James A. Nesbitt

Furnace cyclic oxidation behavior of multicomponent low conductivity thermal barrier coatings

Ceramic thermal barrier coatings (TBCs) will play an increasingly important role in advanced gas turbine engines due to their ability to further increase engine operating temperatures and reduce cooling, thus helping achieve future engine low emission, high efficiency, and improved reliability goals. Advanced multicomponent zirconia (ZrO2)-based TBCs are being developed using an oxide defect clustering design approach to achieve the required coating low thermal conductivity and high-temperature stability. Although the new composition coatings were not yet optimized for cyclic durability, an initial durability screening of the candidate coating materials was conducted using conventional furnace cyclic oxidation tests. In this paper, furnace cyclic oxidation behavior of plasma-sprayed ZrO2-based defect cluster TBCs was investigated at 1163°C using 45 min hot-time cycles. The ceramic coating failure mechanisms were studied using scanning electron microscopy (SEM) combined with x-ray diffraction (XRD) phase analysis after the furnace tests. The coating cyclic lifetime is also discussed in relation to coating processing, phase structures, dopant concentration, and other thermo-physical properties.

Thermal modeling of various thermal barrier coatings in a high heat flux rocket engine

One- and two-dimensional thermal models were developed to predict the thermal response of tubes with and without thermal barrier coatings (TBCs) tested for short durations in a H2/O2 rocket engine. Temperatures were predicted using median thermophysical property data for traditional air plasma sprayed ZrO2–Y2O3 TBCs, as well as air plasma sprayed and low pressure plasma sprayed ZrO2–Y2O3/NiCrAlY cermet coatings. Good agreement was observed between predicted and measured metal temperatures. It was also shown that the variation in the reported values of the thermal conductivity of plasma sprayed ZrO2–Y2O3 coatings can result in temperature differences of up to 180°C at the ceramic/metal interface. In contrast, accounting for the presence of the bond coat or radiation from the ceramic layer had only a small effect on substrate temperatures (<20°C). The thermal models were also used to show that for the short duration test conditions of this study, a 100 μm thick ZrO2–Y2O3 coating would provide a metal temperature benefit of approximately 300°C over an uncoated tube while a 200 μm thick coating would provide a benefit greater than 500°C. The difference in the thermal response between tubes and rods was also predicted and used to explain the previously-observed increased life of TBCs on rods over that on tubes.

Health Monitoring of Thermal Barrier Coatings by Mid‐Infrared Reflectance

Mid-infrared (MIR) reflectance is shown to be a powerful tool for monitoring the integrity of 8wt% yttria-stabilized zirconia (8YSZ) thermal barrier coatings (TBCs). Because of the translucent nature of plasma-sprayed 8YSZ TBCs, particularly at MIR wavelengths (3 to 5 microns), measured reflectance does not only originate from the TBC surface, but contains strong contributions from internal scattering within the coating as well as reflectance from the underlying TBC/substrate interface. Therefore, changes in MIR reflectance measurements can be used to monitor the progression of TBC delamination. In particular, MIR reflectance is shown to reproducibly track the progression of TBC delamination produced by repeated thermal cycling (to 1163 C) of plasma-sprayed 8YSZ TBCs on Rene N5 superalloy substrates. To understand the changes in MIR reflectance with the progression of a delamination crack network, a four-flux scattering model is used to predict the increase in MIR reflectance produced by the introduction of these cracks.

Failure morphologies of cyclically oxidized ZrO2-based thermal barrier coatings

Abstract Advanced and baseline thermal barrier coatings (TBCs) were thermal cycle tested in air at 1163°C until delamination or spallation of the ceramic top coat. The top coat of the advanced TBC’s consisted of ZrO2 with various amounts of Y2O3, Yb2O3, Gd2O3, or Nd2O3 dopants. The composition of the top coat of the baseline TBC was ZrO2-8wt.%Y2O3. All top coats were deposited by air plasma spraying. A NiCrAlY or NiCoCrAlY bond coat was deposited by low pressure plasma spraying onto a single-crystal, Ni-base superalloy. The TBC lifetime for the baseline coatings was approximately 190 cycles (45 minutes at 1163°C per cycle) while the lifetime for the advanced coatings was as high as 425 cycles. The fracture surfaces and sample cross sections were examined after TBC failure by SEM and optical microscopy, and the top coats were further examined by X-ray diffraction. These post-test studies revealed that the fracture path largely followed splat boundaries with some trans-splat fracture. However, there were no obvious distinguishing features which explained the difference in TBC lifetimes between some of the advanced and baseline coatings.

Pit morphology and depth after low-temperature hot corrosion of a disc alloy

Hot corrosion of the low solvus, high refractory (LSHR) disc alloy was studied at 700 °C. The purpose of this study was to determine the conditions which result in a discrete, isolated pit morphology and to examine the influence of SO2 gas additions and various salt concentrations on the depth of those pits. Three salts, pure Na2SO4 and two Na2SO4–MgSO4 compositions, were used. It was found that with a eutectic Na2SO4–MgSO4 salt, there was no significant increase in pit depth between 0 and 30 ppm SO2 when O2 was also present in the gas stream. Gas flow was observed to affect pit formation, but the variation in the position of the corrosion mounds/pits on the sample surface was unexpected. There was limited evidence that pit nucleation was not associated with grain boundaries or grain triple point junctions. An evolution from single, isolated pits, to coalesced pits, to overlapping pits on a single sample was observed. At higher SO2 concentrations, the extent of attack increased, resulting in a uniform type of attack morphology with significant metal loss across the sample surface. It was concluded that hot corrosion attack by pit formation for these conditions is not easily explained or predicted.

High Temperature Stability of Dissimilar Metal Joints in Fission Surface Power Systems

Future generations of power systems for spacecraft and lunar surface systems will likely require a strong dependence on nuclear power. The design of a space nuclear power plant involves integrating together major subsystems with varying material requirements. Refractory alloys are repeatedly considered for major structural components in space power reactor designs because refractory alloys retain their strength at higher temperatures than other classes of metals. The relatively higher mass and lower ductility of the refractory alloys make them less attractive for lower temperature subsystems in the power plant such as the power conversion system. The power conversion system would consist more likely of intermediate temperature Ni‐based superalloys. One of many unanswered questions about the use of refractory alloys in a space power plant is how to transition from the use of the structural refractory alloy to more traditional structural alloys. Because deleterious phases can form when complex alloys are jo...

Metallic Concepts for Repair of Reinforced Carbon-Carbon Space Shuttle Leading Edges

[Abstract] The development of thermal protection systems (TPS) for re-entry vehicles has been a topic of interest since the inception of manned space flight. Re-entry materials systems are critical for a spacecraft to withstand the intense temperature and environmental effects that are associated with the re-entry profile. Recently, intense efforts have been ongoing to develop repair technologies for TPS systems of the Space Shuttle Orbiter to reduce the risk of another catastrophic loss like that of Columbia. In particular, attention has focused on the ability to apply an on-orbit repair of both the wing leading edges and the tiles on the Orbiter’s underbelly. Both large and small areas of damage to leading edges allow hot gases to enter and destroy the reinforced carbon-carbon (RCC), thereby leading to catastrophic failure. Repair of large areas of damage, greater than 25 centimeters in diameter, are particularly of interest since that is the widely accepted reason for the Columbia accident. This paper presents results on a metallic overwrap concept that can conform to the contour of the wing leading edges and prevent internal structural damage from the hot gases during re-entry, which can reach in excess of 1649 o C in a highly oxidizing plasma environment. Sheet samples (380µm thick) of several silicide-coated refractory alloys were evaluated and a series of coated Re-based concepts were down-selected for more rigorous testing in an arcjet testing facility. The Re sheet was coated with an R512E silicide coating to mitigate oxidation. Two additional concepts using Ir were examined. The first examined Ir as a surface layer without the silicide coating while the second added a layer of Ir below an outer layer of Re which also received the R512E silicide coating. The R512E coating, consisting of Fe, Cr, and Re silicides, is known to be brittle and was found to contain numerous fine cracks both perpendicular and parallel to the coating surface. Consequently, a “Type A” coating consisting of a mix of sodium silicate and SiC was applied to the surface to seal fine cracks 2 in the R512E coating as well as for emissivity considerations. The concepts with the R512E coatings (Re+R512E+Type A and Re+Ir+Re+R512E+Type A) survived the arcjet exposure that simulated atmospheric re-entry conditions forming a porous silca scale. However, the Ir coating alone did not sufficiently protect the Re. Part of the failure of the Ir coating, deposited by CVD, may have been due to poor bonding with the Re substrate observed in test samples. Although handling and bending of the coated Re sheet remains a concern, the coated Re materials look promising for repairing Orbiter wing leading edges exposed to temperatures exceeding 1649 o C.

Cyclic oxidation of FeCrAlY/Al2O3 composites

Abstract Three-ply composites consisting of a FeCrAlY matrix and continuous single crystal Al2O3 (sapphire) fibers were cyclically oxidized at 1,000° and 1,100°C for up to 1,000 1-h cycles. FeCrAlY matrix only samples were also fabricated and tested for comparison. Fiber ends were exposed at the ends of the composite samples. Following cyclic oxidation, cracks running parallel to and perpendicular to the fibers were observed on the large surface of the composite. In addition, there was evidence of increased scale damage and spallation around the exposed fiber ends, particularly around the middle ply fibers. This damage was more pronounced at the higher temperature. The exposed fiber ends showed cracking between fibers in the outer plies, occasionally with Fe and Cr-rich oxides growing out of the cracks. Large gaps developed at the fiber–matrix interface around many of the fibers, especially those in the outer plies. Oxygen penetrated many of these gaps resulting in significant oxide formation at the fiber–matrix interface far within the composite sample. Around several fibers, the matrix was also internally oxidized showing Al2O3 precipitates in a radial band around the fibers. The results show that these composites have poor cyclic oxidation resistance due to the CTE mismatch and inadequate fiber–matrix bond strength at temperatures of 1,000°C and above.