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Assessment of the NSTAR ion engine service life is being accomplished through a combination of long duration testing and probabilistic analyses of the credible failure modes. A literature review that examined 65 ion engine endurance tests perfromed over the past 35 years was conducted to compile a list of possible ion engine failure modes.
Assessment of the NSTAR ion engine service life is being accomplished through a combination of long duration testing and probabilistic analyses of the credible failure modes. A literature review that examined 65 ion engine endurance tests perfromed over the past 35 years was conducted to compile a list of possible ion engine failure modes.
Published data for run times exceeding a few seconds in rotating detonation engines (RDEs) is rare, but long duration tests are necessary to understand both transient and persistent phenomena in RDE operation. Previous work has noted different detonation wave modes at different equivalence ratios. In the current study, multiple long-duration runs with constant air flow rate and variable equivalence ratio are conducted in a hydrogen-fueled, water-cooled RDE. Analysis of the capillary tube averaged pressure (CTAP) data reveals a strong correlation with equivalence ratio as well as a small time dependency due to transient thermal processes. Ion probe measurements are used to analyze detonation wave frequencies and modes, both of which are strongly dependent on the equivalence ratio. Mode transitions are observed to occur consistently as the equivalence ratio is traversed between 0.55 and 0.60. These transitions are also found to be a function of the direction of changing equivalence ratio (ascending or descending) and time.
Published data for run times exceeding a few seconds in rotating detonation engines (RDEs) is rare, but long duration tests are necessary to understand both transient and persistent phenomena in RDE operation. Previous work has noted different detonation wave modes at different equivalence ratios. In the current study, multiple long-duration runs with constant air flow rate and variable equivalence ratio are conducted in a hydrogen-fueled, water-cooled RDE. Analysis of the capillary tube averaged pressure (CTAP) data reveals a strong correlation with equivalence ratio as well as a small time dependency due to transient thermal processes. Ion probe measurements are used to analyze detonation wave frequencies and modes, both of which are strongly dependent on the equivalence ratio. Mode transitions are observed to occur consistently as the equivalence ratio is traversed between 0.55 and 0.60. These transitions are also found to be a function of the direction of changing equivalence ratio (ascending or descending) and time.
Biofilm formation in spacecraft wastewater systems is a complex and impactful problem. Biomass flowing in these wastewater systems can cause clogging, higher differential pressures, and flow restrictions impacting equipment reliability and function. To explore this issue, the Marshall Space Flight Center’s biofilm test stand was created and has been utilized to examine multiple biofilm mitigation techniques over a long duration test time. This paper is an update to a previous paper, ICES-2021-122, and will serve to discuss the modifications and enhancements to the biofilm test stand. While the previous configuration tested heat treatments and coatings, this phase of testing will examine differing materials, competing bacteria provided by an external bioreactor, a prototype phytoremediation unit, and additional testing with a nutrient filter. Configuration changes include the addition of float switches, adjustment of the drain assembly architecture and additional support hardware for the competing bacteria bioreactor and phytoremediation units.
Biofilm formation in spacecraft wastewater systems is a complex and impactful problem. Biomass flowing in these wastewater systems can cause clogging, higher differential pressures, and flow restrictions impacting equipment reliability and function. To explore this issue, the Marshall Space Flight Center’s biofilm test stand was created and has been utilized to examine multiple biofilm mitigation techniques over a long duration test time. This paper is an update to a previous paper, ICES-2021-122, and will serve to discuss the modifications and enhancements to the biofilm test stand. While the previous configuration tested heat treatments and coatings, this phase of testing will examine differing materials, competing bacteria provided by an external bioreactor, a prototype phytoremediation unit, and additional testing with a nutrient filter. Configuration changes include the addition of float switches, adjustment of the drain assembly architecture and additional support hardware for the competing bacteria bioreactor and phytoremediation units.
Biofilm formation in spacecraft wastewater systems is a complex and impactful problem. Biomass flowing in these wastewater systems can cause clogging, higher differential pressures, and flow restrictions impacting equipment reliability and function. To explore this issue, the Marshall Space Flight Center’s biofilm test stand was created and has been utilized to examine multiple biofilm mitigation techniques over a long duration test time. This paper is an update to a previous paper, ICES-2021-122, and will serve to discuss the modifications and enhancements to the biofilm test stand. While the previous configuration tested heat treatments and coatings, this phase of testing will examine differing materials, competing bacteria provided by an external bioreactor, a prototype phytoremediation unit, and additional testing with a nutrient filter. Configuration changes include the addition of float switches, adjustment of the drain assembly architecture and additional support hardware for the competing bacteria bioreactor and phytoremediation units.
The concept of a general purpose life support system testbed for use in space grew out of considerations arising from the recent consolidation of NASA's Advanced Life Support (ALS) Systems programs. Both the physical-chemical and the biological approaches to regenerative life support will require significant amounts of in-space testing in order to prepare for the final development of systems for human life support. Considerations of the technical requirements and rationales for in-space testing has led to the concept of a common testbed that will allow faster and less expensive long duration tests.
Aerogel based insulation systems for ambient pressure environments were developed for liquid hydrogen (LH2) tank applications. Solutions to thermal insulation problems were demonstrated for the Space Shuttle External Tank (ET) through extensive testing at the Cryogenics Test Laboratory. Demonstration testing was performed using a 1/10th scale ET LH2 intertank unit and liquid helium as the coolant to provide the 20 K cold boundary temperature. Cryopumping tests in the range of 20K were performed using both constant mass and constant pressure methods. Long-duration tests (up to 10 hours) showed that the nitrogen mass taken up inside the intertank is reduced by a factor of nearly three for the aerogel insulated case as compared to the un-insulated (bare metal flight configuration) case. Test results including thermal stabilization, heat transfer effectiveness, and cryopumping confirm that the aerogel system eliminates free liquid nitrogen within the intertank. Physisorption (or adsorption) of liquid nitrogen within the fine pore structure of aerogel materials was also investigated. Results of a mass uptake method show that the sorption ratio (liquid nitrogen to aerogel beads) is about 62 percent by volume. A novel liquid nitrogen production method of testing the liquid nitrogen physical adsorption capacity of aerogel beads was also performed to more closely approximate the actual launch vehicle cooldown and thermal stabilization effects within the aerogel material. The extraordinary insulating effectiveness of the aerogel material shows that cryopumping is not an open-cell mass transport issue but is strictly driven by thermal communication between warm and cold surfaces. The new aerogel insulation technology is useful to solve heat transfer problem areas and to augment existing thermal protection systems on launch vehicles. Examples are given and potential benefits for producing launch systems that are more reliable, robust, reusable, and efficient are outlined.
Efforts over the past three years have focused on the study of candidate sorbent materials for use in a 4BMS molecular sieve system. The accumulation of knowledge has been invaluable for further decisions and for reflecting on the conclusions of past decisions. The goal of the next generation CO2 removal system is continuous, failure-free operation for nearly 20,000 hours, but no complex life support system has yet reached this lofty goal. In addition to reliability, CO2 removal performance improvements have been intensively studied. The achievements toward this end include highly detailed isotherm measurements which drive system simulations as well as testing physical design improvements. Looking back on the successes and failures of past systems, correlating data from long-duration tests, and carefully projecting future results are all needed for the success of the next system. This work intends to reveal the path we have taken and illuminate the steps to come for CO2 removal life support with the 4BCO2 flight demonstration.
Hall effect thrusters operating at power levels in excess of several hundreds of kilowatts have been identified as enabling technologies for applications such as lunar tugs, large satellite orbital transfer vehicles, and solar system exploration. These large thrusters introduce significant testing challenges due to the propellant flow rate exceeding the pumping speed available in most laboratories. Even with proposed upgrades in mind, the likelihood that multiple vacuum facilities will exist in the near future to allow long duration testing of high-power Hall thrusters operating at power levels in excess of 100 kW remains extremely low. In this article, we numerically explore the feasibility of testing Hall thrusters in a quasi-steady mode defined by pulsing the mass flow rate between a nominal and a low value. Here, our simulations indicate that sub-second durations available before the chamber reaches critical pressure are sufficiently long to achieve the steady-state current and flow field distributions, allowing us to characterize thruster performance and the near plume region.
The NASA s Evolutionary Xenon Thruster (NEXT) program is tasked with significantly improving and extending the capabilities of current state-of-the-art NSTAR thruster. The service life capability of the NEXT ion thruster is being assessed by thruster wear test and life-modeling of critical thruster components, such as the ion optics and cathodes. The NEXT Long-Duration Test (LDT) was initiated to validate and qualify the NEXT thruster propellant throughput capability. The NEXT thruster completed the primary goal of the LDT; namely to demonstrate the project qualification throughput of 450 kg by the end of calendar year 2009. The NEXT LDT has demonstrated 28,500 hr of operation and processed 466 kg of xenon throughput--more than double the throughput demonstrated by the NSTAR flight-spare. Thruster performance changes have been consistent with a priori predictions. Thruster erosion has been minimal and consistent with the thruster service life assessment, which predicts the first failure mode at greater than 750 kg throughput. The life-limiting failure mode for NEXT is predicted to be loss of structural integrity of the accelerator grid due to erosion by charge-exchange ions.
CUBESATS are relatively new spacecraft platforms that are typically deployed from a launch vehicle as a secondary payload, providing low-cost access to space for a wide range of end-users. These satellites are comprised of building blocks having dimensions of 10x10x10 cm cu and a mass of 1.33 kg (a 1-U size). While providing low-cost access to space, a major operational limitation is the lack of a propulsion system that can fit within a CubeSat and is capable of executing high delta v maneuvers. This makes it difficult to use CubeSats on missions requiring certain types of maneuvers (i.e. formation flying, spacecraft rendezvous). Recently, work has been performed investigating the use of iodine as a propellant for Hall-effect thrusters (HETs) 2 that could subsequently be used to provide a high specific impulse path to CubeSat propulsion. 3, 4 Iodine stores as a dense solid at very low pressures, making it acceptable as a propellant on a secondary payload. It has exceptionally high ρIsp (density times specific impulse), making it an enabling technology for small satellite near-term applications and providing the potential for systems-level advantages over mid-term high power electric propulsion options. Iodine flow can also be thermally regulated, subliming at relatively low temperature (less than 100 C) to yield I2 vapor at or below 50 torr. At low power, the measured performance of an iodine-fed HET is very similar to that of a state-of-the-art xenon-fed thruster. Just as importantly, the current-voltage discharge characteristics of low power iodine-fed and xenon-fed thrusters are remarkably similar, potentially reducing development and qualifications costs by making it possible to use an already-qualified xenon-HET PPU in an iodine-fed system. Finally, a cold surface can be installed in a vacuum test chamber on which expended iodine propellant can deposit. In addition, the temperature doesn't have to be extremely cold to maintain a low vapor pressure in the vacuum chamber (it is under 10(exp -6) torr at −75 C), making it possible to 'cryopump' the propellant with lower-cost recirculating refrigerant-based systems as opposed to using liquid nitrogen or low temperature gaseous helium cryopanels. An iodine-based system is not without its challenges. The primary challenge is that the entire feed system must be maintained at an elevated temperature to prevent the iodine from depositing (transitioning from the gas phase directly back into the solid phase), which will block the propellant feed lines. Furthermore, deposition will occur unless the temperature in the lines is not greater than the temperature of the propellant reservoir. The flow rate can be controlled by adjusting the heating applied to the reservoir, but as with any thermal control there is a relatively slow response to changes in the heating rate. In the present paper, we describe the propulsion and propellant feed system for the iodine satellite (iSAT) flight demonstration mission. The system is based around the Busek BHT-200 Hall thruster, which has been modified for chemical compatibility with iodine vapor. While the gross propellant flow rate is maintained by the heated propellant reservoir, the flow to the anode and cathode are adjusted using two heated Vacco proportional flow control valves (PFCV), which provide very fast response on the flow rate adjustment. The flight mission design layout will be presented, showing how the system will be packaged into the overall 12-U spacecraft and the techniques being employed to protect the remaining spacecraft hardware from the propulsion system (e.g., plasma impingement, iodine deposition, thermal loads). In addition to the flight system design, results of testing the thruster and cathode with both operating on iodine propellant are presented. The tests are conducted on a thrust stand (see Fig. 1) in a large vacuum chamber containing a beam dump chilled to below -100 C to 'cryopump' the propellant. The thruster performance during these tests is presented, with these data used to evaluate the feed system and guide further refinements. Results of relatively long duration testing are presented to demonstrate the capability to operate for the length of the iSAT mission and to perform a number of re-starts as will be required by the mission concept of operations.
Flight qualification of ion thrusters typically requires testing on the order of 10,000 hours. Extensive knowledge of wear mechanisms and rates is necessary to establish design confidence prior to long duration tests. Consequently, real-time erosion rate measurements offer the potential both to reduce development costs and to enhance knowledge of the dependency of component wear on operating conditions. Several previous studies have used laser-induced fluorescence (LIF) to measure real-time, in situ erosion rates of ion thruster accelerator grids. Those studies provided only relative measurements of the erosion rate. In the present investigation, a molybdenum tube was resistively heated such that the evaporation rate yielded densities within the tube on the order of those expected from accelerator grid erosion. This work examines the suitability of the density cell as an absolute calibration source for LIF measurements, and the intrinsic error was evaluated.
Flight qualification of ion thrusters typically requires testing on the order of 10,000 hours. Extensive knowledge of wear mechanisms and rates is necessary to establish design confidence prior to long duration tests. Consequently, real-time erosion rate measurements offer the potential both to reduce development costs and to enhance knowledge of the dependency of component wear on operating conditions. Several previous studies have used laser induced fluorescence (LIF) to measure real-time, in situ erosion rates of ion thruster accelerator grids. Those studies provided only relative measurements of the erosion rate. In the present investigation, a molybdenum tube was resistively heated such that the evaporation rate yielded densities within the tube on the order of those expected from accelerator grid erosion. A pulsed UV laser was used to pump the ground state molybdenum at 345.64nm, and the non-resonant fluorescence at 550-nm was collected using a bandpass filter and a photomultiplier tube or intensified CCD array. The sensitivity of the fluorescence was evaluated to determine the limitations of the calibration technique. The suitability of the diagnostic calibration technique was assessed for application to ion engine erosion rate measurements.
Current production of LiOH, which is needed to make Li-ion battery cathode active materials, utilizes a multistep process including solar evaporation, precipitation with Na 2 CO 3 and then conversion to LiOH using Ca(OH) 2 . This process requires a large amount of land area for solar evaporation, the right weather conditions, and chemicals for the conversion process that result in NaCl and CaCO 3 waste products. The production of Ca(OH) 2 is very energy intensive and evolves significant quantities of CO 2 . An alternative process flow utilizing direct lithium extraction techniques, followed by a chemical free conversion process can have benefits in reducing the needed land requirements and chemicals for traditional brine processing. There are many potential direct lithium extraction technologies that are currently being developed. The direct lithium extraction process from typical brine sources will produce a LiCl solution with some impurities including typically high concentrations of Na. This brine then needs to be converted to LiOH for use in battery cathode production. Ideally this conversion could occur without the use of additional chemicals. Electrochemistry can do this conversion either via electrolysis or bipolar membrane electrodialysis (BPED) to produce LiOH and HCl in solution. BPED utilized bipolar membranes to split water, which has a reduced potential as compared to splitting water at electrodes into hydrogen and oxygen gas. This reduced potential required results in a significant energy savings for BPED over electrolysis methods. This CRADA project aimed to develop such an integrated process using direct lithium extraction followed by BPED to produce a LiOH solution. That solution can then be crystallized into battery grade LiOH. In particular, Albemarle utilized a direct lithium extraction process to produce a concentrated LiCl solution that could be used for the BPED process. The BPED process was first tested using various LiCl solutions with impurity ions added at bench scale to understand the effects of impurities and determine processing parameters. Then testing was performed using the direct lithium extracted brine at the bench scale before scaling the process up. After the process was scaled up a long duration test was carried out to estimate the lifetime of the membranes, which is key to the economics of the BPED process.
NASA is currently investigating continuous detonation cycle engines for the application of lander and interplanetary space exploration missions. The performance benefits of a detonation cycle engine may allow for a broader design trade space and more compact geometry required for future missions to the Moon and onwards towards Mars. However, the technology readiness level (TRL) within the US was found to be low with several major risk factors that require understanding prior to full engine system development. One area of uncertainty is the extreme heat loads expected during thermal steady state conditions. To achieve this, an announcement for collaborative opportunity (ACO) partnership between IN Space LLC and NASA Marshall Space Flight Center (MSFC) was established to explore integration of additive manufacturing (AM) processes and the high conductance copper-based alloys, GRCop-42 and GRCop-84. This work outlines the hot fire testing of a 7K lbf thrust class fully AM GRCop-alloy rotating detonation rocket engine (RDRE). Two annular thrust chamber configurations emulating a lander engine system were tested with LOx/GH2 and LOx/LCH4. In both configurations, select hardware was actively cooled using de-ionized water and regeneratively cooled using LCH4. All primary hardware survived long duration tests to thermal steady state up to 133 seconds in a single burn. In total, 18 starts and 802 seconds of duration were achieved with and without visual confirmation of waves present. The proportion of burned propellant, or level of complete combustion, was found to be high compared to the theoretically achievable mean chamber pressure in all cases.
Two long-duration tests were conducted with a surplus Mariner Mars 1969 monopropellant hydrazine reactor in an attempt to induce the washout phenomenon. The Mariner Mars 1969 reactor was chosen because it has a long development history and thus is well characterized. No washout occurred during either of the two 1000-s tests, although slow transients were observed in the reactor operation during what were nominally steady-state conditions. The 2000-s of operating time represents nearly an order of magnitude increase over the rated life of the engine.