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NASA NTRS2012

Thermal Performance Testing of EMU and OSS Liquid Cooling Garments

A test was conducted to evaluate three factors influencing the thermal performance of liquid cooling garments (LCG): (1) the comparable thermal performance of an Oceaneering developed engineering evaluation unit (EEU) prototype LDG, (2) the effect of the thermal comfort undergarment (TCU), and (3) the performance of a torso or upper body only LCG configuration. To evaluate the thermal performance of each configuration a metabolic test was conducted, utilizing suited subjects to generate the metabolic heat. For this study three (3) test subjects of similar health and weight produced a metabolic load on the LDG configuration by either resting (300-600 BTU/hr), walking at a slow pace (1200 BRU/hr), and walking at a brisk pace (2200 BTU/hr), as outlined in Figure 1, the metabolic profile. During the test, oxygen consumption, heart rate, relative humidity, air flow, inlet and outlet air pressure, inlet and outlet air temperature, delta air temperature, water flow (100 lb/hr), inlet water temperature (64 F), delta water temperature, water pressure, core body temperature, skin temperature, and sweat loss data was recorded. Four different test configurations were tested, with one configuration tested twice, as outlined in Table 1. The test was conducted with the suit subjects wearing the Demonstrator Suit, pressurized to vent pressure (approximately 0.5 psig). The demonstrator suit has an integrated ventilation duct system and was used to create a relevant environment with a captured ventilation return, an integrated vent tree, and thermal insulation from the environment.

Rhodes, Richard↗

NASA NTRS2008

Hail Disrometer Array for Launch Systems Support

Prior to launch, the space shuttle might be described as a very large thermos bottle containing substantial quantities of cryogenic fuels. Because thermal insulation is a critical design requirement, the external wall of the launch vehicle fuel tank is covered with an insulating foam layer. This foam is fragile and can be damaged by very minor impacts, such as that from small- to medium-size hail, which may go unnoticed. In May 1999, hail damage to the top of the External Tank (ET) of STS-96 required a rollback from the launch pad to the Vehicle Assembly Building (VAB) for repair of the insulating foam. Because of the potential for hail damage to the ET while exposed to the weather, a vigilant hail sentry system using impact transducers was developed as a hail damage warning system and to record and quantify hail events. The Kennedy Space Center (KSC) Hail Monitor System, a joint effort of the NASA and University Affiliated Spaceport Technology Development Contract (USTDC) Physics Labs, was first deployed for operational testing in the fall of 2006. Volunteers from the Community Collaborative Rain. Hail, and Snow Network (CoCoRaHS) in conjunction with Colorado State University were and continue to be active in testing duplicate hail monitor systems at sites in the hail prone high plains of Colorado. The KSC Hail Monitor System (HMS), consisting of three stations positioned approximately 500 ft from the launch pad and forming an approximate equilateral triangle (see Figure 1), was deployed to Pad 39B for support of STS-115. Two months later, the HMS was deployed to Pad 39A for support of STS-116. During support of STS-117 in late February 2007, an unusual hail event occurred in the immediate vicinity of the exposed space shuttle and launch pad. Hail data of this event was collected by the HMS and analyzed. Support of STS-118 revealed another important application of the hail monitor system. Ground Instrumentation personnel check the hail monitors daily when a vehicle is on the launch pad, with special attention after any storm suspected of containing hail. If no hail is recorded by the HMS, the vehicle and pad inspection team has no need to conduct a thorough inspection of the vehicle immediately following a storm. On the afternoon of July 13, 2007, hail on the ground was reported by observers at the VAB, about three miles west of Pad 39A, as well as at several other locations around Kennedy Space Center. The HMS showed no impact detections, indicating that the shuttle had not been damaged by any of the numerous hail events which occurred that day.

Lane, John E.↗

NASA NTRS2012

Multiple Asteroid Systems: Dimensions and Thermal Properties from Spitzer Space Telescope and Ground-based Observations

We collected mid-IR spectra from 5.2 to 38 microns using the Spitzer Space Telescope Infrared Spectrograph of 28 asteroids representative of all established types of binary groups. Photometric light curves were also obtained for 14 of them during the Spitzer observations to provide the context of the observations and reliable estimates of their absolute magnitudes. The extracted mid-IR spectra were analyzed using a modified standard thermal model (STM) and a thermophysical model (TPM) that takes into account the shape and geometry of the large primary at the time of the Spitzer observation. We derived a reliable estimate of the size, albedo, and beaming factor for each of these asteroids, representing three main taxonomic groups: C, S, and X. For large (volume-equivalent system diameter Deq > 130 km) binary asteroids, the TPM analysis indicates a low thermal inertia (Lambda < or = approx.100 J/1/2 s/K/sq m2) and their emissivity spectra display strong mineral features, implying that they are covered with a thick layer of thermally insulating regolith. The smaller (surface-equivalent system diameter Deff < 17 km) asteroids also show some emission lines of minerals, but they are significantly weaker, consistent with regoliths with coarser grains, than those of the large binary asteroids. The average bulk densities of these multiple asteroids vary from 0.7-1.7 g/cu cm (P-, C-type) to approx. 2 g/cu cm (S-type). The highest density is estimated for the M-type (22) Kalliope (3.2 +/- 0.9 g/cu cm). The spectral energy distributions (SEDs) and emissivity spectra, made available as a supplement document, could help to constrain the surface compositions of these asteroids.

Marchis, F.↗

NASA NTRS2013

Electrodynamic Dust Shield Demonstrator

The objective of the project was to design and manufacture a device to demonstrate a new technology developed by NASA's Electrostatics and Surface Physics Laboratory. The technology itself is a system which uses magnetic principles to remove regolith dust from its surface. This project was to create an enclosure that will be used to demonstrate the effectiveness of the invention to The Office of the Chief Technologist. ONE of the most important challenges of space exploration is actually caused by something very small and seemingly insignificant. Dust in space, most notably on the moon and Mars, has caused many unforeseen issues. Dirt and dust on Earth, while a nuisance, can be easily cleaned and kept at bay. However, there is considerably less weathering and erosion in space. As a result, the microscopic particles are extremely rough and abrasive. They are also electrostatically charged, so they cling to everything they make contact with. This was first noted to be a major problem during the Apollo missions. Dust would stick to the spacesuits, and could not be wiped off as predicted. Dust was brought back into the spacecraft, and was even inhaled by astronauts. This is a major health hazard. Atmospheric storms and other events can also cause dust to coat surfaces of spacecraft. This can cause abrasive damage to the craft. The coating can also reduce the effectiveness of thermal insulation and solar panels.' A group of engineers at Kennedy Space Center's Electrostatics and Surface Physics Laboratory have developed a new technology, called the Electrodynamic Dust Shield, to help alleviate these problems. It is based off of the electric curtain concept developed at NASA in 1967. "The EDS is an active dust mitigation technology that uses traveling electric fields to transport electrostatically charged dust particles along surfaces. To generate the traveling electric fields, the EDS consists of a multilayer dielectric coating with an embedded thin electrode grid running a multiphase low frequency AC signal. Electrostatically charged particles, such as those encountered on the moon, Mars, or an asteroid, are carried along by the traveling field due to the action of Coulomb and dielectrophoretic forces."2 The technical details have been described in a separate article. This document details the design and construction process of a small demonstration unit. Once finished, this device will go to the Office of the ChiefTechnologist at NASA headquarters, where it will be used to familiarize the public with the technology. 1 NASA KSC FO Intern, Prototype Development Laboratory, Kennedy Space Center, University of Central Florida Kennedy Space

Stankie, Charles G.↗

NASA NTRS2013

Cryogenics Research and Engineering Experience

Energy efficient storage, transfer and use of cryogens and cryogenic propellants on Earth and in space have a direct impact on NASA, government and commercial programs. Research and development on thermal insulation, propellant servicing, cryogenic components, material properties and sensing technologies provides industry, government and research institutions with the cross-cutting technologies to manage low-temperature applications. Under the direction of the Cryogenic Testing Lab at Kennedy Space Center, the work experience acquired allowed me to perform research, testing, design and analysis of current and future cryogenic technologies to be applied in several projects.

Toro Medina, Jaime A.↗

NASA NTRS2013

Mars Sample Return Landed with Red Dragon

A Mars Sample Return (MSR) mission is the highest priority science mission for the next decade as recommended by the recent Decadal Survey of Planetary Science. However, an affordable program to carry this out has not been defined. This paper describes a study that examined use of emerging commercial capabilities to land the sample return elements, with the goal of reducing mission cost. A team at NASA Ames examined the feasibility of the following scenario for MSR: A Falcon Heavy launcher injects a SpaceX Dragon crew capsule and trunk onto a Trans Mars Injection trajectory. The capsule is modified to carry all the hardware needed to return samples collected on Mars including a Mars Ascent Vehicle (MAV), an Earth Return Vehicle (ERV) and Sample Collection and Storage hardware. The Dragon descends to land on the surface of Mars using SuperSonic Retro Propulsion (SSRP) as described by Braun and Manning [IEEEAC paper 0076, 2005]. Samples are acquired and deliverd to the MAV by a prelanded asset, possibly the proposed 2020 rover. After samples are obtained and stored in the ERV, the MAV launches the sample-containing ERV from the surface of Mars. We examined cases where the ERV is delivered to either low Mars orbit (LMO), C3 = 0 (Mars escape), or an intermediate energy state. The ERV then provides the rest of the energy (delta V) required to perform trans-Earth injection (TEI), cruise, and insertion into a Moon-trailing Earth Orbit (MTEO). A later mission, possibly a crewed Dragon launched by a Falcon Heavy (not part of the current study) retrieves the sample container, packages the sample, and performs a controlled Earth re-entry to prevent Mars materials from accidentally contaminating Earth. The key analysis methods used in the study employed a set of parametric mass estimating relationships (MERs) and standard aerospace analysis software codes modified for the MAV class of launch vehicle to determine the range of performance parameters that produced converged spacecraft designs capable of meeting mission requirements. Subsystems modeled in this study included structures, power system, propulsion system, nose fairing, thermal insulation, actuation devices, and GN&C. Best practice application of loads and design margins for all resources were used. Both storable and cryogenic propellant systems were examined. The landed mass and lander capsule size provide boundary conditions for the MAV design and packaging. We estimated the maximum mass the Dragon capsule is capable of landing. This and the volume capability to store the MAV was deduced from publically available data from SpaceX as well as our own engineering and aerodynamic estimates. Minimum gross-liftoff mass (GLOM) for the MAV were obtained for configurations that used pump-fed storable bi-propellant rocket engines for both the MAV and the ERV stage. The GLOM required fits within our internal estimate of the mass that Dragon can land at low elevation/optimal seasons on Mars. Based on the analysis, we show that a single Mars launch sample return mission is feasible using current commercial capabilities to deliver the return spacecraft assets.

Mars Ascent Vehicle (MAV)↗

NASA NTRS2014

Recent Ground Hold and Rapid Depressurization Testing of Multilayer Systems

In the development of flight insulation systems for large cryogenic orbital storage (spray on foam and multilayer insulation), testing need include all environments that are experienced during flight. While large efforts have been expended on studying, bounding, and modeling the orbital performance of the insulation systems, little effort has been expended on the ground hold and ascent phases of a mission. Historical cryogenic in-space systems that have flown have been able to ignore these phases of flight due to the insulation system being within a vacuum jacket. In the development phase of the Nuclear Mars Vehicle and the Shuttle Nuclear Vehicle, several insulation systems were evaluated for the full mission cycle. Since that time there had been minimal work on these phases of flight until the Constellation program began investigating cryogenic service modules and long duration upper stages. With the inception of the Cryogenic Propellant Storage and Transfer Technology Demonstration Mission, a specific need was seen for the data and as such, several tests were added to the Cryogenic Boil-off Reduction System liquid hydrogen test matrix to provide more data on a insulation system. Testing was attempted with both gaseous nitrogen (GN2) and gaseous helium (GHe) backfills. The initial tests with nitrogen backfill were not successfully completed due to nitrogen liquefaction and solidification preventing the rapid pumpdown of the vacuum chamber. Subsequent helium backfill tests were successful and showed minimal degradation. The results are compared to the historical data.

Space Environment Simulation↗

NASA NTRS2014

Design and Development of Aerogel-Based Antennas for Aerospace Applications: A Final Report to the NARI Seedling

As highly porous solids possessing low density and low dielectric permittivity combined with good mechanical properties, polyimide (PI) aerogels offer great promise as an enabling technology for lightweight aircraft antenna systems. While they have been aggressively explored for thermal insulation, barely any effort has been made to leverage these materials for antennas or other applications that take advantage of their aforementioned attributes. In Phase I of the NARI Seedling Project, we fabricated PI aerogels with properties tailored to enable new antenna concepts with performance characteristics (wide bandwidth and high gain) and material properties (low density, environmental stability, and robustness) superior to the state of practice (SOP). We characterized electromagnetic properties, including permittivity, reflectivity, and propagation losses for the aerogels. Simple, prototype planar printed circuit patch antennas from down-selected aerogel formulations were fabricated by molding the aerogels to net shapes and by gold-metalizing the pattern onto the templates via electron beam evaporation in a clean room environment. These aerogel based antennas were benchmarked against current antenna SOP, and exhibited both broader bandwidth and comparable or higher gain performance at appreciably lower mass. Phase II focused on the success of the Phase I results pushing the PI aerogel based antenna technology further by exploring alternative antenna design (i.e., slot coupled antennas) and by examining other techniques for fabricating the antennas including ink jet printing with the goal of optimizing antenna performance and simplifying production. We also examined new aerogel formulations with better moisture and solvent resistance to survive processing conditions. In addition, we investigated more complex antenna designs including passive phased arrays such as 2x4 and 4x8 element arrays to assess the scalability of the aerogel antenna concept. Furthermore, we explored the possibility of developing these arrays in thin, flexible form to make conformable antennas.

insulation↗

NASA NTRS2013

An Ultra-Wideband, Microwave Radar for Measuring Snow Thickness on Sea Ice and Mapping Near-Surface Internal Layers in Polar Firn

Sea ice is generally covered with snow, which can vary in thickness from a few centimeters to >1 m. Snow cover acts as a thermal insulator modulating the heat exchange between the ocean and the atmosphere, and it impacts sea-ice growth rates and overall thickness, a key indicator of climate change in polar regions. Snow depth is required to estimate sea-ice thickness using freeboard measurements made with satellite altimeters. The snow cover also acts as a mechanical load that depresses ice freeboard (snow and ice above sea level). Freeboard depression can result in flooding of the snow/ice interface and the formation of a thick slush layer, particularly in the Antarctic sea-ice cover. The Center for Remote Sensing of Ice Sheets (CReSIS) has developed an ultra-wideband, microwave radar capable of operation on long-endurance aircraft to characterize the thickness of snow over sea ice. The low-power, 100mW signal is swept from 2 to 8GHz allowing the air/snow and snow/ ice interfaces to be mapped with 5 c range resolution in snow; this is an improvement over the original system that worked from 2 to 6.5 GHz. From 2009 to 2012, CReSIS successfully operated the radar on the NASA P-3B and DC-8 aircraft to collect data on snow-covered sea ice in the Arctic and Antarctic for NASA Operation IceBridge. The radar was found capable of snow depth retrievals ranging from 10cm to >1 m. We also demonstrated that this radar can be used to map near-surface internal layers in polar firn with fine range resolution. Here we describe the instrument design, characteristics and performance of the radar.

Snow Cover↗

NASA NTRS2012

Bright and Dark Polar Deposits on Mercury: Evidence for Surface Volatiles

Measurements of surface reflectance of permanently shadowed areas near Mercury's north pole reveal regions of anomalously dark and bright deposits at 1064-nanometer wavelength. These reflectance anomalies are concentrated on poleward-facing slopes and are spatially collocated with areas of high radar backscatter postulated to be the result of near-surface water ice. Correlation of observed reflectance with modeled temperatures indicates that the optically bright regions are consistent with surface water ice, whereas dark regions are consistent with a surface layer of complex organic material that likely overlies buried ice and provides thermal insulation. Impacts of comets or volatile-rich asteroids could have provided both dark and bright deposits.

Polar Deposits↗

NASA NTRS2015

Electric Motors for Non-Cryogenic Hybrid Electric and Turboelectric Propulsion

NASA Glenn Research Center is investigating hybrid electric and turboelectric propulsion concepts for future aircraft to reduce fuel burn, emissions, and noise. Systems studies show that the weight and efficiency of the electric system components need to be improved for this concept to be feasible. However, advances in motor component materials such as soft magnetic materials, hard magnetic materials, conductors, thermal insulation, and structural materials are expected in the coming years, and should improve motor performance. This study investigates several motor types for a one megawatt application, and projects the motor performance benefits of new component materials that might be available in the coming decades.

Hybrid propulsion↗

NASA NTRS2015

Electric Motor Considerations for Non-Cryogenic Hybrid Electric and Turboelectric Propulsion

Hybrid propulsion↗

NASA NTRS1961

Lunar Brightness Variations with Phase at 4.3-MM Wave Length

The lunar radio observations and interpretations of Piddington and Minnett (1949) and Gibson (1958) show that the lunar brightness variation with phase at millimeter wave lengths can be used to determine the physical properties of the lunar surface. They found that the millimeter-wave brightness lagged the optical phase, and their interpretation was that the millimeter radiation originates below a surface layer that is a very good thermal insulator. The thickness of this layer could not be determined from observations at one frequency. Observations at different frequencies give different results because of the wave-length dependence of the radio absorption by the surface material. The attenuation in the material increases with decreasing wave length, and therefore it is possible, in principle, to determine surface layer thickness from radio observations at several wave lengths. For this reason, observations of lunar radiation were started at the Naval Research Laboratory at a wave length of 4.3 mm. This is half the wave length used by Gibson (1958) in his earlier studies. The radio telescope used for these observations has been described in detail in a previous publication (Coates 1958). The antenna is a parabolic reflector 10 feet in diameter, and it has a beam width of 6.7 minutes of arc at the wave length of 4.3 mm. This is about one-fifth the diameter of the moon. The receiver was a Dicke-type radiometer.

Coates, Robert J.↗

NASA NTRS2015

Test and Analysis Correlation for a Y-Joint Specimen for a Composite Cryotank

The Composite Cryotank Technology Demonstration (CCTD) project under NASA's Game Changing Development Program (GCDP) developed space technologies using advanced composite materials. Under CCTD, NASA funded the Boeing Company to design and test a number of element-level joint specimens as a precursor to a 2.4-m diameter composite cryotank. Preliminary analyses indicated that the y-joint in the cryotank had low margins of safety; hence the y-joint was considered to be a critical design region. The y-joint design includes a softening strip wedge to reduce localized shear stresses at the skirt/dome interface. In this paper, NASA-developed analytical models will be correlated with the experimental results of a series of positive-peel y-joint specimens from Boeing tests. Initial analytical models over-predicted the experimental strain gage readings in the far-field region by approximately 10%. The over-prediction was attributed to uncertainty in the elastic properties of the laminate and a mismatch between the thermal expansion of the strain gages and the laminate. The elastic properties of the analytical model were adjusted to account for the strain gage differences. The experimental strain gages also indicated a large non-linear effect in the softening strip region that was not predicted by the analytical model. This non-linear effect was attributed to delamination initiating in the softening strip region at below 20% of the failure load for the specimen. Because the specimen was contained in a thermally insulated box during cryogenic testing to failure, delamination initiation and progression was not visualized during the test. Several possible failure initiation locations were investigated, and a most likely failure scenario was determined that correlated well with the experimental data. The most likely failure scenario corresponded to damage initiating in the softening strip and delamination extending to the grips at final failure.

Mason, Brian H.↗

NASA NTRS2016

Liquefaction and Storage of In-Situ Oxygen on the Surface of Mars

ISRU is currently base-lined for the production of oxygen on the Martian surface in the Evolvable Mars Campaign Over 50 of return vehicle mass is oxygen for propulsion. There are two key cryogenic fluid-thermal technologies that need to be investigated to enable these architectures. High lift refrigeration systems. Thermal Insulation systems, either lightweight vacuum jackets of soft vacuum insulation systems.

Liquefaction↗

NASA NTRS2017

Erosion Results of the MISSE 8 Polymers Experiment After 2 Years of Space Exposure on the International Space Station

Polymers and other oxidizable materials on the exterior of spacecraft in the low Earth orbit (LEO) space environment can be eroded due to reaction with atomic oxygen (AO). Therefore, in order to design durable spacecraft, it is important to know the LEO AO erosion yield (E(sub y), volume loss per incident oxygen atom) of materials susceptible to AO reaction. A spaceflight experiment, called the Polymers Experiment, which contained 42 samples, was developed to determine the effect of solar exposure on the AO E(sub y) of fluoropolymers flown in ram, wake, or zenith orientations. The Polymers Experiment was exposed to the LEO space environment on the exterior of the International Space Station (ISS) as part of the Materials International Space Station Experiment 8 (MISSE 8) mission. The MISSE 8 mission included samples flown in a zenith/nadir orientation for 2.14 years in the MISSE 8 Passive Experiment Container (PEC), and samples flown in a ram/wake orientation for 2.0 years in the Optical Reflector Materials Experiment-III (ORMatEIII) tray. The experiment included Kapton H (Registered Trademark) witness samples for AO fluence determination in each orientation. This paper provides an overview of the MISSE 8 mission, a description of the flight experiment with details on the polymers flown, the characterization techniques used, the AO fluence for each exposure orientation, and the LEO E(sub y) results. The E(sub y) of Teflon fluorinated ethylene propylene (FEP) samples flown in ram, wake, and zenith orientations have been compared, and the E(sub y) was found to be highly dependent on orientation and therefore environmental exposure. The FEP E(sub y) was found to directly correlate with the solar exposure/AO fluence ratio showing the effect of solar radiation and/or heating due to solar exposure on FEP erosion. In addition, back-surface carbon painted FEP (C-FEP) flown in the zenith orientation had a significantly higher E(sub y) than clear FEP or Al-FEP further indicating that heating has a significant impact on the erosion of FEP. This experiment provides valuable LEO flight data on the erosion of Teflon FEP, a commonly used spacecraft thermal insulation.

Teflon↗

NASA NTRS2017

Aerogel / Polymer Composite Materials

The invention provides new composite materials containing aerogels blended with thermoplastic polymer materials at a weight ratio of aerogel to thermoplastic polymer of less than 20:100. The composite materials have improved thermal insulation ability. The composite materials also have better flexibility and less brittleness at low temperatures than the parent thermoplastic polymer materials.

Williams, Martha K.↗

NASA NTRS2019

Review of Habitable Softgoods Inflatable Design, Analysis, Testing, and Potential Space Applications

Inflatable space structures have the potential to significantly reduce the required launch volume of large crewed pressure vessels for space exploration missions. Mass savings can also be achieved via the use of high specific strength softgoods materials, and the reduced design penalty from launching the structure in a densely packaged state. Inflatable softgoods structures have been investigated since the late 1950's, and several major development programs at NASA and in industry have helped advance the state-of-the-art in this technology area. This paper discusses the design, analysis, structural testing, and potential applications for inflatable softgoods structures. In particular, this paper will discuss the design of the multi-layer softgoods shell (inner layer, bladder, structural restraint layer, micrometeoroid orbital debris protection layers, thermal insulation layers, and atomic oxygen layer (for low earth orbit) and the results of material and module-level testing that has been conducted over the past two decades at NASA. Finally, the current utilization of expandable spacecraft structures is discussed, as well as potential future applications including airlocks and habitats on the Lunar Orbital Platform-Gateway, and the surface of the Moon and Mars.

Valle, Gerard↗