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21.3.4 Repair proving/testing

This chapter is addicted to the proof of the operation behaviour respectively the capability of a repaired part during a testing. This is also applied for repair processes (chapter 21.3.2 and chapter 21.3.3).

The test of a repair, respectively of the repaired part, must be sufficient near-service conditionstos (Ill. 21.3.4-2). The effort in costs and time must kept as low as possible. For this a step by step approach has proved (Ill. 21.3.2-2 and Ill. 23.3.4-1). Requirement is, that the failure mechanism, together with the causative influences, is sufficient cleared. For this a systematic problem analysis lends itself (volume 4, Ill. 17.1-11). Those findings should enable the definition of a simplified test of samples, based on calculations or a special rig/setup (Ill. 21.3.4-5 and Ill. 21.3.4-7). Crucial is, that its relevance is shown with a sufficient reproduction (volume 1, chapter 4.4).

Usually damages which must be repaired, develop in a complex coaction most different operation influences (Ill. 21.3.4-2). These loads mostly detract a solely „treatment with the computer“. Surely the exclusive analytical theoretical method would be the fastest way and the most cost effective proof. Therefore from experience the danger increases, that such „proofs” are carried out without expensive reproductions. The case specific incapability of this approach is connected with a high safety risk. However in the predominant number of cases, only a test in an aeroengine can give in the end sufficient certainty about the repair success. Also here the maxime „the engine will tell us“ is true.

Those proofs usually demand more than merely a relatively short run at a test bed, even if this run should last 150 operation hours (volume 3, Ill. 11.2.1.3-10). From experience, those time periody are not sufficient near-service to simulate long time deteriorations/damages like oxidation, aging of coatings, erosion and wear. Also for the simulation of influences like landing shocks or gyroscopical forces, test rigs are hardly suitable.

In such cases a proving at the operator offers itself with „fleetleaders”. Perhaps this may be the most expensive way. This with periodic controls respectively inspections monitored aeroengines, must be subject to operation relevant loads. That applies for influences like the mission profile and/or inevitable anomalies. To these belong high amounts of dust or a corrosive atmosphere.

Illustration 21.3.4-1 (Lit. 21.3.4-1 and Lit. 21.3.4-2): Development of processes is discussed in chapter 21.3.2. Here the proof of the capability from the repaired part stands in the foreground.

Understanding the problem respectively the failure which made the repair necessary. To this belong:

  • Failure mechanism (volume 4, Ill. 17.1-11).
  • Temporal sequence.
  • Risks.
  • Causative influences (volume 1, chapter 3).

Not before this knowledge target-oriented action with a high probability of success is possible. It is also the requirement for the definition and execution of the proofs.

For the understanding, failure analysis, statistics and overhaul manuals can provide important informations. Such eyperiences can be especially expected at the OEM and the operator. Because this deals with product specific knowledge, it is not not aiways available for the development of a repair.
Analytical considerations and calculations gain significance for the understanding of deteriorating operation loads. For such analysis the OEM is predestined. He should sufficient know loads an its timely sequences already before the design. Also the effective and meaningful application of the computer requires the „understanding of the failure“.

A problem relevant component test is the best requirement for the evaluation or demonstration of the success to the approving certifying institutions. These important validations take place most effective with comparing tests.

Requirement is a suitable and feasible (costs, time) component test (volume 1, chapter 4.4). Criterion for the applicability is the microscopic and macroscopic reproduction of the failure which must be repaired. This nust be carried out at a relevant component/part, mostly new under defined, repeatable testconditions. Is this highly demanding task solved, the possibility exists to evaluate the behaviour, especially the lifetime of the repaired parts. The limit of such proofs are the operation loads which are not known or not mentioned and therefore can not be simulated in the test set-up. To these belong high frequency vibrations or overtemperatures (e.g., during start).
The proof with tests of aeroengines at the test rig can merely check the behaviour of a repair under short time acting effects. Naturally these must occur in the test program. Usually those tests are, because for cost reasons, limited to short operation times (under 150 hours) compared with the air traffic. With this influences of long time effects like erosion, fretting wear and corrosion can be hardly sufficiently tested (volume 3, Ill. 11.2.1.3-9). Also special loads like landing shock (important for rub of labyrinth seals and tip gaps) or gyroscopic loads during change of the direction (during lift off, military maneuver, volume 2, Ill. 7.1.2-12) elude from those tests.
The most time consuming but also the most meanigful proof takes place at the operator during the typical flights. To minimise risks, some few with the repair equipped aeroengines („fleetleader”) are used. During this the condition of the repair is suitable monitored. For this for example borescope inspections serve (chapter 25.2.2.1). Are test parts removed after significant time periods and tested in the laboratory for the impact of possible deteriorations, we speak about „sampling“.

Illustration 21.3.4-2: The testing of a repair is quite demanding. A useful summary about approaches/sequences and proof problems can be found in the „Department of Defense Handbook” (Lit. 21.3.2-2). It contains the „Engine Structural Integrity Program“ (ENSIP) and got available in the internet.

Frequently for the proof of the seviceability of repairs, extensive cyclic test runs like for example they are carried out for an ETOPS operation, are not available (volume 1, Ill. 3-8). Therefore special proofs

are used (Ill. 21.3.4-1), according the problem respectively the repair and failure type which must be repaired. To the operation influences belong:

Operation limits, flight envelpope and external loads (Lit 21.3.4-2): This deals on the one hand with the flight envelope, i.e., the operation limits depending from the flight height and flight speed. In this region for example ignition and the operation of the combustion chambers as well as an after burner (military) must be certain guaranteed (volume 3, Ill. 11.1-8 and Ill. 11.2.4-4). Also no dangerous effects like an oilfire (volume 2, Ill. 9.2-3) or too high main bearing loads (volume 2, Ill. 7.2.1-2) may occur.
In the operation envelope, also g-loads must be safely tolerated during maneuvers (especially in military use, volume 2, Ill. 7.1.2-12). Here effects like rubbuing of blade tips and labyrinth tips are concerned (volume 1, Ill. 7.1.2-12).

Operator specific missions and service parameters in the operation envelope determine frequency and time laps of parameters like rotation speed, gas temperature and compressor pressures. From these depend loads like thermal fatigue, creep, centrifugal forces and exitations of high frequency vibrations. It can be a big difference if a part is located in an aeroengine for short haul or for long distances. Also the military version usually is significant higher loaded. This can quite be reflected in an about 10-times shorter lifetime than designed. Equivalent behave also the overhaul periods. The repair costs may lay in a contrary relation higher.

Corrosion and erosion conditions are all more harder the longer the operation takes place near the ground. Especially aeroengines of helicopters are highly loaded by these influences. Erosion in the compressor (volume 1, Ill. 5.3.2-9) and the formation of layers/deposites in the turbine (volume 2, Ill. 5.3.2-12.3) are for those aeroengines in many cases lifetime limiting.

Watery corrosion, especially at elder aeroengine types with corrosion sensitive steels (e.g., compressor blading, rotor disks), is markedly promoted by shut down periods and thereby developing condensate water. Locations near the sea and low level flights over sea which are especially frequent during military missions (aircraft carrier!) promote corrosion.

Icing (volume 1, chapter 5.1.4), rain (volume 1, chapter 5.1.1), hail (volume 1, chapter 5.1.2) depend from type and location of the mission/operation. These are influences which especially act at the aeroengines front area. In an extreme case they deteriorate/damage the blading of the fan and/or of the compressor behind.

Lightning strike (volume 1, chapter 5.1.3) is primarily to consicer for turboprop engines and helicopters.

Bird impact represents considerable danger. Potential concerned parts must show in extensive proofs an acceptable behaviour (volume 1, chapter 5.2.2).

An aeroengine should be as resistent as possible against foreign object damage (FOD, volume 1, chapter 5.2.1). The larger the damages in the blading are allowed without the danger of a fatigue fracture, the better. Repairs may at least not deteriorate this behaviour.

The containment behaviour, i.e., the resistance of the casings against the penetration by fragments of rotors (volume 2, chapter 8.2) may not affected by a repair (e.g.,weld). Primarily centrifuged blades and internal fires are concerned (titanium fire, volume 2, chapter 9.1).

Illustration 21.3.4-3 (Lit. 21.3.4-): For the specialist, with some luck it may be possible to extract important conclusions from the microscopic and macroscopic picture of the failure. Helpful are combinations with material investigations, which allow conclusions at the failure mechanism, failure sequence and the causative influences. If this succeeded, the requirements for reproducing component tests (Ill. 21.3.4-1) is created. With this a cost efective possibility for the proving and a risk evalustion exists (Ill. 21.3.4-2). The fracture of a airfoil of a typical low pressure turbine (LPT) rotor blade may serve here as example. Certainly not all features like shown in this sketch are analysable at every fallure case. This rather deals as a summary of possible evidences.

To the rubbing features of the labyrinth tips at the shroud count wear size/depth, distribution and signs of overheating (volume 2, Ill. 7.2.2-3.1). They should allow conclusions if a heavy rubbing, could have stressed the blade dynamically dangerously high.

The degree of fretting wear at the contact surfaces of the shroud (volume 3, Ill. 11.2.3.2-5) gives hints at a reduction of the tension between the neighboring blades. This promotes vibrations of the blade.

The macroscopic wear pattern shows fracture origin, direction of crack propagation and type of the load. From the beach marks it can be suggested at time temporal sequences, e.g., start/shut down cycles. With this the point of time of the crack growth can be in certain limits evaluated. This is of interest if certain operation loads are known. The size of the residual fracture (forced fracture) points at the load level of the blade by the centrifugal force.

A microscopic investigation of the crack surface (scanning electon microscope = SEM, volume 4, Ill. 17.3.2-7) shows feature in the region of the crack origin (e.g., FOD, porosity, crack by creep, fatigue crack).

The condition of the blade surface (e.g., oxidation, condition of the coating, deposits) shows special operation conditions.

A microscopic material investigation (metallography, volume 4, Ill. 17.3.2-6) can answer questions for overheatings, deteriorations (e.g., features of the age hardening phase), material condition and material failures.

Geometrical features like profie aberrations and shortened chordlength (e.g., by rework) can be the cause for a failure triggering changed vibration behaviour (natural frequency/resonance, Ill. 21.3.3-7).

The condition of the labyrinth tips at the root platform can give hints if the turbine stator in front had changed failure influencing its position.

Degree and distribution as well as the structure of the fretting wear at the contact surfaces of the blade root let suggest the type (e.g., bending) and load (microscopical length of the wear marks, wear volume).

Illustration 21.3.4-4: In this summary are hints at sites in the book series „Aircraft Turbine Engine Safety” in which the mentioned effects are discussed in detail.

Illustration 21.3.4-5 (Lit. 21.3.4-3): Labyrinths are functional exposed to deteriorations. Concerned primarily by rub is abrasion (volume 2, Ill. 7.2.2-3.1), formation of heat tears (volume 2, Ill. 7.2.2-9.2), local heating (material structure, internal tension stresses). During assembly damages can occur. Therefore during overhaul, a welding repair takes place (Ill. 21.2.1-5). For this the labyrinth tips are partly ground off and welded up with several layers. After this the weld is machined (grinding) to measure.

The weld is especially problematic for Ni alloys whith tend to heat cracking/tears (Ill-. 21.2.1-3). In contrast to this, in titanium alloys it must be rather reckoned with porosity (volume 4, Ill. 16.2.1.3-29). From such flaws LCF cracks can start during operation (volume 2, Ill. 7.2.2-9 and Ill. 7.2.2-10). Naturally also it must payed attention at a favorable rub behaviour. This means no heat cracking or heating caused deteriorations (changes of material structure, drop of hardness) with crucial decrease in strength.

The sketched example shows the application of the fracture mechanics for the optimising of the weld repair for labyrinth tips. Thereby amongst others, it depends on the finding of the best combination of base material and welding filler material.

At first a so called fishbone diagram (chart below left) was created by a team of design, repair, material technology and stress department (Life Management Engineering). Such a diagram contains the influences at the formation of flaws/weak points on repair welds. This develops with a creative phase („Brain Storming“). Minimum criterion was the operation behaviour of a unrepaired part. For these requirement four superior aspects have been identified: Management in service, properties of the repaired part, repair process and application.

The LCF strength as essential feature of the operation behaviour of repaired labyrinth tips was identified with cyclic tests. For this samples with lengthwise labyrinth tips/grooves (upper frame) have been used for the tension-tension tests (swelling). The geometry of the tested cross section (cross section AA) met the usual LPT interstage labyrinth (sketch below right). The test was carried out till the fracture occurred. The test technolocy did not allow to identify alreadey the initial crack phase.

The test results and its evaluation/analysis shows Ill. 21.3.4-6.

Illustration 21.3.4-6 (Lit. 21.3.4-3): The Ill. 21.3.4-5 shows LCF tests of samples with labyrinth profile allowed for Ti-6Al-4V as weld metal build up on titanium alloy Ti-17 (base material), the following conclusions:

  • Even not build up samples showed a shorter LCF life than round specimens. That means a labyrinth profile itself decreases the LCF life.
  • Samples repaired by welding exhibit a wide scatter of the LCF life.
  • The LCF life of the weld repaired samples was shorter than the unrepaired samples.
  • A thermal treating after the repair welding did not improve the life. This means the internal stresses may be here of minor relevance.

The microscopic follow-up examination of the fracture surfaces from the samples showed, that the typical crack origins have been in the weldseam. Causes for the incipient cracks have been process typical flaws/weak points (Ill. 21.1-2) like pores (volume 4, Ill. 16.2.1.3-31), large material structure inhomogenities and overlaps from machining.
The lifes, evauated by fracture mechanics with the help if the initial flaw size, deviated markedly from the test results (shorter life). This is traced back at the influence of the weld material structure (coarse structure blocks the crack propagation?). With this the fracture mechanics evaluation would be at the `safe side'.

The remaining lifes for the designed stresses and an initial crack size of 0,127 mm have been evaluated by fracture mechanics for Titanium alloys and Ni alloys. This took place for different labyrinths all over the aeroengine (diagrams above). The mean stresses have been related to the highest that will occur during operation. The avaluated lifes from the samples have been compared with the minimum design lifes of the particular component/part. Higher stressed repaired labyrinth tips have in the particular aeroengine types A,B,C a lower <U>relative</U> life expectation. Therefore repairs should preferential carried out at labyrinth rings with low mean stress.
The scatter of the relative lifetimes obviously is for Ni alloys (diagramm above right) markedly longer than for titanium alloys. This may be in connection with the pronounced heat cracking of some Ni alloys. This may also be seen in connection with the wide scatter if the critical crack size in relation to titanium alloys (diagram below). To narrow the scatter there is the requirement of monitoring the strict compliance with optimised process parameters.

Illustration 21.3.4-7: Don't forget besides the serviceability of a repair the proof, respective the optimising of the repair process. Thereby it must be especially considered, that the surface (e.g., oxidation, sketch right) and base material (e.g., material structure) may have been changed during operation. This can change the application of a process, suitable for new parts, without an adaption for repaired parts.

For this reason fore the testing and optimising of repair processses, basically representative parts from service must be used.

References

21.3.4-1 P.C.Patnaik, R.Thamburaj, „Development of a Qualification Methodology for Advanced Gas Turbine Engine Repairs/Reworks”, Paper RTO MP-17, des RTO AVT Workshops on „Qualification of Life Extension Schemes for Engine Components“, Corfu, Greece, 5-6 October 1998, page 12-1 up to 12-11.

21.3.4-2 Department of Defense Handbook, „Engine Structural Integrity Program (Ensip)”, MIL-HDBK-1783B, w/Change 2, 22 September 2004, page 1-176.

21.3.4-3 P.A.Domas, „Fracture Mechanics Evaluation of Weld Repaired Seal Teeth for Life Extension of Aircraft Gas Turbine Engine Components“,Paper des „RTO AVT Workshop on „Qualification of Life Extension Schemes for Engine Components”, Corfu, Greece, 5-8 October 1998, page 16-1 up to 16-8.

© 2021 ITTM & Axel Rossmann
21/213/2134/2134.txt · Last modified: 2020/06/25 22:43 (external edit)

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