Table of Contents
5.3.2 Erosion Damage
Erosion damage can occur on many different engine components (Fig. "Parts effected by erosion") and take many different forms. Typical effects are:
- Worsening of the blading (profile changes, increased roughness, increased tip clearance) and therefore also the operating behavior of the compressor leads to temperature increases with correspondingly high turbine temperatures.
- Erosion in labyrinths can also cause performance decreases, but it can also change the bearing thrust or cause oil to escape from the bearing chambers.
- Erosion wear products can block the cooling air ducts of the hot parts and raise their temperature, considerably reducing the life span of the hot parts (especially turbine rotor blades).
- Dynamic fatigue fractures on the rotor blading in the compressor and turbine due to cross-sectional weakening and notches, in some cases combined with flow stalls due to changed profiles and increased roughness.
- Overheating damage in the combustion chamber due to erosion damage to the injection system caused by coke resulting from overly high fuel temperatures.
- Oxidation damage to turbine blades and stator vanes (Fig. "Erosion of turbine blades") resulting from erosion damage to a protective coating (caused by soot particles, ceramic particles from thermal insulation layers, etc.).
In the compressor inlet dust is usually still distributed evenly across the cross-section. The farther the dust particles are inside an axial compressor (Fig. "Erosion sensitivity of different compressor designs"), the more they are centrifuged outward, resulting in the dust and therefore also the erosion damage being concentrated more and more at the blade tips (Fig. "Extreme erosion stress during sand storms"). A large amount of dust may pass through the bypass duct and not enter the core engine.
In cases where air is extracted from the compressor to cool the hot parts or cabin air, extraction takes place on the rotor side whenever possible, since the air in this area is relatively clean.
Figure "Parts effected by erosion": Many different engine components can be directly or indirectly damaged through erosive particles:
“E1” fan and LPC blading: Roughening through material removal and/or deposits (fouling), changes in blade profiles, increased tip clearance, blade fractures caused by dynamic fatigue fractures originating in erosion notches.
“E2” housing/casing coatings in the fan and LPC: The erosion material removal causes increased tip clearances to the rotor as well as steps and roughness in the path of the gas flow.
“E3” HPC blading (Fig. "Typical erosion on compressor blades"): see “E1”.
“E4” HPC housing coatings ( Fig. "Typical erosion on compressor blades"): see “E2”.
“E5” spacer coatings (Fig. "Singularities in the gas flow"): Erosion of the relatively hard spacer coatings increases the tip clearances between the rotor and compressor stator blades with no inner shroud. Additionally, the roughness of this area of the hub contour can be increased.
“E6” injection system: Erosion of the injection nozzles through coke formation in overheated fuel. This can result in overheating in the combustion chamber area, even including flames escaping from the combustion chamber housing/casing.
“E7” turbine blading (Fig. "Erosion of soft coatings"): Erosion and increased oxidation especially at the leading blade edges caused by dust, coke from the combustion chamber, and spalled ceramic thermal barriers.
Blocking of cooling air ducts and cooling air film bores, which can cause local overheating.
Chemical reactions and erosion material removal in ceramic thermal barriers.
If dust collects in a fir-tree joint, it can unallowably alter the dynamic behavior of the blades (Ref. 5.3.2-1).
“E8” bearing damage (Fig. "Axial face seals"): Erosion in bearing casings and gas ducts caused by circumferentially-traveling trapped particles of ingested dust and labyrinth material wear products. Weakening of wall cross-sections, even causing the walls to fail.
“E9” labyrinths (Fig. "Axial face seals"): Erosion in labyrinths, especially on softer abradable coatings.
“E10” air extraction ducts and air ducts (Fig. "Volcanic ash"): Erosion of walls, especially in the outer area of pipe bends.
“E11” Variable vane adjustment: Dust in the guide vane bearings can cause jamming and wear.
“E12” blade root contact surfaces: In rotor blades and guide vanes, an abrasive process can increase fretting wear due to dust that has penetrated into these areas.
Changes in the friction relationships can alter the dynamic behavior of the blading (Ref. 5.3.2-1).
“E13” Elastomeres: Concerned are rub in coatings from filled silicone rubber. This is used in the front region of the compressor (fan, booster, LPC) at interstage seals. This coating usually is located on the inner rings of the stators opposite the labyrinth tips on the rotor. There is the suspicion, that traces of wear on these coatings are in connection with ice formation. The exact mechanism seams not yet clear. However it is known from overhaul processes, that particles from water ice can be used as blasting abrasive to remove soft contaminations (Ref. 5.3.2-25). In this case however the provider calls it a non abrasive process. An abrasion effect however may merely depend from the process parameters. This is also true if the ice contains abrasive particles (e.g., dust) as contamination.
Figure "Erosion damage on suction an pressure side": Changes to the blade profile at the micro-level, especially the leading edge, can affect the aerodynamic behavior of the blade considerably. The impact of sand grains causes notches and burrs, as well as increased roughness primarily on the pressure side of the blade. On the suction side, especially, burrs promote local flow stalls (bottom diagram).The consequences include a decrease in compressor efficiency and also increased dynamic loads on the individual blades. Combined with the notch effect of the impact notches, this increases the risk of fatigue fractures. Clear signs of high dynamic loads (extrusions, see Fig. "Coating damage I") in erosion notches were verified through use of scanning electron microscopes (SEM).
Figure "Behavior of shaft-power engines" (Ref. 5.3.2-2): The erosion-induced changes to the compressor blades, such as increased roughness and/or changes to the blade profile at the micro- and macro-levels, result in a considerable worsening of the compressor efficiency. This, in turn, affects the operating behavior of the engine and the damage frequency and life spans of other components, especially the hot parts. In order to maintain the same engine performance, the efficiency decrease must be compensated by increasing the gas temperatures. This raises the temperatures of the hot parts (combustion chamber and turbine). In addition, the decreased play to the surge limit can cause it to be crossed, leading to extreme short-term overheating in the turbine with extensive consequential damages.
Example "Resonance" (Ref. 5.3.2-19):
In a 400 kW class helicopter engine with a side inlet, after longer operating times there were cases of dynamic fatigue fractures of individual compressor rotor blades made from the titanium alloy 6Al 4V, resulting in extensive secondary damages. The single-engine helicopters were operated in an environment with dust conditions, that were at least equal to those in a north African desert. Due to costs and efficiency demands, the engines were not equipped with filters or sand separators. The blades, which had been run for long periods, showed serious erosion and notches near the root platform. This is related to the side inlets and the corresponding path of the particles in the inlet flow (Fig. "Blade tip hooking", right diagram). After a full overhaul interval, the blades were eroded down to 50-70 % of the normal profile thickness. This changed the resonant frequency of the blades so much, that it resulted in resonance and dynamic fatigue fractures.
Figure "Typical erosion on compressor blades": The life span of a helicopter engine depends largely on the presence of a dust filter in front of the engine inlet and the separation effectiveness of the filter (bottom right diagram, Ref. 5.3.2-18). It can be seen that only filter systems with high separation rates permit desirable engine life spans.
The top left diagram is a schematic depiction of the typical erosion symptoms in a rear compressor stage. The centrifugal effect of the blading in the air flow concentrates the particles near the casing wall. This makes the erosion effects more pronounced in radially farther outward areas. This especially weakens the root areas of stator blades (“1”), promoting dynamic fatigue fractures in case of flexure. However, the clearance between the rotor and blades without inner shrouds is not affected, minimizing the efficiency decrease in this part of the compressor.
The rotor blades, on the other hand, are worn primarily at the tips (“2”), where the clearance increases are spread over a larger circumference, affecting the compressor efficiency more than the stator blade erosion.
In addition, increased erosion of the relatively soft abradable coatings on the inside of the casing (“3”) further increases the clearance at the rotor blade tips.
With stator blades with inner shrouds and soft abradable labyrinth coatings facing the rotor (“4”), there is a danger of clearance increases due to erosion, although the erosion stress is lower than near the rotor blade tips (lower dust concentration near the hub).
The erosion of the blade profile, with roughening of the leading edge and burr formation on the suction side (“5”), promotes local flow stalls and therefore also compressor stalls.
The erosion behavior (erosion sensitivity) varies considerably between different compressor designs (see diagram bottom left, Ref. 5.3.2-2). Axial compressors (“A”) show the greatest sensitivity. Combined axial-radial compressors (“B”), such as are common in helicopter engines (Fig. "Erosion sensitivity of different compressor designs"), perform better. Side-mounted inlets can result in the blades being damaged especially at the highly dynamically stressed transition to the root platform (sketch above right, Example "Resonance"). Machines with pure radial compressors show the most robust behavior (“C”). This positive performance can be further improved through erosion protection on disks and diffusers (“D”). The horizontal line “E” shows the operating limits of the engine according to regulation AS/AV E-85 930.
Example "Air bleed" (Ref. 5.3.2-3):
Excerpt: “severe trailing edge erosion of an (helicopter engine) 5th and 10th stage compressor blades has been observed in Field Service Evaluations and sand ingestion tests. The erosion typically occurs within 3 millimeters of the blade tip and has been attributed to overtip flow produced by the bleed holes present over the 5th and 10th stage rotors… (measurements of the air flow showed). The flow was found to be unaffected by the bleed holes beyond 3-4 mm from the blade tip.
Comment: An increased erosion at the blade tips of the rotor blades below the air extractions respectively openings for air bleeds are observed on-again (Fig. "Blade tip hooking"). Here obviously the dust acts especially deteriorating, concentrated by the centrifugal force, supported by the disturbed flow.
Figure "Profile change due to erosion" (Ref. 5.3.2-4): “Rollover” as an extreme profile change. The leading edge of a helicopter engine compressor rotor blade has been rolled up by sand erosion.
Example "Tip clearance" (Ref. 5.3.2-7):
Excerpt: “To eliminate compressor stalls caused by deterioration of unit's abradables, (the engine manufacturer) recommended repositioning the compressor's sixth stage vanes and rescheduling the unit's variable geometry through a software change in the engine's electronic control unit.”
Comments: The affected military engine type had massive problems with compressor stalls (surges?) during operation, because the tip clearances in the high-pressure compressor had become too large due to erosion (Example "Rub coating wear"). This evidently contributed to inquiries as to whether the involved aircraft type would be better outfitted with engines from a competing company.
Figure "Delamination of coatings": Coatings on engine parts, especially blades, can locally erode or be damaged by foreign objects so that the flow, combined with blade vibrations, cause the coating to delaminate. This is especially the case with ductile coatings such as organic erosion protection coatings (top diagram) or metallic foils such as damping foils (bottom diagrams), in which the bond with the base material is considerably weaker than the coating strength. This kind of delamination or rolling-up of the coating causes considerably more flow interference than superficial material removal or flaking. The delamination of larger coating areas also poses a threat of damage or blockages in the rear engine area (for example, the cooling ducts of hot parts).
Figure "Blade tip hooking" (Refs. 5.3.2-3, 5.3.2-2, and 5.3.2-17): In blade tips, especially those located in turbulent flow zones such as around areas where air is extracted through bleed valves(Example ""), a typical erosion symptom known as “hooking” can appear in the rotor blade tip area. This causes increased wear of the trailing edge below the corner of the blade.
Figure "Erosion patterns on radial compressor disk" (Ref. 5.3.1-5): Radial compressors are usually particularly erosion resistant (Fig. "Typical erosion on compressor blades"). Depending on the upstream configurations, such as radial or axial inlets and axial or radial compressors, the erosion symptoms on the blading can vary greatly (Fig. "Particle size"). In this case a multiple-stage axial compressor is located upstream. During an erosion test with coarse dust a sudden power decrease of 15% occurred. This was followed by the mechanical failure of several blades.
Figure "Extreme erosion stress during sand storms": This diagram illustrates the technical relationships in the extreme erosion stress mentioned in Example "Sand storms", which can evidently occur during a sand storm. The engines involved were shaft power engines of the 3000 kW class. The blading of the axial compressor in this engine type is made from a high-strength titanium alloy and is relatively filigreed. The involved helicopter is not known to have been equipped with any particle separating devices such as filters. Therefore it can be assumed that the compressor blading experienced considerable profile changes of the blades (similar diagram at bottom left) due to the sand erosion. In this case an unallowable influencing of the engine`s operating behavior and power output was concluded to have occurred.
Example "Sand storms" (Fig. "Extreme erosion stress during sand storms"):
Excerpt 1 (Ref. 5.3.2-5): …“a table… indicated, by location and month, the frequency of suspended dust occurrences. Helicopter pilots, however, were surprised when they encountered the dust, were unprepared to accurately assess its impact on their flight, and stated that they were not advised of the phenomenon.”
Excerpt 2 (Ref. 5.3.2-6): “The remaining (helicopters) were also having problems after running into two totally unexpected and vicious sandstorms. The one flown by the force commander… and No2 chopper made forced landings to ride out the worst of the storms. Luckily, after over 20 minutes on the ground, the storms passed and both crews resumed their journeys…. No5 (helicopter) suffered a major mechanical failure and, unable to carry on with the mission, was forced to return…“
Both excerpts refer to the hostage rescue operation in Tehran on April 24, 1980. The helicopters which were used were a type without filters or separators in front of the engines. Evidently the sand storm danger was estimated to be relatively minor. According to Ref. 5.3.2-6, the crew was forced to land due to poor visibility.
It is also interesting how difficult even short-term prediction or location of a sandstorm is. The bottom right diagram in Fig. "Extreme erosion stress during sand storms" (Ref. 5.3.2-8) shows that in April, the sand storm frequency on the neighboring Arabian Peninsula is roughly one in every 60 days. Evidently this relatively probability was sufficient to select a helicopter type with no particle separator. This decision may have been supported by demands for long flight range, which would not allow the increased fuel consumption caused by the inlet resistance of a filter.
Figure "Singularities in the gas flow": Engine-specific singularities in the gas flow direction can affect the erosion symptoms and the especially erosion-stressed components in characteristic ways. In engines that have a S-shaped housing/casing ducts (swan neck duct, upper figure - left sketch) between the high-pressure and low-pressure compressors, particles are evidently deflected towards the hub. This can be understood by considering the particle deflection in a radial compressor (see Fig. "Erosion sensitivity of different compressor designs"). The first compressor spacer (upper figure - left sketch) after the swan neck duct is especially erosion stressed. This stress should decrease for the following spacers due to the repeated outward centrifuging of the particles as they travel farther through the compressor. Experience seems to confirm this. The erosion symptoms of the spacers usually show the spacing of the upstream rotor blades in a more or less pronounced manner. The blades maintain their position to the spacer. Experience shows that the particles collect on the pressure sides of the blades.
The sketches on the right side are a schematic depiction of typical erosion symptoms on soft housing coatings (such as Ni/graphite spray coatings). The darker zones are the more heavily eroded areas. Similar to the case with rotor spacers, the erosion pattern corresponds to the upstream blades right side upper sketch left), although in this case the blades are the stator vanes, which do not move relative to the housing coating. Evidently the particles can also travel along the stator vane profiles and become concentrated. This typical damage symptom is usually not very pronounced when the rotor blades run-in heavily (bottom - bottom figure).
Example "Rub coating wear" (Ref. 5.3.2-22):
In the moment of lift off at the left aeroengine a loud bang occurred. After this in about one minute further 50 bangs followed. Concerned are surge shocks (flow stalls in the compressor, Ill. 22.214.171.124-1). In the same moment the low pressure rotation speed of this three shaft aeroengine dropped. Contrary the rotation speed of the medium pressure rotor and the high pressure rotor increased. After dumping the fuel the aircraft landed again about 1 hour later at the departure airport.
Already about 3 weeks before the engine health monitoring of the aeroengine (= EHM, Ill. 25.1-5.1) showed an increase of the turbine gas temperature (TGT, Fig. "Behavior of shaft-power engines" and Fig. "Seals and engine performance"). However the following borescope inspection is only limited possible. It produced no alarming indication.
An investigation of the aeroengine by the OEM, together with the responsible aviation authority, determined markedly erosion of the rub in coatings (porous NiCrAl-Bentonit) from the casing rings of the high pressure compressor (HDV, detail sketch in Fig. "Singularities in the gas flow"). An inspection of further aeroengines showed two more cases with a similar result. The enlarged tip gaps of the rear HPC stages caused an efficiency drop and triggered the surge (Ill. 126.96.36.199-8). They also explained the rise of the gastemperature.
In dispute remained the cause for the unusual heavy erosion attack. The aviation authority called the deviations in the quality of the rub in coatings (thermal spray coatings) to account. These promote an increased oxidation and with this the erosion during long time operation. The OEM rather believed at an unusual amount of ingested dust.
As remedy a visual inspection of the high pressure compressor, for coating erosion was ordered, if a step like increase of the gastemperature is registered by the EHM-system.
Figure "Erosion of soft coatings" (Example "Extreme reduction of lifespan", Ref. 5.3.2-7): The erosion of the usually soft abradable coatings in compressor housings/casings is a common problem in modern engines. The high power concentration at the best efficiency levels and good operating behavior demands, among other characteristics, minimization of the compressor rotor blade tip clearance. This can only be achieved by having the blades run-in when the clearance gap is breached. In this case, the blade tips must not be unallowably damaged. This requires soft coatings. Evidently even a multi-stage fan in a bypass engine is not sufficient for centrifuging all the ingested particles and preventing the core engine from experiencing any erosion stress. It can be assumed that, along with the particle concentration which increases near the housing, housings that become conically smaller along with the gas flow promote the erosion of housing coatings as in the depicted case (middle and bottom diagrams).
Figure "Erosion of turbine blades": Particles from the combustion chamber and other engine zones, located upstream in the engine, can cause serious erosion on the pressure side of turbine stator blades and the suction side of turbine rotor blades (also see Figs. "Influence of impact direction", Erosion of turbine blades" and "Thermal barrier coating I"). Typical particles include soot from the combustion chamber (Example "Improved combustion liner") and particles from crumbling ceramic thermal barriers, rub coatings of the seal segments, or hard armor coatings on blade tips and labyrinths.
Even damage to thin diffusion coatings (typical thickness less than 0.1 mm) can result in oxidation of the base material, promoting damaging erosion which can considerably reduce the life spans of the parts.
Figure "Low coke combustor" (Example "Improved combustion liner"): To minimize the life shortening erosion by coke from the combustion chamber of this helicopter aeroengine, a version of the combustion chamber with less coke formation (low coke combustor) was introduced. This combustion chamber has a new designed effusion cooled inner wall. So there was the hope to double or triple the life time of the blades. At the same time also the visible smoke production should be minimized. The effusion cooling enables a more even temperature distribution without hot spots. With this, the lifetime of the inner wall may be markedly extended.
Figure "Dust particle melting": The material removal from a hot part surface can be highly accelerated by a combination of oxidation and erosion. Brittle oxidation layers are especially erosion sensitive at steep impact angles (Fig. "Influence of angle of attack"), such as those expected on leading edges of blades. The erosion material removal from the protective oxide coatings exposes reactive fresh metal surfaces which oxidize correspondingly heavily.
A similar scenario (Ref. 5.3.2-17) occurs when dust particles are melted in the combustion chamber and then strike and stick (glassing, Ref. 5.3.2-4) to the relatively cold cooled surfaces of the hot parts, especially the blades.
The melted dust can cramp and/or react with the relatively coarse oxide layers, creating a powerful bond. Thus especially the flow through cross section at stators (nozzle) of the high pressure turbine can be clogged, so that during a power increase the aeroengine can easy get into a surge (Ref. 5.3.2-4).
When the engine is shut down, these layers harden into brittle coatings with crack initiation and break off partly or completely along with the protective oxide layers (Ref. 5.3.2-20)..
When engine operation is resumed, the fresh metallic surfaces oxide especially heavily, and the process is repeated when dust again enters the engine.
Example "Extreme reduction of lifespan" (Ref. 5.3.2-8, Fig. "Desert environments"):
Excerpt: ”…(in) North American desert environments. Throughout these operations there was no indication of any adverse engine effects associated with erosion or corrosion of turbo machinery components or accessories.
However, problems were already being experienced with the HP and IP Turbine blades, both being cast equiaxed IN100 material.
The HPT blade was suffering shroud loss and airfoil failures due to thermal fatigue and creep. Regular boroscope inspections were carried out to reduce the instances of blade failures.
The IP turbine blade was suffering failures due to High Cycle Fatigue (HCF), thermal fatigue and in some cases airfoil creep.
…(Use in near east desert environment) The first engine rejection occurred at unexpectedly low life due to High Pressure Turbine (HPT) blade Leading Edge (L/E) burning and (later)…a significant number of engines had been rejected due to similar problems.
Laboratory examination of failed… HPT blades showed that both the external leading edge and internal film cooling holes were being restricted. Measurements of L/E cooling hole outlet diameters revealed varying degrees of restriction from the normal size of 0.3 mm width, in some cases total blockage… The inlet diameters of the L/E film cooling holes could be restricted to as little as 30% of normal diameter.”
Comments: These blockages of the cooling air ducts in the turbine blades caused local temperature increases up to 200 °C. If one considers that an increase of about 15 °C can halve the life span, then the life span reduction here is a factor of about 1000! Therefore, life spans of only a few hours can be expected for these components.
Figure "Desert environments" (Ref. 5.3.2-8): The described three-shaft engine type (top diagram) of a twin-engine fighter aircraft was used in a desert environment. The ingestion of sand and dust into the engine is a serious problem. Interestingly, the greatest problems do not occur due to erosion in the compressor, but in the hot parts. Sand particles are carried onto the turbine blading by the main gas flow (bottom left diagram). Here the bores for the cooling air film are blocked up by the sand, which has been made doughy by the combustion chamber. Further damage to the cooling air configuration of the hot parts is caused by blocking of the cooling air ducts (bottom right diagram) inside the blading. The sand layer stuck to the outside was removed with a blasting device. A probe with compressed air was inserted through the boroscope opening in front of the blades and blasted the leading edges of the blades with abrasive particles (Al oxide).
This kind of cleaning method is naturally only recommended in extreme cases, since there is a risk of the bores for the cooling air film becoming blocked by the blasting particles.
Newer developed blasting processes (chapter 19.2.6) can avoid this disadvantage.
Figure "Volcanic ash": This diagram is based on a press photo (Ref. 5.3.2-9, Example "Glassy, sticky ash layers"). The light discolorations in the nose area can be seen clearly.
In the photos it seems that it was not so much an erosion-induced roughening or delamination, but rather thick deposits of volcanic ash. This is supported by the windows which, as far as can be seen, are partly covered, and not only blind due to the erosion roughening.
Dangers of volcanic ash
The danger volcanic ash poses to modern transport aircraft has been concretely known since the early 1980s. In a span of several weeks in the summer of 1982, two large four-engine aircraft over the eastern part of the Indian Ocean passed through the ash cloud of a volcano. All four engines failed. In these cases, the events were similar to those described in Example "Glassy, sticky ash layers". It only became possible to finally restart the engines at low altitude.
The damaging of the engines by fine ash particles that float in the air (typical particle size, 0.005 mm, Ref. 5.3.2-10), but are very hard and sharp-edged, occurs in various ways:
- Erosion of blading and housings/casings in the compressor area, with worsening of compressor efficiency and operating behavior
- Seizing of contacting surfaces (Ref. 5.3.2-16) such as the clappers of fan blades
- Blocking of air bores and gills in the combustion chamber
- Blocking of air ducts in the turbine blades with resulting overheating damage (Fig. "Danger of volcanoes")
- Blocking of the flow cross-sections in the turbine due to sticking doughy and liquid ash. The deposits can be several centimeters (!) thick. Several hundred kilograms of hard-baked ash were removed from the four engines of one of the affected aircraft. The ash caused an unallowable pressure increase in the turbine, which the already erosion damaged compressor could no longer build up due to the thin air at high altitudes.
- Damage to electrical systems such as generators or ignitions.
The blocking of the hot part cross-sections with ash is especially dangerous, since some volcanic ashes have components that have melting points below 1000 °C (the melting point of ash is usually between 1100 °C and 1200 °C, Ref. 5.3.2-16) and the typical gas temperature in and behind the combustion chamber is about 1300 °C. The melt becomes tough or hardens into a glass-like coating on the cooled and therefore colder surfaces of the turbine blading. The smallest cross-section in the gas flow of the core engine is located in the turbine inlet stator. Even relatively small cross-section blockages (in the mentioned cases about 10%) cause the compressor to be unable to provide the necessary pressure increases in the relatively thin air at high altitudes. This results in a compressor surge (Ref. 5.3.2-16) and flame out in the combustion chamber.
Damage due to ash is not the only danger that a cloud of volcanic ash poses for aircraft engines. These clouds contain gas-like concentration of sulfur and sulfuric acid, the corrosiveness of which corresponds to PH values of 5.5 to 2.0 (Ref. 5.3.2-10). For example, volcanic eruptions such as those of El Chichon or Pinatubo (Ref. 5.3.2-11) released 10 to 20 million tons of sulfur dioxide, which resulted in corresponding amounts of acid in the atmosphere. Even if these gases are ingested by engines, it will not necessarily result in spontaneous damage or even engine failure. However, over long operating times serious corrosion damage can occur (such as sulfidation, Ref. 5.3.2-12, Example "Sulphur-laden ash"), especially in the hot parts. It is questionable how practical and technically possible cleaning the damaged parts is.
Preventing damage from volcanic ash:
The technical literature all agrees that damage can only be safely avoided by not flying through clouds of volcanic ash. The onboard radar does not guarantee safe and sufficiently early detection of the danger. For example, as described in Example "Glassy, sticky ash layers", the cloud of volcanic ash was not recognized by the onboard radar. For this reason, early warning systems such as the Darwin Volcanic Ash Advisory Centre (VAAC, Ref. 5.3.2-13) were established especially in the high-risk geographic regions (for example, there are hundreds of active volcanoes in the Asia-Pacific region). These early warning systems also receive current reports of observations from flight crews in these areas.
Engines of aircraft on the ground are also endangered by ash deposits, which can be blown into the engine by wind and/or sucked in during startup. In these situations, engines must be covered in time. If ash has already entered the standing engines, they must be treated in accordance with manufacturer`s specifications before startup due to the danger of ash burning into the hot parts. Intensive cleaning is usually unavoidable.
In case the crew notices having entered an ash cloud during flight, engine power should be reduced as much as possible in order to lower the gas temperatures and prevent the ingested ash from turning into glass on the hot parts. However, this is not entirely risk-free, since it makes compressor stalls more likely, if significant deposits have already built up on the hot parts. The ignition should be activated during the entire flight through the ash cloud (Ref. 5.3.2-10). Activating all airbleeds enables the compressor to reach higher RPM and decreases the probability of a stall/surge.
Checking a suspicion, that vulcano ash got unnoticet into the aeroengines:
Because it is asbsolutely not certain, that there are external signs which point at vulcano ash, in case of doubt an inspection of the aeroengines after shut down is adviced. For this, depending from the aeroengine type the following possibilities offer, which analysis and assessment assumes expertise and experience.
Inspection of filters of the bleed air for the cabin ventilation. This should take place by photo-optical macroscopic (appearance and binoculars) and if the suspicion confirms, microscopic with a scanning electron microscope (=SEM). With a SEM, shape, structure and composition as characteristic features of particles can be identified. With this the expert gets important hints at risks for the aeroengines. Do the particles contain sulfur and phosphor, they can trigger in the hot parts hot gas corrosion as sulfidation. Glass (high SiO2 content) can melt. With this the danger of blocking and overheating of cooled hot parts exists. This is especially true for the blades of the high pressure turbine.
Borescope inspections to identify changes of the compressor blading and of hot parts. To this belong unusual discolourations, roughness as well as deposits.
Healthmonitoring for the examination of the aeroengine behaviour. Especially interresting are effencies of compressor and turbine as well as if in this connection a possible increase of the gas temperatures exists.
Figure "Danger of volcanoes" (Ref. 5.3.2-15): In some regions in the world (in this example, Alaska) there are many active volcanoes. The dangerous ash clouds can threaten aircraft even at great distances (hundreds of kilometers) and altitudes up to 20,000 meters. In the lower sketch, the cross-section of the engine type from Example "Sulphur-laden ash" shows the areas especially threatened by the ingested ash. In this case glassy depositions of molten dust ('glassing') accumulated at the leading edges of the 1st stage high pressure turbine stator (nozzle, sketch in the middle). This detail shows a nozzle vane with pronounced signs of overheating due to blocking of the cooling air ducts. However these did not occur in the case of the example 5.3.5-5.
Example "Glassy, sticky ash layers" (Ref. 5.3.2-14, Fig. "Danger of volcanoes"):
Excerpt : ”…At 11:50 local time, roughly 126 kilometers north of the airport (Anchorage, Alaska), when passing through an altitude of 7700 meters all four engines of the brand-new Boeing 747-400 suddenly failed. Several of the cockpit instruments went crazy or failed completely. The cockpit crew went into a gliding flight and sent out an emergency distress call.
The crew tried to restart the engines a total of seven times with no success. Only after eight minutes following the failure two of the jet engines (fan engines with large bypass ratios) started up again. By this time the aircraft had already lost 3700 meters of altitude. A short while later the other two engines also resumed operation. Even though many of the instruments were unusable, the crew was able to land the jumbo-jet in Anchorage. There it was discovered that the new paint on the engine hull was strangely roughed up (Fig. "Volcanic ash") as if it had been worked over with rough sand paper.
One day before the flight…took place, 3109 meter high (volcano) Mt. Redoubt had erupted after 21 years of dormancy. It spewed ash up to twelve kilometers up into the air for 45 minutes. The ash cloud passed over Anchorage towards the north…the cockpit crew later reported that the onboard radar did not register the ash cloud. From the cockpit windows it was also not possible to discern it from a normal water vapor cloud.“
Comments: Closer looks at the press photos give the impression that it was not only a roughening of the hull area, but that there were also massive deposits of ash in this area.
It can be assumed that the engines failed temporarily not only due to the erosion damage at the compressor blading, but that the hot parts sustained overheating damage due to blockage of the cooling air ducts, and also that sticking ash seriously damaged/changed the blade profiles and flow cross-sections in the turbine. The fact that restart of the engines was only successful at lower altitudes can be largely explained by the higher air density, and therefore sufficient compression in the compressor with correspondingly larger air throughput at the startup RPM.
As later emerged, the erosion in the compressor was relatively light and did not explain the flame out of the aeroengine. Main reason for this was sticking melt of the ash at the 1st stage blades of the high pressure turbine ('glassing'). An about 1,5 mm thick layer developed at the leading edge and pressure side. With this the flow cross sections narrowed at the critical point. So obviously the flame out of the aeroengines occurred. It is interresting, that nothing is reported about a sign for an aeroengine surge (bangs, shocks). This points at a rotating stall in the compressor (Ill. 188.8.131.52-1) without a total flow disruption. The air flowrate however may be such reduced, that the combustion chamber in the high altitude quenched. A restart of the aeroengines was at least not successful, because of a lucky property of the deposits. These glassy layers chip off large-scale during cooling (Fig. "Dust particle melting") and have been transported with the gas stream, without damaging the following turbine blades (Ref. 5.3.2-23). This lead during the start trials to the widening of the critical flow cross section. Helpful may have been also for the restart the high air density during the start trials at low altitude. So a sufficient compression in the compressor with a correspondent higher airflow rate was possible.
Example "Sulphur-laden ash" (Ref. 5.3.2-12):
Excerpt: “Corrosion caused by sulphur-laden ash from the eruption of Mt. Pinatubo may have cut deeply into the Philippine military's force of about 26 fixed-wing combat aircraft.
…As much as 10 in (25 cm) of ash, much of it the consistency of fine sand, `got into everything' after (the) eruption.
…sulphur in the ash combines with moisture to create sulfuric acid, `which gets into every crack and crevice,'..a consultant…said. He was charged with rehabilitating some of the equipment corroded by the eruption.
Comments: It was not possible to take the aircraft to a safe airport. The volcanic ash evidently also affected standing engines.
Figure "Erosion damage on air duct" (Ref. 5.3.2-15): The walls of this air extraction duct in the compressor eroded at the bend behind the extraction hole during erosion tests on an engine of a middle performance class. The increased erosion attack at the outer radius of the pipe bend is characteristic, since centrifugal effects and the redirection of the flow concentrate the erosion particles in this area.
What should be done if a potential danger through vulcano ash arises.
In larger amounts vulcano ash can cause aeroengines to fail/flame out within minutes. This can produce a dangerous flight situation (Fig. "Danger of volcanoes" and Fig. "Erosion damage on air duct"). Tereby are ash concentrations in the air which are not clearly visible with the naked eye. Contrary however seems to exist a connection between `cloud hazes' (e.g. cirrus clouds) and vulcano ashes/gases. Ash componentsm like sulfur (from SO2) and particles form aerosols. Particles act as condensation nuclei for droplets and ice crystals. In this form seemingly little amounts of vulcano ash in the air which by far are below the limit of visibility have unacceptable economical consequences (Ref. 5.3.2-26). Already during a flight of few minutes are for the outsider seemingly harmless deteriorations possible. To these belongs especially the constriction of cooling air holes. However during further operation the locally increase of the component temperature shortens the lifetime of the hot parts up to several orders of magnitude. Thereby a for the flight safety critical spontaneous failure is rather relatively unlikely. So at low ash concentrations no <U>acute</U> safety problem is concerned. Precondition is, that the online monitoring and the specified aeroengine controls take place. Currently for a pilot it is not possible to identify such a situation sufficient certain. At first sight the statement amazing, that at the first glance obviously the compressor blading is smoothened/polished. So even the compressor efficiency improves. This becomes noticeable by slight reductions of the EGT (gas temprature behind the high pressure turbine). This effect is even increased if clogged cooling air holes of the hot parts reduce the cooling air flow and additionally the efficiency of the engine.
With a markedly drop of the efficiency and rise of the EGT is not to be reckoned before a distinct erosion (roughening and cange of the profile) of the bading.
However already at very low concentration of ash/gas a very serious shortening of the hot parts lifetime, especially the very cost-intensive blading of the high pressure turbine must be expected. The danger of a long time deterioration by sulfidation (Fig. "Sulfidation attack") exists already by sulfur deposits which can hardly identified with the borescope. Get these failures visible usually concerned components must be exchanged. Such commercial risks may especially sudenly hit repair shops “Power by the Hour” contracts (flat compensation for every flight hour of the operator/contractual partner). In contracts for repair/overhaul and of insurances such a situation must be considered.
Figure "Indications of volcano ash damage" (Ref. 5.3.2-26): Shown are changes in civil fan engines of a 4-engined airliner. It flew about 7 minutes through a `veil cloud' (cirrus cloud) contaminated with vulcano ash and gases. Subsequently concerned hot parts of the 4 aeroengines had to be repaired for a cost equivalent of the price from about 150 medium-class cars.
The assessment of the influence requires a comparison standard. This enable previous, documented routine investigations at the same aeroengine and/or sufficient experience.
As measures for the evaluation of the consequences offer itself:
- Borescope inspections of the blading from turbine (“S”) and compressors.
- Microscopic investigations of air filters like used in the air conditioning (“F1” and “F2”) or bleed air.
- Health - monitoring (chapter 25.1) of the aeroengine, e.g. observation of a change of the EGT (see above).
- Oil analysis (“O”) spectrometric type (chapter 22.3.4).
5.3.2-1 D.L.Mann, G.D.Warnes, “Future Directions in Helicopter Engine Protection System Configuration”, AGARD-CP-558, proceedings of the conference “Erosion, Corrosion and Foreign Object Damage Effects in Gas Turbines”, Propulsion and Energetics Panel Symposium, Rotterdam, The Netherlands, 25-28 April 1994, Chapter 4, page 8.
5.3.2-2 Rolls-Royce Aero-Engine Conference, contribution to the proceedings, pages 9 to 23.
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5.3.2-4 V.R. Edwards, P.L. Rouse, “U.S. Army Rotorcraft Turboshaft Engines Sand & Dust Erosion Considerations”, AGARD-CP-558, proceedings of the conference “Erosion, Corrosion and Foreign Object Damage Effects in Gas Turbines”, Propulsion and Energetics Panel Symposium, Rotterdam, The Netherlands, 25-28 April 1994, Chapter 3, page 5.
5.3.2-5 Issue 10-14, periodical “Aviation Week & Space Technology, September 22, 1980, pages 143 and 144.
5.3.2-6 “Eagle Claw, The Raid into Iran”, Villard Military Series, ISBN 0-394-74403-9, pages 85 and 86.
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5.3.2-12 “Volcanic Damage to Philippine Aircraft Imperils Anti-Insurgent Operations”, periodical “Aviation Week & Space Technology”, April 13, 1992, pages 60 and 61.
5.3.2-13 R.Cantor, “Complete avoidance of volcanic ash is only procedure that guarantees flight safety”, periodical ” ICAO Journal”, September 1998, pages 18 and 19.
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5.3.2-20 R.A.LeyesII,W.A.Fleming “The History of Norrth american Small Gas Turbine Aircraft Engines”, AIAA, ISBN 1-56347-332-1, 199.page 551.
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5.3.2-22 ATSB Transport Safety Investigation Report, Aviation Safety Incident Report 200403110, “Engine failure - Melbourne Airport, Victoria 25-Aug-2004, Boeing Company 777-312, 9V-SYB”, page 1-81.
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5.3.2-24 T.J.Casadevall, “Volcanic Hazards and Aviation Safety: Lessons of the Past Decade”, Flight Safety Foundation, Zeitschrift “Flight Safety Digest”, May 1993, page 1-9.
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