Engineering Papers⌕ Search

SEARCH · Engineering Papers

Results for “Propellant Tank”

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.

Quote a phrase for an exact phrase match. Source license links do not imply unrestricted reuse.

At least 1,567 records · Page 87
All sourcesNASA NTRSDOE OSTI
NASA NTRS2004

Taming Liquid Hydrogen: The Centaur Upper Stage Rocket, 1958-2002

During its maiden voyage in May 1962, a Centaur upper stage rocket, mated to an Atlas booster, exploded 54 seconds after launch, engulfing the rocket in a huge fireball. Investigation revealed that Centaur's light, stainless-steel tank had split open, spilling its liquid-hydrogen fuel down its sides, where the flame of the rocket exhaust immediately ignited it. Coming less than a year after President Kennedy had made landing human beings on the Moon a national priority, the loss of Centaur was regarded as a serious setback for the National Aeronautics and Space Administration (NASA). During the failure investigation, Homer Newell, Director of Space Sciences, ruefully declared: "Taming liquid hydrogen to the point where expensive operational space missions can be committed to it has turned out to be more difficult than anyone supposed at the outset." After this failure, Centaur critics, led by Wernher von Braun, mounted a campaign to cancel the program. In addition to the unknowns associated with liquid hydrogen, he objected to the unusual design of Centaur. Like the Atlas rocket, Centaur depended on pressure to keep its paper-thin, stainless-steel shell from collapsing. It was literally inflated with its propellants like a football or balloon and needed no internal structure to give it added strength and stability. The so-called "pressure-stabilized structure" of Centaur, coupled with the light weight of its high- energy cryogenic propellants, made Centaur lighter and more powerful than upper stages that used conventional fuel. But, the critics argued, it would never become the reliable rocket that the United States needed.

Dawson, Virginia P.↗
NASA NTRS2011

Robust Control Algorithm for a Two Cart System and an Inverted Pendulum

The Rectilinear Control System can be used to simulate a launch vehicle during liftoff. Several control schemes have been developed that can control different dynamic models of the rectilinear plant. A robust control algorithm was developed that can control a pendulum to maintain an inverted position. A fluid slosh tank will be attached to the pendulum in order to test robustness in the presence of unknown slosh characteristics. The rectilinear plant consists of a DC motor and three carts mounted in series. Each cart s weight can be adjusted with brass masses and the carts can be coupled with springs. The pendulum is mounted on the first cart and an adjustable air damper can be attached to the third cart if desired. Each cart and the pendulum have a quadrature encoder to determine position. Full state feedback was implemented in order to develop the control algorithm along with a state estimator to determine the velocity states of the system. A MATLAB program was used to convert the state space matrices from continuous time to discrete time. This program also used a desired phase margin and damping ratio to determine the feedback gain matrix that would be used in the LabVIEW program. This experiment will allow engineers to gain a better understanding of liquid propellant slosh dynamics, therefore enabling them to develop more robust control algorithms for launch vehicle systems

Wilson, Chris L.↗
NASA NTRS2008

Active Co-Storage of Cryogenic Propellants for Lunar Explortation

Long-term storage of cryogenic propellants is a critical requirement for NASA's effort to return to the moon. Liquid hydrogen and liquid oxygen provide the highest specific impulse of any practical chemical propulsion system, and thus provides the greatest payload mass per unit of launch mass. Future manned missions will require vehicles with the flexibility to remain in orbit for months, necessitating long-term storage of these cryogenic liquids. For decades cryogenic scientific satellites have used dual cryogens with different temperatures to cool instruments. This technology utilizes a higher temperature cryogen to provide a stage that efficiently intercepts a large fraction of the heat that would otherwise be incident on the lower temperature cryogen. This interception reduces the boil-off of the lower temperature cryogen and increasing the overall life-time of the mission. The Active Co-Storage concept is implemented similarly; the 101 K liquid oxygen thermally shields the 24 K liquid hydrogen. A thermal radiation shield that is linked to the liquid oxygen tank shrouds the liquid hydrogen tank, thereby preventing the liquid hydrogen tank from being directly exposed to the 300 K external environment. Modern cryocooler technology can eliminate the liquid oxygen boil-off and also cool the thermal radiation shield thereby reducing the liquid hydrogen boil-off to a small fraction of the unshielded rate. The thermal radiation shield can be a simple conductive shroud or a more sophisticated but lighter Broad Area Cooling (BAC) shroud. The paper describes the design impact of an active co-storage system for the Altair Descent Vehicle. This paper also compares the spacecraft-level impacts of the conductive shroud and the BAC shroud active co-storage concepts with a passive storage option in the context of the different scales of spacecraft that will be used for the lunar exploration effort - the Altair Ascent and Descent Vehicles, the Orion, and the Ares V Earth Departure Stage. The paper also reports on a subscale test of this active co-storage configuration. The test tank is 0.7 m in diameter, approximately one-third the dimension of tanks that would be needed in a lunar ascent module. A thin-walled fiberglass skirt supports and isolates the tank from a 100 K stage. A similar thin-walled skirt supports the lOOK stage from the ambient temperature structure. An aluminum shield with a heavy MLI blanket surrounds the tank and is attached at the 100 K stage. In this initial phase of the project, there is no tank on the 100 K stage, but it is actively cooled by a single-stage cryocooler similar in design to the one used on the RHESSI mission. The test configuration includes a number of innovative elements, including a helical support heat exchanger and an external thermodynamic vent/heat interception system. To avoid the complexity of an explosive gas handling system, testing will be done with liquid helium and liquid neon as simulant fluids. The properties of these fluids bracket the properties of liquid hydrogen. Instrumentation allows tank temperature and shield temperature profiles, tank liquid levels, and pressure drops through the flow lines, to be measured.

Mustafi, S.↗
NASA NTRS1994

Cryogenic on-orbit liquid depot storage acquisition and transfer (COLD-SAT) experiment subsystem instrumentation and wire harness design report

Subcritical cryogens such as liquid hydrogen (LH2) and liquid oxygen (LO2) are required for space based transportation propellant, reactant, and life support systems. Future long-duration space missions will require on-orbit systems capable of long-term cryogen storage and efficient fluid transfer capabilities. COLD-SAT, which stands for cryogenic orbiting liquid depot-storage acquisition and transfer, is a free-flying liquid hydrogen management flight experiment. Experiments to determine optimum methods of fluid storage and transfer will be performed on the COLD-SAT mission. The success of the mission is directly related to the type and accuracy of measurements made. The instrumentation and measurement techniques used are therefore critical to the success of the mission. This paper presents the results of the COLD-SAT experiment subsystem instrumentation and wire harness design effort. Candidate transducers capable of fulfilling the COLD-SAT experiment measurement requirements are identified. Signal conditioning techniques, data acquisition requirements, and measurement uncertainty analysis are presented. Electrical harnessing materials and wiring techniques for the instrumentation designed to minimize heat conduction to the cryogenic tanks and provide optimum measurement accuracy are listed.

Edwards, Lawrence G.↗
NASA NTRS2015

The Zero Boil-Off Tank Experiment Contributions to the Development of Cryogenic Fluid Management

The Zero Boil-Off Technology (ZBOT) Experiment involves performing a small scale ISS experiment to study tank pressurization and pressure control in microgravity. The ZBOT experiment consists of a vacuum jacketed test tank filled with an inert fluorocarbon simulant liquid. Heaters and thermo-electric coolers are used in conjunction with an axial jet mixer flow loop to study a range of thermal conditions within the tank. The objective is to provide a high quality database of low gravity fluid motions and thermal transients which will be used to validate Computational Fluid Dynamic (CFD) modeling. This CFD can then be used in turn to predict behavior in larger systems with cryogens. This paper will discuss the current status of the ZBOT experiment as it approaches its flight to installation on the International Space Station, how its findings can be scaled to larger and more ambitious cryogenic fluid management experiments, as well as ideas for follow-on investigations using ZBOT like hardware to study other aspects of cryogenic fluid management.

Fluid management↗
NASA NTRS1990

Laboratory Simulation of the Effect of Rocket Thrust on a Precessing Space Vehicle

Ground tests of solid propellant rocket motors have shown that metal-containing propellants produce various amounts of slag (primarily aluminum oxide) which is trapped in the motor case, causing a loss of specific impulse. Although not yet definitely established, the presence of a liquid pool of slag also may contribute to nutational instabilities that have been observed with certain spin-stabilized, upper-stage vehicles. Because of the rocket's axial acceleration, absent in the ground tests, estimates of in-flight slag mass have been very uncertain. Yet such estimates are needed to determine the magnitude of the control authority of the systems required for eliminating the instability. A test rig with an eccentrically mounted hemispherical bowl was designed and built which incorporates a follower force that properly aligns the thrust vector along the axis of spin. A program that computes the motion of a point mass in the spinning and precessing bowl was written. Using various RPMs, friction factors, and initial starting conditions, plots were generated showing the trace of the point mass around the inside of the fuel tank. The apparatus will incorporate future design features such as a variable nutation angle and a film height measuring instrument. Data obtained on the nutational instability characteristics will be used to determine order of magnitude estimates of control authority needed to minimize the sloshing effect.

Alvarez, Oscar↗
NASA NTRS1990

Laboratory simulation of the rocket motor thrust as a follower force

Ground tests of solid propellant rocket motors have shown that metal-containing propellants produce various amounts of slag (primarily aluminum oxide), which is trapped in the motor case causing a loss of specific impulse. Although not yet definitely established, the presence of a liquid pool of slag also may contribute to nutational instabilities that have been observed with certain spin-stabilized, upper-stage vehicles. Because of the rocket's axial acceleration - absent in the ground tests - estimates of in-flight slag mass have been very uncertain. Yet such estimates are needed to determine the magnitude of the control authority of the systems required for eliminating the instability. A test rig with an eccentrically mounted hemispherical bowl was designed and built that incorporates a follower force that properly aligns the thrust vector along the axis of spin. A program that computes the motion of a point mass in the spinning and precessing bowl was written. Using various rpm, friction factors, and initial starting conditions, plots were generated showing the trace of the point mass around the inside of the fuel tank. The apparatus will be used extensively during the 1990 to 1991 academic year and incorporate future design features such as a variable nutation angle and a film height measuring instrument. Data obtained on the nutational instability characteristics will be used to determine order-of-magnitude estimates of control authority needed to minimize the sloshing effect.

Source record↗
NASA NTRS2021

Modeling the Effects of Liquid-to-Gas Density Ratio on Slosh Dynamics

Mechanical models are commonly used in Guidance, Navigation and Controls (GN&C) system-level models to represent propellant slosh forces on launch vehicles and spacecraft. The slosh model parameters are typically predicted using semi-empirical analytical methods that are based on experimental data for high liquid-to-gas density ratios (such as water and air at standard sea level conditions). Model parameter calculations typically neglect any contribution to slosh dynamics from the gas phase. However, some cryogenic propellant systems operate in conditions where the density ratio can be orders of magnitude smaller. Analytical and computational modeling was used in this effort to investigate the effects of liquid-to-gas density ratio on slosh dynamics. No experimental data nor previous studies are available at this time that quantify the effect of liquid-to-gas density ratio on slosh dynamics. However, in a recent CFD study, results showed that density ratio can have a significant effect on slosh model parameters. In this study, a dual pendulum slosh model was created that distinctly represents both liquid and gas phase dynamics. Solution of the model, which invokes the Euler-Lagrange equations of motion, demonstrates that slosh frequency is a combination of liquid slosh frequency and gas phase slosh frequency. Additionally, it shows that slosh mass is reduced due to the opposing motion of the gas phase with respect to that of the liquid phase. The slosh model parameter trends were verified using Computational Fluid Dynamics (CFD) analysis results for various liquid-to-gas density ratios, tanks, and fill levels.

Christopher D Moore↗
NASA NTRS2016

NASA's Hydrogen Outpost: The Rocket Systems Area at Plum Brook Station

"There was pretty much a general knowledge about hydrogen and its capabilities," recalled former researcher Robert Graham. "The question was, could you use it in a rocket engine? Do we have the technology to handle it? How will it cool? Will it produce so much heat release that we can't cool the engine? These were the questions that we had to address." The National Aeronautics and Space Administration's (NASA) Glenn Research Center, referred to historically as the Lewis Research Center, made a concerted effort to answer these and related questions in the 1950s and 1960s. The center played a critical role transforming hydrogen's theoretical potential into a flight-ready propellant. Since then NASA has utilized liquid hydrogen to send humans and robots to the Moon, propel dozens of spacecraft across the universe, orbit scores of satellite systems, and power 135 space shuttle flights. Rocket pioneers had recognized hydrogen's potential early on, but its extremely low boiling temperature and low density made it impracticable as a fuel. The Lewis laboratory first demonstrated that liquid hydrogen could be safely utilized in rocket and aircraft propulsion systems, then perfected techniques to store, pump, and cleanly burn the fuel, as well as use it to cool the engine. The Rocket Systems Area at Lewis's remote testing area, Plum Brook Station, played a little known, but important role in the center's hydrogen research efforts. This publication focuses on the activities at the Rocket Systems Area, but it also discusses hydrogen's role in NASA's space program and Lewis's overall hydrogen work. The Rocket Systems Area included nine physically modest test sites and three test stands dedicated to liquid-hydrogen-related research. In 1962 Cleveland Plain Dealer reporter Karl Abram claimed, "The rocket facility looks more like a petroleum refinery. Its test rigs sprout pipes, valves and tanks. During the night test runs, excess hydrogen is burned from special stacks in the best Oklahoma oil field tradition." Besides the Rocket Systems Area, Plum Brook Station also included a nuclear test reactor, a large vacuum tank, a hypersonic wind tunnel, and a full-scale upper-stage rocket stand. The Rocket Systems Area operated from 1961 until NASA shut down all of Plum Brook in 1974. The center reopened Plum Brook in the late 1980s and continues to use several test facilities. The Rocket Systems Area, however, was not restored. Today Plum Brook resembles a nature preserve more than an oil refinery. Lush fields and forests separate the large test facilities. Until recently, the abandoned Rocket Systems Area structures and equipment were visible amongst the greenery. These space-age ruins, particularly the three towers, stood as silent sentinels over the sparsely populated reservation. Few knew the story of these mysterious facilities when NASA removed them in the late 2000s.

Rocket Systems↗
NASA NTRS2006

Analytical Assessment of the Reciprocating Feed System

A preliminary analysis tool has been created in Microsoft Excel to determine deliverable payload mass, total system mass, and performance of spacecraft systems using various types of propellant feed systems. These mass estimates are conducted by inserting into the user interface the basic mission parameters (e.g., thrust, burn time, specific impulse, mixture ratio, etc.), system architecture (e.g., propulsion system type and characteristics, propellants, pressurization system type, etc.), and design properties (e.g., material properties, safety factors, etc.). Different propellant feed and pressurization systems are available for comparison in the program. This gives the user the ability to compare conventional pressure fed, reciprocating feed system (RFS), autogenous pressurization thrust augmentation (APTA RFS), and turbopump systems with the deliverable payload, inert mass, and total system mass being the primary comparison metrics. Analyses of several types of missions and spacecraft were conducted and it was found that the RFS offers a performance improvement, especially in terms of delivered payload, over conventional pressure fed systems. Furthermore, it is competitive with a turbopump system at low to moderate chamber pressures, up to approximately 1,500 psi. Various example cases estimating the system mass and deliverable payload of several types of spacecraft are presented that illustrate the potential system performance advantages of the RFS. In addition, a reliability assessment of the RFS was conducted, comparing it to simplified conventional pressure fed and turbopump systems, based on MIL-STD 756B; these results showed that the RFS offers higher reliability, and thus substantially longer periods between system refurbishment, than turbopump systems, and is competitive with conventional pressure fed systems. This is primarily the result of the intrinsic RFS fail-operational capability with three run tanks, since the system can operate with just two run tanks.

Eddleman, David E.↗
DOE OSTI2017

Ensuring Reliable Natural Gas-Fired Generation with Fuel Contracts and Storage - DOE/NETL-2017/1816

This report finds that natural gas-fired power plants purchase fuel both on the spot market and through firm supply contracts; there do not appear to be clear drivers propelling power plants toward one or the other type. Most natural gas-fired power generators are located near major natural gas transmission pipelines, and most natural gas contracts are currently procured on the spot market. Although there is some regional variation in the type of contract used, a strong regional pattern does not emerge. Whether gas prices are higher with spot or firm contracts varies by both region and year. Natural gas prices that push the generators higher in the supply curve would make them less likely to dispatch. Most of the natural gas generators discussed in this report would be unlikely to enter firm contracts if the agreed price would decrease their dispatch frequency. The price points at which these generators would be unlikely to enter a firm contract depends upon the region that the generator is in, and how dependent that region is on natural gas. The Electric Reliability Council of Texas (ERCOT) is more dependent on natural gas than either Eastern Interconnection or Western Interconnection. This report shows that above-ground storage is prohibitively expensive with respect to providing storage for an extended operational fuel reserve comparable to the amount of on-site fuel storage used for coal-fired plants. Further, both pressurized and atmospheric tanks require a significant amount of land for storage, even to support one day’s operation at full output. Underground storage offers the only viable option for 30-day operational storage of natural gas, and that is limited by the location of suitable geologic formations and depleted fields.

01 COAL, LIGNITE, AND PEAT↗
NASA NTRSDate not supplied

A Crew and Logistics Lander for the Common Habitat Architecture

The Common Habitat Architecture is a conceptual study that explores the use of a large habitat derived from the Space Launch System (SLS) liquid oxygen tank as a core habitation element that can be used for crew missions or training in 0g, 1/6g, 3/8g, and 1g. This is not part of current NASA Artemis mission planning, but instead represents an architecture that could potentially follow after the Artemis missions and initial human Mars landings. This architecture leverages Starship-derived vehicles for crew landing on and ascent from the Moon and Mars in support of long-duration surface missions with pressurized crew transfer baselined as a nominal capability. This is not the Human Landing System variant of Starship but represents a number of modifications to enable long-duration surface missions with an eight-person crew size. A rapid brainstorming study was conducted in February of 2021, using only public data, to identify options to deliver the 90-ton Common Habitat to the surface and emplace it at the intended habitation site. This study compared three lunar lander concepts and three Mars lander concepts before ultimately selecting the SpaceX Starship as the most viable lander. The Common Habitat Architecture assumes that this Starship variant can further be modified for crew and cargo delivery, assuming that the use of a common system will lead to cost benefits. It is assumed that the Starship will expend too much propellant in landing crew and logistics to be able to launch, given the assumed absence of surface propellant production. Consequently, a separable ascent stage is used for crew ascent. The pressurized elements of this modified Starship are discussed: Starship Ascent Module, Airlock, Transfer Tunnel, Pressurized Crew Transfer Module, and Logistics Modules. For each element, a description, dimensions, rough mass estimates, core capabilities, and design features are presented. Key mechanisms and internal structures of the starship are also discussed. This will include flame diverters for the ascend module, the orbital docking hatch, Pressurized Crew Transfer Module Garage door, Starship Ascent Module fairing, flame diverter blow-out panels, Pressurized Crew Transfer Module lift system, logistics module lift system, contingency crew ascent via the logistics module lift system, internal catwalks, and structural interfaces. Concept of Operations will be discussed for both Moon and Mars. This will include in-space crew rendezvous and transfer, crewed landing, shirtsleeve crew transfer, crew departure, contingency surface operations, and final element disposition. Key differences for Mars will be discussed, such as the Martian atmosphere, dust storms, the absence of crew handover, and Deep Space Exploration Vehicle rendezvous. Contingency microgravity maintenance access will also be discussed. This work will demonstrate viable pressurized crew transfer with a Starship-based lander architecture. Forward work includes Garage lighting and camera systems, Starship Ascent Module propulsion system and docking port trades, contingency habitation trades, and mass/power equipment estimation. Finally, forward work includes developing a heavy cargo return system derived from the Crew and Logistics Starship. The goal of this system is the return of at least 10,000 kg payloads from the surfaces of the Moon and Mars.

Lunar Lander↗
NASA NTRS2017

An Overview of Recent Cryogenic Fluid Management Developments

Long-term storage, supply, and transfer of cryogenic fluids are critical capabilities needed to advance the human exploration of space. Technologies and advanced development programs have been pursued to address issues likely to confront the designers and developers of future cryogenic fluid management (CFM) subsystems and propulsion systems. NASA and aerospace industries have continued to conduct research and development for the propulsion applications of cryogenic fluids. From the well known in-space applications, to new applications involving superconducting motors within multiple different aircraft, and a renewed interest in production of oxygen on Mars, NASA continues to probe cryogenic applications within propulsion. This article presents brief reviews of several of the current CFM efforts to support future space missions. NASA GRC is currently focusing on understanding some of the finer points in the application of multilayer insulation. GRC recently finished the activation of a new calorimeter that operates at 20 K with a warm boundary that can operate either around 90 K or at 300 K. Current testing is focused on investigating multiple different seam concepts, initially between temperatures of 300 K and 20 K, and between 20 K and 90 K. One of the larger recent NASA investments has been on the Structural Heat Intercept, Insulation, and Vibration Experiment Rig (SHIIVER). SHIIVER is a 4 m diameter tank that is approximately 1/2 scale of the planned upper stage of the new Space Launch System (SLS) rocket. SHIIVER is focused on demonstrating the thermal benefits of multilayer insulation on the tank domes and boil-off vapor cooling on structural cylinders that hold the tank in-line on the rocket. It will also quantify any damage that may incur during the acoustic environment of over 160 decibels that SLS will experience on its trip to Earth orbit. In support of the possible production of oxygen out of the Mars' atmosphere, a team comprised of four NASA centers (GRC, MSFC, JSC, and KSC) collaborated to investigate multiple different hardware combinations, refrigeration cycles, and integration techniques to minimize power and mass of the storage and liquefaction system. Moreover, high efficiency, high capacity cryocoolers are an element of CFM which is essential for achieving NASA's future long duration missions. Currently in development is a 20W at 20K unit for use with liquid hydrogen. It is scheduled for delivery to NASA in the summer of 2017. Two Phase I SBIRS have recently concluded resulting in the preliminary design of two different 90K units, each having a refrigeration capacity of approximately 150W. Once the development efforts are complete, these high capacity 90K units will be integrated into the design of NASA's Liquid Oxygen (LOX)/Methane applications such as the Lander and Ascent Vehicles, or possibly an in-space stage. CFM technology is critical to the success of missions to Mars, Planetary Exploration, and In-Situ Resource Utilization (ISRU) for cryogenic propellant production. NASA is focusing on the development of CFM technologies needed to provide necessary data and relevant experience to support informed decisions on implementation of design of cryogenic systems for long term space missions. Lessons learned from the described CFM developments would lead to enhanced safety and reliability and enabling technologies which could allow NASA to meet future space exploration goals.

Hedayat, A.↗
NASA NTRS2018

Thermal Modeling of Zero Boil Off Tank Experiment

Understanding fluid behavior in microgravity environments is essential to further development of cryogenic storage in space environments. The Zero Boil Off Tank (ZBOT) experiment was designed to investigate two-phase pressurization and depressurization of a tank in a microgravity environment. The test fluid was the refrigerant Perfluoro-normal-Pentane (PNP). Thermal modeling for the ZBOT model was conducted using Thermal Desktop and SINDA/FLUINT. The temperature distribution within the fluid of the tank is of particular interest. This particular work is centered on ascertaining the thermal behavior of the refrigerant in order to build more complete models of fluid in microgravity. Separate cases were run modeling experiments that were conducted both on the ground and on the International Space Station (ISS) to compare 1g and microgravity environments. The microgravity modeling cases consisted of a fluid lump representing the vapor ullage suspended in a solid to represent the liquid. Mass flow between the liquid and vapor was modeled using the Schrage equation for mass flow. Initial results indicate that the pressure rise and temperature increase within the fluid closely align with the experimental data by matching initial conditions of the experiment. This work is ongoing and will yield further insights into the thermal behavior of fluid mixing in microgravity.

low-gravity fluid modeling↗
NASA NTRS2011

Hazards Induced by Breach of Liquid Rocket Fuel Tanks: Conditions and Risks of Cryogenic Liquid Hydrogen-Oxygen Mixture Explosions

We analyze the data of purposeful rupture experiments with LOx and LH2 tanks, the Hydrogen-Oxygen Vertical Impact (HOVI) tests that were performed to clarify the ignition mechanisms, the explosive power of cryogenic H2/Ox mixtures under different conditions, and to elucidate the puzzling source of the initial formation of flames near the intertank section during the Challenger disaster. We carry out a physics-based analysis of general explosions scenarios for cryogenic gaseous H2/Ox mixtures and determine their realizability conditions, using the well-established simplified models from the detonation and deflagration theory. We study the features of aerosol H2/Ox mixture combustion and show, in particular, that aerosols intensify the deflagration flames and can induce detonation for any ignition mechanism. We propose a cavitation-induced mechanism of self-ignition of cryogenic H2/Ox mixtures that may be realized when gaseous H2 and Ox flows are mixed with a liquid Ox turbulent stream, as occurred in all HOVI tests. We present an overview of the HOVI tests to make conclusion on the risk of strong explosions in possible liquid rocket incidents and provide a semi-quantitative interpretation of the HOVI data based on aerosol combustion. We uncover the most dangerous situations and discuss the foreseeable risks which can arise in space missions and lead to tragic outcomes. Our analysis relates to only unconfined mixtures that are likely to arise as a result of liquid propellant space vehicle incidents.

Osipov, Viatcheslav↗
NASA NTRS2016

Passive Water Assurance Delivery System

This report provides an analysis and evaluation of proposed design improvements to the internal water delivery system of NASA’s passive vegetable growing system. The VEGGIE system is designed to provide fresh vegetables and psychological benefits for astronauts aboard the International Space Station. Preliminary testing conducted by NASA for the VEGGIE system revealed flaws in the water delivery system. In these tests, the interface between the plants and the water reservoir failed to provide passive water delivery through the system both in microgravity experiments and experiments conducted on Earth. Our team’s goal was to fix this interface such that water can be passively delivered from the water reservoir to the plant rooting pillows to grow vegetative crops from seed to harvest. The system requirements outlined by NASA include: minimal total mass, on-demand passive water delivery, same dimensions as existing fixtures, minimal pressure on system to prevent leaking, zero mold growth, minimal swelling or clogging for non-flammable capillary materials, pressure stabilization between the water reservoir and plant pillow bags, even dispersal of water during initial priming, minimal bubble obstruction of capillary interface, maximum gas availability to plant roots, growth in a mixed artificial media (50:50 Arcillite:Fafard #2), and to avoid overwatering and drought conditions for plants. The major proposed design improvements are to use a single-interface capillary cord design to directly connect the water reservoir to the plant-rooting pillows, and to alter the water reservoir to model a propellant management device (PMD) in order to ensure consistent and long term watering for the VEGGIE system. The plant rooting pillow required minimal changes outside of replacing the capillary mat on the bottom of the pillow with O-ring insertion points the single-interface capillary system. - Research was conducted to determine which materials are able to uptake water through capillary action to grow Outredgeous Lettuce plants from seed to harvest. The primary requirements to be met by the team's design and evaluation of the capillary interface were: continuous passive watering for 90 days using non-flammable capillary materials and a peak water delivery rate of 30 mL/hr./0.15 m 2 . The results of the team's experimentation showed that the capillary material Nomex displayed the highest capillary water delivery potential with a maximum flow rate of 3.6 mL/hr. The Nomex capillary systems were the only capillary material to consistently grow Outredgeous Lettuce plants from seed to harvest, and displayed the highest average flow rate for multiple experiment sets. The Nomex material was previously incorporated into the VEGGIE system using a matted version of the capillary material, though the team recommends using a cord configuration in the single-interface capillary system for greater system stability. Nomex has proven to be a promising material, as it has passed both health and fire standards for use aboard the International Space Station. PMDs are made of materials that utilize surface tension and adhesive forces to overcome adverse accelerations to improve stability and ensure fluid delivery. PMDs are typically used in fuel tanks to ensure fuel delivery. The team recommends using a sponge PMD in order to mitigate bubble obstruction, decrease system weight, and ensure reliable water delivery to the capillary interface. The PMD water reservoir requires a rigid water reservoir and a vent tube for pressure stabilization. Since PMD’s cannot be tested in 1-G (Earth conditions), further theoretical modeling and testing is required for the proposed water reservoir design. The intent of this proposed system is to passively water plants in microgravity, though the technology is not limited to microgravity applications. The testing at The Ohio State University has proven that the design is highly effective on Earth, demonstrating that it could serve as a simple water delivery system in home and office applications. This would make vegetative crops more accessible in all indoor applications, thereby improving indoor air quality and occupant comfort. Additionally, this technology has great potential to be utilized in greenhouse plant production, cutting back on more sophisticated watering system energy and time requirements.

Alexandria Jensen↗
NASA NTRSDate not supplied

System Modeling of a Lunar Molten Regolith Electrolysis Plant

Introduction: In-Situ Resource Utilization (ISRU) is the process of extracting local resources to produce commodities for propulsion, life support systems, and off-planet construction rather than transporting consumables from Earth. Molten Regolith Electrolysis (MRE) is a novel ISRU method of extracting oxygen gas and metal alloy from lunar regolith. The MRE process involves placing lunar regolith between two electrodes, through which current is passed, to melt the regolith and reduce the metal oxide constituents by direct electrolysis (e.g. FeO, SiO2, MgO, Al2O3) into oxygen gas and metal alloys. The oxygen is liquefied and used as propellant for landers, while the metals (e.g. Ferro-alloys) are further processed and used in structural building materials and parts manufacturing. A system model was developed that accounted for the major processes of an MRE plant (from excavation of raw materials to storage of products) to assess the feasibility of a lunar MRE plant. The System Engineering and Integration (SE&I) ISRU Modeling and Analysis (SIMA) team utilized its previously documented system sizing model, the Mission Analysis and Integration Tool (MAIT) [1] as framework of the system model. MAIT uses MATLAB/Simulink to integrate subsystem models into a complete system model of the MRE plant. Total mass, volume, and power requirements were computed for numerous iterations of a MRE plant. System Model: Figure 1: MRE Plant Block Diagram The regolith excavation model determines the mass and power needed to excavate sufficient regolith. The preheating auger initiates the regolith heating process before regolith enters the MRE re-actor to reduce the energy required to turn the solid into a molten liquid. The MRE reactor is modeled in COMSOL Multiphysics and based on the research by Dominguez, Sibille, and Schreiner [2, 3, 4]. This preliminary reactor model provides an accurate calculation of thermal equilibrium during electrochemical operation of the reactor system to assess the optimal mass and power required to process the inlet flow of regolith. The model also computes the outlet flowrates of oxygen and molten products. For this analysis, the primary components of the metal alloy considered were iron and silicon. The oxygen is then purified using an Yttrium Stabilized Zirconia (YSZ) electrode, followed by liquefaction using a 90K cryocooler to be stored as liquid oxygen in insulated cylindrical tanks. In future iterations of the system model, the molten metal tapped from the MRE reactor will undergo additional processing or refinement. However, downstream handling of metals is currently a technology gap that is missing a high TRL subsystem model. Therefore, for this analysis, the accumulated metal alloy stream terminates after leaving the MRE reactor. Study Goals: This analysis investigates multiple input variables to the system to determine the sensitivity of a (near) complete plant at full-scale. This preliminary investigation ran parametric sweeps on the MRE reactor geometry, electrical current supply, layers of multi-layer insulation (MLI) on the reactor, size of the electrodes in the oxygen purification model, and regolith composition (based on landing site location). Three production targets of oxygen (1,000, 10,000, and 50,000 kg/yr) were investigated for this analysis. The parametric sweeps conducted in this analysis provide valuable insight into the expected impact of the various model inputs on plant size. This information can be used to identify the most critical components of the plant and guide future decisions on allocating funding for research and development, providing subsystem developers with appropriate interfaces with downstream and upstream processes, and assessing the overall feasibility of MRE when compared to other ISRU plants. References: [1] Carlson, A. et al. (2024) ICES, ICES-2024-53. [2] Dominguez, D.A., and Sibille, L. (2011) AIAA, AIAA-2011-700. [3] Schreiner, S.S. (2015) MIT, Dissertation. [4] Schreiner, S.S. et al. (2016) ASR, 57(7), pp.1585-1603.

ISRU↗
NASA NTRS2017

Repeatability of Cryogenic Multilayer Insulation

Due to the variety of requirements across aerospace platforms, and one off projects, the repeatability of cryogenic multilayer insulation has never been fully established. The objective of this test program is to provide a more basic understanding of the thermal performance repeatability of MLI systems that are applicable to large scale tanks. There are several different types of repeatability that can be accounted for: these include repeatability between multiple identical blankets, repeatability of installation of the same blanket, and repeatability of a test apparatus. The focus of the work in this report is on the first two types of repeatability. Statistically, repeatability can mean many different things. In simplest form, it refers to the range of performance that a population exhibits and the average of the population. However, as more and more identical components are made (i.e. the population of concern grows), the simple range morphs into a standard deviation from an average performance. Initial repeatability testing on MLI blankets has been completed at Florida State University. Repeatability of five GRC provided coupons with 25 layers was shown to be +/- 8.4 whereas repeatability of repeatedly installing a single coupon was shown to be +/- 8.0. A second group of 10 coupons have been fabricated by Yetispace and tested by Florida State University, through the first 4 tests, the repeatability has been shown to be +/- 16. Based on detailed statistical analysis, the data has been shown to be statistically significant.

multilayer insulation↗