High-temperature corrosion is understood to be the oxidative damaging of metals and metal alloys through hot gases, fluids, and solid materials. It includes all damage caused by reactions with the surrounding atmosphere itself or foreign particles in the atmosphere (Ref. 5.4.5-1, Fig. "HTC Classification"). Oxidation and hot gas corrosion (HGC, sulfidation) are important life-span reducing damage mechanisms that occur primarily in hot parts in the hot gas flow, such as combustion chambers and turbine bladings. Contaminants in the intake air, in the main air flow in the engine, and in the air system (e.g. cooling air, barrier air, leakage air) not only affect the outer surfaces of cooled parts that are struck by hot gas, but also affect the back and inside surfaces of parts via the cooling air. In modern engines, which use largely corrosion-resistant materials such as titanium and nickel alloys in the compressor area, the corrosion problems that affected the compressor areas of older engine types have shifted into the hot part zone.
Air contaminants in all states of aggregation affect the hot parts through various corrosion-based damage mechanisms. Corrosion is dominant in the lower temperature range up to about 950°C. Above this temperature, oxidation dominates. This results in flat pittings, pitting-like corrosion notches, grain boundary damage, and deep surface changes. This can alter the geometry through material removal or buildup, reducing the material strength through damage, affecting the plastic strain behavior, and shortening the operating life span. One of the specific damage types in hot parts is damage that can be grouped under the term high-temperature corrosion (HTC).
The hot gas typically contains further oxidation-promoting components from the fuel. Damage due to oxidation can be controlled by applying oxidation-resistant coatings (such as Al diffusion coatings, MCrAlY coatings). If aggressive deposits have collected on the affected surface, then rapidly advancing damaging of the base material occurs. Sulfidation has been shown to be especially aggressive (Fig. "Sulfidation attack"). This type of HGC has unique progress dynamics based on what seem to be relatively minor amounts of sulfur.
If ceramic thermal insulation layers are used on hot parts ( ), then special attention must be given to (substrate) oxidation and/or reactions with deposits.
Types of high-temperature corrosion relevant to engines, and their technical terms:
Oxidation: With the operating temperature ranges of modern engines and their typical structural materials, oxidation leads to creation of conversion coatings made from oxides of the base material, such as Al oxide and Cr oxide.
If these brittle oxides are erosively removed or damaged by strain (e.g. heat strain and volume changes, cracks and spalling; Fig. "Oxidation layer behaviour") then oxidation can progress until it unallowably weakens cross-sections and causes the affected part to fail.
Catastrophic oxidation is a type of oxidation with material loss that occurs due to the creation of liquid or volatile oxides. A typical example of a liquid oxide is vanadium pentoxide with a melting point of about 660 °C. Above 1000°C with high oxygen partial pressure, Cr oxides will largely evaporate. Al oxides, on the other hand, will not noticeably evaporate.
Coking: coking is an internal corrosion type that is based on the formation of carbides that result in a lack of those alloy components that create protective coatings. Damage through coking can occur around injection nozzles and the surrounding combustion chamber areas, especially underneath soot and coke deposits.
Metal Dusting: this special form of coking has been observed in Fe-, Ni- and Co-based alloys under constant coking conditions. It can cause heavy material removal. For this reason, this damage type is also called catastrophic coking (Ref. 5.4.5-1). In an intermediate stage, carbides form as internal corrosion products. Oxides, which are the end product of the damage process, typically fall off of the surface as dust.
Green Rot is also a type of coking which occurs, however, under alternating coking and oxidizing conditions (Ref. 5.4.5-1). The bonding of Cr prevents the creation of a protecting Cr oxide conversion coating, leading to heavy oxidation of Fe components. Extreme material removal was observed in one case in which, under low oxygen exchange, a Ni based alloy was bolted down with bolts made from Fe based material. This resulted in the Ni based alloy being seriously damaged through material removal with clear green discolorations.
Sulfidation (hot gas corrosion HGC, sulfidation): this oxidation process occurs under the effect of sulfur and, in oxidation-resistant materials, this process is several orders of magnitude faster than normal oxidation (Fig. "Sulfidation attack"). Sulfidation can be initiated and further accelerated by small amounts of silver (Ref. 5.4.5-2, Example "Silver coating"). For this reason, some manufacturers only allow the use of non-silver-coated bolts and nuts in certain hot parts.
Figure "Oxidation layer behaviour" (Ref. 5.4.5-1): Based on the mass changes, conclusions can be made regarding protecting or damaging properties of oxide coatings (top diagram). Mass increases that do not level off indicate progressive oxidation due to an oxygen-permeable (diffusion, porosity, and/or cracking) oxide layer.
The lower diagrams show models of the different damaging mechanisms of brittle oxide layers, where either volume increases induce compressive stress or volume losses induce tensile stress that exceeds the coating strength. As soon as crack initiation occurs in the oxide coating, there is a danger of substrate oxidation, which causes the oxide coating to break off, even if there are tension residual stresses present.
Figure "Sulfidation attack": There are two different types of damage processes with sulfidation. Type I occurs above 800°C and Type II occurs between about 600°C to 800°C. These temperatures are not determined by the affected base material, but by the composition of the aggressive deposits. The primary damage mechanism is destruction of the protective oxide conversion coating, which lets in damaging salt melts. The components of damaging salt melts are sodium, vanadium, magnesium, calcium, and sulfur. Because engines burn relatively clean fuel, corrosion-inducing deposits generally come from contaminants in the intake air. Because sodium and sulfur have a special influence, engines that are in a marine atmosphere and/or being acted upon by dusts containing sulfur (such as dusts containing gypsum) are especially threatened. Experience has shown that silver, such as that on silver-plated bolts, can seriously accelerate the sulfidation process (Example "Silver coating"), whereby it probably acts as a catalyst. Silver deposits can evidently be created during standstill in the gas turbine through condensation water causing watery corrosion and acting on surface contaminations. These watery solutions are then transported to other part zones (e.g. flange lugs of the turbine disks). Evidently, corrosion-initiating silver again precipitates from these solutions. Experience has shown this corrosion process to be decisively dependent on the surrounding conditions. Pronounced damage of this type has only been reported in engines that were run in testing rigs in the vicinity of a chemical factory. Here, as well, it should be said that consciously introduced contaminants in the intake air (pollutant burning) must be specifically tested for potential effects such as the ones mentioned. The sulfidation damage mechanism can be explained as follows:
Evidently even very small amounts of sulfur are enough for this damage process, since it only acts on a narrow zone at the transition to the base material and promotes the development of large amounts of Ni oxide. Even small amounts of nickel sulfide are sufficient for transporting the oxygen. Metallographic verification is correspondingly difficult. A non-destructive method for verifying sulfidation in exclusively externally accessible hollow structures (e.g. turbine rotor blades and stator vanes in their assembled state, Fig. "Oxidation symptoms on turbine blades") is magnetoscopic testing. It detects larger amounts of magnetic depleted base material which consists up to 90% from Ni and Co. Nonmagnetic is instead the dense oxide and the unaffected base material.
HTC in turbine rotor blades
Experience has shown that overly high temperatures cannot be completely avoided in limited areas of the hot parts. Even the normal temperature fluctuations in the gas flow over the circumference and in radial direction in the high-pressure turbine (HPT) can be considerably greater than 100°C. This puts extreme thermal stress especially on the inlet stator vanes (Fig. "Oxidation symptoms on turbine blades"). But the downwind HPT rotor blades with high local temperatures will need oxidation protection, despite the temperature-averaging effect of rotation. The life span of these oxidation-resistant coatings, and therefore also the parts, depends on the emaciation due to oxidation. If the oxidation-resistant coating is worn through, the accelerated damaging of the base material ensues.
Sulfidation (Fig. "Sulfidation attack") occurs in part zones with temperature ranges suitable for these damage types. It can affect entire blade cross-sections and thoroughly damage them (Fig. "Internal sulfidation on turbine blades" and Fig. "Perforation by sulfidation"). There is an especially high risk in poorly ventilated hollow structures (Ref. 5.4.5-4) such as, for example, are typical in thick blades which can have “windows” in order to reduce weight. Other susceptible surfaces are those on which dust can easily settle, such as leading edges of stator vane shrouds.
If ceramic thermal barrier coatings are used, they are prone to spalling after longer run times due to substrate oxidation. These coatings, which have necessary functional cracking, can also be burst open and spall by melted deposits that enter into them an then cool (Fig. "Typical problems with thermal barrier layers").
HTC on honeycomb seals
A special problem is long-time damaging of honeycomb seals. These are used in housings across from rotor blade tips and on the inner stator vane shrouds across from the intermediate stage labyrinths of the rotors. They consist of a thin, honeycomb-like structure made from sheet metal and are designed to abrade upon contact with the labyrinth fins in order to minimize clearances, at least during certain operating conditions. The sheet metal honeycombs are only a few tenths of a millimeter thick and subject to long-term influences of hot gas; especially the seals across from the rotor blade tips can be extremely damaged due to oxidation and lose their strength. This causes large surface areas to break out around the circumference, leading to increased clearances and leakages. If this type of damage seems possible, then suitable materials and/or coatings must be used on the honeycomb structure. A suitable pre-oxidation of the new parts offers a certain amount of improvement. Whether or not this is sufficient, however, needs to be determined in each specific situation.
Figure "Oxidation symptoms on turbine blades": The inlet stator vanes of the high-pressure turbine and the blades of the first rotor stage are usually subject to extreme temperatures, despite intense cooling measures. If the normal oxidation-resistant coating is emaciated (“3”) or has suffered considerable thermal damage (e.g. spalling and rolling up), then an accelerated attack on the base material occurs, and can be seen in a rough and cracked surface (“orange peel effect”, arrows “1”) that appears dark and/or green, depending on the oxides that form. If the blade wall has been oxidized through (“2”), hot gas can enter into the cooling air ducts, causing cooling air losses and the danger of serious consequential damages not only to the directly affected part, but also to other parts (such as those that are connected to the same cooling air supply).
Figure "Internal sulfidation on turbine blades": Sulfidation is especially pronounced when the oxygen supply is not sufficient to allow a dense protective oxide layer to develop. This is evidently the case in hollow blades (weight reduction) that are closed on one end. In extreme cases, this results in “windows” opening up in the blades, i.e. both blade walls are broken through. This damage process is made more problematic by the fact that the extent of the damage can only be seen at a very late stage.
In blades with cooling air through-flow, there seem to be limited areas inside the blade where aggressive dusts tend to deposit (Ref. 5.4.5-7), and the part temperature is suitable for this phenomenon (Ref. 5.4.5-4). The temperature ranges necessary for sulfidation to occur determine the location of damages on the engine parts.
Figure "Perforation by sulfidation" (Ref. 5.4.5-8): This picture according to an photo demonstrates the extreme deterioration of a LP-turbine stator after some 1000 operation hours. The sulfidation attack (Fig. "Perforation by sulfidation") starts in the hollow vane profiles. If the stator has an “intermesh function” (Fig. "Intermeshing principle") the question for the remaining function arises.
Excerpt: ”…is planning to replace a silver coating on (a fighter) engine high pressure turbine disk front and rear fasteners after an inspection revealed that small cracks had developed in the disk during intensive testing…
The engine had been run more than 4,300 cycles, which is equivalent to about 2,000 flight hours, before it was torn down for the inspection…
The silver coating used on the nuts and bolts securing the disk's front and rear retainer had reacted with sulfur in the engine's jet fuel, creating a sulfate compound that exposed the disk surface to acid. The acid led to corrosion and about half a dozen cracks in the disk…“
Comments: In this case it is unclear whether corrosive attack occurred due to the creation of sulfuric acid from condensation water during standstill (Ill. 12.4-14, Ill. 22.214.171.124-2.2), and/or if it involved sulfidation in dry, hot conditions. Both mechanisms are described in the literature and have been observed in engines.
Figure "Outer sulfidation on turbine blades": Sulfidation is limited to very specific temperature ranges (Fig. "Sulfidation attack") in which stable aggressive salt melts exist. For this reason, external sulfidation zones develop in very typical and part-specific zones. Aside from suitable temperature, these zones are also determined by a location suitable for deposits. These areas are usually on the pressure side of blades. In turbine blades, these zones are usually behind the bores of the film cooling (top left diagram); in uncooled blades with high edge temperatures, the zones are narrow fields parallel to the edges (top right diagram). Oxides created by the sulfidation are often emaciated due to erosion from the gas flow and particles it carries (bottom left detail), resulting in visible indentations.
If this is the case, it must be assumed that only a very short usable life span remains, relative to the operating time that has already passed.
Figure "Degradation during operation" and Figure "Damage mechanisms" (Ref. 5.4.5-5): The operating damages from atmospheric plasma-sprayed YSZ (yttrium-stabilized zirconium oxide) thermal barrier coatings (TBC) are primarily dependent on the surface temperature of the coating.
Damages in the left part of the diagram, especially those that lead to premature failures, were traced back to manufacturing problems. Curves “1”,”2“, and “3” show the failure behavior due to dust. These dust deposits include, for example, FeO+NiO from seal wear products and MgO+ CaO, Al2O3, SiO2 from dusts ingested from outside the engine. Erosion is more pronounced at low operating temperatures.
During very long operating hours, effects occur that are based on changes in the coating structure:
A special problem which gets more and more attention is the erosion of the thermal barrier coatings (TBC). Especially concerned are obviously thermal sprayed coatings. Concerned are break-outs of small particles from the surface. Besides a deterioration of the TBC itself, the life time of the aeroengine can be unacceptable concerned. The ceramic coating particles have an erosive effect at the blading, following in the gas stream. So for example the typical thin Al-diffusion coatings (ca.0,050 mm) can be removed in few hundred operation hours. With this they offer no more an oxidation protection. The result is an accelerated oxidation of the blading which makes a premature exchange necessary.
Figure "Typical problems with thermal barrier layers" (Refs. 5.4.5-6 and 5.4.5-6): The following are typical damage mechanisms in ceramic thermal barrier coatings made from ZrO2 :
“A”: Erosion through particles and/or the gas flow. Damages to the high-pressure turbine segments are typical (housing-side seal surfaces across from the HPT rotor blade tips).
“B”: Melts of dust deposits can form in the combustion chamber and/or at the heavily heated thermal barrier coating (TBC) of the blading.For example such melts can be low melting silicates from ingested dust (Ref. 5.4.5-10). These can enter into the segmentation cracks (Fig. "Damage mechanisms") and create a bursting effect in the frozen state after cooling.
Salt melts may show this physical effect less pronounced. They shrink, for example vanadium-sulfate, during the freezing process in a manner, that dangerous high compressive stresses are not to be expected
“C”: Chemical reactions with deposits (e.g. fuel remnants). From tests can be identified, that compositions of the elements sodium, sulfur, vanadium and probably also lead and phosphor are primarily responsible for the hot gas corrosion of thermal barrier coatings. The aggressivity develops especially during oxidation of the fuel contaminations in the combustion in the combustion chamber. Thereby develop intense acid ore alkaline acting salt melts (Fig. "Sulfidation attack").
“D”: Delamination of the coating due to poor bonding (e.g. caused by manufacturing problems)
“E”: Oxidation of the contact surface, e.g. the undercoating, caused by oxygen ions being conducted through the hot ceramic coating. This failure mechanism has a relatively long incubation period (up to several 1000 operating hours), making it especially problematic. Also corrosive gases can penetrate the TBC. For example NACL vapour can let the oxidation layer faster grow at the adhesion layer. The so developing thicker oxide layer promotes the spalling.
Against corrosion ba salt melts, usually the zirconium-oxide of the TBC is more resistant as the highly oxidation resistant MCrAlY adhesion coatings. Observations bring close, that even thin ZrO2 coating remains prevent an attack of the adhesion coating by salt melts.
Also solid organic oxides can deteriorate TBCs by acting acid.
“F”: Thermal stresses from manufacture and operation can cause spalling, especially at convex radii and edges. At convex surfaces like at the suction side of turbine blades the high compression stresses in the ceramiclayer act especially flaking. At concave surfaces the contrary is the case (Ref. 5.4.5-9). Simplified the relation is applied: The stress perpendicular to the surface curvature rises direct radius of the surface curvature (i.e., the more even the curvature, the more susceptible for spallings!), the coating thickness and the stress in the coating parallel to the surface.
“G”: Superficial crumbling after longer run times in a hot gas flow. These particles can considerably reduce the life spans of diffusion-coated hot parts (such as turbine blades).
5.4.5-1 Ralf Bürgel, “Handbuch der Hochtemperatur-Werkstofftechnik”, Friedr. Vieweg & Sohn Publishing Co., 1998 ISBN 3-528-03107-7, pages 247 to 305.
5.4.5-2 “F110 Disk Crack Traced to Coating”, periodical “Aviation Week & Space Technology”. August 6, 1984, page 18.
5.4.5-3 R.A. Rudey, J.S. Grobman, “Impact of Future Fuel Properties on Aircraft Engines Fuel Systems”, AGARD-LS-96, Agard Lecture Series No. 96 “Aircraft Engine Future Fuels and Energy Conservation”, October 1978, Chapter 6, page 19.
5.4.5-4 A.J.A. Mom, J.A.M. Bogers, “Simulated Service Test Behaviour of Various Internal and External Coatings Applied on CF6-50 First Stage Turbine Blades”, Proceedings of the conference “Coatings for Gas Turbines and Other Applications”, Liège, Belgium, 6-9 October 1986, Chapter 11.
5.4.5-5 P. König, A. Rossmann, “Ratgeber für Gasturbinen-Betreiber”, Vulkan Publishing Essen, 1999, ISBN 3-8027-2545-X, pages 95 and 114.
5.4.5-6 R.L.Jones, “Some Aspects of the Hot Corrossion of Thermal Barrier Coatings”, periodical, Journal of Thermal Spray Technology“, Volume 6 (1) March 1997, pages 77 to 83.
5.4.5-7 M.M Ratwani, A.K. Koul, J-P. Immarigeon, W. Wallace, “Aging airframes and Engines”, Proceeding Paper in AGARD-CP-600 Vol. 1, of the AGARD symposium on ” Future Aerospace Technology in the Service of the Alliance“, Palaiseau, France, 14-17 April 1997, pages A18-1 to Q18-16.
5.4.5-8 K. Steffens, „Technik der Luftfahrtantriebe”, Vorlesung 2002/03 an der TH-Aachen.
5.4.5-9 D.M. Nissley, „Thermal Barrier Coating Life Modeling in Aircraft Gas Turbine Engines“, Zeitschrift „Journal of Thermal Spray Technology” Volume 6 (1), March 1997, page 91-97.
5.4.5-10 R.L.Jones, „Some Aspects of the Hot Corrosion of Thermal Barrier Coatings“, Zeit-schrift „Journal of Thermal Spray Technology” Volume 6 (1), March 1997, page 77-84.
5.4.5-11 E. Berghof-Hesselbächer, H. Echsler, P. Gawenda, M. Schorr, M. Schütze, „Time and Temperature Dependent Development of Physical Defects in Thermal Barrier Coating Sys-tems“, Zeitschrift „Praktische Metallografie”, 40 (2003) 5, Carl Hanser Verlag, München, page 210-230.