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Search indexed NASA NTRS and DOE OSTI research on propulsion, heat transfer, battery materials and energy systems. Follow report and document links to the original sources.

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At least 109 records · Page 6
All sourcesNASA NTRSDOE OSTI
DOE OSTI2025

Study of the Design and Characteristics of a Modified Pulsed Plasma Thruster with Graphite and Tungsten Trigger Electrodes

The paper presents experimental results for a modified pulsed plasma thruster (PPT) with solid propellant, using a coaxial anode–cathode design. Graphite from pencil leads served as propellant, and a tungsten trigger electrode was tested to reduce carbonization effects. Experiments were performed in a vacuum chamber at 0.001 Pa, employing diagnostics such as discharge current/voltage recording, power measurement, ballistic pendulum, time-of-flight (TOF) method, and a Faraday cup. Current and voltage waveforms matched an oscillatory RLC circuit with variable plasma channel resistance. Key discharge parameters were measured, including current pulse duration/amplitude and plasma channel formation/decay dynamics. Impulse bit values, obtained with a ballistic pendulum, reached up to 8.5 μN·s. Increasing trigger capacitor capacitance reduced thrust due to unstable “pre-plasma” formation and partial pre-discharge energy loss. Using TOF and Faraday cup diagnostics, plasma front velocity, ion current amplitude, current density, and ion concentration were determined. Tungsten electrodes produced lower charged particle concentrations than graphite but offered better adhesion resistance, minimal carbonization, and stable long-term performance. The findings support optimizing trigger electrode materials and PPT operating modes to extend lifetime and stabilize thrust output.

Faraday cup↗
NASA NTRS1985

Rocket thrust chamber thermal barrier coatings

Subscale rocket thrust chamber tests were conducted to evaluate the effectiveness and durability of thin yttria stabilized zirconium oxide coatings applied to the thrust chamber hot-gas side wall. The fabrication consisted of arc plasma spraying the ceramic coating and bond coat onto a mandrell and then electrodepositing the copper thrust chamber wall around the coating. Chambers were fabricated with coatings .008, and .005 and .003 inches thick. The chambers were thermally cycled at a chamber pressure of 600 psia using oxygen-hydrogen as propellants and liquid hydrogen as the coolant. The thicker coatings tended to delaminate, early in the cyclic testing, down to a uniform sublayer which remained well adhered during the remaining cycles. Two chambers with .003 inch coatings were subjected to 1500 thermal cycles with no coating loss in the throat region, which represents a tenfold increase in life over identical chambers having no coatings. An analysis is presented which shows that the heat lost to the coolant due to the coating, in a rocket thrust chamber design having a coating only in the throat region, can be recovered by adding only one inch to the combustion chamber length.

Quentmeyer, R. J.↗
NASA NTRS2004

Discharge Chamber Performance of the NEXIS Ion Thruster

The Nuclear-Electric Xenon Ion System (NEXIS) thruster was designed to produce greater than or equal to 70% efficiency at ISPs in excess of 6500 sec and total power levels in excess of 15 kW. In order to achieve this performance, the thruster requires a large area plasma generator capable of high propellant utilimtion efficiency and low discharge loss while producing a very flat, uniform beam profile. Fortunately, larger thrusters can be made more uniform and efficient due to the higher volume to surface ratio, provided that the magnetic cusp confinement is designed properly and the thruster length to diameter ratio is adequate. This paper describes the discharge chamber performance of the NEXIS Laboratory Model (LM) thruster. The LM discharge chamber is 65 cm in diameter at the grid plane and uses 6 ring-cusps to provide magnetic confinement of the plasma. The thruster was tested with flat carbon-carbon composite grids with the hole pattern masked to 57 cm in diameter and a conventional Type-B "1/2" diameter hollow cathode. During the preliminary "discharge only" tests, the LM thruster demonstrated profile factors of 0.84 and a discharge loss of about 160 eV/ion at 25 V discharge voltage and over 90% propellant utilization efficiency in simulated beam extraction experiments at 3.9 A of beam current. Analysis of the data from these tests used the discharge-only model developed by Brophy. Subsequent beam extraction experiments validated the key variables used in the model to predict the performance from the discharge-only data, and demonstrated 3.9 A of beam current at over 90% propellant utilization efficiency with a flatness parameter of better than 0.8 and a discharge loss of about 185 eV/ion. The slightly higher discharge loss measured during beam extractions was found to be due to a lower screen transparency in the as-manufactured LM grid set. Plasma measurements with a scanning probe internal to the thruster near the screen grid showed plasma densities over l x 10(exp 11) per cubic centimeter and electron temperatures of 3.5 to 5.5 eV depending on the operation parameters. The performance of the NEXIS discharge chamber contributed to the over 78% thruster efficiency measured during beam extraction at 7500 sec ISP and 25 kW of power, and over 81% thruster efficiency measured at 8500 sec ISP.

plasma generation↗
NASA NTRS1975

Plasma core reactor simulations using RF uranium seeded argon discharges

An experimental investigation was conducted using the United Technologies Research Center (UTRC) 80 kW and 1.2 MW RF induction heater systems to aid in developing the technology necessary for designing a self-critical fissioning uranium plasma core reactor (PCR). A nonfissioning, steady-state RF-heated argon plasma seeded with pure uranium hexafluoride (UF6) was used. An overall objective was to achieve maximum confinement of uranium vapor within the plasma while simultaneously minimizing the uranium compound wall deposition. Exploratory tests were conducted using the 80 kW RF induction heater with the test chamber at approximately atmospheric pressure and discharge power levels on the order of 10 kW. Four different test chamber flow configurations were tested to permit selection of the configuration offering the best confinement characteristics for subsequent tests at higher pressure and power in the 1.2 MW RF induction heater facility.

Roman, W. C.↗
NASA NTRS1980

Voyager spacecraft electrostatic discharge testing

The program of environmental testing undergone by the Voyager spacecraft in order to simulate the transient voltage effects of electrostatic discharges expected in the energetic plasma environment of Jupiter is reported. The testing consists of studies of the electrostatic discharge characteristics of spacecraft dielectrics in a vacuum-chamber-electron beam facility, brief piece part sensitivity tests on such items as a MOSFET multiplexer and the grounding of the thermal blanket, and assembly tests of the magnetometer boom and the science boom. In addition, testing of a complete spacecraft was performed using two arc sources to simulate long and short duration discharge sources for successive spacecraft shielding and grounding improvements. Due to the testing program, both Voyager 1 and Voyager 2 experienced tolerable electrostatic discharge-caused transient anomalies in science and engineering subsystems, however, a closer duplication of the spacecraft environment is necessary to predict and design actual spacecraft responses more accurately.

Whittlesey, A.↗
NASA NTRS2002

Evaluation and Characterization Study of Dual Pulse Laser-Induced Spark (DPLIS) For Rocket Engine Ignition System Application

This paper addresses the progress of technology development of a laser ignition system at NASA Marshall Space Flight Center (MSFC). Initial hot-fire tests in a small-scale rocket chamber at MSFC have demonstrated the DPLIS concept having two main advantages over existing laser ignition concepts. First, the DPLIS can be potentially optimized its laser pulse format to maximize the initial plasma volume, the plasma lifetime, as well as the flame kernel growth rate. Characterization studies of the laser pulse format are now underway with experiments of igniting gaseous hydrogen/air in a Hencken burner. Once ignition is achieved, the flame is open to the atmosphere. This open environment allows easy access for diagnostics of the ignition phenomenon. The quick turn-around time of conducting experiments on this burner make it more amenable for conducting a large number of experiments for statistical analysis of the sensitivity of the test parameters. The results from these experiments will help optimize the laser format for future testing in an H2/O2 subscale rocket chamber.

Trinh, Huu P.↗
NASA NTRS2008

Extraction of Thermal Performance Values from Samples in the Lunar Dust Adhesion Bell Jar

A simulation chamber has been developed to test the performance of thermal control surfaces under dusty lunar conditions. The lunar dust adhesion bell jar (LDAB) is a diffusion pumped vacuum chamber (10-8 Torr) built to test material samples less than about 7 cm in diameter. The LDAB has the following lunar dust stimulant processing capabilities: heating and cooling while stirring in order to degas and remove absorbed water; RF air-plasma for activating the dust and for organic contaminant removal; RF H/He-plasma to simulate solar wind; dust sieving system for controlling particle sizes; and a controlled means of introducing the activated dust to the samples under study. The LDAB is also fitted with an in situ Xe arc lamp solar simulator, and a cold box that can reach 30 K. Samples of thermal control surfaces (2.5 cm diameter) are introduced into the chamber for calorimetric evaluation using thermocouple instrumentation. The object of this paper is to present a thermal model of the samples under test conditions, and to outline the procedure to extract the absorptance, emittance, and thermal efficiency from the pristine and sub-monolayer dust covered samples

Gaier, James R.↗
NASA NTRS2008

Extraction of Thermal Performance Values from Samples in the Lunar Dust Adhesion Bell Jar

A simulation chamber has been developed to test the performance of thermal control surfaces under dusty lunar conditions. The lunar dust adhesion bell jar (LDAB) is a diffusion pumped vacuum chamber (10(exp -8) Torr) built to test material samples less than about 7 cm in diameter. The LDAB has the following lunar dust simulant processing capabilities: heating and cooling while stirring in order to degas and remove adsorbed water; RF air-plasma for activating the dust and for organic contaminant removal; RF H/He-plasma to simulate solar wind; dust sieving system for controlling particle sizes; and a controlled means of introducing the activated dust to the samples under study. The LDAB is also fitted with an in situ Xe arc lamp solar simulator, and a cold box that can reach 30 K. Samples of thermal control surfaces (2.5 cm diameter) are introduced into the chamber for calorimetric evaluation using thermocouple instrumentation. The object of this paper is to present a thermal model of the samples under test conditions and to outline the procedure to extract the absorptance, emittance, and thermal efficiency from the pristine and sub-monolayer dust covered samples.

Gaier, James R.↗
NASA NTRS2010

Extraction of Thermal Performance Values from Samples in the Lunar Dust Adhesion Bell Jar

A simulation chamber has been developed to test the performance of thermal control surfaces under dusty lunar conditions. The lunar dust adhesion bell jar (LDAB) is a diffusion pumped vacuum chamber (10(exp -8) Torr) built to test material samples less than about 7 cm in diameter. The LDAB has the following lunar dust simulant processing capabilities: heating and cooling while stirring in order to degas and remove adsorbed water; RF air-plasma for activating the dust and for organic contaminant removal; RF H/He-plasma to simulate solar wind; dust sieving system for controlling particle sizes; and a controlled means of introducing the activated dust to the samples under study. The LDAB is also fitted with an in situ Xe arc lamp solar simulator, and a cold box that can reach 30 K. Samples of thermal control surfaces (2.5 cm diameter) are introduced into the chamber for calorimetric evaluation using thermocouple instrumentation. The object of this paper is to present a thermal model of the samples under test conditions and to outline the procedure to extract the absorptance, emittance, and thermal efficiency from the pristine and sub-monolayer dust covered samples.

Gaier, James R.↗
DOE OSTI2024

Design and Engineering of LUPIN: A Test-Bed Radiofrequency Ion Source for Enhanced Neutral Beam Injection on DIII-D

The Large, Uniform Plasma for Ionizing Neutrals (LUPIN) is an RF inductively coupled plasma (ICP) chamber for demonstrating plasma performance of an RF ICP positive ion source upgrade for the DIII-D Neutral Beam Injection (NBI) system. LUPIN will be used to investigate ion source physics, including neutral gas dynamics, plasma density uniformity, interactions with Faraday shields, and power coupling to novel RF antenna designs. LUPIN has an RF generator capable of delivering 20 kW of power at 2 MHz, which is coupled into a cylindrical quartz vessel measuring 20 cm in length and 10 cm in radius. This configuration matches the power density requirements for a full-scale ion source. Target hydrogen and deuterium plasma densities exceeding 10 18 m -3 would relate to extracted ion current densities of 2100 A/m 2 for 10 s. Vacuum conductance and gas flow calculations predict a maximum achievable neutral gas flow rate of 1675 Pa·L/s at 5 Pa of He, which mimics the gas flow of the DIII-D NBI system. Designs have been developed for an internal Faraday shield to mitigate heat flux and ion sputtering on the dielectric vessel. Thermomechanical finite element simulations demonstrated the Faraday shield design to be capable of withstanding anticipated heat loads from worst-case operation scenarios. Results of upcoming experimental investigations on LUPIN will guide the design of a full-scale prototype for DIII-D integration.

Faraday shield↗
NASA NTRS2016

Solar Array at Very High Temperatures: Ground Tests

Solar array design for any spacecraft is determined by the orbit parameters. For example, operational voltage for spacecraft in Low Earth Orbit (LEO) is limited by significant differential charging due to interactions with low temperature plasma. In order to avoid arcing in LEO, solar array is designed to generate electrical power at comparatively low voltages (below 100 V) or to operate at higher voltages with encapsulated of all suspected discharge locations. In Geosynchronous Orbit (GEO) differential charging is caused by energetic electrons that produce differential potential between coverglass and conductive spacecraft body in a kilovolt range. In such a case, weakly conductive layer over coverglass (ITO) is one of possible measures to eliminate dangerous discharges on array surface. Temperature variations for solar arrays in both orbits are measured and documented within the range of -150 C +110 C. This wide interval of operational temperatures is regularly reproduced in ground tests with radiative heating and cooling inside shroud with flowing liquid nitrogen. The requirements to solar array design and tests turn out to be more complicated when planned trajectory crosses these two orbits and goes closer to Sun. Conductive layer over coverglass causes sharp increase in parasitic current collected from LEO plasma, high temperature may cause cracks in encapsulating material (RTV), radiative heating of coupon in vacuum chamber becomes practically impossible above 150 C, conductivities of glass and adhesive go up with temperature that decrease array efficiency, and mechanical stresses grow up to critical magnitudes. A few test arrangements and respective results are presented in current paper. Coupons were tested against arcing in simulated LEO and GEO environments under elevated temperatures up to 200 C. The dependence of leakage current on temperature was measured, and electrostatic cleanness was verified for coupons with antireflection (AR) coating over ITO layer.

differential charging↗
NASA NTRS1977

Neutron monitoring and electrode calorimetry experiments in the HIP-1 Hot Ion Plasma

Results are presented for two diagnostic procedures on HIP-1: neutron diagnostics to determine where neutrons originated within the plasma discharge chamber and electrode calorimetry to measure the steady-state power absorbed by the two anodes and cathodes. Results are also reported for a hot-ion plasma formed with a continuous-cathode rod, one that spans the full length of the test section, in place of the two hollow cathodes. The outboard neutron source strength increased relative to that at the midplane when (1) the cathode tips were moved farther outboard, (2) the anode diameters were increased, and (3) one of the anodes was removed. The distribution of neutron sources within the plasma discharge chamber was insensitive to the division of current between the two cathodes. For the continuous cathode, increasing the discharge current increased the midplane neutron source strength relative to the outboard source strength. Each cathode absorbed from 12 to 15 percent of the input power regardless of the division of current between the cathodes. The anodes absorbed from 20 to 40 percent of the input power. The division of power absorption between the anodes varied with plasma operating conditions and electrode placement.

Reinmann, J. J.↗
NASA NTRS2000

Enhanced Discharge Performance in a Ring Cusp Plasma Source

There is a need for a lightweight, low power ion thruster for space science missions. Such an ion thruster is under development at NASA Glenn Research Center. In an effort to better understand the discharge performance of this thruster, a thruster discharge chamber with an anode containing electrically isolated electrodes at the cusps was fabricated and tested. Characteristics of this ring cusp ion discharge were measured without ion beam extraction. Discharge current was measured at collection electrodes located at the magnetic cusps and at the anode body itself. Discharge performance and plasma properties were measured as a function of power, which was varied between 20 and 50 W. It was found that ion production costs decreased by as much as 20 percent when the two most downstream cusp electrodes were allowed to float. Floating the electrodes did not give rise to a significant increase in discharge power even though the plasma density increased markedly. The improved performance is attributed to enhanced electron containment.

Foster, John E.↗
NASA NTRS1985

Rocket thrust chamber thermal barrier coatings

A research program was conducted to generate data and develop analytical techniques to predict the performance and reliability of ceramic thermal barrier coatings in high heat flux environments. A finite element model was used to analyze the thermomechanical behavior of coating systems in rocket thrust chambers. Candidate coating systems (using a copper substrate, NiCrAlY bond coat and ZrO2.8Y2O3 ceramic overcoat) were selected for detailed study based on photomicrographic evaluations of experimental test specimens. The effects of plasma spray application parameters on the material properties of these coatings were measured and the effects on coating performance evaluated using the finite element model. Coating design curves which define acceptable operating envelopes for seleted coating systems were constructed based on temperature and strain limitations. Spray gun power levels was found to have the most significant effect on coating structure. Three coating systems were selected for study using different power levels. Thermal conductivity, strain tolerance, density, and residual stress were measured for these coatings. Analyses indicated that extremely thin coatings ( 0.02 mm) are required to accommodate the high heat flux of a rocket thrust chamber and ensure structural integrity.

Batakis, A. P.↗
NASA NTRS1992

H2 arcjet performance mapping program

Work performed during the period of Mar. 1991 to Jan. 1992 is reviewed. High power H2 arcjets are being considered for electric powered orbit transfer vehicles (EOTV). Mission analyses indicate that the overall arcjet thrust efficiency is very important since increasing the efficiency increases the thrust, and thereby reduces the total trip time for the same power. For example, increasing the thrust efficiency at the same specific impulse from 30 to 40 percent will reduce the trip time by 25 percent. For a 200 day mission, this equates to 50 days, which results in lower ground costs and less time during which the payload is dormant. Arcjet performance levels of 1200 seconds specific impulse (lsp) at 35 to 40 percent efficiency with lifetimes over 1000 hours are needed to support EOTV missions. Because of the potential very high efficiency levels, the objective of this program was to evaluate the ability of a scaled Giannini-style thruster to achieve the performance levels while operating at a reduced nominal power of 10 kW. To meet this objective, a review of past literature was conducted; scaling relationships were developed and applied to establish critical dimensions; a development thruster was designed with the aid of the plasma analysis model KARNAC and finite element thermal modeling; test hardware was fabricated; and a series of performance tests were conducted in RRC's Cell 11 vacuum chamber with its null-balance thrust stand.

Source record↗
NASA NTRS2016

Solar Array at Very High Temperatures: Ground Tests

Solar array design for any spacecraft is determined by the orbit parameters. For example, operational voltage for spacecraft in Low Earth Orbit (LEO) is limited by significant differential charging due to interactions with low temperature plasma. In order to avoid arcing in LEO, solar array is designed to generate electrical power at comparatively low voltages (below 100 volts) or to operate at higher voltages with encapsulation of all suspected discharge locations. In Geosynchronous Orbit (GEO) differential charging is caused by energetic electrons that produce differential potential between the coverglass and the conductive spacecraft body in a kilovolt range. In such a case, the weakly conductive layer over coverglass, indium tin oxide (ITO) is one of the possible measures to eliminate dangerous discharges on array surface. Temperature variations for solar arrays in both orbits are measured and documented within the range of minus150 degrees Centigrade to plus 1100 degrees Centigrade. This wide interval of operational temperatures is regularly reproduced in ground tests with radiative heating and cooling inside a shroud with flowing liquid nitrogen. The requirements to solar array design and tests turn out to be more complicated when planned trajectory crosses these two orbits and goes closer to the Sun. The conductive layer over coverglass causes a sharp increase in parasitic current collected from LEO plasma, high temperature may cause cracks in encapsulating (Room Temperature Vulcanizing (RTV) material; radiative heating of a coupon in vacuum chamber becomes practically impossible above 1500 degrees Centigrade; conductivities of glass and adhesive go up with temperature that decrease array efficiency; and mechanical stresses grow up to critical magnitudes. A few test arrangements and respective results are presented in current paper. Coupons were tested against arcing in simulated LEO and GEO environments under elevated temperatures up to 2000 degrees Centigrade. The dependence of leakage current on temperature was measured, and electrostatic cleanness was verified for coupons with antireflection (AR) coating over the indium tin oxide (ITO) layer.

differential charging↗
DOE OSTI2025

Design and Engineering of LUPIN: A Test-Bed Radio-Frequency Ion Source for Enhanced Neutral Beam Injection on DIII-D

The Large, Uniform Plasma for Ionizing Neutrals (LUPIN) is a radio-frequency (RF) inductively coupled plasma (ICP) chamber for demonstrating plasma performance of an RF ICP positive ion source upgrade for the DIII-D neutral beam injection (NBI) system. LUPIN will be used to investigate ion source physics, including neutral gas dynamics, plasma density uniformity, interactions with Faraday shields, and power coupling to novel RF antenna designs. LUPIN has an RF generator capable of delivering 20 kW of power at 2 MHz, which is coupled into a cylindrical quartz vessel measuring 20 cm in length and 10 cm in radius. This configuration matches the power density requirements for a full-scale ion source. Target hydrogen and deuterium plasma densities exceeding 10 18 m -3 would relate to extracted ion current densities of 2100 A/m 2 for 10s. Vacuum conductance and gas flow calculations predict a maximum achievable neutral gas flow rate of 1675 Pa ⋅ L/s at 5 Pa of He, which mimics the gas flow of the DIII-D NBI system. Designs have been developed for an internal Faraday shield to mitigate heat flux and ion sputtering on the dielectric vessel. Thermomechanical finite element simulations demonstrated the Faraday shield design to be capable of withstanding anticipated heat loads from worst-case operation scenarios. Finally, results of upcoming experimental investigations on LUPIN will guide the design of a full-scale prototype for DIII-D integration.

Faraday shield↗
NASA NTRS2003

Investigation of Impinging Stream Vortex Chamber Concepts for Liquid Rocket Engine Applications

This paper addresses the progress of technology development of a laser ignition system at NASA Marshall Spac Flight Center (MSFC). Initial hot-fire tests in a small-scale rocket chamber at MSFC have demonstrated the dual pulse laser-induced spark (DPLIS) technique, which has an advantage over existing single-pulse laser ignition techniques in that it can be optimized in its laser pulse format to maximize the initial plasma volume, the plasm: lifetime, as well as the flame kernel growth rate. The distribution of the total laser energy into two separate pulse also lowers the peak power that would need to be sent through fiber optics to the combustion chamber, making the implementation of this technique more practical than other single-pulse techniques. A first generation prototype of an optic fiber-coupled laser ignition system will be tested a rocket chamber with RP-1/GOX and GH2/GOX propellants systems. Other relevant technology, such as optical windows, flight-qualified laser system etc. will be discussed in this paper.

Trinh, Huu P.↗