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At least 1,477 records · Page 82

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NASA NTRS2006

Recovering Residual Xenon Propellant for an Ion Propulsion System

Future nuclear-powered Ion-Propulsion- System-propelled spacecraft such as Jupiter Icy Moon Orbiter (JIMO) will carry more than 10,000 kg of xenon propellant. Typically, a small percentage of this propellant cannot be used towards the end of the mission because of the pressure drop requirements for maintaining flow. For large missions such as JIMO, this could easily translate to over 250 kg of unusable xenon. A proposed system, the Xenon Recovery System (XRS), for recovering almost all of the xenon remaining in the tank, would include a cryopump in the form of a condenser/evaporator that would be alternatively cooled by a radiator, then heated electrically. When the pressure of the xenon in the tank falls below 0.7 MPa (100 psia), the previously isolated XRS will be brought online and the gas from the tank would enter the cryopump that is initially cooled to a temperature below saturation temperature of xenon. This causes xenon liquefaction and further cryopumping from the tank till the cryopump is full of liquid xenon. At this point, the cryopump is heated electrically by small heaters (70 to 80 W) to evaporate the liquid that is collected as high-pressure gas (<7 MPa; 1,000 psia) in an intermediate accumulator. Check valves between the tank and the XRS prevent the reverse flow of xenon during the heating cycle. The accumulator serves as the high-pressure source of xenon gas to the Xenon Feed System (XFS) downstream of the XRS. This cycle is repeated till almost all the xenon is recovered. Currently, this system is being baselined for JIMO.

Ganapathi, Gani↗

NASA NTRS2010

Apparatus and method for gelling liquefied gasses

The present invention is a method and apparatus for gelling liquid propane and other liquefied gasses. The apparatus includes a temperature controlled churn mixer, vacuum pump, liquefied gas transfer tank, and means for measuring amount of material entering the mixer. The method uses gelling agents such as silicon dioxide, clay, carbon, or organic or inorganic polymers, as well as dopants such as titanium, aluminum, and boron powders. The apparatus and method are particularly useful for the production of high quality rocket fuels and propellants.

Elliott, Adam↗

NASA NTRS1972

Liquid-propellant dynamics and suppression

Technology related to liquid-propellant interactions with space shuttle vehicles is reviewed. Potential problems unique to the shuttle include liquid-structure interactions resulting from coupled lateral and longitudinal deformations, traveling wave phenomena at shallow propellant levels, and liquid impact during abort, staging, or docking. Technology efforts to define the slosh dynamics under shuttle operating conditions are described with emphasis on analytical representations of the liquid by finite-element and marker-and-cell methods. In addition, slosh suppression is discussed and includes baffle damping and pressure loads for tanks fitted with multiple baffles both above and below the undisturbed liquid surface.

Stephens, D. G.↗

NASA NTRS1978

Design and performance verification of a passive propellant management system

This paper describes the design and verification testing of a reusable passive propellant management system. The system was designed to acquire propellant in low- or zero-g environments and also retain this propellant under high axially directed accelerations that may be experienced during launch and orbit-to-orbit transfer. The system design requirements were established to satisfy generally the requirements for a large number of potential NASA and military applications, such as orbit-to-orbit shuttles and satellite vehicles. The resulting concept was a multicompartmented tank with independent surface tension acquisition channels in each compartment. The tank was designed to provide a minimum expulsion efficiency of 98 percent when subjected to the simultaneous conditions of acceleration, vibration, and outflow. The system design has the unique capability to demonstrate low-g performance in a 1-g test environment, and the test program summarized was structured around this capability.

Hess, D. A.↗

NASA NTRS1991

Design of an Advanced Expander Test Bed

The Advanced Expander Test Bed (AETB) is the key element for development of technology for future space engines. The AETB will be used to validate the high pressure expander cycle concept, investigate system interactions and conduct investigations of advanced mission focused components and new health monitoring techniques. The AETB will use oxygen/hydrogen propellants and a split expander cycle with nominal operation at a combustion chamber pressure of 1200 psia, a mixture ratio of 6.0, and an equivalent vacuum thrust of 20,000 lbf. It will function over a wide range of conditions including throttling to 5 percent thrust, operation at a mixture ratio of 12.0, and operation in tank head idle and pumped idle modes.

Masters, Arthur I.↗

NASA NTRS2023

Concepts for the Liquefaction of Hydrogen for In-Situ Operations on the Lunar or Martian Surface

While oxygen will probably be the first fluid liquefied from in-situ developed propellants, hydrogen will also be a sought-after commodity on both the Lunar and Martian surfaces. General hydrogen liquefaction processes are not significantly affected by the atmosphere on Mars and similar processes can be applied on both surfaces. However, hydrogen will require different processes than oxygen (which can be directly liquefied into its final use tank) due to the different spin states the hydrogen nucleus (i.e. parahydrogen and orthohydrogen) can have as well as the significant sensible energy that has to be removed prior to the actual phase change. Due to the low efficiency of 20 K class cryocoolers and high sensible heat removal requirements at higher temperatures, significant benefits can be obtained by precooling the hydrogen gas in a variety of manners. The benefits of three concepts are explored to take advantage of different precooling technics as well as accounting for sensible, latent, and para to ortho conversion heating requirements. Process concepts are shown that can reduce the base input power requirements for heat removal by more than 67% and the radiator area by similar amounts. A majority of the technology development for these processes has already begun, both within NASA and externally, with the few gaps being identified.

ISRU↗

NASA NTRS2023

Concepts for the Liquefaction of Hydrogen for In-Situ Operations on the Lunar or Martian Surface

ISRU↗

NASA NTRS2026

Assessing the Adhesion of Nanofibrous PVDF-HFP as Passive Thermal Control Coatings for the Extraterrestrial Storage of Cryogenic Propellants

Cryogenic propellant passive thermal management techniques in space are essential complements to active thermal management methods, helping reduce absolute dependence on the latter. Particularly, materials with superior optical properties (high solar reflectance and infrared emittance) would be beneficial for passive thermal control to reject most of the incident solar radiation and enhance thermal emission from the storage tank, leading to self-cooling. We developed a nanofibrous, porous passive thermal control material by electrospinning polyvinylidene fluoride-co-hexafluoropropylene (PVDF-HFP) co-polymer onto an aluminum foil substrate. The material exhibits a very high solar reflectance (>99%) and infrared emittance (~86%) when characterized in the ultraviolet-visible-near infrared and mid-infrared wavelengths. This study focuses on the adhesion of the nanofibrous PVDF-HFP to the substrate. We incorporated in-lab fabricated and commercial adhesives, under normal room conditions, between the nanofibers and the aluminum foil substrate to improve their adhesion. Some samples were also exposed to space-like conditions – vacuum and thermal cycling, to investigate their effects on the strengths of adhesives. Subsequently, adhesion peel tests were conducted on all the samples to assess the peel strengths of the bonds. Peel test strengths after space environment conditioning of the adhesives were generally higher than those observed in the pristine, unaged samples. The results are discussed in this study.

Chieloka Ibekwe↗

NASA NTRS2014

Atmospheric Processing Module for Mars Propellant Production

The multi-NASA center Mars Atmosphere and Regolith COllectorPrOcessor for Lander Operations (MARCO POLO) project was established to build and demonstrate a methaneoxygen propellant production system in a Mars analog environment. Work at the Kennedy Space Center (KSC) Applied Chemistry Laboratory is focused on the Atmospheric Processing Module (APM). The purpose of the APM is to freeze carbon dioxide from a simulated Martian atmosphere containing the minor components nitrogen, argon, carbon monoxide, and water vapor at Martian pressures (8 torr) by using dual cryocoolers with alternating cycles of freezing and sublimation. The resulting pressurized CO(sub 2) is fed to a methanation subsystem where it is catalytically combined with hydrogen in a Sabatier reactor supplied by the Johnson Space Center (JSC) to make methane and water vapor. We first used a simplified once-through setup and later employed a H(sub 2)CO(sub 2) recycling system to improve process efficiency. This presentation and paper will cover (1) the design and selection of major hardware items, such as the cryocoolers, pumps, tanks, chillers, and membrane separators, (2) the determination of the optimal cold head design and flow rates needed to meet the collection requirement of 88 g CO(sub 2) hr for 14 hr, (3) the testing of the CO(sub 2) freezer subsystem, and (4) the integration and testing of the two subsystems to verify the desired production rate of 31.7 g CH(sub 4) hr and 71.3 g H(sub 2)O hr along with verification of their purity. The resulting 2.22 kg of CH(sub 2)O(sub 2) propellant per 14 hr day (including O(sub 2) from electrolysis of water recovered from regolith, which also supplies the H(sub 2) for methanation) is of the scale needed for a Mars Sample Return mission. In addition, the significance of the project to NASAs new Mars exploration plans will be discussed.

waste treatment↗

NASA NTRS2014

Atmospheric Processing Module for Mars Propellant Production

The multi-NASA center Mars Atmosphere and Regolith COllector/PrOcessor for Lander Operations (MARCO POLO) project was established to build and demonstrate a methaneoxygen propellant production system in a Mars analog environment. Work at the Kennedy Space Center (KSC) Applied Chemistry Laboratory is focused on the Atmospheric Processing Module (APM). The purpose of the APM is to freeze carbon dioxide from a simulated Martian atmosphere containing the minor components nitrogen, argon, carbon monoxide, and water vapor at Martian pressures (approx. 8 torr) by using dual cryocoolers with alternating cycles of freezing and sublimation. The resulting pressurized CO(sub 2) is fed to a methanation subsystem where it is catalytically combined with hydrogen in a Sabatier reactor supplied by the Johnson Space Center (JSC) to make methane and water vapor. We first used a simplified once-through setup and later employed a H(sub 2)CO(sub 2) recycling system to improve process efficiency. This presentation and paper will cover (1) the design and selection of major hardware items, such as the cryocoolers, pumps, tanks, chillers, and membrane separators, (2) the determination of the optimal cold head design and flow rates needed to meet the collection requirement of 88 g CO(sub 2) hr for 14 hr, (3) the testing of the CO(sub 2) freezer subsystem, and (4) the integration and testing of the two subsystems to verify the desired production rate of 31.7 g CH(sub 4) hr and 71.3 g H(sub 2)O hr along with verification of their purity. The resulting 2.22 kg of CH(sub 2)O(sub 2) propellant per 14 hr day (including O(sub 2) from electrolysis of water recovered from regolith, which also supplies the H(sub 2) for methanation) is of the scale needed for a Mars Sample Return mission. In addition, the significance of the project to NASAs new Mars exploration plans will be discussed.

hydrocarbon fuel production↗

NASA NTRS1987

Independent Orbiter Assessment (IOA): Analysis of the reaction control system, volume 1

The results of the Independent Orbiter Assessment (IOA) of the Failure Modes and Effects Analysis (FMEA) and Critical Items List (CIL) are presented. The IOA approach features a top-down analysis of the hardware to determine failure modes, criticality, and potential critical items. To preserve independence, this analysis was accomplished without reliance upon the results contained within the NASA FMEA/CIL documentation. This report documents the independent analysis results for the Reaction Control System (RCS). The purpose of the RCS is to provide thrust in and about the X, Y, Z axes for External Tank (ET) separation; orbit insertion maneuvers; orbit translation maneuvers; on-orbit attitude control; rendezvous; proximity operations (payload deploy and capture); deorbit maneuvers; and abort attitude control. The RCS is situated in three independent modules, one forward in the orbiter nose and one in each OMS/RCS pod. Each RCS module consists of the following subsystems: Helium Pressurization Subsystem; Propellant Storage and Distribution Subsystem; Thruster Subsystem; and Electrical Power Distribution and Control Subsystem. Of the failure modes analyzed, 307 could potentially result in a loss of life and/or loss of vehicle.

Burkemper, V. J.↗

NASA NTRS1998

Performance of the Spacecraft Propulsion Research Facility During Altitude Firing Tests of the Delta 3 Upper Stage

The Spacecraft Propulsion Research Facility at the NASA Lewis Research Center's Plum Brook Station was reactivated in order to conduct flight simulation ground tests of the Delta 3 cryogenic upper stage. The tests were a cooperative effort between The Boeing Company, Pratt and Whitney, and NASA. They included demonstration of tanking and detanking of liquid hydrogen, liquid oxygen and helium pressurant gas as well as 12 engine firings simulating first, second, and third burns at altitude conditions. A key to the success of these tests was the performance of the primary facility systems and their interfaces with the vehicle. These systems included the structural support of the vehicle, propellant supplies, data acquisition, facility control systems, and the altitude exhaust system. While the facility connections to the vehicle umbilical panel simulated the performance of the launch pad systems, additional purge and electrical connections were also required which were unique to ground testing of the vehicle. The altitude exhaust system permitted an approximate simulation of the boost-phase pressure profile by rapidly pumping the test chamber from 13 psia to 0.5 psia as well as maintaining altitude conditions during extended steady-state firings. The performance of the steam driven ejector exhaust system has been correlated with variations in cooling water temperature during these tests. This correlation and comparisons to limited data available from Centaur tests conducted in the facility from 1969-1971 provided insight into optimizing the operation of the exhaust system for future tests. Overall, the facility proved to be robust and flexible for vehicle space simulation engine firings and enabled all test objectives to be successfully completed within the planned schedule.

Meyer, Michael L.↗

NASA NTRSDate not supplied

Lunar Helium-3: Mining Concepts, Extraction Research, and Potential ISRU Synergies

The M-3 miner was designed to excavate 1258 tonnes/hr with a bucket wheel excavator and process 556 tonnes/hr (during the lunar daytime) of <100 μm regolith through a 12 MW solar concentrator powered heat pipe heat exchanger(HPHX). The HPHX was designed to heat the regolith up to 700°C to release the implanted solar wind volatiles and recuperate 85% of the input heat. Considering a 20 ppb 3He concentration in the regolith, 66kg of 3He would be captured in onboard storage tanks over the course of one year.This mining effort would also result in the release hundreds of tonnes of other valuable volatile by-products, including water and hydrogen, that could be used to support sustainable lunar exploration.In the more near term, a mining operation to demonstrate the ability to collect 15 tonnes water for liquid oxygen/liquid hydrogen propellant from 5% water rich regolith may require ~400 tonnes of regolith to be excavated. This type of pilot scale operation could also demonstrate the ability to extract about 6 g of 3He, assuming the 3He concentration in permanently shadowed regions on the Moon is similar to that of equatorial regions.

Aaron D.S. Olson↗

NASA NTRSDate not supplied

Computational Fluid Dynamics Simulation of Methane Slosh and Drain Experiments

NASA possesses a wealth of historical cryogenic experiments that provide valuable insights into the design, troubleshooting, and understanding involved in the complex fluid and thermodynamics of managing cryogenic propellants. One approach to leveraging these historical datasets is by simulating these experiments to validate the accuracy of simulation environments and constituent models. This study focused on simulating a selection of the K site test series for both static and sloshing pressurized liquid methane draining experiments conducted at NASA in the 1970’s, utilizing computational fluid dynamics. The simulations were performed in the ANSYS FLUENT environment using the Volume of Fluid (VOF) numerical approach. A k-omega turbulence model was used with interfacial turbulence damping, and accurately predicted the amount of pressurant needed to maintain the required tank pressure throughout the static drain. The static simulation predicted the temperature stratification in the ullage observed at the end of the drain. During the methane expulsion with sloshing test, many features were successfully captured using the k-omega turbulence model with interfacial turbulence damping included. The rate of phase change and liquid temperature was overpredicted compared to the experimental measurements. The overprediction may be attributed to uncertainties in the vessel geometry, methane pressurant temperature and composition, and methodological differences in how the sloshing frequency was adjusted during the expulsion.

Cryogenic Propellants↗

NASA NTRS1972

Thermal/structural optimization of cryogenic storage systems.

An analytical method has been developed to optimize a cryogenic storage system designed for long term storage of cryogenic propellants. The type of propellant, mission duration, thermal environment, insulation thermal properties, insulation vent time, and the storage system weight penalty imposed by the design pressure were considered as the dependent parameters. The optimum system insulation thickness and vent pressure were then calculated in terms of these parameters. The method was applied to the design of the Space Shuttle Orbiter oxygen and hydrogen tanks to obtain an optimum design. It was determined that for the mission examined, the hydrogen tank optimum design was close to the initial pressure. A continuously vented hydrogen tank, therefore, was close to the optimum design. Also, it was determined that for this particular mission the oxygen tank should not be vented.

Navickas, J.↗

NASA NTRS1981

The supply of lunar oxygen to low earth orbit

Since oxygen makes up 86% of the total mass of propellants which would normally have been brought up from the earth to LEO, considerable savings are available if this oxygen can be obtained from the moon for little Delta-V penalty. This paper presents a scenario in which 400 T/yr of LOX is delivered to LEO, with the ability for upgrading to 5000 T/yr. In this scenario, cylindrical tanks of liquid oxygen with a mass of 500 kg are launched from the lunar surface by a mass driver and rendezvous with a collection station in a 100-km lunar orbit. The oxygen is removed from each tank and placed into a tanker OTV which later will transfer from low lunar orbit to LEO with an aerobraking maneuver. Launch requirements for aerobraked chemical OTVs using earth oxygen are compared to those using lunar oxygen.

Andrews, D. G.↗

NASA NTRS2008

Behavior of Spinning Space Vehicles with Onboard Liquids, 2nd Edition, Technical Report B8030

Although the fundamental principles of spin stabilization are well established, uncertainty regarding the potential for rapid nutation growth caused by onboard liquids is a continuing concern. NASA and other organizations regularly encounter the issue of rapid nutation growth due to energy dissipation by liquids on spinning vehicles. Of concern is the stability of spinning upper stages and of spacecraft that spin for part or all of their missions. Several missions have required last-minute hardware or operational changes to deal with rapid nutation divergences that were identified late in the program. In some instances, major schedule slips were barely averted. In at least two cases, it was determined that a spinning upper stage was not a viable option. Historically, the "slosh" issue has been addressed by each space vehicle project individually, if it has been addressed at all. Due to budgetary and programmatic constraints, individual projects are unable to address the problem globally. Hence, there has been little effort to collect available test and flight data and use that data to make a coherent, unified picture of the "slosh" effect and how to deal with it. To some extent, each project has had to "reinvent the wheel", which can be both costly and risky. This study is a step toward correcting the situation. Specifically, the goal was to identify and collect available flight and test data for spinning vehicles with onboard liquid propellants. A total of 149 flight data points and 1,692 test points were collected as part of this study. This data was analyzed, correlated, and is presented here in a normalized form. In most cases, the normalization involves a dimensionless nutation time constant that can be used to predict performance of other vehicles with the same type of tank. For some configurations, it was also possible to identify conditions that can lead to resonance between nutational motion and liquid modes. Gaps in the knowledge base are identified and approaches to filling those gaps are outlined. The data presented here has two different but related uses. First, it can be applied directly to current and future spacecraft programs. Second, it can provide truth models for testing analytical techniques. Experience has shown that purely analytical models of the liquid "slosh" effect on spinning vehicles are unreliable unless they are validated against flight or test data. To the author's knowledge, this report contains the most extensive and varied data set available. As such, it should be a good resource for anyone seeking to develop and validate improved analytical techniques. All of the original digital data sets have been archived on disk, with copies provided to NASA/KSC. With some restrictions, many of these data sets can be made available to researchers within the United States. Whenever possible, spacecraft are identified by name in this report. However, several organizations provided access to data with the explicit proviso that their programs not be identified and that parameters be presented only in normalized form. These constraints have been respected.

Hubert, Carl↗

NASA NTRS1971

Space shuttle safety - A hybrid vehicle breeds new problems.

Discussion of a few novel problems raised by the design and flight plan of the space shuttle and by the dangerous cargos it might carry. Among the problems cited are those connected with the inspection of the bearings of the propellant turbopumps, particularly those of the hydrogen pump, for evidence of spalling, as well as problems arising in the inspection of the high-temperature parts of the combustor and turbine section of the airbreathing turbofan for shuttle booster and orbiter, and problems resulting from the possibility of fire hazard due to spontaneous ignition of fuel vapor in the fuel tank vapor space.

Pinkel, I. I.↗