measurements made on axis were used to determine the upper energy limits at which ions created on centerline could be radially accelerated. Wall probes flush-mounted to the thruster discharge chamber anode were used to assess plasma conditions. Langmuir probe measurements at the wall indicated significant differences in the electron temperature in the cylindrical and conical sections of the discharge chamber.

Foster, John E.↗

NASA NTRS2000

Process for Ignition of Gaseous Electrical Discharge Between Electrodes of a Hollow Cathode Assembly

The design and manufacturing processes for Hollow Cathode Assemblies (HCA's) that operate over a broad range of emission currents up to 30 Amperes, at low potentials, with lifetimes in excess of 17,500 hours. The processes include contamination control procedures which cover hollow cathode component cleaning procedures, gas feed system designs and specifications, and hollow cathode activation and operating procedures to thereby produce cathode assemblies that have demonstrated stable and repeatable operating conditions, for both the discharge current and voltage. The HCA of this invention provides lifetimes of greater than 10,000 hours, and expected lifetimes of greater than 17,500 hours, whereas the present state-of-the-art is less than 500 hours at emission currents in excess of 1 Ampere. Stable operation is provided over a large range of operating emission currents, up to a 6:1 ratio, and this HCA can emit electron currents of up to 30 Amperes in magnitude to an external anode that simulates the current drawn to a space plasma, at voltages of less than 20 Volts.

Patterson, Michael J.↗

NASA NTRS1999

Titanium Optics for Ion Thrusters

Ion thruster total impulse capability is limited, in part, by accelerator grid sputter erosion. A development effort was initiated to identify a material with a lower accelerator grid volumetric sputter erosion rate than molybdenum, but that could utilize the present NSTAR thruster grid design and fabrication techniques to keep development costs low, and perform as well as molybdenum optics. After comparing the sputter erosion rates of several atomic materials to that of molybdenum at accelerator voltages, titanium was found to offer a 45% reduction in volumetric erosion rates. To ensure that screen grid sputter erosion rates are not higher at discharge chamber potentials, titanium and molybdenum sputter erosion rates were measured at these potentials. Preliminary results showed only a slightly higher volumetric erosion rate for titanium, so that screen grid erosion is insignificant. A number of material, thermal, and mechanical properties were also examined to identify any fabrication, launch environment, and thruster operation issues. Several titanium grid sets were successfully fabricated. A titanium grid set was mounted onto an NSTAR 30 cm engineering model ion thruster and tested to determine optics performance. The titanium optics operated successfully over the entire NSTAR power range of 0.5 to 2.3 kW. Differences in impingement-limited perveances and electron backstreaming limits were found to be due to a larger cold gap for the titanium optics. Discharge losses for titanium grids were lower than those for molybdenum, likely due to a slightly larger titanium screen grid open area fraction. Radial distributions of beam current density with titanium optics were very similar to those with molybdenum optics at all power levels. Temporal electron backstreaming limit measurements showed that titanium optics achieved thermal equilibrium faster than molybdenum optics.

Soulas, George C.↗

NASA NTRS1998

International Space Station Cathode Life Testing Status

To demonstrate adequate lifetime and performance capabilities of a hollow cathode for use on the International Space Station (ISS) plasma contactor system, life tests of multiple hollow cathode assemblies (HCAs) were initiated at operating conditions simulating on-orbit operation. Three HCAs are presently being tested. These HCAs are operated with a continuous 6 sccm xenon flow rate and 3 A anode current. Emission current requirements are simulated with a square waveform consisting of 50 minutes at a 2.5 A emission current and 40 minutes with no emission current. As of July 1998, these HCAs have accumulated between 1 1,700 and 14,200 hours. While there have been changes in operatin, behavior the three HCAs continue to operate stably within ISS specifications and are expected to demonstrate the required lifetime.

Sarver-Verhey, Timothy R.↗

NASA NTRS1998

Scaling of Ion Thrusters to Low Power

Analyses were conducted to examine ion thruster scaling relationships in detail to determine performance limits, and lifetime expectations for thruster input power levels below 0.5 kW. This was motivated by mission analyses indicating the potential advantages of high performance, high specific impulse systems for small spacecraft. The design and development status of a 0.1-0.3 kW prototype small thruster and its components are discussed. Performance goals include thruster efficiencies on the order of 40% to 54% over a specific impulse range of 2000 to 3000 seconds, with a lifetime in excess of 8000 hours at full power. Thruster technologies required to achieve the performance and lifetime targets are identified.

Patterson, Michael J.↗

NASA NTRS1997

International Space Station Cathode Life Testing

Four hollow cathode assembly (HCA) life tests were initiated at operating conditions simulating on-orbit operation of the International Space Station plasma contactor. The objective of these tests is to demonstrate the mission-required 18,000 hour lifetime with high-fidelity development model HCAS. HCAs are operated with a continuous 6 sccm xenon flow rate and 3 A anode current. On-orbit emission current requirements are simulated with a square waveform consisting of 50 minutes at a 2.5 A emission current and 40 minutes with no emission current. One HCA test was terminated after approximately 8,000 hours so that a destructive analysis could be performed. The analysis revealed no life-limiting processes and the ultimate lifetime was projected to be greater than the mission requirement. Testing continues for the remaining three HCAs which have accumulated approximately 8,000 hours, 10,000 hours, and 11,000 hours, respectively, as of June 1997. Anode and bias voltages, strong indicators of cathode electron emitter condition, are within acceptable ranges and have exhibited no life- or performance-limiting phenomena to date.

Soulas, George C.↗

NASA NTRS1995

Functional Testing of the Space Station Plasma Contactor

A plasma contactor system has been baselined for the International Space Station Alpha (ISSA) to control the electrical potentials of surfaces to eliminate/mitigate damaging interactions with the space environment. The system represents a dual-use technology which is a direct outgrowth of the NASA electric propulsion program and, in particular, the technology development effort on ion thruster systems. The plasma contactor subsystems include a hollow cathode assembly, a power electronics unit, and an expellant management unit. Under a pre-flight development program these subsystems are being developed to the level of maturity appropriate for transfer to U.S. industry for final development. Development efforts for the hollow cathode assembly include design selection and refinement, validating its required lifetime, and quantifying the cathode performance and interface specifications. To date, cathode components have demonstrated over 10,000 hours lifetime, and a hollow cathode assembly has demonstrated over 3,000 ignitions. Additionally, preliminary integration testing of a hollow cathode assembly with a breadboard power electronics unit has been completed. This paper discusses test results and the development status of the plasma contactor subsystems for ISSA, and in particular, the hollow cathode assembly.

Patterson, Michael J.↗

NASA NTRS1994

Multiple Hollow Cathode Wear Testing for the Space Station Plasma Contactor

A wear test of four hollow cathodes was conducted to resolve issues associated with the Space Station plasma contactor. The objectives of this test were to evaluate unit-to-unit dispersions, verify the transportability of contamination control protocols developed by the project, and to evaluate cathode contamination control and activation procedures to enable simplification of the gas feed system and heater power processor. These objectives were achieved by wear testing four cathodes concurrently to 2000 hours. Test results showed maximum unit-to-unit deviations for discharge voltages and cathode tip temperatures to be +/-3 percent and +/-2 percent, respectively, of the nominal values. Cathodes utilizing contamination control procedures known to increase cathode lifetime showed no trends in their monitored parameters that would indicate a possible failure, demonstrating that contamination control procedures had been successfully transferred. Comparisons of cathodes utilizing and not utilizing a purifier or simplified activation procedure showed similar behavior during wear testing and pre- and post-test performance characterizations. This behavior indicates that use of simplified cathode systems and procedures is consistent with long cathode lifetimes.

Soulas, George C.↗

NASA NTRS1994

Status of Hollow Cathode Heater Development for the Space Station Plasma Contactor

A hollow cathode-based plasma contactor has been selected for use on the Space Station. During the operation of the plasma contactor, the hollow cathode heater will endure approximately 12000 thermal cycles. Since a hollow cathode heater failure would result in a plasma contactor failure, a hollow cathode heater development program was established to produce a reliable heater. The development program includes the heater design, process documents for both heater fabrication and assembly, and heater testing. The heater design was a modification of a sheathed ion thruster cathode heater. Heater tests included testing of the heater unit alone and plasma contactor and ion thruster testing. To date, eight heaters have been or are being processed through heater unit testing, two through plasma contactor testing and three through ion thruster testing, all using direct current power supplies. Comparisons of data from heater unit performance tests before cyclic testing, plasma contactor tests, and ion thruster tests at the ignition input current level show the average deviation of input power and tube temperature near the cathode tip to be +/-0.9 W and +/- 21 C, respectively. Heater unit testing included cyclic testing to evaluate reliability under thermal cycling. The first heater, although damaged during assembly, completed 5985 ignition cycles before failing. Four additional heaters successfully completed 6300, 6300, 700, and 700 cycles. Heater unit testing is currently ongoing for three heaters which have to date accumulated greater than 7250, greater than 5500, and greater than 5500 cycles, respectively.

Soulas, George C.↗

NASA NTRS1994

Multiple Hollow Cathode Wear Testing

A hollow cathode-based plasma contactor has been baselined for use on the Space Station to reduce station charging. The plasma contactor provides a low impedance connection to space plasma via a plasma produced by an arc discharge. The hollow cathode of the plasma contactor is a refractory metal tube, through which xenon gas flows, which has a disk-shaped plate with a centered orifice at the downstream end of the tube. Within the cathode, arc attachment occurs primarily on a Type S low work function insert that is next to the orifice plate. This low work function insert is used to reduce cathode operating temperatures and energy requirements and, therefore, achieve increased efficiency and longevity. The operating characteristics and lifetime capabilities of this hollow cathode, however, are greatly reduced by oxygen bearing contaminants in the xenon gas. Furthermore, an optimized activation process, where the cathode is heated prior to ignition by an external heater to drive contaminants such as oxygen and moisture from the insert absorbed during exposure to ambient air, is necessary both for cathode longevity and a simplified power processor. In order to achieve the two year (approximately 17,500 hours) continuous operating lifetime requirement for the plasma contactor, a test program was initiated at NASA Lewis Research Center to demonstrate the extended lifetime capabilities of the hollow cathode. To date, xenon hollow cathodes have demonstrated extended lifetimes with one test having operated in excess of 8000 hours in an ongoing test utilizing contamination control protocols developed by Sarver-Verhey. The objectives of this study were to verify the transportability of the contamination control protocols developed by Sarver-Verhey and to evaluate cathode contamination control procedures, activation processes, and cathode-to-cathode dispersions in operating characteristics with time. These were accomplished by conducting a 2000 hour wear test of four hollow cathodes with different xenon gas purities and activation processes. The following will be presented: a description of the facility and test hardware, testing procedures and operating conditions, a discussion of test results, and conclusions.

Soulas, George C.↗

NASA NTRS1994

Mechanisms of anode power deposition in a low pressure free burning arc

Anode power deposition is a dominant power loss mechanism for arc jets and MPD thrusters. In this study, a free burning arc experiment was operated at pressures and current densities similar to those in arc jets and MPD thrusters in an attempt to identify the physics controlling this loss mechanism. Use of a free burning arc allowed for the isolation of independent variables controlling anode power deposition and provided a convenient and flexible way to cover a broad range of currents, anode surface pressures, and applied magnetic field strengths and orientations using an argon gas. Test results showed that anode power deposition decreased with increasing anode surface pressure up to 6.7 Pa (0.05 torr) and then became insensitive to pressure. Anode power increased with increasing arc current while the electron number density near the anode surface increased linearity. Anode power also increased with increasing applied magnetic field strength due to an increasing anode fall voltage. Applied magnetic field orientation had an effect only at high currents and low anode surface pressures, where anode power decreased when applied field lines intercepted the anode surface. The results demonstrated that anode power deposition was dominated by the current carrying electrons and that the anode fall voltage was the largest contributor. Furthermore, the results showed that anode power deposition can be reduced by operating at increased anode pressures, reduced arc currents, and applied magnetic field strengths and with magnetic field lines intercepting the anode.

Soulas, George C.↗

NASA NTRS1993

Plasma contactor development for Space Station

Plasma contactors have been baselined for the Space Station (SS) to control the electrical potentials of surfaces to eliminate/mitigate damaging interactions with the space environment. The system represents a dual-use technology which is a direct outgrowth of the NASA electric propulsion program and, in particular, the technology development effort on ion thrustor systems. The plasma contactor subsystems include the plasma contactor unit, a power electronics unit, and an expellant management unit. Under this pre-flight development program these will all be brought to breadboard or engineering model status. Development efforts for the plasma contactor include optimizing the design and configuration of the contactor, validating its required lifetime, and characterizing the contactor plume and electromagnetic interference. The plasma contactor unit design selected for the SS is an enclosed keeper, xenon hollow cathode plasma source. This paper discusses the test results and development status of the plasma contactor unit subsystem for the SS.

Patterson, Michael J.↗

NASA NTRS1993

Hollow cathode heater development for the Space Station plasma contactor

A hollow cathode-based plasma contactor has been selected for use on the Space Station. During the operation of the plasma contactor, the hollow cathode heater will endure approximately 12000 thermal cycles. Since a hollow cathode heater failure would result in a plasma contactor failure, a hollow cathode heater development program was established to produce a reliable heater design. The development program includes the heater design, process documents for both heater fabrication and assembly, and heater testing. The heater design was a modification of a sheathed ion thruster cathode heater. Three heaters have been tested to date using direct current power supplies. Performance testing was conducted to determine input current and power requirements for achieving activation and ignition temperatures, single unit operational repeatability, and unit-to-unit operational repeatability. Comparisons of performance testing data at the ignition input current level for the three heaters show the unit-to-unit repeatability of input power and tube temperature near the cathode tip to be within 3.5 W and 44 degrees C, respectively. Cyclic testing was then conducted to evaluate reliability under thermal cycling. The first heater, although damaged during assembly, completed 5985 ignition cycles before failing. Two additional heaters were subsequently fabricated and have completed 3178 cycles to date in an on-going test.

Soulas, George C.↗