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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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At least 289 records · Page 16

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NASA NTRS1990

Characterization of space station multilayer insulation damage due to hypervelocity space debris impact

Four main tasks were accomplished. The first three tasks were related to the goal of measuring the degradation of the insulating capabilities of Space Station multilayer insulation (MLI) due to simulated space debris impacts at hypervelocities. The last task was associated with critically reviewing a Boeing document on the fracture characteristics of the Space Station pressure wall when subjected to a simulated hypervelocity space debris impact. In Task 1, a thermal test procedure for impact damaged MLI specimens was written. In Task 2, damaged MLI specimens were prepared. In Task 3, a computer program was written to simulate MLI thermal tests. In Task 4, the author reviewed a Boeing document describing hypervelocity impact testing on biaxially stressed plates.

Rule, William Keith↗

NASA NTRS2008

Predicting Boundary-Layer Transition on Space-Shuttle Re-Entry

The BLT Prediction Tool ("BLT" signifies "Boundary Layer Transition") is provided as part of the Damage Assessment Team analysis package, which is utilized for analyzing local aerothermodynamics environments of damaged or repaired space-shuttle thermal protection tiles. Such analyses are helpful in deciding whether to repair launch-induced damage before re-entering the terrestrial atmosphere.

Berry, Scott↗

NASA NTRS2016

Combining Passive Thermography and Acoustic Emission for Large Area Fatigue Damage Growth Assessment of a Composite Structure

Passive thermography and acoustic emission data were obtained for improved real time damage detection during fatigue loading. A strong positive correlation was demonstrated between acoustic energy event location and thermal heating, especially if the structure under load was nearing ultimate failure. An image processing routine was developed to map the acoustic emission data onto the thermal imagery. This required removing optical barrel distortion and angular rotation from the thermal data. The acoustic emission data were then mapped onto thermal data, revealing the cluster of acoustic emission event locations around the thermal signatures of interest. By combining both techniques, progression of damage growth is confirmed and areas of failure are identified. This technology provides improved real time inspections of advanced composite structures during fatigue testing.Keywords: Thermal nondestructive evaluation, fatigue damage detection, aerospace composite inspection, acoustic emission, passive thermography

Zalameda, Joseph N.↗

NASA NTRS1993

Micrometeoroid/space debris effects on materials

The Long Duration Exposure Facility (LDEF) micrometeoroid/space debris impact data has been reduced in terms that are convenient for evaluating the overall quantitative effect on material properties. Impact crater flux has been evaluated as a function of angle from velocity vector and as a function of crater size. This data is combined with spall data from flight and ground testing to calculate effective solar absorption and emittance values versus time. Results indicate that the surface damage from micrometeoroid/space debris does not significantly affect the overall surface optical thermal physical properties. Of course the local damage around impact craters radically alter optical properties. Damage to composites and solar cells on an overall basis was minimal.

Zwiener, James M.↗

NASA NTRS1998

Thermal High- and Low-Cycle Fatigue Behavior of Thick Thermal Barrier Coating Systems

Ceramic thermal barrier coatings have received increasing attention for advanced gas turbine and diesel engine applications because of their ability to provide thermal insulation to engine components. However, the durability of these coatings under the severe thermal cycling conditions encountered in a diesel engine (ref. 1) still remains a major issue. In this research at the NASA Lewis Research Center, a high-power laser was used to investigate the thermal fatigue behavior of a yttria-stabilized zirconia coating system under simulated diesel engine conditions. The mechanisms of fatigue crack initiation and propagation, and of coating failure under complex thermal low-cycle fatigue (LCF, representing stop/start cycles) and thermal high-cycle fatigue (HCF, representing operation at 1300 rpm) are described. Continuous wave and pulse laser modes were used to simulate pure LCF and combined LCF/HCF, respectively (ref. 2). The LCF mechanism was found to be closely related to the coating sintering and creep at high temperatures. These creep strains in the ceramic coating led to a tensile stress state during cooling, thus providing the major driving force for crack growth under LCF conditions. The combined LCF/HCF tests induced more severe coating surface cracking, microspallation, and accelerated crack growth than did the pure LCF test. HCF thermal loads also facilitated lateral crack branching and ceramic/bond coat interface delaminations. HCF is associated with the cyclic stresses originating from the high-frequency temperature fluctuation at the ceramic coating surface. The HCF thermal loads act on the crack by a wedging mechanism (ref. 1), resulting in continuous crack growth at temperature. The HCF stress intensity factor amplitude increases with the interaction depth and temperature swing, and decreases with the crack depth. HCF damage also increases with the thermal expansion coefficient and the Young's modulus of the ceramic coating (refs. 1 and 3).

Miller, Robert A.↗

NASA NTRS2006

Orbital Debris Shape and Orientation Effects on Impact Damage to Shuttle Tiles

Taking the damage results from a previous paper as a guide, and using a tile model created for the STS-107 accident investigation, we used the SPHC hydrodynamic code to evaluate the probable worst-case impact effects of flat, rectangular, "flake-shaped," orbital debris particles on Space Shuttle thermal tiles. We compared the damage from flakes with that produced by spheres. The flakes and spheres were sized according to a "characteristic length" (Lc) derived from radar cross-section measurements, and embodied in the NASA Standard Breakup Model (SBM). Impacts were simulated at near-normal obliquity, at 12 km/sec. We modeled the worst-case flake orientation: a corner-on impact, an orientation we term a "Face A-B" impact. Results of our simulations indicate that flake impactors are less damaging than spheres of the same Lc. Since spherical impactors have been assumed in analyses of shuttle orbital debris impact risk, we find that these risks may have been overestimated. This work represents a preliminary second step, i.e., a follow-on to [1], in developing a sensitivity analysis for the expected range of effects on damage considering spherical vs. non-spherical impactors, as recommended by the Institute for Defense Analyses (IDA) report to the Columbia Accident Investigation Board.

Evans, Steven W.↗

NASA NTRS2004

Experimental Hypersonic Aerodynamic Characteristics of the Space Shuttle Orbiter for a Range of Damage Scenarios

Brauckmann, Gregory J.↗

NASA NTRS2007

EVA Hazards due to TPS Inspection and Repair

Tile inspection and repair activities have implicit hazards associated with them. When an Extra Vehicular Activities (EVA) crewmember and associated hardware are added into the equation, additional hazards are introduced. Potential hazards to the Extravehicular Mobility Unit (EMU), the Orbiter or the crew member themselves are created. In order to accurately assess the risk of performing a TPS inspection or repair, an accurate evaluation of potential hazards and how adequately these hazards are controlled is essential. The EMU could become damaged due to sharp edges, protrusions, thermal extremes, molten metal or impact with the Orbiter. Tools, tethers and the presence of a crew member in the vicinity of the Orbiter Thermal Protection System (TPS) pose hazards to the Orbiter. Hazards such as additional tile or Reinforced Carbon-Carbon (RCC) damage from a loose tool, safety tethers, crewmember or arm impact are introduced. Additionally, there are hazards to the crew which should be addressed. Crew hazards include laser injury, electrical shock, inability to return to the airlock for EMU failures or Orbiter rapid safing scenarios, as well as the potential inadvertent release of a crew member from the arm/boom. The aforementioned hazards are controlled in various ways. Generally, these controls are addressed operationally versus by design, as the majority of the interfaces are to the Orbiter and the Orbiter design did not originally account for tile repair. The Shuttle Remote Manipulator System (SRMS), for instance, was originally designed to deploy experiments, and therefore has insufficient design controls for retention of the Orbiter Boom Sensor System (OBSS). Although multiple methods to repair the Orbiter TPS exist, the majority of the hazards are applicable no matter which specific repair method is being performed. TPS Inspection performed via EVA also presents some of the same hazards. Therefore, the hazards common to all TPS inspection or repair methods will be addressed.

Stewart, Christine E.↗

NASA NTRS2015

Mechanical Properties and Real-Time Damage Evaluations of Environmental Barrier Coated SiC/SiC CMCs Subjected to Tensile Loading Under Thermal Gradients

SiC/SiC ceramic matrix composites (CMCs) require new state-of-the art environmental barrier coatings (EBCs) to withstand increased temperature requirements and high velocity combustion corrosive combustion gasses. The present work compares the response of coated and uncoated SiC/SiC CMC substrates subjected to simulated engine environments followed by high temperature mechanical testing to asses retained properties and damage mechanisms. Our focus is to explore the capabilities of electrical resistance (ER) measurements as an NDE technique for testing of retained properties under combined high heat-flux and mechanical loading conditions. Furthermore, Acoustic Emission (AE) measurements and Digital Image Correlation (DIC) were performed to determine material damage onset and accumulation.

Ceramic matrix composites↗

NASA NTRS2006

Thermal Cycle Reliability of PBGA/CCGA 717 I/Os

Status of thermal cycle test results for a nonfunctional daisy-chained peripheral ceramic column grid array (CCGA) and its plastic ball grid array (PBGA) version. both having 560 I/Os. were presented in last year's conference. Test results included environmental data for three different thermal cycle regimes (-55 C/125 C, -55 C/100 C, and -50 C /75 C). Update information on these - especially failure type for assemblies with high and low solder volumes-were presented. The thermal cycle test procedure followed those recommended IPC-9701 for tin-lead solder joint assemblies. Revision A of this specification covers guideline thermal cycle requirements for Pb-free solder joints. Some background information discussed during release of this specification with its current guideline recommendations were also presented. In a recent reliability investigation a fully populated CCGA with 717 I/Os was also considered for assembly reliability. evaluation. The functional package is a field-programmable gate array that has much higher processing power than its previous version. This new package is smaller in dimension, has no interposer, and has a thinner column wrapped with copper for reliability improvement. This paper will also present thermal cycle test results for this package assembly and its plastic version with 728 I/Os. both of which were exposed to three different cycle regimes. Two cycle profiles were those specified by IPC- 9701A for tin-lead, i.e. -55 to 100 C and -55 to 125 C and one was a cycle profile specified by Mil-Std-883, i.e.. -65 C/150 C which is generally used for ceramic hybrid packages. Per IPC-9701 A, test vehicles were built using daisy chain packages and were continuously monitored. The effects of many process and assembly variables-including corner staking commonly used for improving resistance to mechanical loading such as drop and vibration loads--were also considered as part of the test matrix. Optical photomicrographs were taken at various thermal cycle intervals to document damage progress and behavior. Representative samples of these along with cross-sectional photomicrographs at higher magnification taken by scanning electron microscopy (SEM) to determine crack propagation and failure analyses for packages are also presented.

staking↗

NASA NTRS1986

Ceramic-ceramic shell tile thermal protection system and method thereof

A ceramic reusable, externally applied composite thermal protection system (TPS) is proposed. The system functions by utilizing a ceramic/ceramic upper shell structure which effectively separates its primary functions as a thermal insulator and as a load carrier to transmit loads to the cold structure. The composite tile system also prevents impact damage to the atmospheric entry vehicle thermal protection system. The composite tile comprises a structurally strong upper ceramic/ceramic shell manufactured from ceramic fibers and ceramic matrix meeting the thermal and structural requirements of a tile used on a re-entry aerospace vehicle. In addition, a lightweight high temperature ceramic lower temperature base tile is used. The upper shell and lower tile are attached by means effective to withstand the extreme temperatures (3000 to 3200F) and stress conditions. The composite tile may include one or more layers of variable density rigid or flexible thermal insulation. The assembly of the overall tile is facilitated by two or more locking mechanisms on opposing sides of the overall tile assembly. The assembly may occur subsequent to the installation of the lower shell tile on the spacecraft structural skin.

Riccitiello, Salvatore R.↗

NASA NTRS1993

The continuing materials analysis of the thermal control surfaces experiment (S0069)

The long term effects of the natural and induced space environment on spacecraft surfaces are critically important to future spacecraft - including Space Station Freedom. The damaging constituents of this environment include thermal vacuum, solar ultraviolet radiation, atomic oxygen, particulate radiation, and the spacecraft induced environment. The behavior of materials and coatings in the space environment continues to be a limiting technology for spacecraft and experiments. The Thermal Control Surfaces Experiment (TCSE) was flown on the National Aeronautics and Space Administration (NASA) Long Duration Exposure Facility (LDEF) to study these environmental effects on surfaces-particularly on thermal control surfaces. The TCSE was a comprehensive experiment that combined in-space measurements with extensive pre- and post-flight analyses of thermal control surfaces to determine the effects of exposure to the low Earth orbit space environment. The TCSE is the first space experiment to directly measure the total hemispherical reflectance of thermal control surfaces in the same way they are routinely measured in the laboratory. The trend analyses of selected coatings performed as part of the continuing post-flight analysis of the TCSE are described. A brief description of the TCSE and its mission on LDEF are presented. There are several publications available that describe the TCSE, it's mission on LDEF, and initial results in greater detail. These are listed in the TCSE Bibliography.

Wilkes, Donald R.↗

NASA NTRS2003

Experimental Hypersonic Aerodynamic Characteristics of the Space Shuttle Orbiter for a Range of Damage Scenarios

Aerodynamic tests in support of the Columbia accident investigation were conducted in two hypersonic wind tunnels at the NASA Langley Research Center, the 20-Inch Mach 6 Air Tunnel and the 20-Inch Mach 6 CF4 Tunnel. The primary purpose of these tests was to measure the forces and moments generated by a variety of outer mold line alterations (damage scenarios) using 0.0075-scale models of the Space Shuttle Orbiter (approximately 10 inches in length). Simultaneously acquired global heat transfer mappings were obtained for a majority of the configurations tested. Test parameters include angles of attack from 38 to 42 deg, unit Reynolds numbers from 0.26 to 3.0 x10^6 per foot, and normal shock density ratios of 5 (Mach 6 air) and 12 (Mach 6 CF4). The damage scenarios evaluated included asymmetric boundary layer transition, gouges in the windward surface acreage thermal protection system tiles, wing leading edge damage (partially and fully missing reinforced carbon-carbon (RCC) panels), holes through the wing from the windward surface to the leeside, deformation of the wing windward surface, and main landing gear door and/or gear deployment. The aerodynamic data were compared to the magnitudes and directions observed in flight, and the heating images were evaluated in terms of the location of the generated disturbances and how these disturbance might relate to the response of discrete gages on the Columbia Orbiter vehicle during entry. The measured aerodynamic increments were generally small in magnitude, as were the flight-derived values during most of the entry. Asymmetric boundary layer transition (ABLT) results were consistent with the flight-derived Shuttle ABLT model, but not with the observed flight trends for STS-107. The partially missing leading edge panel results best matched both the early aerodynamic and heating trends observed in flight. A progressive damage scenario is presented that qualitatively matches the flight observations for the full entry.

Brauckman, Gregory J.↗

NASA NTRS1980

Thermal performance of a mechanically attached ablator tile for on-orbit repair of shuttle TPS

The reusable surface insulation (RSI) material used in the shuttle thermal protection system is susceptible to damage. If any RSI tiles are damaged or lost during ascent, they must be repaired or replaced prior to entry. One approach to replacing a damaged or missing RSI tile consists of mechanically attaching a tile of ablation material in the place of the RSI tile. The thermal performance of this type of repair tile was evaluated in a simulated entry heating environment. The test specimen consisted of the ablator repair tile mechanically fastened to the strain isolation pad and surrounded by RSI tiles. The evaluation of the thermal performance was based on temperature response, surface integrity, and predicted flight performance. When the ablator tile protruded 1/8 inch above the surrounding RSI tiles, the forward facing steps caused significant inflow of hot gas down the ablator RSI joints and this inflow caused greatly increased back surface temperatures.

Thompkins, S. S.↗

NASA NTRS1963

The effect of micrometeorite impact on the optical properties of metals.

Micrometeorite impact damage effect on solar absorptance to thermal emittance ratio of metal surfaces

METAL SURFACE↗

NASA NTRS2016

Large Area Nondestructive Evaluation of a Fatigue Loaded Composite Structure

Large area nondestructive evaluation (NDE) inspections are required for fatigue testing of composite structures to track damage initiation and growth. Of particular interest is the progression of damage leading to ultimate failure to validate damage progression models. In this work, passive thermography and acoustic emission NDE were used to track damage growth up to failure of a composite three-stringer panel. Fourteen acoustic emission sensors were placed on the composite panel. The signals from the array were acquired simultaneously and allowed for acoustic emission location. In addition, real time thermal data of the composite structure were acquired during loading. Details are presented on the mapping of the acoustic emission locations directly onto the thermal imagery to confirm areas of damage growth leading to ultimate failure. This required synchronizing the acoustic emission and thermal data with the applied loading. In addition, processing of the thermal imagery which included contrast enhancement, removal of optical barrel distortion and correction of angular rotation before mapping the acoustic event locations are discussed.

Joseph N Zalameda↗

NASA NTRS1982

Results of tests of K5NA and a revised formulation of EPDM/cork patch material in MSFC Hot Gas Facility

During prelaunch procedures at Kennedy Space Center some of the EPDM Thermal Protection System material was damaged on the Solid Rocket Booster stiffener stubs. The preferred solution was to patch the damaged areas with a cork-filled epoxy patching compound. Before this was done, however, it was requested that this patching technique be checked out by testing it in the MSFC Hot Gas Facility. Two tests were run in the HFG in 1980. The results showed the patch material to be adequate. Since that time, the formulation of the cork-filled epoxy material has been changed. It became necessary to retest this concept to be sure that the new material is as good as or better than the original material. In addition to the revised formulation material, tests were also made using K5NA as the patch material. The objectives of the tests reported herein were to: (1) compare the thermal performance of the original and the new cork-filled epoxy formulations, and (2) compare the K5NA closeout material to these epoxy materials. Material specifications are also discussed.

Dean, W. G.↗

NASA NTRS2012

NDT and SHM of Carbon Fiber Composites using Linear Drive MWM-Arrays

Carbon fiber composites are used in a wide range of structural applications due to their excellent specific strength and stiffness. However, the anisotropic mechanical and electrical properties associated with the fibers within each composite layer present challenges, and opportunities, for Nondestructive Testing (NDT) methods used to characterize and assess the structure condition. This includes composite condition after manufacture (such as fiber orientation and density, porosity, delamination, and bond strength) and during usage (such as damage from impact, fiber breakage, thermal exposure or applied stress). Ultrasonic and thermographic methods can address some of these challenges, but eddy current methods provide an alternative method for composite structures that contain a conducting material, such as carbon fibers or a metallic liner. This presentation reviews recent advances in the development of eddy current sensors and arrays for carbon fiber composite NDT and Structural Health Monitoring (SHM) applications. The focus is on eddy current sensor constructs with linear drive windings, such as MWM -Arrays, that induce currents primarily within the linear fibers of the composite. By combining this type of sensor construct with micromechanical models that relate composite constituent properties to measurable sensor responses, insight is obtained into the volumetric distribution of electrical properties within the composite and the associated manufacturing, damage, or strain conditions. With knowledge of the fiber layup, this MWM-Array technology is able to detect damage and strain/stress as a function of depth and fiber orientation. This work has been funded by NASA, NA V AIR and the Army for applications ranging from composite overwrapped pressure vessels (COPVs) to aircraft structures and rotorcraft blades. This presentation will specifically present background on the MWM-Array technology, results from the micromechanical modeling effort, and results from ongoing efforts for high resolution imaging and volumetric strain sensing.

Washabaugh, Andrew↗