SEARCH · Search NASA
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.
Surface treatment increases resistance of turbine blades to low-cycle fatigue. Smoothing removes small flaws where cracks start. Intended for blades in turbines subject to thermal shock of rapid starting. No recrystallization occurs at rocket-turbine operating temperatures.
Three experimental research designs investigating candidate materials and processes involved in protective die surface coating procedures by sputter deposition, using ion beam technologies, are discussed. Various pre-test results show that none of the coatings remained completely intact for 15,000 test cycles. The longest lifetime was observed for coatings such as tungsten, platinum, and molybdenum which reduced thermal fatigue, but exhibited oxidation and suppressed crack initiation only as long as the coating did not fracture. Final test results confirmed earlier findings and coatings with Pt and W proved to be the candidate materials to be used on a die surface to increase die life. In the W-coated specimens, which remained intact on the surface after thermal fatigue testing, no oxidation was found under the coating, although a few cracks formed on the surface where the coating broke down. Further research is planned.
Two new techniques have been developed to characterize Pb(Mg1/3Nb2/3)O3-PbTiO3 (PMN-PT) based ferroelectric single crystals: (i) electro-thermal imaging, and (ii) single crystal x-ray diffraction in the transmission mode. (i) Electro-thermal imaging is a remote sensing technique that can detect the polarization direction and poling state of a whole crystal slice. This imaging technique utilizes an IR camera to determine the field induced temperature change and does not require any special or destructive sample preparation. In the resulting images it is possible to distinguish regions of 180 deg domains. This powerful technique can be used remotely during poling to determine the poling state of the crystal to avoid over-poling that can result in inferior properties and/or cracking of the crystals. Electro-thermal imaging produced the first direct observations of polarization rotation. Under bipolar field, the domains near the corners were the first to switch direction. As the field increased above the coercive field, domains at the center part of the crystals switched direction. (ii) X-ray diffraction in the transmission mode has long been used in structure determination of organic crystals and proteins; however, it is not used much to characterize inorganic systems. 0.7Pb(Mg1/3Nb2/3)O3-0.3PbTiO3 single crystals were examined by this XRD technique for the first time, and a never-before-seen super-lattice was revealed with a doubling of the unit cell in all three directions, giving a cell volume eight times that of a traditional perovskite unit cell. The significance of the super-lattice peaks increased with poling, indicating a structural contribution to ordering. Lack of such observations by electron diffraction in the transmission electron microscope examinations suggests the presence of a bulk effect.
Tests were made in the 20 MW arc jet facility at the NASA ARC to determine the suitability of sapphire and fused silica as window materials for the Aeroassist Flight Experiment (AFE) entry vehicle. Twenty nine tests were made; 25 at a heating rate about 80 percent of that expected during the AFE entry and 4 at approximately the full, 100 percent AFE heating rate profile, that produces a temperature of about 2900 F on the surface of the tiles that protect the vehicle. These tests show that a conductively cooled window design using mechanical thermal contacts and sapphire is probably not practical. Cooling the window using mechanical thermal contacts produces thermal stresses in the sapphire that cause the window to crack. An insulated design using sapphire, that cools the window as little as possible, appears promising although some spectral data in the vacuum-ultra-violet (VUV) will be lost due to the high temperature reached by the sapphire. The surface of the insulated sapphire windows, tested at the 100 percent AFE heating rate, showed some slight ablation, and cracks appeared in two of three test windows. One small group of cracks were obviously caused by mechanical binding of the window in the assembly, which can be eliminated with improved design. Other cracks were long, straight, thin crystallographic cracks that have very little effect on the optical transmission of the window. Also, the windows did not fall apart along these crystallographic cracks when the windows were removed from their assemblies. Theoretical results from the thermal analysis computer program SINDA indicate that increasing the window thickness from 4 to 8 mm may enable surface ablation to be avoided. An insulated design using a fused silica window tested at the nominal AFE heating rate experienced severe ablation, thus fused silica is not considered to be an acceptable window material.
Thermal and environmental barrier coatings have been applied for protecting Si based ceramic matrix composite components from high temperature environment in advanced gas turbine engines. It has been found that the delamination and lifetime of T/EBC systems generally depend on the initiation and propagation of surface cracks induced by the axial mechanical load in addition to severe thermal loads. In order to prevent T/EBC systems from surface cracking and subsequent delamination due to mechanical and thermal stresses, T/EBC systems reinforced with nano-composite architectures have showed promise to improve mechanical properties and provide a potential crack shielding mechanism such as crack bridging. In this study, a finite element model (FEM) was established to understand the potential beneficial effects of nano-composites systems such as SiC nanotube-reinforced oxide T/EBC systems.
A simple two-layer plasma-sprayed thermal barrier coating system was developed which has the potential for protecting high temperature air-cooled gas turbine components. Of the initially examined coatings, the most promising system is an Ni-16Cr-6Al-0.6Y (in wt%) bond coating (about 0.005 to 0.010 cm thick) and a ZrO2-12Y2O3 (in wt%) thermal barrier coating (about 0.025 to 0.064 cm thick). This thermal barrier substantially lowered the metal temperature of the air-cooled airfoil. The coating withstood 3200 cycles (80 s at 1280 C surface temperature) and 275 cycles (1 hr at 1490 C surface temperature) without cracking or spalling. No separation of the thermal barrier from the bond coating or the bond coating from the substrate was observed.
Among the various techniques available, ultrasonic Lamb waves offer a convenient method of evaluating composite materials. Since the Lamb wave velocity depends on the elastic properties of a structure, an effective tool exists to monitor damage in composites by measuring the velocity of these waves. Lamb wave measurements can propagate over long distances and are sensitive to the desired in-plane elastic properties of the material. This paper describes two studies which monitor fatigue damage and two studies which monitor thermal damage in composites using Lamb waves. In the fatigue studies, the Lamb wave velocity is compared to modulus measurements obtained using strain gage measurements in the first experiment and the velocity is monitored along with the crack density in the second. In the thermal damage studies, one examines samples which were exposed to varying temperatures for a three minute duration and the second includes rapid thermal damage in composites by intense laser beams. In all studies, the Lamb wave velocity is demonstrated to be an excellent method to monitor damage in composites.
A Solder Committee designated to investigate a solder cracking phenomena occurring on the SATURN electrical/electronic hardware found the cause to be induced stress in the soldered connections rather than faulty soldering techniques. The design of the printed circuit (PC) board assemblies did not allow for thermal expansion of the boards that occurred during normal operation. The difference between the thermal expansion properties of the boards and component lead materials caused stress and cracking in the soldered connections. The failure mechanism and various PC boards component mounting configurations are examined in this report. Effective rework techniques using flanged tubelets, copper tubelets, and soft copper wiring are detailed. Future design considerations to provide adequate strain relief in mounting configurations are included to ensure successful solder terminations.
A simple, two-layer plasma-sprayed thermal barrier coating system was developed which has the potential for protecting high temperature air-cooled gas turbine components. Of those coatings initially examined, the most promising system consisted of a Ni-16Cr-6Al-0.6Y (in wt%) thermal barrier coating (about 0.005 to 0.010 cm thick) and a ZrO2-12Y2O3 (in wt%) thermal barrier coating (about 0.025 to 0.064 cm thick). This thermal barrier substantially lowered the metal temperature of an air-cooled airfoil. The coating withstood 3,200 cycles (80 sec at 1,280 C surface temperature) and 275 cycles (1 hr at 1,490 C surface temperature) without cracking or spalling. No separation of the thermal barrier from the bond coating or the bond coating from the substrate was observed.
The effects of welding, five selected surface coatings, and stress relieving on the thermal fatigue resistance of H-13 Die Steel for aluminum die casting dies were studied using eleven thermal fatigue specimens. Stress relieving was conducted after each 5,000 cycle interval at 1050 F for three hours. Four thermal fatigue specimens were welded with H-13 or maraging steel welding rods at ambient and elevated temperatures and subsequently, subjected to different post-weld heat treatments. Crack patterns were examined at 5,000, 10,000, and 15,000 cycles. The thermal fatigue resistance is expressed by two crack parameters which are the average maximum crack and the average cracked area. The results indicate that a significant improvement in thermal fatigue resistance over the control was obtained from the stress-relieving treatment. Small improvements were obtained from the H-13 welded specimens and from a salt bath nitrogen and carbon-surface treatment. The other surface treatments and welded specimens either did not affect or had a detrimental influence on the thermal fatigue properties of the H-13 die steel.
For long duration missions of solar sails, the sail material needs to survive harsh space environments and the degradation of the sail material controls operational lifetime. Therefore, understanding the effects of the space environment on the sail membrane is essential for mission success. In this study, we investigated the effect of simulated space environment effects of ionizing radiation, thermal aging and simulated potential damage on mechanical, thermal and optical properties of a commercial off the shelf (COTS) polyester solar sail membrane to assess the degradation mechanisms on a feasible solar sail. The solar sail membrane was exposed to high energy electrons (about 70 keV and 10 nA/cm2), and the physical properties were characterized. After about 8.3 Grad dose, the tensile modulus, tensile strength and failure strain of the sail membrane decreased by about 20 ~ 95%. The aluminum reflective layer was damaged and partially delaminated but it did not show any significant change in solar absorbance or thermal emittance. The effect on mechanical properties of a pre-cracked sample, simulating potential impact damage of the sail membrane, as well as thermal aging effects on metallized PEN (polyethylene naphthalate) film will be discussed.
This paper explores the methods used to examine Mars Curiosity Rover’s flight temperature data to the Package Qualification and Verification (PQV) accelerated thermal cycle life test data in response to flight anomalies. The redundant Rover Compute Element (RCE-A, B) has experienced an inability to mount the flash memory several times. The leading cause of these anomalies is likely the loss of electrical connectivity, which is theorized to be due to cracks on the memory component solder joints by thermal cycle fatigue. To investigate this theory, the number of thermal cycles accumulated in flight was compared to PQV accelerated thermal cycle life test data. The goal of this study was to inform mission operations of the risk associated with thermal cycling damage on the RCE. However, since the RCE-B does not have continuous temperature data from the Platinum Resistance Thermometer (PRT) closest to the flash memory, a model was made to correlate continuous data from two other nearby sensors on the RCE-B to the noncontiguous sensor. Using the Rainflow Counting Algorithm, the reconstituted data was converted into the number of equivalent thermal cycles. The Coffin-Manson Equation was used to compare flight thermal cycles to the PQV test. The accelerated life test showed cracking on the memory component solder joints after 500 cycles of 80°C temperature excursions. This study found that the RCE-A and RCE-B have undergone an equivalent of 144 and 430 cycles of 80°C temperature excursions during flight, respectively.
This paper explores the methods used to examine Mars Curiosity Rover’s flight temperature data in relation to the Package Qualification and Verification (PQV) accelerated thermal cycle life test data in response to flight anomalies. The redundant Rover Compute Element (RCE-A, B) has experienced an inability to mount the flash memory several times. The leading cause of these anomalies is likely the loss of electrical connectivity, which is theorized to be due to cracks on the memory component solder joints by thermal cycle fatigue. To investigate this theory, the number of thermal cycles accumulated in flight was compared to PQV accelerated thermal cycle life test data. The goal of this study was to inform mission operations of the risk associated with thermal cycling damage on the RCE. However, since the RCE-B does not have continuous temperature data from the Platinum Resistance Thermometer (PRT) closest to the flash memory, a model was made to correlate continuous data from two other nearby sensors on the RCE-B to the noncontiguous sensor. Using the Rainflow Counting Algorithm, the reconstituted data was converted into the number of equivalent thermal cycles. The Coffin-Manson Equation was used to compare flight thermal cycles to the PQV test. The accelerated life test showed cracking on the memory component solder joints after 500 cycles of 80°C temperature excursions. This study found that the RCE-A and RCE-B have undergone an equivalent of 144 and 430 cycles of 80°C temperature excursions during flight, respectively.
SSME components such as the turbine blades of the high pressure fuel turbopump are subjected to rapid and extreme thermal transients that contribute to blade cracking and subsequent failure. The objective was to develop thin film sensors for SSME components. The technology established for aircraft gas turbine engines was adopted to the materials and environment encountered in the SSME. Specific goals are to expand the existing thin film sensor technology, to continue developing improved sensor processing techniques, and to test the durability of aircraft gas turbine engine technology in the SSME environment. A thin film sensor laboratory is being installed in a refurbished clean room, and new sputtering and photoresist exposure equipment is being acquired. Existing thin film thermocouple technology in an SSME environment are being tested. Various coatings and their insulating films are being investigated for use in sensor development.
Thermal control coatings with high thermal emittance and low solar absorptance are needed for Space Station Freedom (SSF) solar dynamic power module radiator (SDR) surfaces for efficient heat rejection. Additionally, these coatings must be durable to low earth orbital (LEO) environmental effects of atomic oxygen, ultraviolet radiation and deep thermal cycles which occur as a result of start-up and shut-down of the solar dynamic power system. Eleven candidate coatings were characterized for their solar absorptance and emittance before and after exposure to ultraviolet (UV) radiation (200 to 400 nm), vacuum UV (VUV) radiation (100 to 200 nm) and atomic oxygen. Results indicated that the most durable and best performing coatings were white paint thermal control coatings Z-93, zinc oxide pigment in potassium silicate binder, and YB-71, zinc orthotitanate pigment in potassium silicate binder. Optical micrographs of these materials exposed to the individual environmental effects of atomic oxygen and vacuum thermal cycling showed that no surface cracking occurred.
Thermal control coatings with high thermal emittance and low solar absorptance are needed for Space Station Freedom (SSF) solar dynamic power module radiator (SDR) surfaces for efficient heat rejection. Additionally, these coatings must be durable to low earth orbital (LEO) environmental effects of atomic oxygen, ultraviolet radiation and deep thermal cycles which occur as a result of start-up and shut-down of the solar dynamic power system. Eleven candidate coatings were characterized for their solar absorptance and emittance before and after exposure to ultraviolet (UV) radiation (200 to 400 nm), vacuum UV (VUV) radiation (100 to 200 nm) and atomic oxygen. Results indicated that the most durable and best performing coatings were white paint thermal control coatings Z-93, zinc oxide pigment in potassium silicate binder, and YB-71, zinc orthotitanate pigment in potassium silicate binder. Optical micrographs of these materials exposed to the individual environmental effects of atomic oxygen and vacuum thermal cycling showed that no surface cracking occurred.
The fiber-matrix interface requirements in an Al2O3/NiAl composite were examined from theoretical considerations. Several factors that influence the interface bonding requirements were analyzed. These include: (1) residual stresses due to fiber-matrix coefficient of thermal expansion (CTE) mismatch; (2) matrix cracking stress at room temperature; (3) fracture toughness at room temperature; (4) load transfer from the matrix to the fiber and ultimate tensile strength at the use temperature; and (5) creep resistance at high temperature. A relatively weak fiber-matrix bond, with an interfacial shear strength of approximately 15-20 MPa, might be sufficient for attaining the desired mechanical properties in the fiber direction at the use temperature. A weak fiber-matrix bond is also beneficial for increasing the fracture toughness of the composite at room temperature. In contrast, a strong fiber-matrix bond is required to withstand some of the residual stresses resulting from the fiber-matrix CTE mismatch, which are not likely to be reduced significantly by interface coatings. A relatively strong bond is also beneficial in increasing the matrix cracking stress at room temperature. Various interface coating options to accommodate the conflicting bonding requirements were reviewed. One viable coating option is to incorporate a thick, ductile interface layer well bonded to both the fiber and the matrix.
The most common material system currently used for low thrust, radiation-cooled rockets is a niobium alloy (C-103) with a fused silica coating (R-512A or R-512E) for oxidation protection. However, significant amounts of fuel film cooling are usually required to keep the material below its maximum operating temperature of 1370 C, degrading engine performance. Also the R-512 coating is subject to cracking and eventual spalling after repeated thermal cycling. A new class of high-temperature, oxidation-resistant materials are being developed for radiation-cooled rockets, with the thermal margin to reduce or eliminate fuel film cooling, while still exceeding the life of silicide-coated niobium. Rhenium coated with iridium is the most developed of these high-temperature materials. Efforts are on-going to develop 22 N, 62 N, and 440 N engines composed of these materials for apogee insertion, attitude control, and other functions. There is also a complimentary NASA and industry effort to determine the life limiting mechanisms and characterize the thermomechanical properties of these materials. Other material systems are also being studied which may offer more thermal margin and/or oxidation resistance, such as hafnium carbide/tantalum carbide matrix composites and ceramic oxide-coated iridium/rhenium chambers.