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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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All sourcesNASA NTRSDOE OSTI
NASA NTRS1975

Particulate contamination spectrometer. Volume 1: Technical report

A laser particulate spectrometer (LPS) system was developed to measure the size and speed distributions of particulate (dusts, aerosols, ice particles, etc.) contaminants. Detection of the particulates was achieved by means of light scattering and extinction effects using a single laser beam to cover a size range of 0.8 to 275 microns diameter and a speed range of 0.2 to 20 meter/second. The LPS system was designed to operate in the high vacuum environment of a space simulation chamber with cold shroud temperatures ranging from 77 to 300 K.

Schmitt, R. J.↗
NASA NTRS1975

A laser particulate spectrometer for a space simulation facility

A laser particulate spectrometer (LPS) system was developed to measure the size and speed distributions of particulate contaminants. Detection of the particulates is achieved by means of light scattering and extinction effects using a single laser beam to cover a size range of 0.8 to 275 microns diameter and a speed range of 0.2 to 20 meters/second. The LPS system was designed to operate in the high-vacuum environment of a space simulation chamber with cold shroud temperatures ranging from 77 to 300 K.

Schmitt, R. J.↗
NASA NTRS1982

Development of a spinning wave heat engine

A theoretical analysis and an experimental investigation were conducted to assess the feasibility of developing a spinning wave heat engine. Such as engine would utilize a large amplitude traveling acoustic wave rotating around a cylindrica chamber, and it should not suffer from the inefficiency, noise, and intermittent thrust which characterizes pulse jet engines. The objective of this investigation was to determine whether an artificially driven large amplitude spinning transverse wave could induce a steady flow of air through the combustion chamber under cold flow conditions. In the theoretical analysis the Maslen and Moore perturbation technique was extended to study flat cylinders (pancake geometry) with completely open side walls and a central opening. In the parallel experimental study, a test moel was used to determine resonant frequencies and radial pressure distributions, as well as oscillatory and steady flow velocities at the inner and outer peripheries. The experimental frequency was nearly the same as the theoretical acoustic value for a model of the same outer diameter but without a central hole. Although the theoretical analysis did not predict a steady velocity component, simulaneous measurements of hotwire and microphone responses have shown that the spinning wave pumps a mean flow radially outward through the cavity.

Zinn, B. T.↗
NASA NTRS1992

Chamber For Testing Infrared Imaging Detectors

Chamber provides cold, vacuum, low-background-radiation environment and optical and electrical feedthroughs needed for testing and calibration of planar arrays of long-wavelength infrared photodetectors. Equipped with multiple ports for vacuum pumps, vacuum gauges, cryogenic fluids, sources of radiation, and camera heads. Accommodates variety of infrared sources, detectors, and associated equipment for testing at temperatures down to about 4.2 K in low-background-radiation environment.

Staller, Craig O.↗
NASA NTRS2003

Spectral Characteristic of Tholin Produced from Possible Early Earth Atmospheres and its Role in Antigreenhouse Effect on Early Earth

We have produced organic material simulating a methane photochemical haze in a CO2- rich atmosphere of the early Earth by irradiating gas mixtures in an inductively coupled cold plasma chamber with pressure approx. 0.25 mbar at 100 W total power. The flow rate was 24 cm3 min. We added progressively higher levels of CH, by combining gas mixtures of N2/CH4 (9/1) and N2/CO2 (9/1) to change the ratio of CH4/CO2. Tholin was accumulated for 5 hours in each experiment; the onset of tholin formation is in the range CH4/CO2 = 0.5 to 1. As the mixing ratio of CH, is increased, the production rate of the brownish tholin film increases. IR spectra showed the C-H and N-H bands similar to that of Titan tholin and closely resemble Titan tholin made at 0.13 mbar pressure. A decrease in the CH bonds on decreasing CH4/CO2 is noted. Ether bands (-(2-O-C) were tentatively detected, but no detectable carbonyl (C=O) band was found. The absorption in the UV region for the early Earth tholin is found to be substantially greater than the Titan tholin. Quantitative values of the optical constants of early Earth tholin are currently being measured.

Khare, B. N.↗
NASA NTRS2006

In-Vacuum Photogrammetry of a Ten-Meter Square Solar Sail

Solar sailing is a promising, future in-space propulsion method that uses the small force of reflecting sunlight to accelerate a large, reflective membrane without expendable propellants. One of two solar sail configurations under development by NASA is a striped net approach by L'Garde, Inc. This design uses four inflatably deployed, lightweight booms supporting a network of thin strings onto which four quadrants of ultrathin aluminized membranes are attached. The NASA Langley Research Center (LaRC) provided both experimental and analytical support to L'Garde for validating the structural characteristics of this unique, ultralightweight spacecraft concept. One of LaRC's responsibilities was to develop and apply photogrammetric methods to measure sail shape. The deployed shape provides important information for validating the accuracy of finite-element modeling techniques. Photogrammetry is the science and art of calculating 3D coordinates of targets or other distinguishing features on structures using images. A minimum of two camera views of each target is required for 3D determination, but having four or more camera views is preferable for improved reliability and accuracy. Using retroreflective circular targets typically provides the highest measurement accuracy and automation. References 3 and 4 provide details of photogrammetry technology, and reference 5 discusses previous experiences with photogrammetry for measuring gossamer spacecraft structures such as solar sails. This paper discusses the experimental techniques used to measure a L Garde 10-m solar sail test in vacuum with photogrammetry. The test was conducted at the NASA-Glenn Space Power Facility (SPF) located at Plum Brook Station in Sandusky, Ohio. The SPF is the largest vacuum chamber in the United States, measuring 30 m in diameter by 37 m in height. High vacuum levels (10(exp -6) torr) can be maintained inside the chamber, and cold environments (-195 C) are possible using variable-geometry cryogenic cold walls. This test used a vacuum level of approximately 1 torr (sufficient for structural static/dynamic characterization) and instead of using the cryogenic cold walls, used local LN2 cold plates underneath each of the four cold-rigidizable solar sail booms instead.

Pappa, Richard S.↗
NASA NTRS2008

Design and Development of an In-Space Deployable Sun Shield for the Atlas Centaur

The Centaur, by virtue of its use of high specific-impulse (Isp) LO2/LH2 propellants, has initial mass-to-orbit launch requirements less than half of those upper stages using storable propellants. That is, for Earth escape or GSO missions the Centaur is half the launch weight of a storable propellant upper stage. A drawback to the use of Liquid oxygen and liquid hydrogen, at 90 K and 20 K respectively, over storable propellants is the necessity of efficient cryogen storage techniques that minimize boil-off from thermal radiation in space. Thermal blankets have been used successfully to shield both the Atlas Centaur and Titan Centaur. These blankets are protected from atmospheric air loads during launch by virtue of the fact that the Centaur is enclosed within the payload fairing. The smaller Atlas V vehicle, the Atlas 400, has the Centaur exposed to the atmosphere during launch, and therefore, to date has not flown with thermal blankets shielding the Centaur. A design and development effort is underway to fly a thermal shield on the Atlas V 400 vehicle that is not put in place until after the payload fairing jettisons. This can be accomplished by the use of an inflatable and deployable thermal blanket referred to as the Centaur Sun Shield (CSS). The CSS design is also scalable for use on a Delta upper stage, and the technology potentially could be used for telescope shades, re-entry shields, solar sails and propellant depots. A Phase I effort took place during 2007 in a partnership between ULA and ILC Dover which resulted in a deployable proof-of-concept Sun Shield being demonstrated at a test facility in Denver. A Phase H effort is underway during 2008 with a partnership between ULA, ILC, NASA Glenn Research Center (GRC) and NASA Kennedy Space Center (KSC) to define requirements, determine materials and fabrication techniques, and to test components in a vacuum chamber at cold temperatures. This paper describes the Sun Shield development work to date, and the future plans leading up to a flight test in the 2011 time frame.

Dew, Michael↗
NASA NTRS1997

Transpiring Cooling of a Scram-Jet Engine Combustion Chamber

The peak cold-wall heating rate generated in a combustion chamber of a scram-jet engine can exceed 2000 Btu/sq ft sec (approx. 2344 W/sq cm). Therefore, a very effective heat dissipation mechanism is required to sustain such a high heating load. This research focused on the transpiration cooling mechanism that appears to be a promising approach to remove a large amount of heat from the engine wall. The transpiration cooling mechanism has two aspects. First, initial computations suggest that there is a reduction, as much as 75%, in the heat flux incident on the combustion chamber wall due to the transpirant modifying the combustor boundary layer. Secondly, the heat reaching the combustor wall is removed from the structure in a very effective manner by the transpirant. It is the second of these two mechanisms that is investigated experimentally in the subject paper. A transpiration cooling experiment using a radiant heating method, that provided a heat flux as high as 200 Btu/sq ft sec ( approx. 234 W/sq cm) on the surface of a specimen, was performed. The experiment utilized an arc-lamp facility (60-kW radiant power output) to provide a uniform heat flux to a test specimen. For safety reasons, helium gas was used as the transpirant in the experiments. The specimens were 1.9-cm diameter sintered, powdered-stainless-steel tubes of various porosities and a 2.54cm square tube with perforated multi-layered walls. A 15-cm portion of each specimen was heated. The cooling effectivenes and efficiencies by transpiration for each specimen were obtained using the experimental results. During the testing, various test specimens displayed a choking phenomenon in which the transpirant flow was limited as the heat flux was increased. The paper includes a preliminary analysis of the transpiration cooling mechanism and a scaling conversion study that translates the results from helium tests into the case when a hydrogen medium is used.

Choi, Sang H.↗
NASA NTRSDate not supplied

Cryobotics: Extreme Cold Environment Testing

The extreme cold environment test chamber was designed to conduct research in cryobotics; an area of study that focuses on robotic systems and rotating machinery operating in extreme cold environments including Earth, low Earth orbit, Mars, Moon, asteroids, Solar orbit, planetary orbit, or those encountered during travel among these destinations. The test chamber incorporates a modular dynamometer, consisting of a variety of brakes, torque sensors and motors to be easily interchanged between tests. Each test employs a unique test profile that incorporates different setpoints of applied torques and velocities for a given period or number of revolutions. The modularity of the dynamometer setup allows for any combination of motor, gearbox to be tested. The chamber has been used to run tests for various projects including Bulk Metallic Glass Gears (BMGG), Volatiles Investigating Polar Exploration Rover (VIPER), Intelligent Payload Experiment (IPEX), and Pilot Excavator. Various upgrades have been made to the extreme cold environment test chamber for the use of cryobotic research. These upgrades greatly increased the autonomous capabilities of the test set up by providing redundancies in the hardware and software. The redundancies were primarily added to protect the integrity of the cryohead. A new strapping and insulation method was performed to create the thermal conductive path from the actuators to the cryohead. The software was upgraded to include temperature setpoint control, further increasing the autonomous capabilities of the test. This paper goes into detail regarding the upgrades made to the extreme cold environment test chamber, as well as highlights the results from a COLDArm acceptance test.

Jonathan Drew Smith↗
NASA NTRSDate not supplied

Experimental Capabilities and Achievements of the Space Environment Dynamometer (SED)

The Space Environment Dynamometer (SED) chamber was designed to conduct research in cryobotics; an area of study that focuses on robotic systems and rotating machinery operating in extreme cold environments including Earth, low Earth orbit, Mars, Moon, asteroids, Solar orbit, planetary orbit, or those encountered during travel among these destinations. The test chamber incorporates a modular dynamometer, consisting of a variety of brakes, torque sensors and motors to be easily interchanged between tests. Each test employs a unique test profile that incorporates different setpoints of applied torques and velocities for a given period or number of rev- olutions. The modularity of the dynamometer setup allows for any combination of motor, gearbox to be tested. This chamber is one of a kind and resides at the Swamp Works facility within the Granular Mechanics and Regolith Operations laboratory (GMRO) at NASA Kennedy Space Center. Other cold chambers exist, however they are project specific and do not simultaneously output real time torque, temperature and efficiency data. The modularity of this extreme cold environment test chamber, coupled with the custom software and instrumentation, makes it one of a kind. The chamber is capable of stabilizing cryogenic temperatures and pressures to commensurate moon environments. Tests have successfully been conducted on motors and gearboxes for various cryogenic temperature set points, torques, and angular velocities. Relevant internal temperatures of the test article and chamber were recorded using a variety of temperature sensors. The temperature setpoints, on the motors and gearboxes tested, were stabilized by using PID gain scheduling of the PWM signal for the various heaters. The heat removal for the motor was provided by creating a thermally conductive path from a cryohead directly to the test article using copper straps. A variety of strainwave gears (SWGs), also known as Harmonic drives, planetary gear systems and DC motor actuator configurations have been tested in the chamber. The experiments performed were for various projects including Bulk Metallic Glass Gears (BMGG), Volatiles Investigating Polar Exploration Rover (VIPER), In-Situ Resource Utilization Pilot Excavator (IPEX) . Various upgrades have been made to the extreme cold environment test chamber for the use of cryobotic research. These upgrades greatly increased the autonomous capabilities of the test set up by providing redundancies in the hardware and software. The redundancies were primarily added to protect the integrity of the cryohead. A new strapping and insulation method was performed to create the thermal conductive path from the actuators to the cryohead. The software was upgraded to include temperature setpoint control, further increasing the autonomous capabilities of the test. This paper goes into detail regarding the upgrades made to the extreme cold environment test chamber, as well as highlights the results from a COLDArm acceptance test.

Casey J. Clark↗
NASA NTRS1984

High temperature acoustic levitator

A system is described for acoustically levitating an object within a portion of a chamber that is heated to a high temperature, while a driver at the opposite end of the chamber is maintained at a relatively low temperature. The cold end of the chamber is constructed so it can be telescoped to vary the length (L sub 1) of the cold end portion and therefore of the entire chamber, so that the chamber remains resonant to a normal mode frequency, and so that the pressure at the hot end of the chamber is maximized. The precise length of the chamber at any given time, is maintained at an optimum resonant length by a feedback loop. The feedback loop includes an acoustic pressure sensor at the hot end of the chamber, which delivers its output to a control circuit which controls a motor that varies the length (L) of the chamber to a level where the sensed acoustic pressure is a maximum.

Barmatz, M. B.↗
NASA NTRS2004

Sliding Mode Control of a Thermal Mixing Process

In this paper we consider the robust control of a thermal mixer using multivariable Sliding Mode Control (SMC). The mixer consists of a mixing chamber, hot and cold fluid valves, and an exit valve. The commanded positions of the three valves are the available control inputs, while the controlled variables are total mass flow rate, chamber pressure and the density of the mixture inside the chamber. Unsteady thermodynamics and linear valve models are used in deriving a 5th order nonlinear system with three inputs and three outputs, An SMC controller is designed to achieve robust output tracking in the presence of unknown energy losses between the chamber and the environment. The usefulness of the technique is illustrated with a simulation.

Richter, Hanz↗
NASA NTRS2004

Ignition Characterization Tests of the LOX/Ethanol Propellant Combination

A series of contracts have been issued by the Marshall Space Flight Center (MSFC) of the National Aeronautics and Space Administration (NASA) to explore candidate technologies considered to be important for the Next Generation Launch Technology (NGLT) effort. One aspect of the NGLT effort is to explore the potential of incorporating non-toxic propellants for Reaction Control Subsystems (RCS). Contract NAS8-01109 has been issued to Aerojet to develop a dual thrust Reaction Control Engine (RCE) that utilizes liquid oxygen and ethanol as the propellants. The dual thrust RCE incorporates a primary thrust level of 870 lbf, and a vernier thrust level of 10 - 30 lbf. Aerojet has designed and tested a workhorse LOX igniter to determine LOX/Ethanol ignition characteristics as part of a risk mitigation effort for the dual thrust RCE design. The objective of the ignition testing was to demonstrate successfid ignition from GOX to LOX, encompassing potential two-phase flow conditions. The workhorse igniter was designed to accommodate the full LOX design flowrate, as well as a reduced GOX flowrate. It was reasoned that the initial LOX flow through the igniter would flash to GOX due to the inherent heat stored in the hardware, causing a reduced oxygen flowrate because of a choked, or sonic, flow condition through the injection elements. As LOX flow continued, the inherent heat of the test hardware would be removed and the hardware would chill-in, with the injected oxygen flow transitioning from cold GOX through two-phase flow to subcooled LOX. Pressure and temperature instrumentation permitted oxygen state points to be determined, and gas-side igniter chamber thermocouples provided chamber thermal profile characteristics. The cold flow chamber pressure (P(sub c)) for each test was determined and coupled with the igniter chamber diameter (D(sub c)) to calculate the characteristic quench parameter (P(sub c) x D(sub c)), which was plotted as a function of core mixture ratio, MR(sub c). Ignition limits were determined over a broad range of valve inlet conditions, and ignition was demonstrated with oxygen inlet conditions that ranged from subcooled 173 R LOX to 480 R GQX. Once ignited at cold GOX conditions, combustion was continuous as the hardware chilled in and the core mixture ratio transitioned from values near 1.0 to over 12.5.

Popp, Christopher G.↗
NASA NTRS2006

Ignition Characterization Test Results for the LO2/Ethanol Propellant Combination

A series of contracts were issued by the Marshall Space Flight Center (MSFC) of the National Aeronautics and Space Administration (NASA) under the auspices of the Exploration Systems Mission Directorate to develop and expand the maturity of candidate technologies considered to be important for future space exploration. One such technology was to determine the viability of incorporating non-toxic propellants for Reaction Control Subsystems (RCS). Contract NAS8-01109 was issued to Aerojet to develop a dual thrust Reaction Control Engine (RCE) that utilized liquid oxygen and ethanol as the propellants. The dual thrust RCE incorporated a primary thrust level of 870 lbf, and a vernier thrust level of 10 - 30 lbf. The preferred RCS approach for the dual thrust RCE was to utilize pressure-fed liquid oxygen (LOX) and ethanol propellants; however, previous dual thrust feasibility testing incorporated GOX/Ethanol igniters as opposed to LOX/Ethanol igniters in the design. GOX/Ethanol was easier to ignite, but this combination had system design implications of providing GOX for the igniters. A LOX/Ethanol igniter was desired; however, extensive LOX/Ethanol ignition data over the anticipated operating range for the dual thrust RCE did not exist. Therefore, Aerojet designed and tested a workhorse LOX igniter to determine LOX/Ethanol ignition characteristics as part of a risk mitigation effort for the dual thrust RCE design. LOX, encompassing potential two-phase flow conditions anticipated being present in real mission applications. A workhorse igniter was designed to accommodate the hll LOX design flowrate, as well as a reduced GOX flowrate. It was reasoned that the initial LOX flow through the igniter would flash to GOX due to the latent heat stored in the hardware, causing a reduced oxygen flowrate because of a choked, or sonic, flow condition through the injection elements. As LOX flow continued, the hardware would chill-in, with the injected oxygen flow transitioning from cold GOX through two-phase flow to subcooled LOX. permitted oxygen state points to be determined in the igniter oxidizer manifold, and gas-side igniter chamber thermocouples provided chamber thermal profile characteristics. The cold flow chamber pressure (P(sub c)) for each test was determined and coupled with the igniter chamber diameter (D(sub c)) to calculate the characteristic quench parameter (P(sub c) x D(sub c)), which was plotted as a function of core mixture ratio, m. Ignition limits were determined over a broad range of valve inlet conditions, and ignition was demonstrated with oxygen inlet conditions that ranged from subcooled 210 R LOX to 486 R GOX. Once ignited at cold GOX conditions, combustion was continuous as the hardware chilled in and the core mixture ratio transitioned from values near 1.0 to over 12.5. Pulsing is required in typical RCS engines; therefore, the workhorse igniter was pulse tested to verify the ability to provide the required ignition for a pulsing RCE. The minimum electrical pulse width (EPW) of the dual thrust RCE was 0.080 seconds.

Popp, Christopher G.↗
NASA NTRS2007

Ignition Characterization Test Results for the LO2/Ethanol Propellant Combination

A series of contracts were issued by the Marshall Space Flight Center (MSFC) of the National Aeronautics and Space Administration (NASA) un der the auspices of the Exploration Systems Mission Directorate to de velop and expand the maturity of candidate technologies considered to be important for future space exploration. One such technology was to determine the viability of incorporating non-toxic propellants for R eaction Control Subsystems (RCS). Contract NAS8-01109 was issued to A erojet to develop a dual thrust Reaction Control Engine (RCE) that ut ilized liquid oxygen and ethanol as the propellants. The dual thrust RCE incorporated a primary thrust level of 870 lbf, and a vernier thru st level of 10 - 30 lbf. The preferred RCS approach for the dual thru st RCE was to utilize pressure-fed liquid oxygen (LOX) and ethanol pr opellants; however, previous dual thrust feasibility testing incorporated GOX/Ethanol igniters as opposed to LOX/Ethanol igniters in the de sign. GOX/Ethanol was easier to ignite, but this combination had syst em design implications of providing GOX for the igniters. A LOX/Ethan ol igniter was desired; however, extensive LOX/Ethanol ignition data over the anticipated operating range for the dual thrust RCE did not e xist. Therefore, Aerojet designed and tested a workhorse LOX igniter to determine LOX/Ethanol ignition characteristics as part of a risk m itigation effort for the dual thrust RCE design. The objective of the ignition testing was to demonstrate successful ignition from GOX to LOX, encompassing potential two-phase flow conditions anticipated being present in real mission applications. A workhorse igniter was desig ned to accommodate the full LOX design flowrate, as well as a reduced GOX flowrate. It was reasoned that the initial LOX flow through the igniter would flash to GOX due to the latent heat stored in the hardwa re, causing a reduced oxygen flowrate because of a choked, or sonic, flow condition through the injection elements. As LOX flow continued, the hardware would chill-in, with the injected oxygen flow transitioning from cold GOX through two'phase flow to subcooled LOX. The Workh orse igniter was well instrumented: Pressure and temperature instrumentation permitted oxygen state points to be determined in the igniter oxidizer manifold, and gas-side igniter chamber thermocouples provide d chamber thermal profile characteristics. The cold flow chamber pres sure (Pc) for each test was determined and coupled with the igniter chamber diameter (De) to calculate the characteristic quench parameter (Pc x Dc), which was plotted as a function of core mixture ratio, MRc . Ignition limits were determined over a broad range of valve inlet conditions, and ignition was demonstrated with oxygen inlet conditions that ranged from subcooled 210 deg R LOX to 486 deg R GOX. Once ign ited at cold GOX conditions, combustion was continuous as the hardwar e chilled in and the core mixture ratio transitioned from values near 1.0 to over 12.5. Pulsing is required in typical RCS engines; therefore, the workhorse igniter was pulse tested to verify the ability to pr ovide the required ignition for a pulsing RCE. The minimum electrical pulse width (EPW) of the dual thrust RCE was 0.080 seconds. Igniter pulse tests were performed at three conditions: (1) an EPW of 0.080 se conds at 25% duty cycle for 400 pulses; (2) an EPW of 0.160 seconds a nd a 5% duty cycle for 124 pulses; (3) an EPW of 0.160 seconds and a 50% duty cycle for 380 pulses. Successful ignition of LOX/Ethanol was demonstrated over a broad range of valve inlet conditions, with the empirically determined LOX/Ethanol ignition limits extending the previous database established for GOX/Ethanol ignition limits. Although th e observed chill-in characteristics of the hardware varied significan tly with flowrate, ignition was readily achieved. Combustion was marg inal at extremely fuel-rich conditions, and it fluctuated as the oxygen passed rough the twophase flow regime during the period of hardware chill-in. Pulse testing showed good repeatability with 100 percent r e-ignition for all pulses. Certain pulse-to-pulse repeatability requirements for actual RCS operation may necessitate establishment of mini mum oxygen flow rates and engine thrust levels for satisfactory engin e performance.

Robinson, Philip J.↗
NASA NTRSDate not supplied

An Experimental Study on Low Pressure Frost Formation for Lunar Polar Water Capture

In-situ resource utilization (ISRU) is a vital component of NASA’s mission to the Moon and beyond, as the extraction of resources from the environment can reduce payload weight and the frequency of resupply missions. Since the discovery of water in the regolith of the permanently shadowed regions (PSR) of the Moon, its extraction and transport has become an area of increasing interest for NASA. One proposed method is heating icy regolith to free the water before desublimating and transporting it as ice in a tanker. However, little is known regarding the dynamics of frost growth at low pressures, and an understanding of the heat transfer process is required to properly size the heat exchanger for the tanker. To investigate this phenomenon, a cold plate was placed in a vacuum chamber with water vapor directed at its surface. The chamber pressure (300 and 500 Pa) and cold plate temperature (–18 to –5 °C) were kept below the triple point and varied throughout the experiment to examine their impact on frost layer heat transfer. As water flowed into the chamber and deposited on the cold plate surface, the heat flux and temperature were measured along with the frost layer thickness and/or mass. Density calculated at the conclusion of each test suggests the frost layer is denser than that of frost grown at atmospheric pressure. The results demonstrate unexpected density and heat transfer characteristics and require development of a new model of frost growth for lunar conditions.

In situ Resource Utiliztion↗
NASA NTRSDate not supplied

An Experimental Study on Low Pressure Frost Formation for Lunar Polar Water Capture

In-situ resource utilization (ISRU) is a vital component of NASA’s mission to the Moon and beyond, as the extraction of resources from the environment can reduce payload weight and the frequency of resupply missions. Since the discovery of water in the regolith of the permanently shadowed regions (PSR) of the Moon, its extraction and transport has become an area of increasing interest for NASA. One proposed method is heating icy regolith to free the water before de-sublimating and transporting it as ice in a tanker. However, little is known regarding the dynamics of frost growth at low pressures, and an understanding of the heat transfer process is required to properly size the heat exchanger for the tanker. To investigate this phenomenon, a cold plate was placed in a vacuum chamber with water vapor directed at its surface. The chamber pressure (300 and 500 Pa) and cold plate temperature (–18 to –5 °C) were kept below the triple point and varied throughout the experiment to examine their impact on frost layer heat transfer. As water flowed into the chamber and deposited on the cold plate surface, the heat flux and temperature were measured along with the frost layer thickness and/or mass. Density calculated at the conclusion of each test suggests the frost layer is denser than that of frost grown at atmospheric pressure. The results demonstrate unexpected density and heat transfer characteristics and require development of a new model of frost growth for lunar conditions

Lunar polar↗
NASA NTRS2023

Regenerable Liquid Desiccants for High Efficiency Humidity Control in Microgravity

The NASA X-Hab project aims to design, manufacture, test, and prove functionality of an air humidity control subsystem to dehumidify and re-humidify air from a space cabin. The method of regulating cabin air humidity utilizes vortex phase separation, which uses an ionic liquid (IL) desiccant for air-water phase separation. This subsystem is intended to be integrated with a CO 2 removal system requiring de-humidified air to operate efficiently. The air humidity control subsystem is composed of a cold-desiccant or cold-side Vortex Phase Separator (VPS) that dehumidifies the cabin air. The dehumidified air exits the cold-side (CS) VPS chamber to flow into the CO 2 removal module, and the cold, water-laden liquid desiccant flows to a regenerative heat exchanger. From the heat exchanger the liquid desiccant continues to the heater. After heating, the desiccant flows into a hot-desiccant or hot-side (HS) VPS as the dehumidified air from CO 2 removal module enters through the HS VPS air inlet. The rehumidified air exits the HS VPS into the space cabin. One pump installed at the liquid exit of the CS VPS and one pump installed at the liquid exit of the HS VPS transport the fluid through the system. To evaluate the system’s efficiency and effectiveness, the temperature, pressure, flow rate and relative humidity are read, recorded, and analyzed at critical points along the module.

Gerardo Castro↗