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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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NASA NTRS1979

Xenon component organization in 14301

Xe in 14301 is a superposition of in situ spallation and presumably also fission Xe, of trapped solar wind, and of trapped fission Xe and Xe-129 redistributed after origin elsewhere on the moon. Some and probably most of the trapped fission originates from Pu-244. All of the components of various origins are released independently in stepwise heating, indicating that they were acquired by different means and remain in different sites within the sample. This applies to the excess fission and to the excess Xe-129, which fail to correlate quantitatively with each other; they must thus have been trapped by different mechanisms or at different times. In spite of their independent behavior in stepwise heating, the various trapped components contribute in constant proportion to the total Xe contents of the various size fractions. This quantitative correlation between excess Xe and solar wind Xe indicates that the same circumstances control the acquisition of both. The variation of the ratio of excess to solar wind Xe among different breccias thus can be ascribed more confidently to compositional variation in total Xe available for trapping rather than to the details of the trapping process.

Bernatowicz, T. J.↗
NASA NTRS1994

Cratering and penetration experiments in teflon targets at velocities from 1 to 7 km/s

Approximately 20 sq m of protective thermal blankets, largely composed of Teflon, were retrieved from the Long Duration Exposure Facility after the spacecraft spent approximately 5.7 years in space. Examination of these blankets revealed that they contained thousands of hypervelocity impact features ranging from micron-sized craters to penetration holes several millimeters in diameter. We conducted impact experiments to reproduce such features and to understand the relationships between projectile size and the resulting crater or penetration hole diameter over a wide range of impact velocities. Such relationships are needed to derive the size and mass frequency distribution and flux of natural and man-made particles in low-earth orbit. Powder propellant and light-gas guns were used to launch soda-lime glass spheres into pure Teflon targets at velocities ranging from 1 to 7 km/s. Target thickness varied over more than three orders of magnitude from finite halfspace targets to very thin films. Cratering and penetration of massive Teflon targets is dominated by brittle failure and the development of extensive spall zones at the target's front and, if penetrated, the target's rear side. Mass removal by spallation at the back side of Teflon targets may be so severe that the absolute penetration hole diameter can become larger than that of a standard crater. The crater diameter in infinite halfspace Teflon targets increases, at otherwise constant impact conditions, with encounter velocity by a factor of V (exp 0.44). In contrast, the penetration hole size in very thin foils is essentially unaffected by impact velocity. Penetrations at target thicknesses intermediate to these extremes will scale with variable exponents of V. Our experimental matrix is sufficiently systematic and complete, up to 7 km/s, to make reasonable recommendations for velocity-scaling of Teflon craters and penetrations. We specifically suggest that cratering behavior and associated equations apply to all impacts in which the shock-pulse duration of the projectile is shorter than that assigned a unique projectile size, provided an impact velocity is known or assumed. This calibration seems superior to the traditional ballistic-limit approach.

Horz, Friedrich↗
NASA NTRS1995

Cratering and penetration experiments in Teflon targets at velocities from 1 to 7 km/s

Approximately 20 sq m of protective thermal blankets, largely composed of Teflon, were retrieved from the Long Duration Exposure Facility (LDEF) after the spacecraft had spent approximately 5.7 years in space. Examination of these blankets revealed that they contained thousands of hypervelocity impact features ranging from micron-sized craters to penetration holes several millimeters in diameter. We conducted impact experiments in an effort to reproduce such features and to -- hopefully -- understand the relationships between projectile size and the resulting crater or penetration-hole diameter over a wide range of impact velocity. Such relationships are needed to derive the size- and mass-frequency distribution and flux of natural and man-made particles in low-Earth orbit. Powder propellant and light-gas guns were used to launch soda-lime glass spheres of 3.175 mm (1/8 inch) nominal diameter (Dp) into pure Teflon FEP targets at velocities ranging from 1 to 7 km/s. Target thickness (T) was varied over more than three orders of magnitude from infinite halfspace targets (Dp/T less than 0.1) to very thin films (Dp/T greater than 100). Cratering and penetration of massive Teflon targets is dominated by brittle failure and the development of extensive spall zones at the target's front and, if penetrated, the target's rear side. Mass removal by spallation at the back side of Teflon targets may be so severe that the absolute penetration-hole diameter (Dh) can become larger than that of a standard crater (Dc) at relative target thicknesses of Dp/T = 0.6-0.9. The crater diameter is infinite halfspace Teflon targets increases -- at otherwise constant impact conditions -- with encounter velocity by a factor of V0.44. In contrast, the penetration-hole size is very thin foils (Dp/T greater than 50) is essentially unaffected by impact velocity. Penetrations at target thicknesses intermediate to these extremes will scale with variable exponents of V. Our experimental matrix is sufficiently systematic and complete, up to 7 km/s, to make reasonable recommendations for the velocity-scaling of Teflon craters and penetrations. We specifically suggest that cratering behavior and associated equations dominate all impacts in which the shock-pulse duration of the projectile (tp) is shorter than that of the target (tt). We also demonstrate that each penetration hole from space-retrieved surfaces may be assigned a unique projectile size, provided an impact velocity is known or assumed. This calibration seems superior to the traditional ballistic-limit approach.

Hoerz, Friedrich↗
NASA NTRS2015

Radiant Heating Tests on Meteoritic Material

A new effort geared toward modeling the physics of meteor entry and break-up is underway at NASA Ames Research Center. This is part of a broader interdisciplinary effort on providing physics-based risk assessment models for potentially hazardous objects. As part of the entry modeling task we are seeking to improve our understanding of, among other things, the ablation of meteoric material during high speed entry into earths atmosphere. Meteoroid entry differs greatly in some key respects from spacecraft entry modeling. First, the aerothermal environment at these high velocities (18 km/s) is dominated by radiation. Second, meteoroids less than, say, 50m in size, will likely lose a significant portion of their mass during the high-speed atmospheric entry process due to vaporization as well as melting and spallation. The mass of the object, in turn, directly affects the amount of energy deposited in the atmosphere, and therefore the amount of damage done. Thus it is important for us to understand and be able to model this process in greater detail in order to assess the hazard posed by these objects.In this presentation, we first give an overview of the simple ablation models that are typically used in meteor entry calculations. Additionally, we will present a new model which utilizes a similar approach to what is typically done for spacecraft TPS response modeling. This uses an equilibrium assumption near the surface to compute the ablation rate, as is done in heritage material response codes. Next we describe the radiant heating experiment which uses a high-powered laser to emulate the radiation dominated heating environment experienced by meteoroids during atmospheric entry. The facility the Laser Hardened Material Evaluation Laboratory (LHMEL) permits us to expose samples of meteoritic material to heating rates in excess of 100kW/sq.cm. An overview of the experimental set-up and test plan for the initial exploratory campaign at this facility will be given. Then we present both qualitative and quantitative results from this initial test series. Comparisons between predicted and measured ablation rates suggest that there may be significant blockage of the incident beam by the ablation plume. Furthermore, melt is shown to be a significant ablation mechanics, even at high heating rates (16 kW/sq.cm). Finally, comparisons between the phenomenology of the ablation of terrestrial rocks namely, basalt -- to that of meteorites show very different behavior. This is shown to likely be due in part to the effect of composition on the melt viscosity.

Stern, Eric↗
NASA NTRS2021

Pyrolysis Rate and Yield Strength Reduction Incarbon Fiber and Glass Fiber Composites Underreentry Heating Conditions

The behavior of composite materials, specifically carbon fiber reinforced plastic (CFRP) and glass fiber reinforced plastic (GFRP), under reentry conditions poses a problem for space debris reentry risk modeling. Since these materials pyrolyze rather than melt and their different components demise at different rates, modeling their destruction to determine ground impact risk is complex. Modern spacecraft are using these materials in ever-greater quantities owing to their superior strength-to-weight characteristics, and this has required that the orbital debris community improve its understanding of how these materials demise on reentry. The NASA Orbital Debris Program Office undertook an extensive test campaign to better understand the rate at which several types of GFRP and CFRP materials pyrolyze under reentry heating conditions and how that pyrolysis affects the ultimate strength of the material. GFRP with a polyester resin (G10/FR-4) and CFRP with epoxy, cyanate ester, vinyl ester, and phenolic resins were tested. The test campaign was carried out at the Inductively Coupled Plasma (ICP) Torch Facility at the University of Texas at Austin. Because the ICP facility operates in a shirt-sleeve environment, test samples can be changed within seconds or minutes, allowing many samples to be tested in a short period. Two heat flux rates, 20 W/m2 and 30W/m2, and two oxygen concentration conditions, 0% and 2% of atmospheric, were applied to all five types of material. To measure both the char rate and the effect of pyrolysis on the ultimate material strength, two types of test were carried out for each material: a char rate test on a ~10 mm thick sample of material and an in-situ bending stress test of a ~2 mm thick sample of material. Measurements of the char rate showed very similar average pyrolysis front velocity in epoxy resin CFRP as in G10 at 3.6 mm/min and 3.4 mm/min, respectively. However, the total mass loss rate in theG10 was nearly double that of the CFRP at 3.8 g/min and 2.2 g/min, respectively. This result represented a slow ablation rate of carbon fibers in the CFRP at the temperatures encountered in low Earth orbit reentry and a comparatively rapid removal of the glass fibers in G10 due to melting and spallation. Pyrolysis front velocity was more significantly affected by the type of plastic than the type of fiber, with the cyanate ester CFRP samples displaying an average pyrolysis front velocity of only 1.9 mm/min. Similarly, the effect of thermal exposure on the ultimate strength of the material depended heavily on the type of plastic and very little on the type of fiber in the material. Epoxy, vinyl ester, and polyester resins all behaved very similarly, with complete structural failure at between 400 J/g and 600 J/g of specific absorbed heat. Phenolic and cyanate ester resins, on the other hand, displayed a change in structural properties that was only barely measurable with the current apparatus even after the maximum exposure time tested. These data are being incorporated into a numerical model of the ablation and demise of composite materials that will be used to more accurately calculate the ground casualty risk of future spacecraft.

Benton R. Greene↗
NASA NTRS1998

Residual Stresses Modeled in Thermal Barrier Coatings

Thermal barrier coating (TBC) applications continue to increase as the need for greater engine efficiency in aircraft and land-based gas turbines increases. However, durability and reliability issues limit the benefits that can be derived from TBC's. A thorough understanding of the mechanisms that cause TBC failure is a key to increasing, as well as predicting, TBC durability. Oxidation of the bond coat has been repeatedly identified as one of the major factors affecting the durability of the ceramic top coat during service. However, the mechanisms by which oxidation facilitates TBC failure are poorly understood and require further characterization. In addition, researchers have suspected that other bond coat and top coat factors might influence TBC thermal fatigue life, both separately and through interactions with the mechanism of oxidation. These other factors include the bond coat coefficient of thermal expansion, the bond coat roughness, and the creep behavior of both the ceramic and bond coat layers. Although it is difficult to design an experiment to examine these factors unambiguously, it is possible to design a computer modeling "experiment" to examine the action and interaction of these factors, as well as to determine failure drivers for TBC's. Previous computer models have examined some of these factors separately to determine their effect on coating residual stresses, but none have examined all the factors concurrently. The purpose of this research, which was performed at DCT, Inc., in contract with the NASA Lewis Research Center, was to develop an inclusive finite element model to characterize the effects of oxidation on the residual stresses within the TBC system during thermal cycling as well as to examine the interaction of oxidation with the other factors affecting TBC life. The plasma sprayed, two-layer thermal barrier coating that was modeled incorporated a superalloy substrate, a NiCrAlY bond coat, and a ZrO2-8 wt % Y2O3 ceramic top coat. We examined the effect on stress during burner rig thermal cycling of the following independent variables: creep in the bond coat and top coat, oxidation, bond coat coefficient of thermal expansion, number of thermal cycles, and interfacial roughness. All these factors were suspected of influencing TBC failure. The model showed that all the material properties studied had a significant effect on the coating's residual stresses if the interface of the bond coat was rough. Bond coat expansion, bond coat oxidation, and bond coat creep had the highest effect on coating stresses, and these were highly interactive. The model also showed that the mechanism of stress generation during thermal cycling changed with the number of thermal cycles. Bond coat and top coat creep dominated stress generation during early thermal cycles, greatly increasing delamination stresses at the peaks of the bond coat. Therefore, creep is the prime driver for delamination cracking early in life, but cracking is limited to the bond coat peak region. Oxidation of the bond coat, on the other hand, tended to dominate stress generation during later cycles by greatly increasing delamination stresses over bond coat valleys. These results indicate that oxidation is the driver for the continued cracking necessary to cause ceramic layer spallation.

Freborg, A. M.↗