Fretting damage can be found in many different engine parts (see
Figure "Fretting threatened engine parts" (Ref. 6.2-12): Engine components that are typically affected by fretting (compare with Fig. "Fretting zones in an aeroengine"). These indicate the multifariousness and frequency of this problem.
Fan blades with short chord lengths are used in older engine types. These blades are strengthened by clappers and snappers in the top half of the blade (“1”). The surface pressure on the clappers during impact fretting is between 7 and 70 MPa, which can cause fatigue fractures with serious consequential damages. This oscillation stress can cause the clapper plug or blade leaf to fail.
The same is true for turbine rotor blade shrouds. The slim blades of the low-pressure turbine that support one another with shrouds are especially sensitive (“2”). Wear occurs at the lay-on surfaces with the most tension (Fig. "Fretting wear at turbine blade shrouds") and can therefore lead to a drop in tension and increased oscillation stress. Worn out shroud assemblies are susceptible to the shrouds gliding over one another (shingling), which can cause blade failures.
The classic fretting damage zone is in the lay-on surfaces of the compressor and turbine rotor blades (“3”) and the corresponding disk slots ( ).
Fretting can also be observed in the lay-on surfaces of guide vanes. Metal removal in the housing (Fig. "Fretting wear at compressor guide vanes") and the mutual contact surfaces of the blades (“4”) lead to a change in the blade angle (danger of a worsening of compressor performance, stalls, oscillations) and in extreme cases the blade breaking out of the housing slot (Fig. "Fretting wear at compressor guide vanes").
To prevent axial travel, rotor blades must be secured in the disk slots. If this is done with metal sheets, wire rings, or labyrinth mounts (“5”), the lay-on surfaces at the edge of the ring as well as the front of the blade can be affected by fretting and cracks (Fig. "Fretting by vibration of blade fastener").
Threaded connections (“6”) can be subjected to fretting wear on all lay-on surfaces, depending on stiffness. This can be especially dangerous if the highly dynamically stressed screw shaft is damaged by fretting (e.g. in fitted bolts) or if fretting occurs in the highly stressed disk bore and creates an dynamic crack (Fig. "Fretting damage at turbine rotor").
Cushioning damping wires (damping rings ) are used to avoid high-frequency oscillations. These are inserted into an internal slot of the rotating part, and the centrifugal force provides sufficient pressure. A further type (e.g. in axial turbine disks of turbochargers and older industrial gas turbine types) of friction damping is drawing a slitted or segmented damping wire through the blade leaf. The intended friction between the damping wire and the applied part leads to strong fretting in the slot, i.e. the bore and its ring. This stress should be taken into account by the designer.
Threaded flange behaves similarly to threaded connections (“7”, ). Insufficient stiffness leads to pronounced friction wear, loosening, and fatigue fractures.
Multiple splinings are used to connect rotor shafts and as plug connections for auxiliary drives (Ills. 6.2-19 and 6.2-20). If the splining is not sufficiently lubricated and/or the connection is poor (e.g. misaligned), it will result in considerable wear. In extreme cases there is a complete failure of the momentum transfer - the connnection “runs through”.
Heavy fretting is common in roller bearing seats (“9”, ) with insufficient stiffness, poor fastenings, or a oil-damped snap-through bearing suspension. Configurations in which the bearing ring (steel) is seated directly in a titanium surface are especially prone to wear.
Combustion chambers and gas baffles (“10”) are usually subject to strong vibrations and relative movements caused by thermal strain. This puts so much wear stress on the plug connections (Fig. "Fretting at combustion chambers"), that the amount of metal removal becomes a deciding factor in the life span of the part.
Clamp fastenings can wear through and break pipes if they are loose or the material combination is poorly chosen (“11”).
Guide vane adjusters (“12”) have a multitude of constantly lubricated joints and bearings that vibrate strongly during operation and/or make small adjustment movemnts. Because these parts are outside the engine, they are prone to corrosion. These conditions promote fretting.
Figure "Failed rotor blade inside the engine": In an older, single-engine tactical aircraft, a rotor blade in the front compressor stage suffered a fatigue fracture. The crack initiated in the fretting zone in the blade root (details “A” and “B”). The material, a type of 13% Cr steel, is relatively resistant to dynamic strength losses due to fretting, but the damage occured due to other factors that promoted it:
Unfavorable engine operation. In all likelihood, there was a long idle during the start phase. Additionally, the crack inspection during an overhaul was not optimal, making recognition of initial cracks difficult.
In similar cases, signs of improper shot peening (the angle was too shallow) during overhaul were observed. During shot peening, crack-like overlapping was created with insufficient hardening (lower compressive stress and less peen hardening)
It is interesting to note that the primary, hook-shaped bent (detail “B”) blade fragment was caught in the following guide apparatus in many cases, preventing it from traveling through the engine and minimizing consequential damages.
The size of the remaining fractured surface “FG” in relation to the fatigue fractured surfaces “FS” corresponds to the static stress (mean stress of centrifugal force and gas flexural stress) of the blade root. If the remaining surface is relatively small, then the static stress was low. This increases the chance of success of remedies that optimize the tribo-system.
The size of the remaining fractured surface in rotor blades is relatively small (especially in titanium alloys). This can be explained by the designed low levels of static stress (Fig. "Fretting surface cracks at contact") and must take fretting damage into account.
Figure "Fretting influence seen at crack types": A typical weak point of compressor rotor blades are the lay-on surfaces and the areas near a dovetail root in a disk slot (Ref. 6.2-12). High-frequency blade oscillations (“A”) and/or low-frequency expansions with centrifugal force changes (“B”) can create micro-movements with typical wear signs, so-called Fretting wear (oscillation friction wear, fretting corrosion, etc.). This wear is especially dangerous because it weakens the blade material. It promotes fretting fatigue with crack initiation (arrows) and/or an increase in stress levels (equivalent stress during thrust and normal stress, see Fig. "Fretting load parameters in dovetail") in the areas surrounding the contact surface that do not seem directly affected (arrow).
Unusual fretting types can indicate unusual operating conditions. These wear images give important clues as to the damage causes and thus help in designing remedies.
A sign of the damage intensity is the length of the wear tracks (a) in the micro-region. The total width of the wear track must not be confused with the movement amplitude. It can be discerned by microscopically analyzing the damaged surface.
The location of fatigue cracks in relation to the fretting zone gives clues as to the damaging of the surface, stress, and load distibution:
Figure "Fretting surface cracks at contact" (Ref. 6.2-6): Under relatively low bending loads in a blade root, unexpanding micro-cracks of about 0.1 mm (“1” detail) can form in the wear surfaces. These micro-cracks are often observed in practice (see Ref. 6.2-1). The crack expansion depends on the stress levels as determined by design and centrifugal force. The diagram (the ordinate is the bending amplitude, the abscissa is the static centrifugal stress, i.e. mean stress) is based on calculated estimates and spot tests. It shows the capability for expansion of a 0.05 mm crack in a 13% Cr steel blade with typical root geometry. It is shown that micro-cracks initiate at low bending fretting loads before they expand into macro-cracks (detail “2”). Evidently no macro-cracks occur in the region between the two curves.
This insight also corresponds to new thoughts about the dynamic strength-increasing effect of pressurized residual stresses near the surface (Ref. 6.2-7), for example those achieved through shot peening.
“…the non-expanding cracks have been known for a long time…the crack closed as force decreased but before it reached zero. In this manner, the crack only opened again when the force was considerably above zero during following small amplitudes. This also means, however, that the effective stress amplitude or…fracture mechanical…the effective stress intensity factor was minimized as the most important parameter for the dynamic strength…The cause of the crack closing is the residual stress created by the treating process.”
This also explains the observation that overlappings of the type created shot peening at a low angle do not always lead to dangerous crack expansion. For the same reason, damaged blade roots can be regenerated through shot peening, even if micro-cracks (that could not be found during serial non-destructive inspection) cannot be removed.
Understandably, the more a material tends to hardening and creation of pressurized residual stresses, the greater the dynamic strength-increasing effect is. In line with experience, steels behave more advantageously than titanium alloys (Fig. "Fretting at titanium alloys").
If the operating loads in the wear zone (temperature and stress) are large enough that the residual stresses are noticeably reduced by creeping effects over operating time, the protective influence of hardening is reduced accordingly. This is followed by the micro-cracks growing. For this reason, it may be important to renew the hardening at specific time intervals (maintenance intervals), which is common practice. It is important that the intervals are sufficiently short. Comparable residual stress measurements during the development stage can assist in determining the interval length.
Interestingly, the usable dynamic strength is not dependant upon centrifugal stress at levels above those characteristic for the configuration and the tribo-system (in this case, roughly 100 N/mm2).
From this it can be concluded that dangerous crack expansion in the fretting zone requires limiting of the bending loads. This relatively low value depends on the individual design.
If the centrifugal stress in the blade root is relatively high, the blade root loads must be lowered considerably below the critical value to avoid fractures due to fretting (see also Fig. "Failed rotor blade inside the engine").
Figure "Fretting by vibration of blade fastener": Rotor blades in compressors and turbines are affected by axial forces consisting of gas loads, disk vibrations, thermal distortion, and misaligned lay-on surfaces (see Fig. "'Blade walking' by axial forces"). Therefore, the blades must be fixed axially in both directions. This results in limited areas of the blade laying against fasteners (e.g. rings, spacers, labyrinth mounts, disk annulus), at least for a short time. In this case, differences in stiffness will result in fretting movements. These movements can be caused by oscillations in the blade and/or the surface it is in contact with. The diagram depicts a case in which vibrations of a labyrinth mount (top left diagram) led to fretting wear in the area laying against the blade root (top right diagram). This resulted in sharp-edged depressions forming in the labyrinth mount, and the stress concentration in these combined with the oscillation stress initiated cracks (bottom diagrams). A microscopic analysis of the fretting marks on the blade root indicates the oscillation amplitude (“S2”) and thus gives important clues as to the nature of the oscillations (size, form) in the labyrinth mount. This is an important requirement for the designing of remedies and calculation of risk assessment.
Example "Design changes due to cracks in compressor blades and disks" (Ref. 6.2-11): “…cracks were discovered in the first-stage compressor blades and disks. The (owner) placed a 200-hour flying life limit on the disks and blades while the problem was investigated. (The OEM) believes that the cracks result from a design fault which imposes abnormally high stress on blade roots.
The company is redesigning the disk and blade interface to reduce the blade root crushing stress.
Comment: Cracks forming in the lay-on surfaces of the disk and blade after such a short time is unusual. Usually, the cracks initiate in one of the components and cause damages before other components are cracked. The fast crack formation definitely indicates especially high stress in the entire root region. It is odd that this crack formation was misjudged or simply not found during the obligatory trial runs during certification of the engine before it was cleared for serial production. A design change in the blades and disks means that the OEM believes there to be a design flaw, which has grave consequences for the project.
Figure "'Blade walking' by axial forces": Damages of the type in Fig. "Fretting surface cracks at contact" show that, under certain conditions, turbine rotor blades can travel against the flow (bladewalking) or, as the case may be, the pressure loss in the disk slot (in axial direction). In turbine blades, this movement is against the gas flow (forwards, see also Fig. "Coefficient of friction during relative motion"), in compressor blades in direction of the gas flow (rearwards)
The diagrams are of two hypotheses that explain this phenomenon. It may be that both of these possibilities act in combination. In both cases, high frequency oscillations of the blade leaf are assumed (base flexure).
The hypothesis shown in the top right diagram assumes coupled oscillations of a turbine disk and blades. Gas and mass forces take effect, depending on the oscillation deflection, which periodically causes a movement against the gas flow. The mechanism is similar to that in the top left diagram. The figure climbs up the periodically stressed string against gravity through alternating clamping force (frictional force).
The deflection of the blade annulus should be between 102 - 103 mm. A certain combination of forces and movements in the root region causes the oscillating blade (Fig. "Fretting by vibration of blade fastener") to lay against the labyrinth ring (bladewalking). This is where the observed friction wear occurs. If, under certain circumstances, the coupled blade/disk oscillations “stand still” and run around with the rotor, corresponding fretting wear on the hold surface will appear periodically around the circumference.
The bottom diagram depicts a hypothesis that is based only on blade oscillations. It has been observed that, if these are powerful enough, they can lower the friction force in the blade root so much that even a small axial force will cause the oscillating blade to travel. This axial force can be explained by the disk warping due to temperature gradients, even if they are very minor. In compressor disks, the rear side of the disk is assumed to be warmer, which causes the disk to warp and offsets the annulus to a sufficient degree that the blade experiences a rearwards pull against the increased gas pressure.
In the (uncooled) turbine disk it is typically the front side that is warmer, warping the disk so that the pull is in a forward direction, i.e. against the gas pressure. In this hypothesis, the blades stay in their axially changed position. Therefore only the individual oscillating blades should leave fretting marks on the hold surface.
Example "Fretting damage on titanium disk" (Ref. 6.2-1, Fig. "Fretting damage on titanium disk"):
Case of engine failure caused by fretting initiated fatigue failure of Ti-6Al-4V compressor disk
”…When investigating an engine failure of a … engine, which failed after only 306 hours service, it was found that the failure originated from an area on the compressor disk just under the attachment part for a blade. The wear path coincides with the edge of the contact surface between the compressor ring and the compressor disk and the wear path has a counterpart on the compressor ring. The depth of the wear path was measured to 0.04 mm. Wear marks were also visible near the bolt holes, from the bolts holding disk and ring together.“
Comment: This case concerns an older-model fan engine with a relatively low bypass ratio that has found wide use in civil commercial aircraft. It is interesting to note that the weak point was not in the blade, but in the disk itself. The fretting path in the region of the disk fittings indicates a structural flaw (stiffness, connection). Disk oscillations may also have played a part. At least one of the oscillation cracks most likely initiated on the back side of the disk outside of the bolt bore. Evidently a crack also ran into the bolt bore. There are no statements analyzing this crack in connection with the damage causes. The damage symptoms indicate that fretting occurred inside the bore, as well (even burrs seem to have formed, although it is not clear if these were consequential damages). Assessments concerning this heavily stressed zone would be interesting.
Figure "Fretting damage on titanium disk" (Example "Fretting damage on titanium disk", Ref. 6.2-1): Fretting in disks usually occurs in high-stress zones such as threaded bores, flange overlays, and centering collars. The high stress levels in these zones create the possibility of dangerous crack expansion. The limited possibilities for inspection of bores add to the difficulties. Because of this, the dynamic strength-reducing effect in the fretting zones is especially important in this case and must be taken into account by the designer.
Coating areas that are susceptible to fretting (Ref. 6.2-5) with plasma spray-coatings or galvanized coatings is possible, but very problematic. First it must be proven that the coating itself does not lower the dynamic strength of the part to a unallowable degree. This effect can result from the application process (e.g. strength/stiffness drops due to thermal changes, hydrogen-embrittlement) or the coating itself (e.g. brittleness, crack initiation, low dynamic strength).
A tribo system with outward-facing open cracks is clearly affected by intruding media. Typical damaging media are sea atmosphere and condensation water.
Remnants from the manufacturing process or lubricants that break down into aggressive components at high operating temperatures must be taken into account. Various oxides and other reaction products can form inside cracks, fundamentally changing the behavior of a tribo-system (see also Fig. "Wear products influencing fretting"). These factors must be considered during technical trials and verification of the operational capability of tribo-systems.
Bores are usually hardened by shot peening or “piercing” (chapter 6.3)
Figure "Fretting damage at turbine rotor" (Ref. 6.2-1): This case concerns an older engine type. The areas affected by fretting are the flange lay-on surfaces and the bolt bores, both of which are most likely subject to relative micro-movements. In this situation, it is not a screw connection that firmly fastens parts together, but rather an anti-twist device. Warping occurred internally through the hollow shaft that acts as a bushing.
Fatigue cracks initiated in the bolt bores of a low-pressure turbine disk. The disk material is 13% Cr steel.
The cracks started in fretting zones of the flange supporting surface and from cylindrical surfaces of the bolt bores. There were two typical areas where cracks initiated (bottom detail): the edge of bevel “1” and in the axial center of the cylindrical surface “2”. Along with the micro-cracks, areas with macro-cracks and a typical bow-shaped progression also appeared in the fretting zone.
This example shows that this type of anti-twist device without braced flange surfaces are susceptible to fretting and should therefore be avoided.
Example "Fretting in the hub zone" (Ref. 6.2-5): ”…The compressor last-stage disk wheel hub exhibits fretting and wear. These include the mating interface with the shoulder-on shaft, and also the retaining ring of the Z cross section. In some cases, radial cracks have originated in the fretting region of the hub. Applying a sacrificial protective coating of copper-nickel-indium alloy to the hub has proven effective in reducing fretting and preventing cracks in this disk wheel.
Comment: If fatigue cracks occur in the fretting zone of steel parts, the danger of this happening in titanium alloy disks, as are common in modern engines, is many times greater.
This example shows that fretting must be avoided in the hub zone by design from the very start.
This type of situation can occur in radial vents or in fixtures to direct the flow between compressor disks, for example.
Example "Fretting wear at compressor guide vanes" (Ref. 6.2-5): “…An example of a persistent fretting problem is the leading edge wear at contact surfaces of compressor stator vane bases in the last few stages, and particularly in the outlet guide vanes. These stator vane bases are supported from T-shaped grooves in the casings (similar to Fig. "Fretting wear at compressor guide vanes")….The forward ledge is 0.088 in. (approx. 2.2 mm) with no wear evident. The aft ledge has worn until the remaining ledge thickness is 0.024 in (approx. 0.6 mm). High-frequency vibration has been demonstrated by strain gages applied to outlet guide vanes. Dynamic pressure gages located in the compressor discharge diffuser reveal high-frequency fluctuations of the flow and static pressure.
Comment: This excerpt from 1967 concerns a still-recurring problem with guide vanes fastened individually or in segments to the inside of casings (compare Fig. "Fretting wear at compressor guide vanes"). The affected parts are the T-connections of individual blades and blade segments. The fretting stress high enough that even steel blades are affected, as in this case. This shows the importance of searching for fretting traces as early as the engine trial stages, so that suitable remedies can be applied if necessary (e.g. coatings, see chapter 6.3)
Figure "Fretting wear at compressor guide vanes": The connection between guide vanes and the casing or supporting intermediate structures are typical fretting zones. Not only the unavoidable blade oscillations must be observed, but also differences of thermal expansion between casings and blades. Different thermal expansion and/or temperature differences cause the connection to loosen, if only for a short time, allowing glide wear and impact wear to occur. These influences must be considered when testing a tribo-system and can affect the selection of possible coating materials. Although dynamic cracks are less likely in the roots of guide vanes than in rotor blade roots, they can still occur, and the geometric wear can offset the blade. If this happens, the blade may come into contact with the rotor. In extreme cases the blade or blade segment may fall out.
For this reason, even new parts made of titanium alloys in fretting-susceptible zones should be armored with an anti-wear coating (e.g. tungsten carbide).
Figure "Failing of retaining bolts at contact areas" (Ref. 6.2-8): Turbine guide vanes are affected by high perimeter forces from gas deflection. The rotation of the guide apparatus in the casing is avoided through detachable fastenings. These fastenings are subject to strong static forces, part-specific vibrations, and thermal strain. This leads to micro-movements at the contact areas to the guide apparatus. There are several structural solutions to this problem. In the case at hand, 44 so-called “anti-rotation pins” were used. In modern engines, there are catches on the shrouds of the guide vane segments that fit into corresponding recesses in the casing.
If these fastenings fail, engine failure can be expected. In extreme cases, the guide apparatus rotates in the housing, grinds through the casing wall, and throws fragments out of the engine (Example "Failing retaining bolts 1" and Example "Failing retaining bolts 2").
Excerpt (Ref. 6.2-8): ”…. The company (OEM) action is in response to a NTSB recommendation made to the FAA last week that urged prompt inspection of the engines and replacement of affected parts. As many as 700 engines could be involved.
The problem area has been identified as the third-stage stator assembly of the low pressure turbine. This stationary assembly is made up of 22 subassemblies or clusters. Each cluster is held to the turbine case by two locking pins. Failure of some pins and subsequent failure of additional pins due to increased loads allowed the clusters to work loose into the rotating parts of the engine.
Two contained engine failures, one in 1985 and one last year (1986), are believed to have been caused by this. A third failure … occurred on Mar. 23 ….This failure was uncontained. (The OEM) plans to make new pins manufactured from higher strength alloy…“
Comment: Evidently the first and second damage instances were not considered dangerous enough to warrant immediate replacement of the locking pins.
Even though fretting is not explicitly mentioned, the function of this element and the damage sequence combined with insufficient creep-resistance indicate that fretting probably contributed to the damage occurring. There are indications of this in Refs. 6.2-9 and 6.2-10. In Ref. 6.2-10, “excessive wear” and “anti rotation pins” in a second engine type from the same manufacturer are mentioned. The relationship is explained in Ref. 6.2-9: ”..inspections of turbine vane clusters in engines (the second engine type) with similar problems“.
In the second engine type, this type of damage had been noticed at least 15 years (!) earlier. Evidently the information gained from this did not lead to the implementation of effective remedies.
Excerpt (Ref. 6.2-10): ”…Some anti-rotation pins showed excessive wear, but… (OEM) officials claim this was corrected with the newer….engines which have the thrust frame, or stiffeners. But none of these has reached 1,000 hrs. in operation yet and has not had the required tear-down inspection.“
Comment: This concerns a large, first generation fan engine where similar wear problems occur in the anti-rotation pins as occurred in the old engine type.
Figure "Fretting wear at turbine blade shrouds": Turbine rotor blades are equipped with shrouds, which support and dampen to prevent vibrations. Fretting wear at the bearing surfaces of the shrouds (top left diagram) can affect their life spans. Therefore, the wear resistance of the bearing surfaces is increased by welding on armor (usually a stellite).
The damage mechanism can be either gliding wear and/or hammering wear (Fig. "Load-specific wear types").
Experience has shown that a change in the blade material can have a strong influence on shroud wear, especially if the change is from a polycrystalline cast material to a single crystal material. This may be due to different wear behavior of the new material in unarmored bearing surfaces. The wear could also be connected to increased blade oscillations and/or stiffness changes due to the change in the modulus of elasticity.
Heavy fretting wear causes tight shrouds to loosen, “overriding” (top right diagram), and fatigue crack initiation with shroud breakouts (bottom diagram). The reduction in shroud tension and increased relative movements accelerate the wear damage and promote fatigue fractures in the blade leaf or shroud.
Figure "Fretting at combustion chambers" (Ref. 6.2-4): Combustion chambers are subjected to vibrations around 100 Hz due to the pressure changes from combustion. Hot parts, especially combustion chambers, also have many plug connections with the surrounding parts to counterbalance changes in thermal expansion. The top diagrams show a pipe combustion chamber from an older military engine with typical plug connections:
The problem of fretting in combustion chambers and hot gas pipes is especially serious in “Low-NOx” combustion chambers due to their typical flame instability (flickering) and corresponding pressure changes with vibrations.
Figure "Fretting wear induced fir tree cracks": Wear is an important factor for the life span of engine parts and a determining factor for the cost of overhauls and repairs. Usually, the repairable wear damages are exactly specified in overhaul manuals and allowable repair methods also given. Determining these specifications and manuals requires a profound knowledge of the stress levels and repair processes, as well as their possible effects on operating properties.
An important criterion is the repair and permissibility of geometric changes through metal removal. It must be remembered that overhaul and repair involves fusing new parts to run-in ones. The contact surfaces must guarantee safe functioning of the part during the new interval of operation. If a new surface is fused with an unevenly worn one, there must be no unpermitted tension concentrations or load changes.
In the case shown, the turbine blade roots were slightly shorter than the disk width (tolerances of the blade root width and the fastening elements, mounting tolerances). A “nose” formed on the disk slot due to metal removal (bottom diagram). After the old blading was replaced with a new assembly, the blade roots lay against these “noses”. Centrifugal force then created strong bending loads in the fir-tree slots of the disk, causing cracks to initiate in the upper disk slot (top right diagram). It is also important to ensure that, when setting overhaul guidelines (manuals, specifications), the metal removal from contact surfaces is given extra attention, if new contact surfaces are to be combined with old ones (e.g. during a modulus replacement).
This problem occurs especially in the following parts:
Excerpt: ”…The top removal reason is the HP compressor where wear in the variable stator vane (VSV) bushings has caused the most removals. (The OEM) is now recommending these bushings be replaced at shop visit and is investigating new materials for a replacement bushing.
Comment: This concerns a late-model medium-sized fan engine that is in wide commercial use. This example shows the “art” of finding suitable bushings for a guide vane bearing assembly. The demands for long operating times at tight tolerances and good glide qualities make this a very challenging task. Evidently the experience or testing were insufficient for serial implementation (see Fig. "Jamming by fretting").
Figure "Jamming by fretting": Guide vane adjustment systems are affected by fretting wear. This is due primarily to their filigreed design and the large number of moveable parts with joints and bearings. The system can be made to vibrate at high frequency by the blading and casing. Because the adjustment systems are mounted in the bypass duct or outside of the engine, corrosive atmosphere (e.g. ocean air) can affect them even during operation. Because they require good glide qualities, strict tolerances (guaranteeing zero-backlash), and abrasion resistance, the joint elements must be made of steel. These are not corrosion-proof, not even 13% Cr steels. The various operating loads that act on a guide vane adjustment require combining very different materials, including steel, titanium alloys, Mg or Al alloys, and glide materials, such as synthetics and bronze. The use of unsuitable auxiliary materials during assembly and maintenance (cleaners, lubricants) or unusual operating conditions (e.g. long standing times in corrosive environments with large temperature changes and/or high dust levels) can promote the creationg of reaction products at the contact areas. In this way, the fretting behavior of a tribo-system can become worse over time.
Experience has shown that fretting can result in glide surfaces jamming and getting hung up. This can result in fractures of part regions under high stress (bottom diagram).
Similar wear conditions can also affect different engine systems (e.g. thrust reverser, Fig. "Fretting wear at titanium with steel contact").
A special problem is posed by the glide bearings of the central pin in guide vanes being rubbed out or sticking fast (Example "Variable stator vane bushings"). Synthetic bushings can swell and jam due to condensation water or unsuitable cleaning materials. If they are not properly lubricated, synthetics filled with inorganic materials (e.g. fiberglass) are susceptible to the trunnion being worn down severely, increasing play to dangerous levels. Overly high operating temperatures cause embrittlement and crack initiation or shrinking of the bushing.
The suitability of a part must be verified in trials that sufficiently simulate operating conditions before it is serially implemented. These trials are very time-consuming, since combined factors such as wear, corrosion, and aging (temperature cycles) cannot be temporally accelerated.
Figure "Wear in protective sleeves": Under unfavorable conditions, guides of control cables and flexible shafts are subjected to strong abrasive wear by gliding and hammering movements. When transferring large amounts of momentum, flexible shafts tend to offset and vibrate inside the conduit. In extreme cases, glide layers or lubricants are broken through and wear soon begins to affect the part`s ability to operate. It was observed that a vibrating flexible shaft of steel wire exerts a powerful hammer stress on the softer cladding tube (e.g. 18 9 CrNi steel or titanium alloy). The cladding tube was drawn out to the point that it lengthened several millimeters. The resulting warping and vibrations can lead to pipes fracturing and the system failing.
Figure "Fretting wear at lever joints": As mentioned in Fig. "Jamming by fretting", adjustment- and blocking systems in thrust reversers are subject to high fretting stress and corrosive influence; the latter can be observed most frequently on adjustment bearings (bottom diagram). If wrong (unsuitable) cleaning materials and lubricants are used in this area during maintenance, the joints can lock up. One situation that has been shown to be problematic is when glide coatings containing MoS2 come into contact with ocean atmospheres. Degreasing constantly-lubricated joints and bearings during maintenance (e.g. with unsuitable sprays) or overhaul (e.g. hot degreasing-baths) can gravely accelerate this process.
The top diagram is a schematic of a mechanical blocking system in a military thrust reverser. Extra attention should be paid to the material combination of the lock pin and the lead-through. The hammer stress typical to this part can, in a very short time, lead to abrasive wear in many different areas (“1” to “5”), causing the part to weaken and fracture or dangerously increase the amount of play. An effective remedy is an elastic bracing of the opened flap over a flexible seal.
Figure "Fretting wear at titanium with steel contact" and Figure "Fretting wear between bearing and titanium shaft": Even under seemingly ideal lubricating conditions, titanium alloys tend to extreme metal removal during high-frequency vibrations against steel surfaces. Typical situations include roller bearing seats in casings (Fig. "Fretting wear at titanium with steel contact", Ref. 6.2-12) or on titanium-alloy shafts (Fig. "Fretting wear between bearing and titanium shaft"). In only a few hours, several milimeters of metal were removed from a titanium-alloy bearing casing in which an oil-damped bearing ring was set (see also Ref. 6.2-14; top diagram). However, the steel ring showed no measurable wear. It is typical that markings stamped into the front surface of the bearing ring were negatively embossed in the titanium surface (center diagram). The problem was finally solved by applying an anti-wear coating in the form of a tungsten carbide cobalt plasma spray coating to the contact surfaces of the titanium part.
Under similar circumstances, an inner bearing ring on a titanium shaft loosened. Fretting wear combined with an oil feed led to the bearing ring loosening and running through, destroying the shaft. In this case, as well, the bearing ring seat in the shaft was successfully protected with a WC-Co-layer.
These damages indicate that the abrasion process is actually promoted by the oil between the contact surfaces. Whether or not this effect is due to the wear products being transported out of the “pumping” gap between the contact surfaces has not been determined.
Example "Loss of aircraft due to alternator shaft wear" (Ref. 6.2-2): “….During takeoff the engine lost thrust and the pilot ejected. Just after the ejection the engine recovered, thrust increased and the aircraft crashed into the ground on government property. Investigators believe that wear on an engine alternator shaft caused the alternator slip, and the powerplant's control interpreted this as an engine overspeed. As a result, the controller kept reducing fuel to the point where thrust was severely compromised.
Eventually, just after the pilot ejected, the controller determined that an overspeed was not a problem and it switched to secondary control, which corrected the problem, prompting an increase in engine thrust.'This was a known problem, and it was being addessed by an order that called for all …(of the affected fighter type ) with electronic controls to undergo an alternator shaft wear inspection as a part of normal aircraft maintenance. We thought there was a low probability that the failure would happen during takeoff',…
A longer-term corrective action includes a software improvement to permit a switch to secondary control more quickly. The ultimate correction will focus in redesigning the engine's alternator shaft or its attachements”.
Comment: The weak point was evidently known, but the usual inspection was unable to sufficiently detect and prevent the failure. Typically, it was only hope that led to the assessment, that a decisive failure would not occur during takeoff (against Murphy`s law).
In the end, applied measures (evidently there had been similar cases before) were not sufficient and redesigning was unavoidable. This also means that this was primarily a design flaw.
Figure "Fretting wear at spline couplings": Detachable shaft couplings in the form of interleaved plugs are often used in engines. This design principle is applied in main shafts and auxiliary units (top diagram, “1” to “6”). These shaft connections are extremely reliable if the following conditions are met:
If any of these conditions are not met, wear in the interleaving can be expected. This often remains unnoticed until the coupling ceasses to mesh (Fig. "Failure due to fretting wear at spline coupling", Refs. 6.2-3 and 6.2-12).
Figure "Failure due to fretting wear at spline coupling" (Ref. 6.2-3, Example "Increased wear caused by incorrect machining"): Auxiliary components such as hydraulics, fuel pumps, oil pumps, and starter generators are usually powered via a so-called radial shaft from the main engine shaft to the auxiliary transmission. These often use tight and moveable interleaved plugs to connect the gears to the shafts and auxiliary components. The above mentioned example describes the typical damage symptoms when one of these connections fails due to extreme fretting in the splining.
“The failed splined shaft and gear assembly forms an integral part of the ….(single engine fighter-) engine accessory drive gear train. The gear train provides power from the main engine to the engine accessories, including hydraulic pumps and an electrical system.
During landing from a routine flight, an engine flame-out occurred. The aircraft was recovered safely. Stripping of the engine revealed that the accessory drive failed.
No previous instances of accessory angle drive gear failures had been reported…
The bevel gear was able to rotate freely on the shaft. Disassembly of the components revealed that the splines of the shaft and of the bevel gear had failed completely from severe fretting wear.. The splines of the shaft had been plated with a silver alloy to provide fretting wear resistance. The bevel gear is normally held to the shaft by a specially castellated nut, support ring, and a nut locking mechanism.
Most probable cause: The wear was caused by incorrect machining of the shaft splines.
…It was concluded that the splines of the shaft had been incorrectly machined, preventing correct tightening of the bevel gear to the shaft on assembly. In fact, much of the torque generated in tightening the nut had been expended in forcing the support ring into the ends of the shaft splines….it is clear that the pressure on the bevel gear was insufficient to prevent the gear moving relative to the shaft.
It was recommended that all accessory angle drive splined shaft and bevel gear assemblies…in service…be examined for correct machining of the shaft splines. A modified manufacturing and assembly procedure for the component was introduced.”
This case shows the importance of tuning the structural design with the available production methods. This must be ensured by close cooperation between design and production. The alteration of the assembly process proves that this would have been possible from the start. If sufficient coordination exists, the risk of this type of assembly defects is minimized.
It is probably no coincidence that the failure of the gear/shaft connection was noticed during landing. It is plausible that the additional performance required, e.g. through use of the landing gear (powering up the hydraulic pumps and the electrical system) and/or the increased electricity demanded by operation of the control flaps, could have demanded an increased torsion moment.
6.2-1 “Fatigue Crack Initiation and Propagation under Fretting Conditions” (Lit 458)
6.2-2 S.W. Kandebo, “USAF Targets Engine Mishaps”, periodical “Aviation Week & Space Technology”, Page 84, March 29, 1999.
6.2-3 S.R. Lamb, G. Clark, “Engine Accessory Angle Drive Gear Failure”, “Handbook of Case Histories in Failure Analysis”, Chapter: Transportation Components 17, ASM international.
6.2-4 M.Nakao, M. Ikeyama, S. Abe, “Analytical Condition Inspection and Extension of Time Between Overhaul of F3-30 Engine”, ASME-Paper 91-GT-277, Proceedings of the “International Gas Turbine and Aeroengine Congress and Exposition” Orlando,FL June 3-6,1991.
6.2-5 J.S. Alford, “Design Criteria and Configuration for Long-life Aircraft Gas Turbines”, Flight Propulsion Div. General Electric Co. SAE Paper 670344 of the National Aeronautic Meeting, New York, April 24-27, 1967.
6.2-6 H.A. Jergus, “Microslip Initiated Cracks in Compressor Blade Attachements”, Paper des Romforsa Research Seminar on Material Science, 15-16 August 1977.
6.2-7 W.Schütz, “Sind Druckeigenspannungen die Ursache von Schwingfestigkeitssteigerungen”, Deutscher Verband für Materialforschuing und -prüfung e.V., DVM-Report 122 “Leichtbau durch innovative Fertigungsverfahren”-1996, Page 95-101.
6.2-8 “Pratt Developing Inspection, Repair Methods For JT8D-200s Following Three Failures”, periodical “Aviation Week & Space Technology”, May 4, 1987, Page 21.
6.2-9 “NTSB Urges FAA to Inspect Pratt and Whitney JT - 8D Engines for cracks”, Turbine Intelligence, DMS Volume 12, Number3 May 4, 1987.
6.2-10 “Longer Inspection Times Sought After JT9D Hot Section Cracks”, periodical “Aviation Week & Space Technology, July 6, 1970.
6.2-11 S. Hoadley, “Rolls-Royce to redesign faulty Viper blades”, periodical “Jane's Defence Weekly, 2 December 1995, page 14.
6.2-12 R.L.Jonson, R.C. Bill, U.S.Army, “Fretting in Aircraft Turbine Engines”, AGARD-CP-161, Oct. 1974, pages 5-1 to 5-13.
6.2-13 “The CFM56 in service” periodical “Aircraft Technology Engineering & Maintenance, June/July 2001, page 86-96.
6.2-14 R.L Johnson, R.C. Bill, “Fretting in Aircraft Turbine Engines”, Specialists Meeting on Fretting in Aircraft Systems, Proceedings AGARD-CP-161, page 5-10.