This chapter explains typical damage sequences. A damage sequence in an aircraft accident is a causal chain (Fig. "Phases of the damage") made up of several phases (initial process until impact). It is not necessary for all phases to be present in a sequence. It is probable that it will not be clear until the investigation has begun, which phase of the damage sequence is being dealt with. The next step is to attempt to determine the direction of the sequence and understand the order of the phases in the damage sequence.
Understanding the entire damage sequence is extremely important for comprehending all damage-relevant influences, for risk assessments for other affected machines, and for developing corrective measures. The typical phases of a typical damage sequence are:
Figure "Damage system" This diagram is taken from Ref. 4.5-1, and is based on a descriptive analogy from the FAA used to explain how an aircraft accident occurred.
Every independently rotating disk represents one step (in this case a component) of the damage process. The holes are typical failure mechanisms of these components. Catastrophic damage to the engine only occurs if holes in each disk are aligned in a straight line. Complex processes can have more than 20 such steps (disks). If even one disk does not have an aligning hole, then catastrophic damage will not occur, since a link in the causal chain will have been broken. The objective is to keep the number of holes in the disks as low as possible.
Figure "Stop causes" Even in aircraft with multiple engines, the standstill of an engine in flight is treated as a serious occurrence and is classified as an “incident” (see Chapter 2). This also includes shutting down an engine because of an actual or assumed malfunction. Consciously shutting down an engine is referred to as an “inflight shutdown”.
Understandably, engines of single-engine aircraft, usually military aircraft, require considerably lower failure rates than, for example, engines of the same type in twin-engine aircraft (also see Fig. "Operating experience").
In order for an aircraft design to be certified, its safe operation must be verified with occupancy and maximum fuel supply even during the start phase. Engine failure is especially important in twin-engine commercial aircraft that have to cover longer distances over water. These are covered by what are called ETOPS regulations (Chapter 3), which allow the maximum flight times with a single engine. These times must be achieved through extensive verifications during testing and/or flight operation.
This diagram shows typical causes of engine stoppage during flight. These are certainly not all damage to the engine itself. For example, fuel starvation and accidental engine shutdown must also be included in these considerations. The following text briefly covers the typical causes for engine stoppage during flight.
Fuel starvation can be caused by fuel system failure (e.g. malfunctioning pumps or valves) or improper fuel management (Figure "Double engine failure"). According to Ref. 4.5-2, roughly 10% of all incidents concerning flight safety can be traced back to inadequate fuel management due to improper actions. If the incident is not due to failure of the fuel system or the fuel supply to one engine, but rather a complete lack of fuel or problems with the nacelle-side fuel supply (see example 5.4-7), then all engines are affected. This is an emergency situation that can result in a crash in extreme cases (Example "Engine shutdown mistake II" and Example "Engine separation").
Fuel system failure:
Aside from the tanks, fuel systems consist mainly of pipes, pumps, regulators, filters, and the injection systems in the engines. In some cases there is a redundancy that compensates for malfunctions of a component. On the other hand, there are also cases in which there is only one fuel pump (see examples in Chapter 3 and Example "Torque sensor casing") or other vital component. However, if the entire system is affected by freezing of (unsuitable) fuel, for example, the redundancies will no longer be effective.
Stoppage or shutdown of a single engine can usually be traced back to a malfunction or damage.
The pilot can shut down the affected engine after malfunctions or damage. It is a fairly common occurrence that engines are shut down due to malfunctioning indicator instruments or sensors.
Experience has shown that in emergency situations, such as after serious engine damage or due to malfunctioning indicator instruments, the wrong engine, i.e. the still functioning engine, can inadvertently be shut down (Example "Engine shutdown mistake I" and Example "Engine shutdown mistake II").
Excerpt: “..a rotorcraft…was substantially damaged upon impact with terrain during an auto-rotative emergency landing following a total loss of power….the roller bearings were observed to be heavily deformed and the forward (geared) end of the accessory drive shaft had separated from the remainder of the shaft thus depriving the fuel pump of rotational input.”
Comment: The Fuel starvation is caused by a chain of failures. It is a consequence of the fuel pump outage. This is a secondary failure of the fracture from the drive shaft to the auxiliary devices. This again is the result of a main bearing failure.
Excerpt 1: (Ref. 4.5-4): “…Engine failure due to a failed throttle cable…crashed.”
Excerpt 2: (Ref. 4.5-5): “…during an air-to-ground attack after interception. Engine failure due to a fractured throttle cable.”
Excerpt 3: (Ref. 4.5-6): “…crash resulting from an engine failure caused by interruption of the power supply to the fuel valve.”
Comments: This is a single-engine fighter aircraft. Therefore, engine failure will result in a crash. The frequency of throttle transmission system failures mentioned in the literature is astounding.
After the start the left aeroengine lost suddenly power in about150 meters hight above ground. The pilot feathered at once the propeller. He had big problems to hold the airplane under control. During the landing the right aeroengine pulled the airplane severe to the right. It deviated from the runway and got on the grass.
An investigation of the concerned aeroengine at the OEM under supervision of the responsible aviation authority showed as most probable accident cause:
The casing of the torque sensor had a primary failure. This lead to the decoupling of the gears to the fuel control unit. Thus the control unit in the fuel pump dropped out, then the aeroenginen followed.
Comment: At shaft aeroengines, especially turboprops, the released driving torque and with this the power is an important indicator. So the mechanical influence of the fuel control drive can be understood.
Figure "Causes for shaft separation" Separated shafts are a serious occurrence that must be considered as early as the design and construction stages, in order to ensure that no unallowable consequential damages can occur and threaten the entire aircraft (see Fig. "Detached fan rotors"). Shaft failures are rare, but when they do occur, then they are usually consequential damage after the failure of a rotor system component or directly neighboring part groups such as bearing casings or labyrinth seals. Typical causes of separated shafts include (not in any specific order):
Softening of the shaft due to a fire:
Shaft overheating can be caused by several different factors. In multiple-shaft systems, oil fires can occur and stabilize between the shafts. They heat up the outer lying shaft from the inside. Depending on the intensity of the fire, dangerous engine part temperatures can be reached in minutes.
Oil fires in the bearing chambers can overheat the shaft or the flange from the outside (see Example "Power-turbine retention system" and Example "Damage to the fan shaft").
It is also possible that overheating can occur after combustion chamber damage, if hot gases penetrate the heat protection and come into contact with the shafts.
Main bearing damage:
Main bearing damage can cause extreme temperatures (up to the melting point of the steels in the engine) in a very short time (seconds). In engines with multiple hollow shafts, this can also overheat the inner shafts (Example "Power-turbine retention system" and Example "Damage to the fan shaft"). Because the main bearing determines the axial position of the shaft system, main bearing failure always results in shaft offset.
Labyrinth failure can directly or indirectly cause shaft failure. A labyrinth carrier can cut or burn through the shaft cross-section after extreme rubbing (e.g. after failure of the opposite stator assembly) or due to bearing failure.
If the labyrinth seal failure occurs as a typical self-increasing process resulting from material buildup and expansion of the rotating part, then high temperatures and intense spark development can burn through hollow shafts and disk lugs located concentrically inside or outside.
Since disks are usually an integral part of the shaft system, burst disks usually cause the shaft to separate. Also, disks can also burst due to uncontrolled overspeed following a shaft failure.
Shaft separation due to operating loads without overtemperatures:
Shafts can be made to separate by various operating damages, even if this occurrence is very rare:
If the shaft itself serves as a rub-tolerant surface (e.g. opposite shroudless blade tips of a compressor stator) for a spacer ring, then unusually heavy contact (e.g. during a premature restart of an engine with rotor bow) can result in overheating and/or the shaft being ground through.
If the threaded rotor connection fails, possible consequential damages include loose flying fragments such as bolt heads or nuts that were knocked off (Example "Forgotten tool").
There may also be unusual operating damages, such as embrittlement due to inward diffusion of metals (e.g. silver, cadmium) or corrosion (e.g. sulfidation, crack corrosion). Failure of threaded connections usually occurs as a type of “unbuttoning” process. This makes it possible to detect the damage when unbalance begins to set in, before shaft separation occurs.
Figure "Consequences of shaft separation" Separation of a main shaft is always an extremely dangerous situation and has very different damage sequences, depending on the construction.
Main shafts usually connect the compressor with the power turbine. In triple-shaft engines, they are referred to as the high-, middle-, and low-pressure shafts. The low-pressure shaft is also referred to as the “fan shaft”. In two-shaft engines, the fan is located on the same shaft as the booster. In triple-shaft engines, the turbine only powers the fan. If a shaft fails, the entire power of the turbine can contribute to an RPM increase. This can directly lead to the turbine rotor bursting due to overspeed.
The location of the fixed bearing relative to the shaft fracture determines which components can become axially offset, e.g. the fan, compressor, or turbine. Usually, the pressure ratios around the disks are designed so that the axial forces of the turbine (rearwards) and the compressor (forwards) largely cancel each other out across the shaft and only relatively small axial forces affect the fixed bearing. If the shaft separates, these forces are released and deflect the section of the shaft that is no longer held by the fixed bearing.
Naturally, the section of the shaft held by the fixed bearing can temporarily experience correspondingly large axial forces.
The consequential damages to a runaway low-pressure turbine are partially dependent on whether or not the fixed bearing prevents axial offset. If the shaft breaks behind the fixed bearing, then the rotor will also rub against static parts axially , especially against the turbine stators. This can be useful if it brakes the rotor before it bursts. However, the rubbing can also damage (overheating, sectional weakening) other rotor components (e.g. the disks) so that even moderate overspeed result in fragments breaking off. Therefore, in some engine types the rotor is designed specifically so that the braking process occurs through the blading (Fig. "Intermeshing principle").
Shaft failure can result in excessive oil leakage from the damaged bearing seals. This oil can ignite and the resulting oil fire can further endanger the aircraft.
Example "Forgotten tool" (Ref. 4.5-17):
Excerpt: “…The NTSB report said that bolts holding the rotor disks together had been hammered by the foreign material when the engine was operated. The fragments destroyed the capability of the bolts to hold the disks together. When the bolts failed, the first stage low-pressure rotor disk overspeed and disintegrated…“
Comment: During an assembly obviously a tool was left in the low pressure turbine rotor drum (Ill. 20.2-7.1/-7.2). This damaged the heads of the disk bolting. Those could then fracture a fast sequence ('zipper effect', Ill. 188.8.131.52-7). The 'load shedding' at the fan side caused the runaway of the rotor. Thereby developed fragments with a kinetic energy above the design limits of the casing containment.
Example "Power-turbine retention system" (Ref. 4.5-18):
Excerpt: ”…The biggest reason for removals - more than 50 percent - had been problems with the power rotor turbine, a one-piece bladed disc. The…(FAA) has now certificated a redesigned multiple, insertable blade unit which will be introduced on production engines in December, along with a new Power-turbine retention system. Such engines are exempt from the 50-hour inspection required under FAA directive.
A new redundant overspeed system will be fitted to new engines next year, replacing the mechanical form with electronics to cut the fuel flow if power turbine limits are exceeded. A redesigned rear-bearing support housing is scheduled for production engines, to eliminate strut cracking in about a years' time.”
Comments: The fact that the measure taken was to change from an integral bladed disk (blisk) to a disk with insertable blades indicates that thermal fatigue cracks damaged the power turbine disk, which resulted in extreme unbalances and separation of the shaft. The most likely result would be a runaway power turbine with further uncontained fragments escaping the engine.
It is interesting that there was a change from a mechanically triggered fuel shut-off to an electronic. This is in contrast to later experiences of other OEMs which after tests remained with a mechanical shut-off (Fig. "Overspeed prevention").
Figure "Detached fan rotors" (Ref. 4.5-8) Damage caused by main bearing failure and the resulting shaft separation. The rotation energy stored in the released fan section caused it to climb up the intake duct.
Excerpt: “… The latest incident occurred….when (the aircraft) was climbing…At 15 000 feet the six-foot diameter fan of the center engine detached from the shaft, and spun up the S-band centre intake. Fragments punctured the rear pressure bulkhead of the aircraft and parts of the titanium fan penetrated one of the aircrafts toilets.
…(before) suffered a similar failure… In addition to actual failures of the fan module where the blades and disc actually broke away, the…early versions of the engine have suffered failures of the low-pressure location bearing which is just behind the fan” (Fig. "Shaft failure caused by bearing damage").
Remedies (also see Example "Damage to the fan shaft"):
Example "Damage to the fan shaft" (Ref. 4.5-9, also see Example "Power-turbine retention system"):
Damage to the fan shaft of a large civilian transport aircraft! The damage occurred about 1.5 years after several parallel damages (e.g. Fig. "Shaft failure caused by bearing damage").
Excerpt: “The failure …shortly after the aircraft took off….was the first to occur since the engines were fitted with a fan retention device more than a year ago after a series of in-flight fan shaft failures.
The fan shaft retention system …kept the fan disk inside the engine casing. In some earlier incidents, the fan either had flown out of the engine nacelle or had been destroyed when it attempted to do so.
The fan, however is the primary speed control for the engines turbine section. After the fan shaft failed, the turbine reached overspeed and suffered a catastrophic and uncontained failure. Turbine blades thrown from the engine damaged the aircraft's.. engine pylon, the fuselage and portions of the wing and wing flaps systems.
Preliminary investigation indicated that the crew received an indication of above normal vibration in the engine shortly before the incident.
…engineers determined that the first series of fan shaft failures came about when the engines low pressure bearing was starved on oil and ran dry, leading to a sequence of events culminating in the failure of the fan shaft. Higher than normal vibration was caused by these events, and the vibration indicators were installed to warn the crews that an engine fan shaft failure might be imminent.”
Comment: The retrofit of a retention system prevented the exit of the fan disk to the front (Fig. "Detached fan rotors" and Fig. "Shaft failure caused by bearing damage"). Because the shaft fracture is positioned axial in front of the locating bearing, the low pressure rotor can not shift to the rear. Thus a braking of the rotor by contact with the stators (intermesh, Fig. "Intermeshing principle") is not possible. Therefore further a dangerous overspeed of the released low pressure turbine with fracture exit is possible.
Measures for preventing overspeed due to shaft separation.
During development and certification, engines have to prove that they perform safely in the event of shaft separation. There must be no unallowable overspeed that could lead to a rotor burst. Several possible measures can be successfully implemented in order to ensure this, .
A basic principle is that the fuel control unit must be connected to the engine in way that ensures shaft failures will not cause the fuel flow to be increased. This can happen if the regulator is attached to a power take-off shaft in a shaft power engine, where it can misinterpret shaft failures and increase the fuel flow because the engine RPM seem to have dropped.
This can be corrected in older engine types by retrofitting a fast electronic regulator that directly monitors the turbine RPM (Ref. 4.5-18).
A further measure is shown in Fig. "Overspeed prevention". Here, the axial rearward movement of the turbine is used to give the regulator a mechanical signal.
In large engines, the low-pressure turbine rotor is braked by a stator vane stage with specially designed blades (intermesh design, Fig. "Intermeshing principle"). In case of axial offset, the stator vanes rub against the rotor, braking it. The difficult task in this case is to keep the inevitable resultant blade damage to a minimum in order to keep the effect this process has on the engine within the limits specified in the regulations.
Figure "Shaft failure caused by bearing damage" This diagram is a detailed view of the damaged area of an engine.
The damage process is complicated (Ref. 4.5-8). Evidently, the first step was the failure of the oil supply in the front main bearing of the fan shaft. This caused the bearing to overheat and ignited an oil fire which softened the shaft (see description of Fig. "Consequences of shaft separation"). The result was shaft separation with a load drop in the low-pressure turbine, which reached overspeed and burst, creating uncontained fragments.
Example "Worn out ball bearing" (Fig. "Possible consequences of bearing damage", Ref. 4.5-10):
Excerpt: “The crash of an …airliner…was caused by a chain reaction disintegration of its left engines, damaged elevator controls, severed electrical cables and an uncontrolled fire.
The aircraft, which is powered by four aft-mounted…turbofans had climbed about 8 200 meters when a worn ball bearing in the left inside engine caused eccentric rotation of a high-pressure turbine shaft.
A faulty support for the crankshaft gave way under this stress, causing the shaft to smash into a low-pressure turbine that shared the same support.
The blow of the low-pressure turbine's crankshaft severed the compressor that it drove and caused the turbine's fan to spin out of control. Centrifugal force disintegrated the fan.
Pieces from the fan shredded the outboard engine and also pierced the aircraft's fuselage..”
The phrase “low-pressure turbine's fan” is assumed to mean the engine fan.
The term “crankshaft” most likely refers to the shaft support. Evidently, damage to the low-pressure compressor, which is located on the same shaft as the fan, decreased the power received by the compressor until the fan was accelerated to overspeed by the turbine and disintegrated. It is still unclear why the damage to the low-pressure compressor did not reduce the power output from the turbine due to a decrease in the air flow, which would have prevented fan overspeed.
Example "Fan decoupler" (Ref. 4.5-15):
Excerpt: “CFM's new technology includes a fan “decoupler” to protect the pylon and airframe from out of balance forces in the event of a fanblade departure. The decoupling would probably be achieved by fuse bolts.”
Comment: Here the fuse bolts/shear bolts at the aeroengine mounting may be concerned. These act at overload as predetermined breaking points. Thereby arises the big problem to keep these sufficient safe for the normal to bear high dynamic loads (Ill. 10-8). This demands an oversizing against short time loads. At the other hand the connection react at these with a forced fracture. In contrast to the, primarily as bending acting operation forces, the forced overload acts as shear at the bolts (see volume 2, example 10-13).
Figure "Overspeed prevention" There are several strategies for preventing runaway turbines after shaft failures. This diagram shows a mechanical variation which makes use of the axial offset that follows shaft separation. Naturally, this assumes that the shaft failure occurs behind the fixed bearing. During offset, the shaft moves a bolt (B) rearward, which actuates the fuel control unit by means of a linkage assembly.
One problem with this design is the possibility that lubricating oil could get into the clearance gap of the bolt, which juts into the bearing chamber. If the oil cokes up, it could block the bolt.
For this reason, and also to make use of the faster responses of modern electronic control units, attempts are made to use the axial movement of the shaft in this case, as well. One possibility would be to have the shaft movement split a conductor with through-flow.
Figure "Intermeshing principle" If a rotor is braked by rubbing against a closed ring surface, then a lubricating pool crater may form and diminish the braking effect. This can be prevented if the surface provides non-continuous contact, such as that between blade rows (intermesh-construction).
One example is breaking the low-pressure turbine rotor after a shaft failure. The axial offset causes the outer rotor blades to rub against a specially designed stator vane stage. The contact forces and the decrease in power output due to the damaged blading must be sufficient to prevent the rotor from reaching uncontrollable overspeed. To ensure this, the blades must absorb sufficient energy and must not fail, e.g. through a rapid haircut. The braking effect is increased by the damage to the blades, which decreases the drive torque. Additionally, the circumferential stress on the blade decreases, lowering the risk of a the disk bursting. The blade design and material selection are very important for minimizing brittle fractures and ensuring sufficient braking energy. The above factors must also be considered when implementing single-crystal and directional-hardened materials.
Figure "Housing deformations" (Ref. 4.5-11) The number of fragments resulting from a blade failure leaves typical deformations and holes in housings/casings. The above diagrams are from burst tests in spin test rigs. The most common scenario is depicted in the top diagram. It is a symmetrical disk split due to a hub fracture. The movement of disk fragments has different translatory and rotatory components, depending on the fragment type. These determine the damage symptoms of the housing. In some cases, this makes it possible to make conclusions about the fragment size and shape based on the damage symptoms (see Fig. "Damage symptoms of the housing").
Figure "Damage symptoms of the housing" (Refs. 4.5-12 and 4.5-13) The top diagram shows the typical fragments of a failed integral turbine disk of the type used in smaller low-power engines. The cracks labeled “1” are the primary cracks that led to the hub bursting. Cracks “2” are consequential damage caused by the fragments striking the housing. The typical damage symptoms of this type of turbine housing are shown in the bottom left diagram. The symmetrical nature of the damage should be noted. The symptoms of fan damage in the engine of a passenger aircraft are similar, leading to the conclusion that this case was disk failure in which the disk broke into two roughly equal parts.
Figure "Predetermined breaking point" Predetermined breaking points are used in cases where damage limitation is the goal (Ref. 4.5-14). They must break at a certain defined load limit. Broken predetermined breaking points are therefore always an important indicator of high load levels, i.e. that the defined load levels were exceeded due to a malfunction or damage. Naturally, it must be shown that overstress due to the designated loads is plausible (e.g. shaft overstress due to overly high torsion; suspension bolt failure due to shear stress) and is not the result of the crash or other factors that are separate from the intended function of the predetermined breaking point.
Predetermined breaking points are found in different parts of the engine. Engineers are considering whether or not to use predetermined breaking points in large fan engines so that the fan could be separated from the engine by explosive bolts in case of fan blade damage followed by unbalance and vibrations (e.g. similar to rocket technology, Example "Fan decoupler").
There have also been discussions about predetermined breaking points on engine suspensions, which would allow controlled jettisoning of the engine in case of overstress. Predetermined breaking points are also used in the power systems of auxiliary components (e.g. pumps, starters, generators, regulators). When these devices are blocked, their connection is severed at a predetermined point in order to protect other functions and/or to prevent uncontrollable consequential damages. When fuel-carrying components are blocked (e.g. due to broken gear teeth, bearing damage, or DOD such as oil nozzle fractures), there is a risk of seals and housings being damaged so that fuel could leak out uncontrollably. The escaping fuel can cause fires. If the starter is blocked or the engine is sluggish during start-up, predetermined breaking points can prevent greater damages. Predetermined breaking points are always a challenge for the designer. While they must be able transfer all normal operating loads with sufficient safety, they must also guarantee separation at the break point at a certain defined overstress level. The operating loads are dynamic and act over long operating times. Therefore, the predetermined breaking point must have a high dynamic strength. This demand can be met through sufficient fatigue resistance of the surface. This can be accomplished by plastic strengthening (e.g. shot peening or rolling) or by surface hardening (e.g. nitriding or carburizing). The ability to shear off in case of overstress is ensured by a sufficiently hard and tough core section.
Figure "Deposits" The illustrated example shows a main engine bearing after capital damage. Usually, this type of damage would destroy the bearing so completely that the causes could only be speculated about, unless there are recorded data available (e.g. temperatures, oil flow, vibrations, etc.).
However, in some cases, additional information can be gained from the structure, composition, and layering of the wear products that were created during the damage process.
“A” is the heavily damaged balls of the bearing. The balls and the bearing tracks show extensive wear that has changed their contours considerably.
In section “B” of the bearing casing, there is a thick coat of oil coke which is full of shavings. This coat is made up of several layers which correspond to the times at which the oil coke and shavings were created. The oil content rises considerably together with the temperature increases in the bearing.
“C” is a detail of the coating cross-section near the bearing casing wall, i.e. the beginning of the deposits with fatigue-induced outbreaks from the roller bearing tracks.
The next layer, “D”, contains bonded shavings and the top layer (the end of the damage process) contains slugs from the contact surfaces of the bearing that reached extreme overtemperatures.
This all explains the damage sequence, in which pittings formed in the tracks of the main bearing, then the bearing cage became blocked, and finally the bearing function was similar to that of a friction-type bearing.
Deposits that can explain damage sequences:
The buildup of deposits can indicate the temporal progression of damage. This is true both for deposits that were carried onto a surface by the air or gas flow, as well as deposits in flowing fluids (e.g. fuel and oil).
Deposits from the air and gas flows:
Many damages in the air and gas flow result in solid wear products and/or melted particles. Solid particles are created by processes such as labyrinth rubbing or blade tips rubbing into abradable coatings. These particles then stick to the surfaces they strike, either through local fusing, or they melt when traveling through the combustion chamber and solidify on the relatively cool turbine blades. The composition of these particles can be very different and have characteristics that determine their origin.
Solidified drops or splashes of melted metal can be analyzed and often traced back to their point of origin. The material may have melted due to the rubbing process itself, due to a fire (e.g. titanium fire), or by passing through the through the combustion chamber.
Deposits in the oil flow:
Deposits in the oil flow (Fig. "Deposits") can be expected in “eddys” and suitably shaped surfaces in the back oil flow. Also, due to their functions, oil filters and magnetic plugs are predestined locations for analyzable deposits.
Deposits in the fuel flow:
These deposits can be organic overheating products, wear products (e.g. from a fuel pump, see Fig. "Statistics"; or a galling regulator component), or separated synthetic materials (e.g. unsuitable O-rings). Deposits found in a fuel system could, for example, explain heavy wear in an injection system (e.g. injection nozzle).
Figure "Statistics" (Example "Fuel with poor lubrication properties", Ref. 4.5-16): In order to do a statistical analysis, it is important that there is a sufficiently large number of “suitable” cases of damage. With isolated occurrences or first damages, no sufficiently meaningful statistics are likely to exist. Damage statistics can be helpful for understanding damage mechanisms and risk assessment, i.e. determining inspection intervals.
The Weibull diagram shows the distribution of leakages over the life (up to the time of damage) of an axial piston pump.
The scatter can be seen in the angle of the line, which is the Weibull parameter. The greater this parameter is, the steeper the line becomes, and the scattering is smaller. Fatigue and wear processes usually have large scatterings (small Weibull parameters) compared with other damage mechanisms. The fact that the lines are parallel shows that both have the same scattering, which indicates a causal relationship between them.
The top right diagram is a Woehler diagram for the life span of dynamically stressed parts. Evidently, the life span scattering is relatively small at high load amplitudes (LCF zone). If the loads are slightly greater than the fatigue strength (HCF zone), then the same load differences result in a considerably greater scattering of the life span. One can also make conclusions regarding the size of the dynamic loads and specific damage causes based on the life span scattering of dynamically stressed parts.
For example, if abbreviated tests (i.e. harsher test conditions than during actual operation) of an improved pump version show the same scattering (i.e. angle of Weibull parameter) as the original part, it shows that the tests are suitable for verifying that the part was improved.
Example "Fuel with poor lubrication properties" (Ref. 4.5-16): Premature wear of the afterburner fuel pumps was found in the engine of a military fighter aircraft. The engine had to be disassembled within 150 to 800 operating hours. This considerable effort was undertaken in order to understand the damage mechanism and ensure that any improvements to the pumps would be successful.
The damage occurred during the energy crisis of the 1970s, when highly refined fuel with poor lubricating properties was used. The poor lubrication caused problems in the afterburner fuel pumps:
The pistons and piston rings of these axial piston pumps wore heavily. This frequently led to broken piston rings, which allowed the fuel to escape from the high-pressure area into the pump housing, where it caused strong vibrations. The vibrations caused housing fractures and fuel leaks that posed a great risk for the aircraft. Because of this, the pumps were inspected every 30 hours for signs of a drop in pressure which could indicate serious wear on the piston ring and groove. The scattering was very large, at m= 2.6 (36 failures in 800 engines with 11 external fuel leaks due to housing fractures), and was typical for a serious wear process (Fig. "Statistics"). The scattering of the run times with fuel leaks was m= 2.9, which would correspond to a damage mechanism with dynamic fatigue. The difference between the characteristic life span (see references for calculations) of the drop in pressure and the housing fractures was about 640 operating hours, which could be seen as the safety interval until the constructively modified pumps were available.
Figure "Incubation time" Similar to natural processes (e.g. in medicine and biology), the term incubation time refers to the time span between the start of an influence and its characteristic damaging effect. During the incubation time, damaging changes occur that do not appear alter the macroscopic outer behavior of the engine in any damaging way. Typical damage mechanisms with pronounced incubation times include dynamic fatigue, erosion, and corrosion processes.
During dynamic fatigue in metals, microscopic changes take place in the structure, including strengthening, weakening, and crack initiation, which cause macroscopic crack growth at the end of the incubation time.
During the incubation period of an erosive process, parts may begin to shatter or erosion particles can become stuck in the surface. These processes can considerably decrease material removal for a short time, which is a characteristic sign of erosion. There are even reported cases where the sticking particles increase the weight of the eroded sample (Fig. "Unexpected behavior").
Often, protective oxide coatings form during the incubation time of corrosive processes. The incubation period ends when the medium can directly damage the base material.
Incubation time is used in order to achieve acceptable life spans or reliability for components with limited life spans. One example is the proportion of the life span that is taken up by the incubation time of engine parts under LCF stress.
Threshold value (Fig. "Cause of damage") refers to the limit value of a damaging influence, beyond which damage actually occurs. Damage mechanisms with one or more damaging influences with different threshold values are quite common. With fatigue processes, fatigue resistance is a dynamic load threshold value. Dynamic fatigue fractures will only occur above this value. Similar threshold values can be observed in parts with internal and external notches (internal: e.g. micro-cracks, cavities, flaws; external: e.g. grooves, structural changes) with regard to a certain stress concentration (see Fig. "Cause of damage").
A classic example of threshold behavior is stress corrosion cracking (see Chapter 5.4.2). The typical crack initiation occurs only after a certain tensile stress level has been reached with material conditions that are sensitive to the corrosive medium.
During erosion, a threshold value may be dependent on the hardness of the surface relative to the hardness of the erosion particles that are acting on it (Fig. "Influence of hardness of particles").
Threshold behavior is of foremost importance for engine technology. This effect is used when designing fatigue resistant parts, estimating the acceptability of certain (unavoidable) weaknesses and flaws characteristic of the technology during design, and conducting risk assessments.
The way in which the temporal damage sequence occurs after the incubation period can depend on various parameters, which must not necessarily by the same as the parameters that determine the threshold value.
For example, if the damage sequence does not involve corrosion, but rather cyclical crack growth due to dynamic loads, other structural characteristics will most likely be the determining factors.
4.5-1 “Making Safety Global”, periodical “Aerospace Engineering” , December 1998, pages 35 and 36.
4.5-2 G. Warwick, “Pilot error still main cause of US accidents”, periodical “Flight International” 10-16 March 1999, page 24.
4.5-3 NTBS Identification SAE96FA076, Index for April 1996.
4.5-4 G. Fischbach, “916 Deutsche F-104 Starfighter, ihre Bau- und Lebensgeschichten”, page 681.
4.5-5 G. Fischbach, “916 Deutsche F-104 Starfighter, ihre Bau- und Lebensgeschichten”, page 721.
4.5-6 G. Fischbach, “916 Deutsche F-104 Starfighter, ihre Bau- und Lebensgeschichten”, page 724.
4.5-7 NTBS Identification NYC96LA188, Index for September 1996.
4.5-8 Interavia Air Letter, “RB.211 Service Bulletin Follows Fan Failures”, No. 9849, October 6, 1981.
4.5-9.1 “Rolls Investigating RB.211 Fan Shaft Failure”, periodical “Aviation Week & Space Technology”, January 31, 1983, page 30.
4.5-9.2 “RB. 211 Incidents Prompt New L-1011 Procedures”, periodical “Aviation Week & Space Technology”, October 5, 1981, page 36.
4.5-9.3 “Rolls Accelereates RB. 211 Modification Plan”, periodical “Aviation Week & Space Technology”, September 28,, 1981, page 29.
4.5-9.4 “Rolls-Royce Tests RB.211 Engine Modification Kits”, periodical “Aviation Week & Space Technology”, December 14, 1983, page 30.
4.5-10.1 “Investigators Cite Engine Failure In Polish Ilyushin Il-62M Crash”, periodical “Aviation Week & Space Technology”, August 3, 1987, page 57.
4.5-10.2 “LOT grounds Ilyushin Il-62Ms”, periodical “Flight International”, 22 August 1987, page 6.
4.5-11.1 C.J. Mangano, “Studies of Enginerotor Fragment Impact on Protective Structure”, AGARD CP-248, 1978, Chapter 10.
4.5-11.2 E.A. Witmer, T.R. Stagliano, J.A. Rodal, “Engine Rotor Burst Containment/Control Studies”, AGARD CP-248, 1978, Chapter 15.
4.5-11.3 J.T. Salvino, G.J. Mangano, R.A. DeLucia, “Rotor Burst Protection: Design Guidelines for Containment”, AGARD CP-248, 1978, Chapter 19.
4.5-12 L.M. Jenkins, S.E. Crow, “ RB-211-524B Disc and Drive Cones Hot Cyclic Spinning Test”, AGARD-AR-308, Published September 1992, pages 67-70.
4.5-13 J.Ott, “JT8D Hub Failure Sparks Intense Inquiry”, periodical “Aviation Week & Space Technology”, July 15,, 1996, page 29-41.
4.5-14 “CFM56-7 failures spark FAA action”, periodical “Flight International”, 8-14 July, 1998, page 8.
4.5-15 “CFM56 replacements”, periodical “Aerospace International”, April 1998, News Index, page 8.
4.5-16 C.H. Medlin, F.L. Elsaesser, “Weibull/Weibayes Analysis of Hydraulic Pump Malfunction Data”, SAE Technical Paper 831542, presented at Aerospace Congress& Exposition Long Beach, California, October 3-6, 1983.
4.5-17 periodical: “Aviation Week & Space Technology” April 12, 1982, page 34.
4.5-18 “Lycoming spends $ 30 million on LTS101 turbine”, periodical “Flight International” 14. October 1989, page 10.