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7.1.3 Blade Tip Rubbing Damage

Rubbing damage occurs in the tribo-system itself and/or develops from the rubbing process. The prerequisite of the demand for a high degree of effectiveness over long periods of operation is that the clearance gap is minimized in all operating situations. This makes rubbing of rotor components much more likely, most of all in blade tips and the hub contour with housings and guide assemblies. Rubbing occurs with high relative movements and causes to high thermal and mechanical loads on the contact surfaces. If the maximum load levels given by the designer are exceeded, damage will occur.
The complexity of the damage sequences that can lead to rubbing is especially problematic. It is very probable that rubbing problems are not recognized in the development stage of the engine. After the engine is in serial operation, specific remedies are expensive and drawn-out (Fig. "Honeycomb seals problems", Example "Drastic performance loss caused by rubbing"). In order to create an successful remedy for rubbing damage, the damage mechanism must be correctly determined, along with all of its causes. Damage symptoms give important evidence and clues. Therefore, in this chapter, special emphasis is placed on the explanation of damage symptoms and conclusions that can be drawn from them.

Figure "Blade tip rubbing damage symptoms": This diagram shows the typical characteristics of compressor blade tip rubbing. The top diagram gives the external characteristics:

Burring and dynamic fatigue crack initiation (“1”): There are various types of burring (see Fig. "Rub coating and blade stresses"). The material in the burrs is usually damaged (loss of strength, embrittlement). Also, burrs have typical cracks from which fatigue cracks can spread into the blade leaf (Fig. "Compressor blade cracks by rubbing").

Planed material removal (“2”): This wear occurs, for example, through cutting by the coating particles. Tip wear can also be caused by plastic deformation caused by the heating and softening of the blade tip. The cutting properties of the coating vs. those of the blade tip determine which of these wear processes is the primary one. Even soft abradable coatings can wear down the blades to a certain degree if, for example, the aging process has resulted in some of their components hardening (similar to a soft-bonded grinding disk). If this process occurs in a rub-tolerant system with hard coatings, material removal will be primarily from the blade tip.

Grooves (“3”): Large grooves are usually caused by the edges of outbreaks from hard coatings or their are due to a material that has been smeared onto the coating.

Deposited material (“4”): Metallic coating material (e.g. aluminum) or blade material is smeared (deposited) on the coating or the blade tip. These deposits are created at high rubbing temperatures and can therefore be expected to bond with the base material and/or diffuse into it. This creates local anomalies in the blade tip material and its properties.
Deposits are often embrittled (oxidation) and have high tension residual stresses from the cooling process. The residual stresses and/or the micro-cracks can significantly lower the dynamic strength of the material.

Overheating zones (“5”): If the coating does not have outstanding cuttable properties, the friction heat created by rubbing will locally overheat the blade tip. The temperatures in the contact area will reach the melting points of the metallic rubbing partners. The consequences are structural changes with dynamic strength loss, embrittlement, changes in hardness, etc. (bottom diagram).

Local traces of fire (“6”): With titanium-alloy blades, rubbing temperatures can reach the ignition temperature (roughly 1500°C). The diagram shows the first indications of a titanium fire as an undercut with oxide formation and fused drops of material.

Erosion of the leading edge (“7”): Erosion of the tips of compressor rotor blades can be caused by dust, removed material, or material outbreaks from the rub coatings in the front stages.

The bottom diagram schematically illustrates the findings of a metallographic cross-cut through the typical heating zone of a rubbing area. Material changes and damage are spread out corresponding to the radial temperature gradients. Various damage zones can be defined. Naturally, the extent of these zones depends heavily on the blade material and the chronology of the temperature changes. Martensitic steels, for example, can harden at high temperatures with rapid cooling, whereas lower temperatures have a tempering effect with hardness loss. Titanium alloys tend to absorb oxygen and embrittle at high temperatures. Melted deposits, structural changes, hardness distribution, gas absorption, and alloying are examples of symptoms that enable conclusions as to the rubbing temperatures, the extent of the heated area, and the extent of damages.

Figure "Rotor spacer rings damages": This diagram depicts typical damage characteristics of a rotor spacer with a hard ceramic spray coating that rubbed against stator vane tips (top diagram).

Material removal, wear (“A”): Material removal is minimal due to the hardness of the coating. Considerable material removal could only be expected if the tip was armored (unusual in stator vanes).

Corrosion (“B”): If the base material is made of Cr steel (e.g. due to the low heat strain), then ocean atmosphere/condensation water can penetrate through micro-cracks and come into contact with it. If the bonding layer is not sufficiently thick, conditions for a corrosion attack below the coating are present. If corrosion products escape outward through cracks in the ceramic coating, they leave distinctive marks on the surface (detail).

Erosion (“C”): erosion caused by particle-loaded air flow (dust, removed material) becomes noticeable after only several hundred hours of operation. Due to the redirection of the flow, the affected area is the first spacer ring after a swan-neck canal (see Volume 1). Particles strike this area most heavily. A typical characteristic of erosion is periodic roughening on the surface of rotor blade pitch with structuring in direction of the flow.

Coating outbreaks (“D”): coating outbreaks can cause dangerous consequential damages on the blading (notches, deformation). They usually occur at the edge of a coating. This typical weak point is due to coating imperfections from poor spray conditions and gap formation, as well as differing expansion rates (thermal strain, differing stiffness) between the coating material and base metal. The first indication of the cause of insufficient bond strength is detaching of the bond layer. Discolorations of the detached surface (tarnishing, corrosion, deposits) indicate the age of the damage and operating influences.

Crack initiation (“E”): Cracks in the coating can occur during the manufacturing process (e.g. due to thermal stress). It is more likely, however, that crack initiation occurs during operation due to differences in the expansion rates of the coating and base material. Coating cracks do not necessarily mean that the operational suitability of a part is limited. Segmentation can even be a prerequisite for positive performance during operation (e.g. in ZrO2 coatings).

Outbreaks from the layer (“F”): these outbreaks can stem from crack initiation if the coating strength is insufficient. Unless the cracks are in a zone subject to heavy rubbing, they are probably due to a manufacturing error.
On the other hand, if the outbreaks are in a heavily rubbed area with considerable metallic material deposits (“I”), then they are more likely to be a result of the rubbing process. The cooling of the hot metallic deposits creates high tensile stress with crack initiation and outbreaks (“H”).

Grain outbreaks (“G”): The breaking-out of individual coating crystals near the surface (bottom detail) indicates a fatigue process in the coating. Fatigue is due to dynamic stressing of the coating through the rubbing blade tips.

Figure "Damages at housing rub coatings": Depicted are typical damage characteristics of (soft) abradable coatings such as thermally sprayed Ni-graphite coatings (top detail) and metal powder-filled synthetic resin (bottom detail). These coatings are used primarily in compressor housings.

Material removal, wear (“A”): Soft abradable coatings are intended to experience considerable material removal during rubbing. The surface of these coatings is relatively smooth at first, but after long periods of operation, it is usually roughened by erosion and becomes difficult to identify as a rubbing surface.

Erosion (“B”): Particles carried by the air flow that have been sucked in or created within the engine (e.g. material removed during rubbing) can erode away soft abradable coatings within a few hundred operating hours. The eroded surface is usually rough with a structure oriented in the direction of flow and a periodic pitch that corresponds to the stator vanes. Altered (oxidized) coatings can even be eroded by the airflow alone.

Material deposits (“C”): Metallic deposits from rotor blade tips are unusual on soft porous coatings and indicate a poor abradable tribo-system. If the coating is rubbed through to the base material or bond layer, it can result in considerable deposits with traces of high friction temperatures (tarnishing, buckling).

Impact damage (“D”): Large foreign bodies that have been sucked into the engine, as well as fragments from inside the engine (outbreaks from hard coatings, blade fragments, probe fragments), can cut large notches into the coatings and help identify the primary damage (location of the part failure).

Coating outbreaks (“E”): There are various causes of coating outbreaks, depending on operating conditions: Dynamic fatigue (sonic fatigue) due to the blade passing frequency, oscillating housings, heat strain, and strain from mechanical loads. Typical production flaws include high residual stresses and insufficient bond strength. The location of the separation (from the base material, from the bond layer, or within the coating itself) is the first indication of the cause of the outbreak.

Separation symptoms (“F”): If the coating has crack initiation parallel to the bond surface, a large outbreak can be expected. This type of crack initiation can be caused by manufacture and/or operation. A microscopic inspection of the cracked surface can yield important evidence as to the cause of the damage (e.g. “microscopic spheres”).

Coating changes (“G”): If, for example, components of a multi-phase coating (top detail: G1 = graphite, G2 = nickel particles) are removed by oxidation or reactions occur, the inner strength of the coating compound can be weakened and erosion may increase. Further problems are coating embrittlement and/or heating-up. The result is an unsuitable tribo-system.

Corrosion (“H”): Some coatings/coating components are sensitive to corrosion. A typical example is polyester resin (H3) filled with Al powder (H2). Ocean air penetrating through gaps and cracks (H1) corrodes the Al particles, creates lumps, and/or initiates cracks that cause the coating to spall.

Figure "Compressor blade cracks by rubbing" (Ref. 7.1.3-1): Similar to a violin, rubbing blades can begin to vibrate, at which point the basic flexural mode ceases. This causes fractures near the blade root platform in case of dynamic overstress (bottom diagram). It is assumed that temporary overloads cause damages without macro-crack initiation. This damage can later lead to blade failure under normal dynamic loads (Fig. "Blade tip rubbing damage symptoms"). Measurements of the dynamic loads in rubbing blades show that in a rubbing rotor blade, the vibration amplitude increases with every rotation while the blade is still making contact (bottom right). In this way, the system “vibrates itself up”. The strength of this effect depends upon the damping of the blade and the restoration of the clearance gap due to material removal from the rubbing components. The process is strengthened by the thermal lengthening of the blade due to friction heat (Fig. "why the shortest blades seem to rub first"). Integrally bladed rotors (blisk, bling) are more easily brought up to dangerous vibration levels during rubbing than conventional disks with inset blades (Example "Compressor blade cracks by rubbing"). Therefore, the rubbing process should be given special attention with blisk configurations.
Naturally, the friction-induced dynamic stressing of the blades also depends on the rubbing time, infeed force, and the length of the rubbing contact around the circumference as well as the total friction travel.
The dynamic strength loss in the blade tip due to damage following rubbing (Fig. "Blade tip rubbing damage symptoms") leads to radial crack initiation and corner fractures (Fig. "Dynamic fatigue at blade tips by rubbing") in case of high-order, high-frequency vibrations that stress this damaged area (e.g. Lyra-mode vibrations, top diagram, Example "Blade rubbing in a turboprop engine"). It is not completely clear to what extent these outbreaks are also caused by vibrations from the rubbing process, or if the dynamic fatigue only occurs later due to the normal dynamic loads on the blade (Example "Blade fracture long time after heavy rub").
Blade failure due to vibrations caused by rubbing is not limited to the compressor. This type of damage also occurs with turbine blades (Example "Optimizing of the tip clearance gaps by engine trails")
Evidently, stator vanes are less frequently affected by fatigue fractures due to rubbing. Compressor rotor spacer rings show rubbing symptoms with chatter marks that indicate vibrations in the stator vanes during rubbing (Fig. "Rub coating and blade stresses").

Example "Blade fracture long time after heavy rub" (Ref. 7.1.3-2):

Excerpt: “…the cause for a fourth-stage booster blade failure during final flight tests of the powerplant….
Once a vibration developed in the engine, power was reduced and the aircraft landed safely…with the …(engine) still running. Subsequent examination revealed that the blade in the fourth stage of the engine booster had broken.
Although it is early in the investigation, … officials believe the blade broke as a result of a “heavy rub” it received early in the flight test program. …engineers are confident the problem is a mechanical one that will be resolved relatively easily.”

Comments: This case indicates a damage sequence in which a blade was damaged by rubbing, but the (fatigue?) fracture did not occur until considerably later (Fig. "Compressor blade cracks by rubbing"). The cause of this type of damage is hard to verify; there are usually only indications.

Example "Optimizing of the tip clearance gaps by engine trails" (Ref. 7.1.3-3):

Excerpt: “….(the) turbofan has suffered a low-pressure compressor (LPC) blade failure on its last test flight….The fourth stage LPC blade failed towards the end of the 19th, and final, test flight planned for 1994. ` It wasn't an uncontained failure and, although it did damage other blades in the stage, it did not damage any other parts of the engine,' says the company.
The incident is the latest in a series of mishaps to hit the engine during its test programme, and is the third related to blade-tip clearance problems.Early mishaps included a high pressure turbine (HPT) rub, which resulted in the module being returned to …(the OEM). HPT Clearances have since been increased and drawings modified.
… The latest problem was initially caused by a (low pressure turbine) LPT rub which occurred in a recent testflight. After the incident, …(the OEM) increased clearances in the compressor and revised the design, but a blade had been stressed in the incident. …(the OEM) believes that it was the blade which probably failed on 13 April. The company says that the blade-tip rub and failure problems are an inevitable part of testing. `It's a clearance issue: we're trying to optimise the clearances and see how close you can get.'“

Comments: This large turbofan engine with an especially large bypass ration evidently has rubbing problems in the compressor and high- and low-pressure turbines. A reason for this may be the, relative to the small engine core, large fan module (backbone bending?).
It is interesting that the engine company states that the tip clearance gaps must be optimized through engine trials. This emphasizes the limited applicability even of modern computer modeling and calculating systems.

Figure "Compressor rotor damages by rubbing": Drum-type, titanium alloy compressor rotors are especially susceptible to overheating and melting if the spacer rings (bottom diagram) or intermediate racks rub (Ref. 7.1.4-11). This is due to the relatively thin walls and comparatively low heat conductivity of titanium alloys. The one-sided heating-up of the rotor hub causes the rotor to bow, and the rubbing process increases itself, along with the imbalances. This process is self-increasing. Areas that have been ground through, heavily oxidized, and have deposits fused to them are typical (Fig. "Compressor rotor perforated by rubbing"). This type of damage can increase to the point of catastrophic rotor failure in a matter of seconds. Experience has shown that smooth ceramic coatings can not safely prevent this type of damage. If a compressor rotor has pronounced rubbing traces along with oxidized material deposits (top diagram) and possible outbreaks of the ceramic coating (top right detail), an intense inspection of the extent of damage is necessary. The inside of the rotor drum (this usually requires disassembly) must be inspected for signs of overheating (tarnishing, oxidation, crack initiation, loss of hardness, etc.) The operating suitability must then be determined in accordance with specifications.

Figure "Compressor rotor perforated by rubbing": After intensive rubbing caused by rotor bow and imbalances, these damage symptoms were found upon disassembly. This is a drum rotor made from the titanium alloy TiAl6V4 with an Al2O3 spray coating. The typical damage characteristics are:

“A”: Hole burned-through the rotor drum in the rubbing zone. Oxidized areas with yellow and white oxides are typical.

“B”: Deposits of stator vane material in the rubbing zone.

“C”: A coating on the drum rotor hub.

“D”: Damaged and burned (ignition of a local titanium fire, Ill. 9.1.2-2) stator vane edges (material TiAl6V4).

Figure "Rub coating damage mechanism on rotors": This example is a coating outbreak on the rotor spacer ring of a compressor rotor. Typical damage symptoms (top diagram) in an area of intensive rubbing with material deposits.

A”: Rubbing compressor stator vane with heating-up zone.

B”: Blade material deposits on the ceramic rub-tolerant hub coating.

C”: Crack initiation in the metal coating, which has been embrittled by oxidation and had material deposited on it. These cracks run primarily in an axial direction corresponding to the induced residual stresses (see Fig. "Spalling at spacer coatings").

D”: Large surface outbreak from the coating. Depending on the weak points, the fracture lies in the transition zone to the bond layer or the base material. Separations within the rub-tolerant coating are typical.

The radial thermal strain of the blade during rotation while rubbing resulted in layers being deposited (smeared; bottom diagram). These thin layers oxidize at the high rubbing temperatures and become brittle.

Figure "Spalling at spacer coatings": The following explanation is given for coating outbreaks in rubbing zones with material deposits (top diagrams):
Stator vane material (“A”) is melted or becomes doughy during rubbing and combines with the hot coating. Simultaneously, the rub coating becomes very hot. It expands and high internal compressive stress is created. This causes crack initiation and weakens the coating`s bond layer (“B”). During cooling, high tensile residual stress (“C”) forms in the brittle deposits and transfer themselves into the rub coating, which becomes detached in the area (“D”) that was weakened during phase “B”.
The bottom diagram shows typical results of large outbreaks from a ceramic rub-tolerant coating on a spacer ring in the rear compressor region (filigreed blading).
During the separating process, heavy local rubbing occurred that damaged (strength loss, crack initiation) the stator vane tip. If the damage is especially pronounced, then the guide vane is plastically bowed and begins to oscillate due to the flow displacement. In some cases, unusually high dynamic loads cause dynamic crack initiation and blade failure. If coating fragments strike the leading edges of the blades of the following stator assembly, it is possible that the notches this creates cause fatigue fractures.

Figure "Ceramic rub coating damage at turbine seals": The high-pressure turbine has the tightest flow cross-section in the engine with especially demanding operating conditions such as high gas temperature and gas pressure. Even very small changes in clearance gap size in this area have a strong effect on engine performance. In order to minimize the gap between the high-pressure turbine rotor blade tips and the housing in modern engines, assemblies that radially infeed the side of the housing are used (top diagram, Fig. "Ceramic rub coatings"). This demands segmentation of the turbine ring in the housing. The extreme temperature gradients heavily stress the rub coatings on the segments.
Even the thermal spraying process during manufacture must be optimized in a comprehensive development process (pre-treatment of the bond layers, thermal conductivity, porosity/structure, coating synthesis, etc.). It is important to avoid dangerous residual stresses, which can add to the operating loads and cause the coating to delaminate.
The segments of turbine rings must be well sealed in order to avoid hot-gas leaks, which could damage other components (e.g. overheating the housing). Radial gaps between the segments should be as small as possible. On the other hand, bridging of these gaps can result in high stress levels in the brittle ceramic coatings and cause them to break out. Especially susceptible to this are corners (bottom diagram) that are already subject to high stress levels from the thermal stress between the ceramic layer and the metal substrate.
Ceramic rub-tolerant coatings must be relatively hard in order to withstand the high erosion stress from the gas flow and the particles it carries (e.g. coke from the combustion chamber; see Fig. "Ceramic rub coating damage at turbine rings").
The high hardness lowers abradability and necessitates correspondingly hard blade tips. Therefore, shroudless blades are armored with particles of hard material (Fig. "Optimizing blade tip coatings" ). If uncoated blade tips undergo heavy rubbing, the blade material may smear (deposit) onto the opposing surface. Mechanical overstressing (bending loads, vibrations) of the blade during rubbing can lead to extensive damages (Fig. "Rotor spacer rings damages").

Figure "Cracks in thermal insulation coatings": The top left diagrams show typical characteristics and damage symptoms of ceramic spray coatings on turbine segments (see also Fig. "Crack initiation at thermal insulation coatings"). The most strongly affected coatings are zirconium oxide thermal barriers on metallic substrates.

A”: Large spalling in the ceramic coating. There is a pattern of dark lines that follows the former segmentation cracks (see Fig. "Turbine ring segments damages"). Diagram “1” at top right depicts the probable damage sequence:
The ceramic coating is struck by hot gas and heats up to a high degree due to its thermal insulating effect. High compressive stress is created in the ceramic coating and, around the segment cracks, this stress runs into the bond layer and then curves back into the ceramic coating. This causes the coating to spall from the substrate (Fig. "Crack initiation at thermal insulation coatings"). Discolorations along the former segmentation cracks are in a line pattern.
Probable cause: insufficient bond strength in the ceramic coating. This is related to unoptimized manufacturing processes and oxidation of the substrate.

B”: “Giraffe skin”. Around the segmentation cracks in a ceramic coating with large spalling, there remain small raised lines of coating remnants. These appear as light lines against the dark oxidized metal surface. The most probable damage mechanism is depicted in diagram “2”:
The ceramic coating is struck by hot gas and heats up to a high degree due to its thermal insulating effect. High compressive stress is created in the ceramic coating, and around the segment cracks, this stress runs into the bond layer and then curves back into the ceramic coating. The segments spall along the bond layer. Raised ridges of the ceramic coating remain along the former segmentation cracks. This process is accelerated if the bond strength of the coating has been weakened by oxidation (see Chapter 5.4.5).
Probable cause: High compressive residual stress in the ceramic coating. This can be causally related with poor manufacturing procedures and/or rapid heating-up to high operating temperatures.

C”: Individual light-colored islands are all that remain of the coating where it spalled from the metallic bond layer/base material.

D”: Corner outbreaks. The corners of the ceramic coating have completely separated from the metallic base material/bond layer. A fissure between the ceramic coating and substrate can often be seen under a microscope.
Probable cause: Problems during thermal spraying, which lower the bond strength of the ceramic coating. These include high residual stress due to unfavorable temperature changes, “pellet effect” from impacts, oxidation of the bond surfaces, and poor edge machining.

E”: Spalling of the corners. Cracks run into the ceramic coating. The structure of the cracks indicates a crack that spreads out from the corner and/or the narrow edge of the segment. This usually affects the corners of adjacent segments opposite the crack.
Probable cause: Bridging of the gap between segments. If the coating edges come into contact with one another, powerful forces between the segments are created.

Corners tend to spall (see “C”). In References 7.1.3-6 and 7.1.3-7 (bottom diagrams), calculations show that corners are especially heavily stressed by residual stress and operating stress at manufacturing and/or operating temperatures. The shearing stresses this creates drop sharply inwards along the bond surface. This explains why spalling is often limited to corners, even after long operating times, i.e. many operating cycles.

Figure "Crack initiation at thermal insulation coatings" (Ref. 7.1.3-12): When one inspects crack growth in a ceramic thermal insulation coating under two-axis stress loads, the following typical damage mechanisms can be observed:
Under compressive stress (top diagram), such as occurs during rapid heating-up (e.g. engine start-up) or an internally cooled wall (e.g. cooled hot parts such as turbine blades or combustion chamber walls), the coating compresses and bulges. In case of overloads, this leads to spontaneous spalling along the bond surface or crack growth followed by spalling. It is natural that this cracking process is promoted by convex curved surfaces such as, for example, the leading edges of turbine blades that are subject to large thermal loads.
If the hot, coated wall cools rapidly from the coating side (e.g. after engine shut-down), shear stress builds up in the thermal insulation layer (bottom diagram). This causes crack initiation perpendicular to the layer and crack growth parallel to the layer. This growth can occur at the transition to the bond layer (adhesive delamination) or in the thermal insulation layer (cohesive delamination). These different crack processes can be observed in connection with typical damage symptoms of substrate surfaces where the coating has spalled off (see Fig. "Cracks in thermal insulation coatings")

Figure "Ceramic rub coating damage at turbine rings": Ceramic coatings of turbine ring segments are vulnerable to damage from particles in the gas flow. The relatively hard coatings are noticeably eroded over the long operating times of modern civilian aircraft engines (bottom diagram). Clearance gap expansion of several tenths of a millimeter can occur after only a few thousand hours of operation. The structure/roughness of the eroded surface indicates the direction of the stress.
The erosion process is promoted by other factors. Sand particles can be melted in the combustion chamber and react with the ceramic coating. This damages it and lowers its resistance to erosion.
Sand escaping from the rotor blade can damage the coating above the dust-removal bore (top diagram). One damage mechanism is based on the melted sand entering into the segmentation gaps of the ceramic coating, which are necessary to compensate for the strain differences between the substrate and coating. If these gaps are blocked by the fused sand after cooling, strong compressive forces are created in the ceramic rub-tolerant coating. In extreme cases, large sections of coating spall and separate.

Example "Drastic performance loss caused by rubbing" (Fig. "Honeycomb seals problems", Ref. 7.1.3-5):

Excerpt: ”…(the OEM) has made substantial design changes to its …engine after numerous…operators complained to the engine's lack of staying power.
…(the customer told of).a “complete” engine failure of one engine in flight. The engine, which had less than 500 hours of use, had “extensive turbine damage.”
As a result of these most recent complaints from operators about the …(engine type's) performance, coupled with the results of its own factory tests, ..(the OEM) will upgrade the materials in the stage one and stage two turbine blades, the stage two turbine nozzle, and will replace the honeycomb Bradeloy shroud with a solid metallic shroud made of Conicraly, a multi-metal composition
Turbine blade tips rub up against the shroud, maintaining a clearance. If the shroud deteriorates, as the honeycomb variety has been known to do, a gap is formed between the blade tip and the shroud, causing the engine to lose performance. To alleviate the problem, ..(the OEM) replaced the shroud with a solid material.
…(the OEM) says the new material is harder and less susceptible to erosion and or damage by heat. Consequently the operating life of the engine should increase.

Comments: This example documents the consequences of the high operating loads on high-pressure turbine turbine rings. Erosion and high temperature damage can lead to drastic performance/efficiency losses, especially in relatively small engines that are extremely sensitive to clearance gap changes. The switch from a honeycomb seal to a massive metallic rub-coating attests to the known oxidation problems in honeycomb seals (Fig. "Oxidation of seals coatings" ).

Figure "Honeycomb seals problems" (Ref. 7.1.3-5, Example "Drastic performance loss caused by rubbing"):
If proven engines are used in a new application with different operating conditions (e.g. from a helicopter to a propeller turbine, top diagram), it can result in unexpected damage to “proven parts”. If, for example, the operating times between overhauls in civilian operation are increased, or relatively high output levels are demanded over long periods of time, the “wear” of the blade tip seals in the hot part area can become a deciding factor for the engine`s operating life span.
The thin metal ridges of honeycomb seals are weakened by oxidation (bottom left diagram) until they break off and increase the tip clearance gap in the high-pressure turbine to impermissibly large size (bottom right diagram).

Figure "Oxidation of seals coatings" (Refs. 7.1.3-10 and 7.1.3-11): The right diagram shows the life span, i.e. the decrease in tensile strength, of typical honeycomb materials (Hastelloy X und Haynes 188). The limit of long-term operation for these materials is at temperatures around 700°C. FeNiCrAlY-materials can be used at operating temperatures over 1000 °C. The left diagram (Ref. 7.1.3-9) shows the increase of erosion-induced material removal from a honeycomb wall with an Al-diffusion coating. In oxidizing atmospheres, this coating forms a protective Al2O3 layer at the surface, which minimizes carburization. The Al-diffusion coating changes thin honeycomb walls into an intermetallic phase that is highly resistant to oxidation but ovious not to particleerosion. This good oxidation behavior has proven itself in practice. Fears that the function of the honeycomb seals during rubbing would be considerably worsened (due to embrittlement) were unfounded.
Even in honeycomb walls made of MCrAlY-materials, considerable improvements can be expected. In this case, problems are more likely to arise from large-scale availability and the workability of such thin metal sheets.

Figure "Honeycomb seals oxidation": The thin ridges of the honeycomb seals are usually made of the alloy Hastelloy X. They are susceptible to oxidation at the high operating temperatures and long operating times in modern engines. A crack-like attack along the grain boundary starts from the surface, also causing plane emaciation. Pockets of oxidation are created through internal oxidation (Ref. 7.1.3-8). “Carburization”, if present, promotes the oxidation process. This occurs when sufficient carbon, in gaseous compounds or as a solid material, comes into contact with the reactive metal surface. This causes the honeycombs to break out and leads to considerable leak losses (Example "Drastic performance loss caused by rubbing", Fig. "Honeycomb seals problems").
The bottom diagrams show a typical design of a honeycomb seal. The honeycomb is made up of thin molded metal strips connected with spot welds. When the honeycomb is soldered on, the (high-temperature) solder is sucked between the honeycomb walls due to capillary attraction. Because the solder is more susceptible to oxidation than the honeycomb, it suffers increased oxidation damage (bottom left diagram).
If the new honeycomb seals are pre-oxidized in an oven for several hours before being used in operation, the incubation period is lengthened considerably and the life span increased.

Figure "Filled honeycomb seals problems": Filling a honeycomb structure with metallic or ceramic materials strongly affects its behavior:

  • Thermal insulation of the substrate from rapid heating-up during unsteady operation and reduction of cooling air use during steady operation.
  • Changed rubbing behavior and minimized material removal. However, the increased wear resistance may damage the blade tip.
  • Delays oxidation damage
  • Improves erosion resistance

The honeycomb structure is usually filled after it has been soldered onto the substrate. Filling can be conducted in various ways:

  • Thermal spraying
  • Sintering loose powder

These process have their own specific problems. Separation of the sinter- or spray powders before processing is especially problematic. Vibrations and shocks during handling, transport, and storage of the powder containers causes the powders to separate according to grain size and specific weight. Powder mixtures with various components are especially vulnerable to separation. The more a powder tends to trickle, the more pronounced the separation is. Spitting powders are therefore less likely to separate.
Due to risk of separation, it is especially important that the powder is mixed well in the containers immediately before processing (top right diagram). If this is not done sufficiently, the honeycomb filling may tend to create cavities in segregated areas and result in an impermissible structure (middle diagram). Concentration of a low-melting powder component in a large area promotes local deposits during sintering of the filling (top left diagram).

Evidently, thermal spraying does not react to separation of the powder so strongly that mistakes could be detected during application. Therefore, it is more likely that damaged parts would be put into operation unnoticed and only later come to attention due to their unsuitability for operation. This would leave only expensive, time-consuming corrective measures.

More frequent than damages due to separation are problems due to the thermal spraying process during filling of the honeycomb structure. These problems are connected with the spray jet and bouncing particles (bottom diagrams). If the honeycomb walls are disturbing the spray jet, it leads to poor bonds between the filling and the honeycomb, as well as typical porosity. Inhomogeneous application of the filling also promotes crack initiation.

Figure "Turbine ring segments damages": There are two coating technologies that are primarily used on ring segments in high-pressure turbines: ceramic-based thermal spray coatings and filled honeycomb.

Ceramic coatings: The thermal insulation behavior differs strongly from that of the metallic substrate. The desired low heat conductivity has the drawback that rapid heating-up results in the creation of large temperature gradients in the coating as well as high surface temperatures. The temperatures during manufacture create residual stresses, which combine with operating stresses. The coating must have a network of segmentation gaps in order to absorb the different expansions of the thermal insulation coating and the substrate (top left diagram). These can compensate for differing expansion rates by widening and narrowing. If the “coating islands” this creates spall off, typical damage symptoms can be observed (top right diagram, compare Fig. "Blade tip rubbing damage symptoms"0).

Filled honeycomb also has a thermal insulation effect. It is more resistant to large-scale surface outbreaks than ceramic coatings. The weak points of filled honeycomb are oxidation of the honeycomb walls (Ref. 7.1.3-9) and the problem area at the transition between the filling and honeycomb. If the fillings break out, there is danger of local overheating of the segment base material with crack initiation (bottom left diagram).

Example "Blade rubbing in a turboprop engine" (Ref. 7.1.3-4, Fig. "Compressor blades rubbing control"):

Excerpt: ”…(The OEM) is making additional modifications to the second stage of its …turboprop engines following a compressor blade failure…(the aircraft type was)… involved in three incidents in a four-day period….
The Nov. 3 incident… occurred during a landing approach on a training flight…, and apparently stemmed from a blade failure in the second stage.
Three of the …incidents occurred on the ground- two during taxi and one in the early stages of the takeoff roll as the engine reached full takeoff power.
One of the … blade failures came during an approach to landing.
Following the first three blade failures, the…(OEM) instituted a fix that increased the compressor blade clearance by machining material from the engine casing. The objective was to eliminate blade scrubbing as a source of blade stress, a company official said.
Since the engines that sustained blade failures on the …aircraft Oct. 11 and on the …aircraft Nov. 3 had been fitted with machined cases for increased blade clearance, …(the OEM) has had to implement other fixes aimed at improving the vibratory stress margins in the second stage blades.
The failures have occurred at various power settings, but the maximum aerodynamic excitation effect on the blades seems to be occurring at off-design intermediate settings in the 80-90% range. The fixes, which have been designed to have the most effect on that opening point, include:

  • Reducing the blade angle of incidence by about 2 deg. at the tip to ease aerodynamically induced vibration on them.
  • Shot-peening all second stage blades to reduce residual material stresses in the blades and tailor the metal to withstand higher levels of vibration.
  • Changing the blade anti-corrosion coating from one that is put on by the deposition method to a painted coating.

The shift in the method of coating the blades was necessary because it was found that the deposition process of applying the anti-corrosion material negated the conditioning effects of the shot-peening treatment…“

Comments: Evidently the compressor blades fail due to high dynamic loads. One is given the impression that the rubbing of the rotor blades was an important contributor to the damage, but was not the only cause. This is supported by the fact that further damage occurred even after the clearance increases. The measures taken on the blades themselves are meant only to increase the dynamic strength of the affected blading. This indicates a certain “helplessness” towards the damage causes.
It could be speculated that housing distortion may have created an uneven circumferential clearance gap and therefore caused rotor and blade vibrations (Figs. "Tip clearance change after shut-down" and "Compressor blade cracks by rubbing"). Housing distortion could be explained by the relatively small engine with the relatively large gear box it supports. This could cause vibrations and/or shearing force transmitted into the engine to ovalize the housing.

Figure "Compressor blades rubbing control" (Ref. 7.1.3-4): Example "Blade rubbing in a turboprop engine" comments on the contents of the diagram. It concerns a relatively small shaft-power engine, another version of which is widely used as a helicopter power plant (Ref. 7.1.3-5), but which has never experienced comparable problems when used in helicopters.

Figure "Dynamic fatigue at blade tips by rubbing": Rubbing can cause corners of blade tips to vibrate under high loads (Fig. "Compressor blade cracks by rubbing"). The risk of a fatigue fracture that originates in the blade tip increases with the amount of damage in the corners. Ridges have been shown to lower dynamic strength considerably. The damage-mechanical assessment by H. Huff (top diagram) gives an idea of the effect that crack/notch depth has on dynamic strength in the burred area. This assumes a threshold stress intensity (Kth), under which no crack growth can be expected. This value was derived from the formula in the bottom diagram. “a” is the crack length under which no further crack growth occurs. One can also recognize the effect that has been observed in practice, where Ti- and Ni-alloys, which normally have similar dynamic strengths, have considerably different degrees of sensitivity once cracks have formed. Titanium alloy blades are much more sensitive to crack initiation after burr formation than nickel alloy ones. For example, if a crack is only about 100 m deep (0.1 mm), Ni-alloys have twice the dynamic strength of titanium alloys. This shows that even relatively small burrs can be very dangerous. Therefore, it is important to give special attention to the blade, not only concerning the base material, but also the geometric shape and any possible coatings (armoring) that may be applied to it.

Figure "Rubbing overstressing due to blade design": Rubbing of a rotor blade can be subject to a dangerous self-increasing effect: if rubbing occurs with an extremely twisted blade with a long chord length, such as is characteristic for “wide chord fans”, the leading edge is affected by powerful “untwisting” loads. The blade is subject to a powerful untwisting moment. This causes the leading edge to tilt against the direction of rotation and axially forward. In fan blades, this deformation increases the radial extension of the leading edge by centimeters. This increases the rubbing process at the leading edge and further untwists the blade. This untwisting tilts the trailing edge in direction of rotation and also increases rubbing. However, the rubbing forces at the trailing edge act against this motion and are not self-increasing. If the abradable coating is not sufficiently thick to absorb this radial movement, the blade tip rams the metallic housing wall (top diagram). This results in extreme loads on both the housing and blade with the potential for catastrophic damages. The axial movement of the blade forward requires sufficiently wide coatings (bottom right diagram).

Figure "Rubbing traces indicate influences": The self-increasing rubbing process described in Fig. "Rubbing overstressing due to blade design" causes typical damage (wear) symptoms (top diagram. If the increased rubbing process at the blade edges is continuous, it causes pronounced “scalloping” (arrow).
The shape and distribution of the wear zone also depends on the constructive stiffness distribution of the housing. The lower housing has axial flange. This causes centrally asymmetrical deformation due to operating influences. The greater stiffness in this zone minimizes expansion and causes ovalization. The smaller local diameter (waist) increases rubbing during clearance gap bridging. Heat strain can cause the housing wall to cave in at the waist and increase rubbing accordingly.

Figure "Elastomer rubbing coatings": In low-pressure compressors, boosters, and in the fan, filled elastomers are used as rub coatings. They are typically used on housings opposite rotor blade tips and compressor stator vane shrouds opposite labyrinth tips (of the intermediate stage seals) on the hub. The consistency of these coatings is similar to that of a gum eraser. These coatings are sensitive to thermal aging (the same is true for epoxide- and polyester based filled synthetic resins; see Fig. "Resin rub-coatings technology") and aggressive media (e.g. improper solvents and cleaners, degreasers, fuel). These influences can lead to shrinking, separation, embrittelement, hardening, crack initiation, and chalking (left diagram). For this reason, considerable experience is required in order to properly use these coatings over long operating times.
If the bond layer is insufficient, such as with casing treatments with deep circumferential grooves running through the coating, the elastomer coating may separate. In this case, the forces of the air flow and rubbing are damaging in combination with weakening of the bond strength due to aging.
For this reason, the elastomer coatings should also be secured by form fitting (right diagram).

Figure "Rubbing behaviour of elastomer coatings": Elastomer rub-coatings are subject to specific damage mechanisms. In case the blade edge does not have sufficient cutting capacity, the desired chipping of the coating, as shown in the top diagram, does not occur.
Instead, buckling, shear, bulging, and local separation of the coating occur (middle diagram).
In the final stage, the damaged coating separates in large sections around the circumference (bottom diagram). This process can heavily stress the blades and result in extensive blade damage.

References

7.1.3-1 B.Tudes, “Rub Induced Rotor/Stator Vibration Analysis on CF700 engine”, National Research Council Canada, Institute for Mechanical Engineering, Technical Report 1991/01, TR-ENG-007, pages 1 - 71.

7.1.3-2 periodical “Aviation Week & Space Technology”, April 25, 1994, page 15 “GE90 Blade Failure Cause Eyed”.

7.1.3-3 G. Norris, “Blade failure setback hits GE90 testing ”, periodical “Flight International “27. April - 3. May 1994, page 5.

7.1.3-4 “GE Modifying CT7 Second Stage Following Operational Failure”, periodical “Aviation Week & Space Technology”, November 12, 1984 page 32.

7.1.3-5 “Power degradation target of CT7 improvements”, periodical”Air Transport World” 3 88, page 108.

7.1.3-6 H.A.Nied, “Edge Stress Concentrations in Layered Ceramic-Metal Composites Due to Thermal Mismatch”, Proceedings of the Fourth National Spray Conference, Pittsburgh, PA, USA, 4-10 May 1991, pages 315-322.

7.1.3-7 J.D. Lee, H.Y. Ra, K.T. Hong, S.K. Hur, “Analysis of deposition phenomena and residual stress in plasma spray coatings”, periodical “Surface and Coatings Technology”, 56 (1992) pages 27-37.

7.1.3-8 HAYNES Co. specifications for HAYNES Alloy No. 214.

7.1.3-9 R.C. Bill, L.T. Shiembob, “Some Considerations of the Performance of Two Honeycomb Gas-Path Seal Material Systems”, periodical “Lubrication Engineering”, April 1981, pages 209-216.

7.1.3-10 G.P. Jarrabet, L. Lu, “Low Leakage Fiber Metal Seals”, ASME-Paper 92-GT-141 of the International Gas Turbine and Aeroengine Congress and Exposition, Cologne, Germany, June 1-4, 1992, pages 1-6.

7.1.3-11 R.P. Tolokan, M.S.Beaton, “Fiber Metal Abradable Seals”, ASME-Paper 84-GT-67, pages 1-7.

7.1.3-12 U.Rettig, U. Bast, D. Steiner, M. Oechsner, “Characterization of Fatigue Mechanisms of Thermal Barrier Coatings by a Novel Laser-Based Test”, Journal of Engineering for Gas Turbines and Power, April 1999, Vol 121, page 260.

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