Search NASA⌕ Search

SEARCH · Search NASA

Results for “Multilayer shielding”

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 37 records · Page 2
All sourcesNASA NTRSDOE OSTI
NASA NTRS2021

A Study of Layered Structural Configurations as Thermal and Impact Shielding of Lunar Habitats

The focus of this study is the analysis of a proposed multilayer material configuration as lunar habitat shielding so as to assess its ability to resist temperature extremes and impact. In order to resist this environmental hazard, the analysis will compare various configurations of potential shielding layers using both in-situ resource utilization (ISRU) and Earth-based materials. The technology for implementation of in-situ resources in construction is developing, though not yet able to produce a functional lunar habitat. As the lunar infrastructure is developed, ISRU techniques for manufacturing and construction may become more viable. Therefore, both types of materials are considered due to potential for application towards both near-term and long-term commencements of habitat construction. The thermal analysis in this paper utilizes the nonlinear heat equation to produce a depth-dependent temperature gradient from the habitat exterior to the interior that changes continuously with time during the lunar diurnal cycle. Static loading conditions meant to simulate the impact force of a 77.9 g meteorite were used to analyze the ability of a proposed multilayer shielding configuration to resist impact. A thermal stress analysis was performed to observe the effects of heat transfer through the proposed multilayer configuration. The results of this study should be useful in determining the potential for implementation of efficient shielding layers for the initial lunar base.

Layered systems↗
NASA NTRS1992

Operability of Space Station Freedom's meteoroid/debris protection system

The design of Space Station Freedom's external structure must not only protect the spacecraft from the hazardous environment, but also must be compatible with the extra vehicular activity system for assembly and maintenance. The external procedures for module support are utility connections, external orbital replaceable unit changeout, and maintenance of the meteoroid/debris shields and multilayer insulation. All of these interfaces require proper man-machine engineering to be compatible with the extra vehicular activity and manipulator systems. This paper discusses design solutions, including those provided for human interface, to the Space Station Freedom meteoroid/debris protection system. The system advantages and current access capabilities are illustrated through analysis of its configuration over the Space Station Freedom resource nodes and common modules, with emphasis on the cylindrical sections and endcones.

Kahl, Maggie S.↗
NASA NTRS2013

Space Environmental Testing of the Electrodynamic Dust Shield Technology

NASA's exploration missions to Mars and the moon may be jeopardized by dust that will adhere to surfaces of (a) Optical systems, viewports and solar panels, (b) Thermal radiators, (c) Instrumentation, and (d) Spacesuits. We have developed an active dust mitigation technology, the Electrodynamic Dust Shield, a multilayer coating that can remove dust and also prevents its accumulation Extensive testing in simulated laboratory environments and on a reduced gravity flight shows that high dust removal performance can be achieved Long duration exposure to the space environment as part of the MISSE-X payload will validate the technology for lunar missions.

Calle, Carlos I.↗
NASA NTRS2017

Cooling the Origins Space Telescope

The NASA Astrophysics Division has commissioned 4 studies for consideration by the 2020 Decadal Survey to be the next flagship mission following WFIRST (Wide Field Infrared Survey Telescope). One of the four studies is the Origins Space Telescope (OST), which will cover wavelengths from 6 microns to 600 microns. To perform at the level of the zodiacal, galactic, and cosmic background, the telescope must be cooled to 4 degrees Kelvin. 4 degrees Kelvin multi-stage mechanical cryocoolers will be employed along with a multilayer sunshield/thermal shield to achieve this temperature with a manageable parasitic heat load. Current state-of-the-art cryocoolers can achieve close to 4 degrees Kelvin, providing about 50 megawatts of cooling at 4 degrees Kelvin with an input power of 500 watts. Multiple coolers at this power level will be used in parallel. These coolers also provide extra cooling power at intermediate temperature stages of 15-20 degrees Kelvin and 50-70 degrees Kelvin . This upper stage cooling will be used to limit the heat conducted to 4 degrees Kelvin . The multi-layer sunshield will limit the radiated thermal energy to the 4 degrees Kelvin volume. This paper will describe the architecture of the cryogenic system for OST along with preliminary thermal models.

cryogenics↗
NASA NTRS2005

Apparatus for Testing Flat Specimens of Thermal Insulation

An apparatus has been developed to implement an improved method of testing flat-plate specimens of thermal-insulation materials for cryogenic application. The method includes testing under realistic use conditions that could include vacuum and mechanical loading at a pressure up to 70 psi (=0.48 MPa). The apparatus can accommodate a rigid or flexible specimen having thickness up to 1.25 in. (=3.2 cm) and diameters between 6 and 10 in. (about 15.2 and 25.4 cm, respectively). Typical test conditions include boundary temperatures between 77 K and 373 K and vacuum/interstitial gas filling at a pressure between 10(exp -6) torr (=1.3 x 10(exp -4) Pa) and 760 torr (atmospheric pressure =0.1 MPa). The interstitial gas could be N2, He, CO2, or any other suitable gas to which the insulation is expected to be exposed in use. Relative to prior apparatuses and testing methods, this apparatus and the testing method that it implements offer advantages of relative simplicity and ease of use. The basic principle of operation of the apparatus is that of boil-off calorimetry, using liquid nitrogen or any other suitable liquid that boils at a desired temperature below ambient temperature. Comparative rates of flow of heat through the thicknesses of the specimens (heat-leak rates) and apparent-thermal-conductivity values are obtained from tests of specimens. Absolute values of heat-leak rates and apparent thermal conductivities are computed from a combination of (1) the aforementioned comparative values and (2) calibration factors obtained by testing reference specimens of materials that have known thermal-insulation properties. The apparatus includes a full complement of temperature sensors, a vacuum pump and chamber, a monitoring and control system, and tools and fixtures that enable rapid and reliable installation and removal of specimens. A specimen is installed at the bottom of the vacuum chamber, and a cold-mass assembly that includes a tank is lowered into position above and around the specimen (see figure). A spring-based compensating fixture helps to ensure adequate thermal contact with possibly irregular specimen surfaces. For a high-compression test, the springs can be replaced with spacers. A flat circular load cell at the bottom of the chamber measures the compressive load on the specimen. Once the desired compressive-load, temperature, and vacuum/gas-filling conditions are established, testing begins. During a test, all measurements are recorded by use of a portable data-acquisition system and a computer. The total heat-leak rate is measured and calculated as the boil-off flow rate multiplied by the latent heat of vaporization. The parasitic heat leak (to the side of the specimen and to the top and side of the cold-mass tank) is reduced to a small fraction of the total heat leak by use of a combination of multilayer-insulation (MLI) shield rings, reflective film, a fiberglass/epoxy centering ring, and a bulk fill of aerogel beads. This combination eliminates the need for a cryogenic guard chamber used in a typical prior apparatus to reduce the parasitic heat leak.

Fesmire, James E.↗
NASA NTRS1988

Hypervelocity impact damage assessment for Space Station

To inhibit damage and limit the probability of penetration of the Space Station pressure wall by micrometeoroids and orbital debris, a shield placed away from the wall is used to form a double wall. To determine shield effectiveness and assess impact damage, existing test data were reviewed and additional testing was performed for Space Station double wall designs. Empirical spallation and penetration functions derived from the data show that shield thickness and impact angle affect the damage to the wall. Thick shields reduce wall damage for low angle impacts but increase damage for oblique impacts. Multilayer insulation between the shield and wall reduces impact damage to the wall. A relationship between impact velocity and spall damage to the wall is demonstrated. Preliminary test results on Li-Al shield material indicate possible improved effectiveness over Al shields.

Coronado, Alex R.↗
NASA NTRS1991

Shielding analysis for a manned Mars rover powered by an SP-100 type reactor

Shield design is one of the most crucial tasks in the integration of a nuclear reactor power system to a manned Mars rover. A multilayered W and LiH shield is found to minimize the shield mass and satisfy the dose rate limit of 30 rem/y to the rover crew. The effect on dose rate of tungsten layers thicknesses and position within the lithium hydride shields is investigated. Due to the large cross section for the W (n,gamma) reaction, secondary gammas become a significant radiation source. The man-rated shield mass for the Mars rover vehicle is correlated to the reactor thermal power. The correlation fits to within 9 percent of the calculated shield mass and results in an uncertainty of below 4 percent in the overall rover mass. The shield mass varied from 8600 kg to 20580 kg for a reactor thermal power of 100 to 1000 kW(t), respectively.

Morley, Nicholas J.↗
NASA NTRS1982

Progress report on the infrared astronomical satellite cryogenic system

The main cryogen tank is sized to hold 70 kg of superfluid helium with 12 percent ullage at a temperature of 1.8 K. The insulation system surrounds the main cryogen tank with four blankets of multilayer insulation spaced by means of three vapor-cooled shields. It is noted that the multilayer insulation is 6.4 micron double-aluminized Mylar separated by polyester net. The main shell provides the primary structural integrity of the system. In orbit the main shell is cooled to 170 K using insulation blankets on one side and second surface paint on the opposite side. The aperture cover subassembly is the vacuum seal for the main shell during operation on the ground. It is also a gas condensation trap before and during cooldown of the main cryogen tank during launch hold. It contains 6 kg of supercritical helium, which permits 14 days on orbit with a 48-hour launch pad hold.

Urbach, A. R.↗
NASA NTRS1975

Design of a heat pipe governed thermal control system for the Solar Electric Propulsion Stage /SEPS/

A 2200-w capacity spacecraft heat rejection system designed for the SEPS and utilizing heat pipe radiator panels has been investigated. The total thermal control system consists of two radiator panels connected to the heat source by variable conductance heat pipes (VCHP's). The system was designed to operate in the 223 to 333 temperature range. The radiators have an emittance of 0.88 at their operational temperature and a fin efficiency of approximately 80 percent. The radiators are thermally isolated from the SEPS and environment by multilayer insulation and thermal shields. Butane was selected as the working fluid for the VCHP because of its low freezing point (135), which is necessary to prevent diffusion freezeout of the liquid during the cold outbond missions. Helium was selected for the control gas. This paper describes the VCHP system, discusses the system design parameters and presents the results of the analyses.

Ruttner, L. E.↗
NASA NTRS1986

Mars transit vehicle thermal protection system: Issues, options, and trades

A Mars mission is characterized by different mission phases. The thermal control of cryogenic propellant in a propulsive vehicle must withstand the different mission environments. Long term cryogenic storage may be achieved by passive or active systems. Passive cryo boiloff management features will include multilayer insulation, vapor cooled shield, and low conductance structural supports and penetrations. Active boiloff management incorporates the use of a refrigeration system. Key system trade areas include active verses passive system boiloff management (with respect to safety, reliability, and cost) and propellant tank insulation optimizations. Technology requirements include refrigeration technology advancements, insulation performance during long exposure, and cryogenic fluid transfer system for mission vehicle propellant tanking during vehicle buildip in LEO.

Brown, Norman↗
NASA NTRS1987

Advanced long term cryogenic storage systems

Long term, cryogenic fluid storage facilities will be required to support future space programs such as the space-based Orbital Transfer Vehicle (OTV), Telescopes, and Laser Systems. An orbital liquid oxygen/liquid hydrogen storage system with an initial capacity of approximately 200,000 lb will be required. The storage facility tank design must have the capability of fluid acquisition in microgravity and limit cryogen boiloff due to environmental heating. Cryogenic boiloff management features, minimizing Earth-to-orbit transportation costs, will include advanced thick multilayer insulation/integrated vapor cooled shield concepts, low conductance support structures, and refrigeration/reliquefaction systems. Contracted study efforts are under way to develop storage system designs, technology plans, test article hardware designs, and develop plans for ground/flight testing.

Brown, Norman S.↗
NASA NTRS1993

Pressurized modules for Space Station Freedom

Crew workstations, storage, and facilities in the SSF Habitation module along with station housekeeping and supporting subsystems (power, thermal, and life support) are considered. The U.S. Laboratory and Habitation modules are based on a common structural design. The pressurized logistics modules (PLMs) have maximum commonality with the other pressurized modules including common ring frames and waffle grid skin; multilayer insulation and debris shield; and one common endcone.

Hopson, George D.↗
NASA NTRS2004

Flexible Shields for Protecting Spacecraft Against Debris

A report presents the concept of Flexshield a class of versatile, lightweight, flexible shields for protecting spacecraft against impacts by small meteors and orbiting debris. The Flexshield concept incorporates elements of, but goes beyond, prior spacecraft-shielding concepts, including those of Whipple shields and, more recently, multi-shock shields and multi-shock blankets. A shield of the Flexshield type includes multiple outer layers (called bumpers in the art) made, variously, of advanced ceramic and/or polymeric fibers spaced apart from each other by a lightweight foam. As in prior such shields, the bumpers serve to shock an impinging hypervelocity particle, causing it to disintegrate vaporize, and spread out over a larger area so that it can be stopped by an innermost layer (back sheet). The flexibility of the fabric layers and compressibility of the foam make it possible to compress and fold the shield for transport, then deploy the shield for use. The shield can be attached to a spacecraft by use of snaps, hook-and-pile patches, or other devices. The shield can also contain multilayer insulation material, so that it provides some thermal protection in addition to mechanical protection.

Christiansen, Eric L.↗