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21.2.6 Rework of damages by blending/levelling.

The specification conform blending of small FODs at the compressor blading is an important repair taski. It gets more important with the increased application of integral components like blisks, brazed, welded or cast stators in turbine and compressor. These systems distinguish itself with a lower ;

(no the friction like inserted blades) and a pronness for resonance vibrations for „;
“ effects (volume 3, Ill. 12.6.3.4-6.2 and volume 4, Ill. 16.2.2.9-3). Just small changes of the vibration behaviour can get dangerous. This demands extra tight specified component dimensional blending limits (Ill. 21.2.6-1.2).

After a damage at the blading is identified with a direct visual check or with the help of a borescope, the blending can be carried out on wing in the certified limits during the maintenance of the aeroengine. For such tasks fan blades offer itself. Here may be pointed at the probably in the future usable possibilities of „invasiveblending respectively „in situ“ blending (chapter 19.2.4).

During an overhaul blendings are carried out at damaged blades, according the specifications in the overhaul manual (Ill. 21.2.6-2).

The blending of flaws/faults is a quite demanding, safety relevant task. Also seemingly small deviations from the specified process respectively parameters are later no more to verify with non destructive testing and not allowed. They can lead to the fracture of the component with catastrophic consequences (Ill. 21.2.6-3).

Ill. 21.2.6-1.1: The rework at small, operation caused notches at the compressor blading is damanding. This is not alway sufficiently aware. Rework is necessary, because from such notches vibration fatigue cracks can start (sketches below). It is difficult to say in which case the fatigue strength or the damaged blade is no more sufficient to withstand the dynamic loads. Therefore the rework limits are a precaution as far as possible „limited to the safe side” and accordant restrictive. The ability for a rework of a notch depends not only from its size and type („A“). Important for a rework is, if it is also located at the airfoil, accordant with the specifications in the manual (Ill. 21.2.6-3). This is complied with distribution of the vibration load in the airfoil. It dependes from experience with the dangerous vibration modes about the blades of this stage (volume 3, Ill. 12.6.3.1-5).

Does a blade fracture, at least the shut down/failure of the engine must be expected. In an extreme case at some aeroengine the danger of an uncontained failure from a titanium fire exists (volume 2, chapter 9.1.1).

Reworkable notches occur as small foreign object damages (FODs), mostly by ingested stones. They are frequently identified during an external inspection at blades which are visible from the entrance. In the rear compressor region borescope inspections serve for identification (Ill. 25.2.2.1-7). Than, if there is accessibility in the assembled condition, a rework according to the specifications in the maintenance manual can take place.

In the future it can be expected, that so called invasive technologies will also allow the rework of FOD notches in the not manual accessible compressor region (Ill. 19.2.4-5).
During an overhaul the disassembled during an external inspection blading will be inspected and if necessary reworked.

Important is the specification of the rework limits by the OEM and its accurate compliance. Development and proof of the rework limits (volume 1, Ill. 5.2.1.1-9 and Ill. 5.2.1.1-10.2) is demanding. These limits usually rely on experience with relevant parts under comparable conditions. Anyway the proneness for FODs, relating to the likelihood of impacts and typical dynamic loads in operation, are extremely part specific.Therefore it must be assured with sufficient operation experience.

The rework of a notch at an airfoil usually is limited at the „blending”. This is carried out by a machining process („B1“, „B2”). Thereby attention must be payed not only at the specified macro geometry, good chamfering and the tolerable surface roughness of the blending.

The rework process itself can dangerously deteriorate the part by carelessness, deficient handling of the tool or unsuitable tool (volume 1, Ill. 5.2.1.1-9).

Usually for bending mode vibrations of the airfoil axial oriented machining grooves are dangerous (detail B1). Are these scores not sufficient removed or smeared („C“) by a following finishing process, an alarming drop of the fatigue strength must be expected.

A similar situation occures, if the machining temperatures have been too high. Potential deteriorations are changes of the materials structure, crack formation and internal tensile stresses (volume 4, Ill. 16.2.1.1-2). Indications for this are annealing colours or a particular burr formation, which can be monitored by the technician during the machining. However, are those removed again (detail at „B2”) a sufficient safe proof can be ruled out. Such an overheating can decrease the fatigue strength by drop of the strength/hardness and/or induced internal tensile stresses (volume 4, Ill. 16.2.2.7-6). This deterioration has, depending from the intensity of the temperatures a certain depth effect. Also a following polishing is possibly not enough to remove this damage.

To ease such problems a subsequent strain hardening peening (shot peening) offers itself if the airfoil geometry is suitable. If necessary it increases a dropped fatigue strengh („D2“, „D1”). Naturally the peened surface may not once more damaged by deformations and notches. Peening with glass beads is often the suitable processs. Therefore such an after treatment must be also developed, proved and approved/certified by the responsible person, especially the OEM.

Illustration 21.2.6-1.2 (Lit 21.2.6-8): Even seemingly small mistunings of the natural frequencies of the blades from a blisk can rise a multitude of the vibration load and so trigger fatigue fractures. This behaviour must not only considered during the new parts production. Just repairs by blending of notches from FODs or the replacement of whole damaged blades (e.g., by linear friction welding) seem especially concerned. So it must be rekoned, that this effect occures at blisks much more intense as at inserted blades with friction damping at the root contact surfaces. Literature about this effect in its dangerous form is not available for inserted blades.

Illustration 21.2.6-1.3: The rework of surfaces, deteriorated during operation is a quite demanding task. Such a rework can be carried out with hand held grinding tools (Ill. 21.2.6-1.1), etching processes (chapter 21.2.4), electrochemical processes (electro polishing) or abrasive blasting (volume 4, Ill. 16.2.1.6-18). Often these processes are combined. Usually the depth of the rework is limited by tolerance requirements. Are these not met, they can dangerously decrease the operation safety. For example the vibration behaviour can change unacceptable, caused by a cross section weakening from rework. The danger are resonance vibrations (Ill. 21.3.3-7) which lead to the fraction of the blade.

On the other side, if deteriorations are not fully removed, they can grow accelerated during the reuse (diagram obove right). With such a behaviour must be reckoned for sulfidation, if the nickel sulphide containing zone remains in the surface of the part (volume 1, Ill. 5.4.5.1-4). Corrosion attack and/or oxidation can grow along the grain boundaries and, is than difficult to identify. So the danger exists, that this deterioration zone was not fully removed (sketch above).

This also depends from a sufficient confident testing, to identify with non destructive testing remained deteriorations (Ill. 21.2.6-4). This is at least true for the development of repair processes. The procedure respectively the parameters of the removing process must be sufficent documented/established and monitored. An adequate check can be damage specific an etching test.

The testability complicated if the rework process closes faults to the surface by smearing or blocking (grinding, abrasive blasting). A further problem is the possibility of an unnoticed material damage by too intense material remove with machining (Ill. 21.2.6-1.1).

In this case nonmachining removing processes like etching and electro polishing have an advantage (volume 4, Ill. 17.3.1-7). However the potential danger exists to damage the surface with a crack like attack (chapter 21.2.4).

In cases of blocked surface faults/flaws (pores, cracks) a thermal cycle has proved. This must not be fast, because it causes the opening by creep with the bending of thin, heavy strain hardened surface latches/burrs respectively of sharp crack edges (volume 4, Ill. 17.3.1-7).

Example 21.2.6-1 (Lit. 21.2.6-6): After the fatal flight accident this single engined helicopter type was temporary grounded because of safety deficits of the aeroengine. A similar failure had occurred at an other helicopter on ground. Obviously the failure concerned the 4th stage turbine wheel (2nd stage of the power turbine, Ill. 21.3.3-5). The supposed cause was:

Citation: „“…fourth-stage turbine wheel, which was found to have an overblended, thin trailing edge. The blade was blended beyond standard requirements….”
The OEM recommended as measure a singular inspection of the trailing edge thickness at wheels of the suspicious aeroengines. This must be carried out during the next 150 operation hours at turbine wheels with more than 1000 operation hours. Wheels with less operation hours must be checked in time intervals of 300 operation hours.

Comment: This failure may deal with fatigue by high frequency vibrations. Such a crack propagates too fast and can not be intercepted respectively identified in time. An inspection in time periods appears therefore not resonable. Rather the cracks my caused by thermal fatigue and/or a creep effect. A blade fracture with the drop out of the whole aeroengine made the crash of the helicopter likely.

Illustration 21.2.6-2(Lit. 21.2.6-2 and Lit.21.2.6-3): The rework of an operational damage at a compressor blade must be carried out specification conform. This are instructions in the manual, from the OEM and/or of the responsible aeronautical authority.

Usually the part is divided (sketch right) into zones with different rework intensity (chart below left). Thereby the main criteria are hight and type of the operational load. Especially considered are high frequency vibration loads in the HCF region of the blade airfoil „A-E“ and combined dynamic HCF and LCF loads in the blade root („X”, „Y“). The stress level depends from influences like vibration modes, exitations (frequecy, intensity, volume 3, Ill. 12.6.3.3-21), size, weight and geometry of the blade as well as of its aerodynamic characteristics. A special influance has the profile geometry (details left). At flat profiles as they are typical for rotor blades, during cross bending vibrations occur also markedly stresses in the midth of the chord („B”). In contrast at stator blades which are mostly more chambered, the high loaded regions can be expected at the edges (volume 1, Ill. 5.2.1.1-7).

Besides considerations about the strength, the rework of an airfoil is also limited by aerodynamic requirements. For the blade of an aerodynamically high loaded stage with the susceptibility for a stall (rotating stall, volume 3, Ill. 11.2.1.1-1) may prevail more restrictions as for the blade of a stage which tolerates from experience FOD notches without failing.
Sometimes in modern compressors there are stages of rotor and stator at which all corners from leading edge and trainling edge are removed (similar zone „A“). Usually blades with thin profiles and long chords are concerned („wide chord”). Sometimes the trailing edge of long stator blades (vanes) in the fan region from fighter aeroengines is removed so, that the chord is shortened about several millimeters. These changes are normally not in connection with FODs. Admittedly they are rather a remedy to avoid crack formation caused by high frequency vibrations (lyra mode, tremline mode, volume 3, Ill. 12.6.3.1-6) that seems a little „helpless“.

Illustration 21.2.6-3 (Lit. 21.2.6-5): This comparing test at the fan of a big „first generation” three shaft aeroengine (sketch above) shows in a convincing manner the importance of the edge optimizing and the smoothening of the airfoils. This is for fan blades of special importance for the (specific) fuel consumption (pecific fuel consumption = SFC), in this case related at the thrust (thrust specific fuel consumption = TSFC).

During the normal operation, especially fan blades are erosion loaded, depending from the operation conditions (volume 1, chapter 5.3.2) and damaged by foreign objects (volume 1, chapter 5.2.1.1). With this frequent starts and landings in short distance operation are more failure effective than in long haul operation. Additionally the arrangement of the aeroengines at the airplan (fuselage, wings), distance from the ground (volume 1, Ill. 5.2.1.2-7) and the location of the airports (e.g., desert region) influences the ingestion of particles.

Fanblades rum near the tips in the supersonic region and have therefore very sharp profiles. Those edges are especially erosion loaded. Its geometry influences the aerodynamic quality markedly more (volume 4, Ill. 16.2.2.9-5) then subsonic blade profiles with a larger edge radius.

Also the surface roughness of the airfoil (volume 3, Ill. 11.2.1.1-8, Ill. 11.2.1.1-10 and Ill. 11.2.1.1-9.1) determines the efficiency of the fan. Thereby the absolute roughness cross to the flow direction is rather relevant than the average roughness Ra (volume 4, Ill. 16.2.2.1-1). However this was used as a criterion in the available literature.

With the test the trend of the TSFC was diagnosed on a calibrated test rig from an aeroengine with a longer operation time under typical conditions (diagram below left, dashed curve). Then the blades have been disassembled and, compared with an usual overhaul reworked/refurbishment with an improved process.

1. The airfoil surfaces have been smoothened/polished to a roughness R< 5µm (20rms µ in) smoothened/ polished.
2. The leading edge was rounded and thinned over 5-10% of the chord length (detail middle).
3. Also the reworked leading edge was smoothened to R<1µ.

After this the blades have been again assembled into the same aeroengine of whith the TSFC was determined on the same test rig (diagram below left, continous curve). This showed for see level conditions a decrease of the TSFC about 0,7 %. At the first glance this seems not much, but should not be underestimated. It should be considered , that 1% SFC of an airline fleet can mean a considerable profit for the operator.

During the tests also the exhaust gas temperature (= EGT) was measured. This showed a decrease of about 3°C for the reworked case. Also this can lower the operation costs markedly. Keep in mind, that an 12°C increase of the material temperature of the high pressure turbine blades halfs the creep life of those expensive parts (volume 3, Ill. 12.5-4).
An evaluation showed, that 70-80 % if the improvements can be assignet the thinning and rounding of the reworked edges. In contrast the smoothening of the airfoil surfaces claimed only 20 %. Thereby it must kept in mind, that the roughness may again increase after a relative short operation time and reduce the benefit further.

Note: Especially the geometry of the fan blades leading edge influences the efficiency and with this the fuel consumption of an aeroengine.

Illustration 21.2.6-4 (Lit. 21.2.6-9): At the rear side of the LP turbine 4th stage disk and the frontside of the 5th stage (detail middle right) during the last overhaul several (40-50!) blendings of mechanical damages have been carried out. About 3 years later, after about 12 200 operation hours and 2 300 cycles a radial fracture of the 5th stage disk occurred. The inital crack was in a intercrystalline fracture zone about 10 mm wide. It is located in the outer area of the disk membrane (sketch below left), about 12 mm radial below the transition to the hollow shaft. The crack obviously propagated radial inward and outward.

The disk surface did not show the expected circumferential oriented machining grooves, but looked polished (detail below middle). This is typical for a blending process. Here the cross section of the disk was 0,025-0,075 mm thinner than the thickness on the drawing of 2,75 mm but was inside the tolerance. The materials structure correlated at this position and other checked disk zones widely the the specification of the OEM.

Through the intercrystallin crack origin a metallographic cross section was made. It showed grain boundary cracks perpendicular to the surface.

Further investigations of other blendings on the disk showed not fully blended/removed mechanical damages. Here oxidation pointed at the fact, that these are not secondary failures. In the remaining rest of the damages, metallographic features of plastic deformations (detail low right) could be identified.

With such a deterioration, a cyclic test with five minutes dwell time at 537°C at samples of the disk material IN718 was carried out. Thereby intercrystalline fracture surfaces developed like at the fracture origin of the disk.

From these findings the following conclusions have been made:

  • Not sufficiently blended damages with plastic deformations and/or small cracks triggered in this highly loades disk zone a LCF crack. Such deteriorations for example occur also by a tool fracture during machining and show their danger (see volume 4, Ill. 16.2.2.5-5 and Ill. 17.5-1).
  • Such damages can not be identified with the specified non destructive testing methods
  • After that the responsible authority recommended to the OEM to investigate the blending process at its applicability for such disks.
  • Review of the design of the disk.

Illustrations 21.2.6-5.1 and 21.2.6-5.2 (Lit. 21.2.6-7): This case shows the high risk of repairs, respectively rework of a surface, damaged by rubbing. This is especially true for highly LCF loaded rotor components (Ill. 21.2.6-4).

During rubbing, high amounts of frictional heat are produced and introduced into the contact surface (volume 2, Ill. 7.1.3-5 and volume 3, Ill. 12.2.4-10). The lower the heat conductivity, the higher the heating temperature and with this the damaging potential. The depth of a damage can be unexpected large and penetrate cross sections in the range of centimeters.

Titanium alloys are especially endangered because of their extremely bad heat conductivity. Thereby not the worse is a materials structure caused drop of strength. Exeeds the heating temperature a limit at which oxygen absorption and oxidation occur, it must be reckoned with a dangerous embrittlement (volume 4, Ill. 16.2.1.3-20). This embrittlement worsens also the fracture mechanics variables for crack propagation and critical crack length (critical fracture toughness), from which a spontaneous fracture (forced fracture) will occur.

For nickel alloys besides a drop in strength by „solution annealing“ the danger of hot tears exists (volume 2, Ill. 7.2.2-9.2).

In every case it must be reckoned during rubbing with local heating and remaining internal tensile stresses. With the increase of the mean stress, these lower the bearable cyclic strength and so promote so crack formation by HCF and LCF with a markedly shortened lifetime.

References

21.2.6-1 Australian Transportation Safety Board, Aviation Safety Investigation Report 200205780, „In-flight uncontained engine failure and air turn-back, Boeing 767-219ER, ZK-NBC” page 1-43.

21.2.6-2 Lufthansa Technische Schule, Training Manual, Kapitel „ Nacharbeit von Schaufelbeschädigungen“, Schulungsunterlage „Technik, 7.2 Triebwerke, Turbinentriebwerke”, Ausgabe 01.84, Bestell-Nr.: 0047011 page 60.

21.2.6-3 I.E.Traeger, „Aircraft Gas Turbine Engine Technology“, Second Edition, Glencoe Verlag, ISBN 0-07065158-2, 1994, page 337.

21.2.6-4 J.T.McKenna, „Industry Team Pushes Focused Safety Plan”, Zeitschrift „Aviation Week & Space Technology“ February 16, 1998, page 30 and 31.

21.2.6-5 W.B.Roberts, „Advanced Turbofan Blade Refurbishment Technique”, Zeitschrift „ Journal of Turbomachinery“, (Transactions of the ASME), Vol. 117, October 1995, page 666 and 667.

21.2.6-6 J.Wastnage, „Offshore operator pots Bell 407 back to service”, Zeitschrift „ Flight International“, 25 December 2003 - 5 January 2004, page 6.

21.2.6-7 NTSB Identification LAX92IA307, microfiche Number 48059A, „Schedules 14 CFR 121 operation of United Airlines. Incident occurred Sep-23-92 at San Francisco, CA, Aircraft Mc.Donnell Douglas DC 10-10, registration N1845U”, page 1.

21.2.6-8 T.Klauke, „Schaufelschwingungen realer integraler Verdichterräder im Hinblick auf Verstimmung und Lokalisierung“, Dissertation, Dezember 2007, page 129 and 130.

21.2.6-9 National Transportation Safety Committee, Republic of Indonesia 2001, Aircraft Incident Report, „Japan Airlines Flight JL726, B747-300 JA8178, Tangerang. West Java, Indonesia, 5. September 2000”, page 1-14.

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21/212/2126/2126.txt · Last modified: 2020/06/25 22:43 (external edit)

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