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19.2.6 „Invasive" (in situ) repair.

The term „invasive-repair“ ( in situ repair) is chosen in dependence on modern surgery technics. Here are invasions carried out in comparatively short time and with relatively low expense. So this deals with maintenance procedures and repair measures inside the aeroengine without opening the engine and without considerably preparatory work.

In a similar way alteady today at the earoengines on wing internal work during maintenance is carried out (Ill. 19.2.6-1 and Ill. 19.2.6-2). To these can count the washing of the compressor and hot parts (chapter 19.2.3). Different to the compressor which accessible from the entrance/intake, hot parts need an entrance by a suitable opening. Besides borescope openings/holes the ducts/holes of removed injection nozzles or spark plugs are used (Ill. 19.2.3-12 and Ill. 19.2.3-16).

Already today surface damages on blades which are approachable from the intake (compressor/fan) or the exhaust (low pressure turbine) will be mechnically reworked (polished). Also this can be seen as an „invasive repair”.

As a sort of „invasive maintenance process“ can count non destructive testings at an aeroengine on wing. To those belong x-ray displays (Ill. 25.2.2.2-4 and Ill. 25.2.2.2-5 , Lit. 19.2.6-15), crack detection/testing with eddy current, penetrant inspection or ultrasonic testing and magnetic measurements (sulfidation inside hollow turbine blading, volume 1, Ill. 5.4.5.2-2.1).

The increasing number of borescope holes (Ill. 19.1.4-7), which enable the visual inspection of important inner components offer a requirement for „invasive repairs”. An especial advantage is, that it is not necessary to open the aeroengine after an expensive demounting of the accessories or even to send it to the shop. Also the trend to integral constructions, represented in an impressive manner by the blisk design allows the use of advantages. For example if a small FOD can be invasive repaired, it is not necessary to exchange the whole component. This saves demanding assembly work and logistics.

In the following examples are shown which are already in routine use (Ill. 19.2.6-2 and Ill. 19.2.6-4). Beyond this there are developments in progress (Ill. 19.2.6-5 and Ill. 19.2.6-6) and new technologies offer further chances (Ill. 19.2.6-7).

Illustrations 19.2.6-1 and 19.2.6-2 (Lit. 19.2.6-1): This example shows in an impressive manner the benefits of an „invasiven repair“. But also the problematics of the process development get noticeable. Like here they must be adapted to special situations. This is the reason why in the following those will be considered in detail.

Aeroengines of fighters (Ill. 19.2.6.2 sketch above right) can suck in desert environments large amounts of sand. How mutch sand the air carries depends on the season with predominant winds (volume 1, Ill. 5.3.2-9). The consequence is a contamination of the hot parts, especially the turbine blading. The cooling air holes to the blade surface were in the shown case partial clogged by fused and recast sand. The composition of the deposits differs depending on if they passed the combustor or were transported by the cooling air.

  • Deposits at the outside of the high pressure turbine rotor blades consisted of SiO2 (silica) and calcium-magnesium-silicate.
  • In the cooling air holes of the high pressure turbine rotor blades the deposits consisted of silica and calcium sulfate (anhydrite= gypsum without crystallisation water).
  • The guide vanes of the 1st turbine stage (HP nozzle guide vanes) were coated with calcium-magnesium-silikate and calcium sulfate.

A reduction of the cooling channels cross section leads to an increase of the high pressure turbine (HPT) guide vanes (nozzles) and rotor blades temperature. Thereby it must be considered , that at the common operation temperatures in the region of those components about 12°C increase bisects the creep life. So at the turbine rotor blades of the forward stages more frequent fractures of the outer shrouds and the blades occurred . A result of creep and thermal fatigue. Also typical was a „burning” (heavy oxidation, volume 3, Ill. 11.2.3.1-10) at the leading edges of the HPT rotor blades. A follow-up examination showed bore holes of the film cooling which were partly clogged from inside and outside. This lead to metal temperatures of 1100 -1150°C instead 950 - 1000°C. By such an increase of temperature the creep life is some decimal powers reduced. So it must be reckoned with a operation lifetime of few hours.

Sulfur contents in the gypsum deposits (operation in „gypsum deserts“) let expext besides clogging of the cooling air channels/bores also an especial intense hotgas corrosion attack as sulfidation. This was later confirmed by follow-up investigations of the bladings.

Necessary are special maintenance measures with specific requirements.

  • Applicable in the assembled condition at the aircraft.
  • Minimalexpenditure of time for a fast supply of the airplane.
  • No sticking media on other components.

Initial method (Ill. 19.2.6-2, sketch above left): Experimental abrasivee dry blasting with 25 my alumina (aluminium oxide, corundum) and air as transport medium was used.
Sketch above left shows the function chart. Thereby the turbine rotor is rotated step by step and the jet nozzle automatically guided radial along the leading edge of the particular blade. The monitoring of the cleaning process happens on a display ossociated with a borescope plugged into the combustor. A radial further inside positioned nozzle for compressed air should avoid that abrasives get into the radial cooling air bore holes inside the blade.

However this method proved as very time consuming and its effectivity was unsufficient. For this reason an improved method was introduced. This was a high pressure water jet process. The water jet exits with a pressure of 620 bar (62 MPa) against the leading edge of the blade. Thus the labour time could be shortened about 90% compared with dry blasting. To feed and purge the water an approved facility is needed, of which several hoses lead to the aeroengine (Ill. 19.2.6-1).
This cleaning process has proved its worth and is applied in the described application in intervals of 32 operation hours if the „sand coats”exceed the specified limits.
Comment: Fused glassy coatings bond similar to an high temperature enamel with the benefitial protective oxides at the blade surface (volume 1, Ill. 5.3.2-12.3). Are those glassy coatings removed by the blasting process, it must be assumed that also the oxides will be stripped (Lit. 19.2.6-2). Possesses the surface like often turbine blades diffusion coatings as oxidation protection, they may be stripped after few cleaning processes. The result is an accelerated oxidation. This effect will be intensified after the cleaning by the reactive metallic surface. Thus it must be reckoned with a considerably reduction of the blade wall thickness, especially at the thermally high loaded leading edge. The cleaning process may considerably reduce the blade lifetime. However, there exists obviously in desert environments an advantage compared with the risks as consequence of a deteriorated cooling.

Illustration 19.2.6-3 (Lit. 19.2.6-3 up to Lit. 19.2.6-6): It is understandable that inside aeroengines blasting processes with abrasive particles are problematic. It is hardly to ensure, that particles don't get into the cooling air channels of hot parts or through labyrinth seals into the oil circuit. With this failures of sealings, gears and bearings would be pre-programmed.

Those problems are not to expect from a blasting process whose particles evaporate/sublimate. Used are ice particles of frozen CO2 („dry ice“). Those are compacted from CO2 snow and introduced in the transporting air stream. Shortly after the impact on the surface respectively the cleaning effect, those particles evaporate without residues. The effectivity of the CO2- blasting process nearly correlates with conventional abrasive blasting processes. The cleaning effect depends not on abrasion but on thermal and kinetic effects. This preserves the substrat. Therefore this process does not count as abrasive. A further advantage is, that no for the environment hazardous respectively toxic abrasion remains (Ill. 19.2.3-7). Already Ni-oxide from the hot parts belongs to such critical materials.

CO2 blasting already is used during the overhaul of aeroengines at removed/disassembled parts/components instead of chemical cleaning processes and stripping processes for coatings.
The relatively low kinetic energy of the in fact fast, but very light CO2 particles will be equalised by other effects. If the dry ice particles hit a surface, extreme temperature gradients with corresponding thermal stresses develop („thermal shock”, sketch) between coating and base material. Additionally the very low temperature can short-term embrittle ductile coatings what e.g., is used in an other application also for the mechanical removing of chewing gum. The immediately evaporation of the CO2 particles leads to high flow velocities and gas forces (thermal-cinetic-effekt). This supports the blasting effect and leads to chipping of the contaminations/coatings.

With the CO2 blasting process coke and corrosion deposits, greaes, oils and sticky coatings/residues can be removed. With this process may also suited for the „invasive cleaning“ of compressors. Because there is no abrasion of the base material it must not be reckoned with an erosion effect at the blad profiles and/or the relatively soft rub in coatings in the casings.
It is not clear why till now this process obviously did not emerge in an „invasive” application. That could be seen in connection with the size of the jet nozzle for which possibly the borescope holes are too narrow.

Illustration 19.2.6-4 (Lit. 19.2.6-7): With such a washing equipment fuel incection nozzles can be cleaned in the assembled position. For this a cleaning nozzle is inserted into the combustor at a suitable opening (e.g., opening for a borescope or a spark plug). So for different aeroengine types coke deposits must be removed. Such deposits can deflect and/or deform the the fuel jet. So the danger exists that the walls of the combustion chamber will be overheated and in an extreme case the catastrophic bursting of the combustor casing will occur (volume 3, Ill. 11.2.2.2-9). An unfavorable shape of the flame can deteriorate the temperature distribution and so lead to a local overheating. To guarantee a sufficient cleaning effect the pressure of the washing fluid will be increased.

This shows how the washing device can be adapted for special applications.

Ill. 19.2.6-5 (Lit. 19.2.6-8, Lit. 19.2.6-9 and Lit. 19.2.6-9): This mashining technique to even/blend small FODs inside the aeroengine is carried out through the openings for the borescopes. That process is used when a damage without rework, e.g., corresponding to the maintenance manual is no more allowed (pictures above left). The damage can be assessed with the help of of special, layed in gauges and a borescope. In this case an „invasive repair“ (in-situ-repair) can avoid a disassembly of the blading. Implemented are micro-grinding devices or a small high speed milling tool. The process is monitored by a simultaneously inserted borescope (sketch above right). The rework naturally takes place within the limits, specified by the OEM. The process is composed of chipping/grinding followed by polishing. Since years (about from 1990) it is tried to apply this repair process (blending) at aeroengines on wing also on not through the entrance reachable compressor blades (e.g., fan, booster). In the second tool generation sufficient robust, elbowed milling heads were applied (sketch below right). The processing is observed by a insensitive borescope with a lense system which guarantees sufficient high resolution (sketch of the facility obove right). The newest version uses a movable gripping head in which different tools such as millers, grinding and polishing devices with ard metal edges or diamond tips can be fixed. For the insertion of the tool through the opening for the borescope, the tool head will be put straight. After this it can be elbowed into a 90°-position. The generated chips are very small and are obviously not regarded as a problem. By experience about 30 minutes are needed for the blending of one damage.

Comment: The high aerodynamic load of the blades of modern compresors demands a fatigue strength as high as possible. Even if the blending is very shallow in this region must be reckoned with a small increase of stress as result of the notcheffect. Because the blending of blade damages in the end is carried out by hand, deviations in the process can not be ruled out. Changed processing parameters (condition of the tool, temperature of the machining surface, smeared scratches/marks) can alarming decrease the fatigue strength. Therefore it is obvious to use after the blending a strain hardening process, which increases the material strength and induces protecting compression stresses. For this laserpeening offers itself which is described in Ill. 19.2.6-7 (Lit. 19.2.6-14, Ill. 19.2.6-7).

Ill. 19.2.6-6 (Lit. 19.2.6-1, Lit. 19.2.6-2, Lit. 19.2.6-8, Lit. 19.2.6-9 and Lit. 19.2.6-10): Integral rotor stages of the fan and the compressor, so called blisks (sketch below right) are increased used in modern civil and military aeroengines (sketch above). For this reason also small FODs have here especially an influence at the operation safety of the components.

  • The lack of friction damping at the blade roots influences the vibration behaviour.
  • High costs of the components.
  • Logistics in case of an exchange.

The requirements for the fatigue strength let it appear desirable to introduce a work hardening after the blending process. Here the laserpeening almost offers itself (comment in Ill. 19.2.6-5).

„Invasive repairs” have especially for blisks a high potential for savings and improving of the safety. The application of „invasive processes“ for the blending of FODs at blisks, concerning a strain strengthening, blendung emerged already (Lit. 19.2.6-14, Ill. 19.2.6-7).

Illustration 19.2.6-7 (Lit. 19.2.6-11 up to Lit. 19.2.6-14): Laserpeening is a process to strengthen the surface by strain hardening, using the energy of light which also produces high compression stresses (diagram below). It can be seen that the internal stresses are in the magnitude of the shot peenung. Astonishing the depth effect is even higher. This effect needs a light semi-permeable (opaque) film and it applied on a light conducting (transparent) film (sketch middle left). The laser impulse penetrates the transparent film. The opaque film evaporates under the extreme temperatures. The developing plasma creates a violwent pressure impulse. The shock wave exits the film/coating into the base material. The transparent covering prevents, that the pressure impulse escapes external. During this process the opaque film prevents that the laser beam hits directly the surface, damaging fusing it. Such up to 25 my thick fusing zones would decrease the fatigue strength. Suitable are different dry or liquid opaque films/coatings like

  • dried liquid paints/lacquers,
  • black adhesive tape,
  • metal foils.

The easiest and economic transparent cover film is flowing water (sketch above). It can be applied with a nozzle. As already mentioned, water serves as a transparent coat to cover the plasma but not for cooling.

The process can be locally used in the region of notches as they are also represented by FODs. Thereby it offers itself especially for the curing of already blended damages (Ill. 19.2.6-5 and Ill. 19.2.6-6). Obviously it's at least in the military field just before an application at from the outside accessible fan blades.

Thinkable to treat are components which are not from the outsice accessible components like compressor blades with the help of light cables and/or mirrors through openings for borescopes.
Also efforts are reported which deal with a regenreation of blade tips, which are deteriorated by rubbing.

It is imaginable that this process could be also benificial applied in other regions of an aeroengine. Possibly persistent, hard glassy coatings on the hot parts can so be removed.

References

19.2.6-1 M.G.Down, M.J. Williams, „Out of Area Experiences with the RB 199 in Tornado”, AGARD Konferenz „Erosion, Corrosion and Foreign Object Damage Effects in Gas Turbines“, AGARD-CP-568, Proceedings des „Propulsion and Energetics Panel (PEP Symposium Rotterdam, The Netherlands, 25-28 April 1994, page 1-1 bis 1-7.

19.2.6-2 P.König, T.Miller, A.Rossmann, „Damage of High Temperature Components by Dust-Laden Air”, AGARD Konferenz „Erosion, Corrosion and Foreign Object Damage Effects in Gas Turbines“, AGARD-CP-568, Proceedings des „Propulsion and Energetics Panel (PEP Symposium Rotterdam, The Netherlands, 25-28 April 1994, page 25-1 bis 25-12.

19.2.6-3 „What is Dry Ice Blasting”, Fa. COLDJET, www.coldjet.com., 17.03.2006, page 1-3.

19.2.6-4 „Dry Ice Blasting Process“, Fa. ICE SONIC, RSG-Technologies, www.rsg-technologies.com., 17.03.2006, page 1 und 2.

19.2.6-5 „Carbon Dioxide Blasting Operations”, www.p2pays.org., 17.03.2006, page 1-6.

19.2.4-6 „Interactive JTEG Depot Maintenance Projects“, Air Force Application: Jet Engine Components, Project Number: 42566 : „Carbon Dioxide Pellets”, Project Number 41421: „Carbon Dioxide Technology“, www.jdmag.wpatb.af.mil, 17.03.2006.

19.2.6-7 „Kell-Strom PT6 Engine Wash Rigs”, „ Customized Aircraft Turbine Engine Wash Rigs“, Prospekt der Fa. Kell-Strom, www.Kell-strom.com., 1.07.2006, page 1und 2.

19.2.6-8 Y.Tanaka, S.Nagai, M.Ushida, T.Usui, „Large Engine Maintenance Technique to Support Flight Operation for Commercial Airlines”, Fa. Mitsubishi Heavy Industries, Ltd.,“Technical Review”. Prospekt der Fa. Kell-Strom, www.Kell-strom.com., 1.07.2006, page 1-5.

19.2.6-9 W.Ohnesorge, „Boroscope Maintains Gas Turbines“, Zeitschrift „Diesel & Gas Turbine Worldwide”, April 1999, page 12-14.

19.2.6-10 „Through the looking glass“, Zeitschrift „Aircraft Technology & Maintenance”, April/May 2001, page 1-26.

19.2.6-11 „Section 1.6 Manufacturing Technology for Affordable LSP“, HCF 2002 Annual Report, www.pr.afl.af.mil, April/May 2001, page 1-4.

19.2.6-12 M.R.Hill, A.T.DeWald, J.E.Rankin, M.J.Lee, „Measurement of laser peening residual stresses”, Zeitschrift „Materials Science and Technology“, 2005, Vol 21, No.1, page 13-9.

19.2.6-13 L.Hackel, „Engineering residual stress”, Engine Yearbook 2006“, 2005, page 80-83.

19.2.6-14 R.D.Tentaglia, D.F.Lahrman, „Preventing Fatigue Failures with Laser Peening”, Zeitschrift „The Amptiac Quarterly“, 2005, Volume 71, Number 2, 2003, page 3-7.

19.2.6-15 R.M.McCord, „Turbine Engine Inspection without Disassembly”, Proceeding Paper AIAA-80-1152, der AIAA/SAE/ASME 16th Joint Propulsion Conference, June 30-July2, 1980/ Hartford, Connecticut, page 1-4.

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