The components of combustion chambers are subject to widely varying loads:
All components in the combustion chamber area exhibit specific damages. Typical examples include the housing, which is subject to mechanical loads and thermal fatigue like a pressure cooker; the inner wall, which is subject to thermal fatigue; the fuel nozzle, which can become clogged by coking. New technologies have advantages, but also come with new damage mechanisms and risks. This is also true for the combustion chamber (Chapter 14). Ceramic thermal barriers on the inside of the combustion chamber can break out or erode, causing extensive consequential damages. The flame can burn through the side of the combustion chamber, causing the housing to burst. If parts of the inner combustion chamber wall or shingles break out, they can behave like foreign objects and/or disturb the flow. They may excite vibrations that cause serious turbine damage.
Long-term damage to the turbine blades is often caused by poor temperature distribution in the gas flow at the combustion chamber exit. Particles from the combustion chamber can cause unallowably heavy erosion of the blading and accelerate oxidation.
Figure "Typical combustion chamber damages": Damage mechanisms that affect the inner wall of a combustion chamber:
The large temperature gradients between cooled and uncooled combustion chamber zones cause plastic deformations when expansion is restricted. A typical symptom is buckling of the combustion chamber wall. The thin-walled, ring-shaped lips that create the cooling air film are especially susceptible to buckling (Fig. "Combustion chamber deformation"). If the temperature changes cause cyclical loads in the plastic zone, it is referred to as thermal fatigue (see Chapter 12.6.2). This results in fatigue cracking (LCF, see Chapter 12.6.1). Openings for the cooling air film in the combustion chamber inner wall are commonly affected. In older engine types, the edges of combustion chamber gills (bulging slits, Fig. "Thermal fatigue in combustion chambers") crack. Newer combustion chamber models are outfitted with rows of holes that tend to “unbutton”. Thermal fatigue cracks also form at larger bores for the combustion air. All of these cracks usually occur in cross-sections with large stress gradients. The average stress in these areas is relatively low, however. For this reason, these cracks are generally controllable for longer operating times/startup-shutdown cycles. The maximum allowable crack lengths can vary greatly, depending on the specific parts and engines involved. Unallowable crack lengths are reached when the crack growth rate accelerates under additional forces such as gas bending loads or centrifugal force. Fast growth rates are especially dangerous in weakened wall sections with high-frequency vibrations. These vibrations can be caused by the typical pulsations during combustion, for example. For this reason, tolerable crack lengths should be limited in low NOx combustion chambers, which are known to have a strongly pulsating combustion process. In shingle combustion chambers (Ill. 184.108.40.206-1), dynamic fatigue fractures can cause shingles to break out.
Consequential damages from a wall breakout occur when, for example, the flame escapes through the side of the engine. This can cause the combustion chamber liner to explode or block part of the stator. This type of flow disturbance causes dangerous vibrations of the turbine rotor and/or its blading.
Cracking in the combustion chamber wall can create unallowably large temperature gradients in the hot gas flow at the combustion chamber exit (Fig. "Temperature variation at the combustion chamber outlet"). This type of cracking is usually the result of asymmetrical combustion and/or localized overheating of the combustion chamber wall due to hot streaks (Fig. "Hot gas streaks as combustion chamber problem") caused by the failure of the cooling air film. Suitable measures such as timely boroscope inspections of the combustion chamber and high-pressure turbine make these damages controllable.
Unavoidable vibrations caused by the combustion process lead to fretting wear (see Volume 2, Chapter 6) in the socket connections (for withstanding thermal strain differences) typically used in combustion chambers. Especially pronounced dynamic stress in low NOx combustion chambers can cause unallowably high wear rates.
Unusually high temperatures accelerate oxidation and emaciation of the walls. In colder combustion chamber areas, sulfur can cause sulfidation (Volume 1, Chapter 5.4.5). Coking can minimize damages caused by overheating, but has other, damaging, effects (corrosion).
If the combustion chamber wall has a thermal insulation coating, the latter can be expected to erode through the delamination of small particles. These particles can wear down the thin diffusion-protection coatings of the turbine blades and considerably reduce their oxidation life spans. Larger coating sections delaminate especially through cyclical plastic deformation at the edges of the combustion chamber wall. This results in a temperature increase and accelerated overheating damages. The metal lips that generate the cooling air film are especially susceptible to this type of damage.
Figure "Thermal fatigue in combustion chambers": This older engine type of various fighter aircraft is outfitted with can-type combustion chambers. The inner combustion chamber liner of these combustion chambers has pronounced thermal fatigue cracks at the gill edges after only a few hundred hours of operation (detail). After several cases in which these cracks grew uncontrollably quickly due to high-frequency vibrations and caused large sections of the wall to break out (bottom diagram), the maintenance interval was reduced considerably.
The danger of thermal fatigue cracks in the inner combustion chamber liner is also made evident by damage that occurred in an older civilian engine type (Volume 2, Ill. 9.3-5). In this case, cracks were initiated in the circumferential roll seam and flames escaped. The flames overheated the combustion chamber mantle. This caused an explosive bursting of the mantle and the forced ejection of several combustion chamber tubes (Ref. 220.127.116.11-9).
Example "Splash plate redesign" (Ref. 18.104.22.168-3):
Excerpt: “…(the OEM) has found that the splash plates (which direct air to the ignition area) inside the low emission combustors mounted in a recently built series of engines are prompt to fatigue cracking and must be replaced before completing 1,000 cycles…The splash plate has now been redesigned to eliminate the problem, and the improved plates can be retrofitted into the original combustors.”
Comments: The life span reduction corresponding to a specific number of cycles indicates a thermal fatigue problem. The mentioned plate is probably necessary for eliminating ignition problems in the low NOx combustion chamber with its high excess oxygen levels.
Example "Pounding vibration during taxiing" (Ref. 22.214.171.124-11):
Excerpt: “…One of the …(business jet-) engines also has combustor instability, company officials said, that emits a pounding low-frequency vibration at about 130 dB. as the aircraft taxies…(the OEM is modifying the) design to eliminate the low-rpm combustion instability on the flight test engine, that instability should not occur on the production (engine)…“
Comments: Since this is a low-emission combustion chamber of a new engine type, the described problems are probably related to this. The described volume and vibrations caused by the pressure shocks in the combustion chamber must have been extremely impressive and are unallowable from the standpoint of engine safety, since they indicate powerful dynamic loads on the engine parts.
Figure "Problems by soot and coke in combustion chamber": Soot and coke formation can promote damage in many different ways.
Soot in the flame means increased thermal radiation compared with a blue flame (Fig. "Combustion chamber life dependence from flame radiant energy"). A surface struck by radiation heats up correspondingly faster. These surfaces are primarily combustion chamber walls and the inlet edges of the turbine inlet stator. A smoking flame also always causes higher engine part temperatures with considerable life span reduction and/or increased repair costs ( ).
Under coke deposits in the combustion chamber dome (Fig. "Turbine damage by soot from the combustion chamber"), increased corrosive attack can occur (Ref. 126.96.36.199-13). These damages especially occur near the fuel atomizers. The different types are catastrophic oxidation, coking, metal dusting, and green rot (Volume 1, Chapter 188.8.131.52). A typical damage mechanism is the reaction of coke with the wall material. This creates carbides from alloy components that are necessary for the formation of protective oxide layers. If the coke also contains corrosion-promoting components such as vanadium or sulfur, then sulfidation is accelerated.
Thicker coke deposits can constrict the flow through bores and gills that carry cooling air and combustion air, locally increasing the fuel/air ratio and compromising combustion. The temperature in the downstream combustion chamber wall increases (hot streaks) and the temperature profile at the combustion chamber exit worsens.
Overheating of the wall accelerates deformation, oxidation, and cracking (Fig. "Typical combustion chamber damages")
Spalling coke particles can locally deform the inlet edge area of rotor blades of the first high-pressure stage (carbon impact, Volume 1, Chapter 184.108.40.206). The result is usually a constriction of the inner cooling air ducts and local overheating of the blades. Very small coke particles and even soot can cause erosion wear on the turbine blading (Refs. 220.127.116.11-6 and 18.104.22.168-7). When the thin diffusion layer that acts as oxidation protection is eroded, the blades will most likely fail prematurely due to accelerated oxidation of the blade wall.
An important role is played by the fuel supply system and especially the fuel nozzle. Coke deposits near the inflow of the atomizing air and/or in the fuel duct of the nozzle or the nozzle exit (Fig. "Damage by coke build up at the fuel nozzle") change the spray pattern (Ref. 22.214.171.124-6). The behavior of atomizers (Ref. 126.96.36.199-8) is also sensitive to coke deposits. The temperature distribution in the combustion chamber (hot streaks, Fig. "Hot gas streaks as combustion chamber problem") and at its exit becomes less even. Larger droplets make combustion difficult. This can result in increased soot development with increased thermal radiation and/or unstable combustion. In this case, unallowably heavy pressure fluctuations can be expected (Ill. 188.8.131.52-4). This creates a danger of dynamic fatigue in the HCF range for combustion chamber components. Examples of this are breakouts from walls that have been damaged by thermal fatigue, and the separation of shingles.
Fuel can already exceed its thermal stability in the supply system. Cracking occurs at fuel temperatures above 163°C (Ref. 184.108.40.206-6). If extremely fine, hard carbon particles are formed, it can quickly lead to unallowable erosion in the jet duct (Volume 1, Chapter 5.3.1). If the injection jet and flame are deflected, in extreme cases they can overheat the combustion chamber and housing wall to failure and escape from the side of the engine.
Figure "Unstable combustion in 'low NOx' combustion chambers": The combustion chambers of modern engines are operated with very high amounts of excess air, i.e. very lean combustion, in order to minimize pollutant emissions (especially nitrous oxides; low NOx combustion chambers). The excess air compromises the stability of the combustion (Fig. "Combustion chamber development for low emission"). The result is a self-increasing process (flickering of the flame) with powerful pressure pulsations (Ill. 220.127.116.11-4). This subjects the combustion chamber components to especially high dynamic loads. Typical known (in stationary gas turbines) and thinkable damage mechanisms are:
Instability of combustion occurs due to an alternating effect between the air flow and the spray cone (Ill. 18.104.22.168-4). This affects the air supply system and gas flow, as well as the expansion and size of droplets in the spray cone. Larger droplets promote soot development and therefore also increase deposits on the fuel nozzles. If the spray cone is deformed and/or deflected, it can create hot streaks that locally overheat the combustion chamber wall and/or the turbine blading. Increased coke deposits form as the combustion chamber wall is wetted with fuel, since larger droplets are more likely to strike it and turn into coke on the spot (Ref. 22.214.171.124-8). Coke deposits drastically hinder atomization and protect the combustion chamber wall from thermal radiation. Temperature gradients increase, resulting in thermal strain with thermal fatigue damage near the deposits. Inspections have shown that the deposits on the fuel nozzle can evidently be traced back to particles from the flame (Fig. "Damage by coke build up at the fuel nozzle"), which are carried by the recirculating gas flow in the primary zone. If this circulation is increased during the phases of pulsation, more coke deposits can also be expected.
The combustion chamber walls are subject to increased dynamic loads during pulsating combustion. This explains the danger of wall sections that are already damaged by thermal fatigue, breaking out. If shingles come loose in shingle combustion chambers, extensive consequential damages to the combustion chamber and turbine can be expected. Overheating of the walls occurs when pressure pulses in the hot gas disturb the protective cooling air film. Ref. 126.96.36.199-18 includes references to a poor temperature profile in the combustion chamber of a modern civilian fan engine causing damage. The created thermal strain evidently resulted in cracks in the turbine housing.
Pressure pulsations in the combustion chamber have a vibration-exciting effect on the turbine (blades, Ref. 188.8.131.52-18; rotor) and the compressor. Dynamic fatigue fractures and intense fretting wear in socket connections and fasteners become more important. If particles or coating fragments break out of a combustion chamber wall coated with a ceramic thermal barrier, the erosion loads on the turbine blading increase.
Increased coking in the combustion chamber increases the probability of impact damage through coke particles and erosion of the blading (carbon impact, Ref. 184.108.40.206-13, Fig. "Temperature caused damages at high pressure turbine vanes").
Example "Contanimated fuel lowering combustion chamber efficiency" (Ref. 220.127.116.11-19):
Excerpt:“The aircraft was descending…with the throttles at idle. As power was added…the left engine surged and exceeded max EGT (gas inlet temperature inside/ahead of the turbine). 18 seconds later the right engine surged and exceeded its max EGT. The left and right engines were restarted…The inability of the engines to accelerate after the manually induced surge was due to contaminated fuel nozzles which significantly reduced combustion chamber efficiencies and which resulted in a subidle stall.”
Comments: The problems with the injection nozzles on both engines indicate that the fuel was the cause of the problem. It is not known if the contaminants formed in the fuel nozzle (coking) or were already present in the fuel (Fig. "Damage by coke build up at the fuel nozzle"). It is interesting that the blocking of the fuel nozzles resulted in a compressor surge.
Figure "Damage by coke build up at the fuel nozzle" (Ref. 18.104.22.168-8): The fuel nozzle and the area around it most frequently have coke deposits. This is most likely due largely to the relatively cool wall temperatures. As shown in the bottom diagram, the composition of the fuel is of decisive importance in coking. Coking is promoted by fuels containing large amounts of high-boiling components. For example, while the boiling point of kerosine is between 140°C and 235°C, that of diesel oil is between 180°C and 380°C. This explains the relatively heavy coke deposits that occur when burning diesel fuel.
There are two different fundamental coking mechanisms:
Coke deposits near the fuel nozzle and on its front end are composed of about 98% carbon and 1-2% oxygen. They have an amorphous microstructure of spheres with diameters of about 10-3 mm. These particles are considerably larger than normal soot particles. Some of them are graphitized, which indicates that they were created at temperatures above 900°C. Because these temperatures do not occur in the coked walls, the particles must have formed in the flame. One explanation is that they are relatively large coked particles from the spray cone. After the coked particles have been carried by the recirculating hot gases around the fuel nozzle and in the combustion chamber dome, they settle on the relatively cold surfaces (Ill. 22.214.171.124-6). If the deposits contain corrosive components such as sulfur, they may corrode the area around them (Ref. 126.96.36.199-13).
The coke deposits on the downstream combustion chamber walls are formed by an entirely different mechanism than those at the nozzle. This is indicated by their relatively high oxygen content of about 20 % and an, albeit small, hydrogen content of several percent. During flight, autooxidation of relatively large fuel droplets occurs at temperatures below 300°C. Higher temperatures reduce autooxidation and cause the droplets to evaporate more quickly. The deposits can be explained by fuel droplets striking hot combustion chamber walls and coking up on impact (thermal stability exceeded). This process requires wall temperatures between 180°C and 460°C. The larger the diameter of the fuel droplets, the more likely they are to make it to the combustion chamber wall, and the greater the coke buildup. The increased evaporation time of fuels with higher boiling points also promotes the formation of coke after the fuel strikes the wall. This creates several times the amount of deposits.
Contaminated or non-specified fuels can block fuel nozzles that do not usually tend to this phenomenon (Example "Contanimated fuel lowering combustion chamber efficiency"). Either these contaminants promote coking in the nozzle at the relatively high temperatures, or they block the nozzles themselves. These contaminants can include corroded synthetic materials (e.g. seals) that are unsuited to the fuel system.
Leaks in fuel nozzles can be expected to cause dangerous consequential damages such as the failure of overheated pressure housings (rear compressor housing, combustion chamber housing; Example "Leak in fuel nozzle case 1"). For this reason, these leaks must absolutely be avoided. If a fuel leak occurs, it must not only be repaired, but its cause must be determined and targeted corrective measures taken.
Figure "Turbine damage by soot from the combustion chamber": The diagram above shows the view into a combustion chamber dome in a can-type combustion chamber of a small shaft-power engine. It is interesting that the coke deposits on the heat shield, which is located concentrically around the fuel nozzle, increase towards the outer edges.
There are no coke deposits around the air-feed pipes. This is explained by the direction of the air flow. It prevents soot carried in the recirculating flow from depositing (Fig. "Damage by coke build up at the fuel nozzle"). Between these air ducts, near the nozzle, there are thick coke deposits that have already partially spalled. These coke particles can erode the stator vanes and running blades of the front turbine stage (bottom left diagram, Ref. 188.8.131.52-13 and 184.108.40.206-14). In extreme cases, this causes deformation of the inlet edge of turbine rotor blades of the first stage (carbon impact, bottom left diagram).
Figure "Carbon erosion at high pressure Turbine blades and vanes" (Ref. 220.127.116.11-13): Carbon erosion can be a serious problem (Ref. 18.104.22.168-24). The depicted case is a marine application. Hot gas corrosion was certainly also involved and promoted erosion. The increased erosion in the outer areas of the blades and near the root platform and shroud is interesting. Contrary to expectations, the middle of the blades is evidently less eroded. A possible remedy is putting heat-resistant hard coatings on the blade surfaces (Fig. "Temperature caused damages at high pressure turbine vanes").
Figure "Overheating at turbine vanes and blades": Near rotor blades, the radial temperature distribution (RTD) at the combustion chamber exit is less balanced by the rotation in the gas flow. The orbital temperature distribution (OTD) is more balanced (top right diagram). If the maximum temperature moves radially outward, then the corners of the rotor blade shroud can curl (tip curling) or suffer a creep fracture (bottom diagram, Example "Thermal impact by combustion chamber").
The high-pressure turbine stator experiences the temperature distribution with no balancing effects. This means that the individual blades can be subject to very different temperature levels. This causes overheating damages that vary widely in their extent and distribution (top left diagram, also see Fig. "Temperature variation at the combustion chamber outlet").
Example "Thermal impact by combustion chamber" (Ref. 22.214.171.124-12):
Excerpt: ”…The second…..incident….occurred in early January, when a third-stage low-pressure turbine blade failed. This too is a well known type of failure. Since the engine runs slightly hotter than originally planned, third-stage low-pressure turbine blades suffer from creep. When this occurs, the tip shroud on the elongated blade can curl, then break…Since the problem was first identified about three years ago, it has been controlled in the field through boroscope inspections, which are required every time a blade is subjected to a specific temperature level for more than 10 hrs. The long-term correction for this problem is to add a redesigned low-pressure turbine module to the powerplant when it returns to the depot for regular maintenance.”
Comments: This case concerns a two-shaft engine from a fighter aircraft with two-stage high-pressure and low-pressure turbines. The low-pressure turbine is usually cooled considerably more poorly than the high-pressure turbine (rotor blades without a cooling air film). It is evidently especially sensitive to the radial temperature distribution at the combustion chamber outlet. It seems that the radial temperature distribution behind the combustion chamber was not given sufficient attention during the design phase.
Example "Leak in fuel nozzle case 1" (Ref. 126.96.36.199-20):
Excerpt:“Approximately one minute after takeoff, the crew observed left engine fire warning…An inspection of the left engine revealed a leak at the No. 7 fuel nozzle, which was the origin of the fire. A hole was burned into the diffusor case, which allowed hot gases to escape and contact the oil lines. This resulted in extensive fire damage to the engine and cowling.
Comments: Reports of leaks in fuel nozzles are not uncommon. There are many causes for such leaks: cracks in the nozzles and fasteners, seal failures, installation errors, etc. As the example illustrates, leaks and seal failures must be taken very seriously, since they can cause dangerous consequential damages (Fig. "Combustion chamber housing failure consequences").
Example "Leak in fuel nozzle case 2" (Ref. 188.8.131.52-1):
Excerpt: “Approximately one minute after take off, the crew observed left engine fire warning indications in the cockpit. The left engine power lever was retarded and a turn towards the airport was initiated. The fire warning indications actuated intermittently until the crew shut down the engine and activated the fire extinguisher. An uneventful landing was performed. Ground personnel informed the captain of visible fire damage to the left engine, and an emergency evacuation was performed. An inspection of the left engine revealed a leak in the No. 7 fuel nozzle, which was the origin of the fire. A hole was burned in the diffusor case, which allowed the hot gases to escape and contact oil lines.This resulted in extensive fire damage to the engine and cowling.”
”…An examination of the number 7 fuel nozzle revealed the nozzle support mount flange, nozzle support heat shield, and nozzle head swirl vanes had partially melted. A leak and flow check confirmed leakage at the nozzle support-to -head interface.
Comments: It is unclear how the fuel leak in the fuel nozzle occurred, but the danger of a flame escaping sideways is clear (also see Volume 2, Chapter 9.3). This type of temperature overload usually overstresses the combustion chamber housing.
Example "Fatigue fracture of the combustion chamber outer case" (Ref. 184.108.40.206-2):
Excerpt: “During takeoff, an uncontained failure of the left engine occurred. An examination revealed a fatigue fracture through a rear flange bolt position of the engine. The engine had been inspected in accordance with an applicable airworthiness directive; however, the fracture occurred 2,060 cycles prior to the next scheduled inspection. Additional inspection requirements resulted from the examinations of the engine and the fracture.
Probable cause: The fatigue fracture of the combustion chamber outer case. A factor was the manufacturer's inadequate inspection criteria.”
Comments: This exemplary case shows the danger of the explosion of a combustion chamber due to thermal fatigue (also see Volume 2, Ill.9.3-5). In this case, and in two previous parallel cases, the cracking originated in a flange bore. After these cases, both the FAA and the manufacturer issued directives.
A large number of additional bores were also cracked, despite instructions that should guarantee a safe number of load cycles (start-up/shut-down). The cracks evidently reached a critical length before the next inspection took place. The crack growth was also apparently underestimated. This shows, that estimation of crack growth rates should be done very conservatively. One must use only a fraction of the cycles that will most likely lead to failure (usually estimated based on failure investigations and operating data from parallel cases).
Figure "Weak points at housings by thermal fatigue": Combustion chamber housings are highly-stressed pressure cookers. They must withstand internal pressure that can reach up to 40 bar in modern engines. Additionally, normal operation subjects them to cyclically-recurring temperatures of several hundred °C. This can induce dangerously high thermal fatigue strain. In the case of combustion chamber damage, possible consequences include excessive localized overheating (Volume 2, Chapter 9.3) with buckling (“7”) and cracking, culminating in explosive bursting (Volume 2, Ill. 9.3-5).
The corners (“1”, bottom right diagram) and edges (“ 2”, Fig. "Combustion chamber housing failure consequences") of axial flanges are especially susceptible to cracking due to thermal fatigue.
Experience has shown that welded connections and flanges tend to cracking (“3”). Gradual transitions to the housing wall are less threatened than cylindrical eyes (bottom left diagram). Welded connections with pronounced notch effects in the transition must absolutely be avoided.
There have also been reports of catastrophic damages in housings that are not split axially. In this case, the weak point was near the threaded connections (“6”) of the circumferential flange (Ref. 220.127.116.11-25).
Figure "Combustion chamber housing failure consequences": The bursting of combustion chamber walls due to hot gases escaping radially from the combustion chamber is rare, but exceptionally dangerous.
There are two primary reasons for flames escaping from combustion chambers:
If the combustion chamber housing is perforated, an extremely energy-rich hot gas jet can escape and cause extensive consequential damages (Volume 2, Chapter 9.3). In a fighter aircraft with two parallel engines in the hull or in aircraft with two engines in a nacelle, the probability of the hot gas jet damaging the second engine is high. Experience has shown that normal firewalls are not always sufficient.
Axial bursting of the housing is more likely than a hole in the housing wall (bottom diagram, Ref. 18.104.22.168-9).
In the depicted case, the housing of the can-type combustion chamber burst axially. Experience has shown that the explosion will throw the combustion chamber tubes in the area of the damage out of the housing. In the case shown at bottom, these tubes struck the wing and were an immediate fire danger to the nacelle.
Most cases in which the combustion chamber burst were caused by thermal fatigue (Fig. "Weak points at housings by thermal fatigue"). In the above case of a two-engine fighter aircraft, a thermal fatigue crack formed along the transition of the axial flange (Fig. "Weak points at housings by thermal fatigue") due to a notch specific to the design and production process. The crack grew slowly at first, and then rapidly progressed to instability. It was fortunate that the direction of the explosion and the escape of the combustion chamber tubes occurred towards the outside of the fuselage. Despite a gaping hole in the fuselage with a size of about 1 m2 , the pilot was able to land safely with the remaining engine.
In the case of a commercial aircraft (Ref. 22.214.171.124-25), the combustion chamber housing burst axially after cyclical fatigue and creep evidently caused a crack in the annular flange at a bolt bore.
Example "Complications of low emission combustion chamber" (Ref. 126.96.36.199-10):
Excerpt: “…The dual annular combustor has suffered problems with cracking of the outer liner near the combustor exit and similar cracking problems with the inner liner. In addition to correcting these issues, the dual annular combustor…product improvement package is aimed `…at reducing emissions to get closer to what we had promised'
…The improvements include a modified cooling system, redesigned forward inner nozzle support and extended splashplates to improve emissions at lower power settings…“
Comments: Evidently the cracks were not sufficient to cause the pressurized combustion chamber housing to fail. It can be assumed that this combustion chamber type, which is designed for low emissions, dynamically overloaded the liner. It is not clear if the cracking was thermal fatigue cracking and/or HCF cracking due to high-frequency vibrations. One can speculate that low NOx conditions (unstable combustion) at the very least accelerate damage progression.
Example "Cracked combustor casings" (Ref. 188.8.131.52-21):
Excerpt: “The … engines…are to be inspected following the discovery of cracked combustor casings during routine maintenance. A total of 452 aircraft will be sent…for inspection and the replacement of faulty components. A hundred engines have been examined since December 13, revealing 34 cracked casings. Aircraft passing this inspection will be returned to service but re-examined at regular intervals which have yet to be determined…“
Comments: These are engines of a two-engine fighter aircraft type. The inspection intervals indicate that these cracks were related to the start-up/shut-down cycles, and probably caused by thermal fatigue.
Fig. "Danger by testing Engines in vertical position": During a run in the course of developing a small shaft-power engine, dangerous overspeed occurred. The drainage of fuel remnants after a previous aborted start was evidently not sufficient. This example shows that functioning drainage systems are extremely important, especially in engines that are started in various positions. The regulator can generally not completely compensate for collected fuel and is therefore unable to safely ensure that overspeed and the danger of a burst rotor do not occur.
Figure "Sensitiveness of 'Spark plugs' to poor design" (Refs. 184.108.40.206-23 and 220.127.116.11-24): The spark plugs in the combustion chamber of an engine must generate considerably greater sparks without damage than the spark plugs of automobiles (Fig. "Combustion chamber development for low emission"). For safety reasons, combustion chambers have at least two spark plugs. The distance between electrodes is greater than in automotive spark plugs because the air pressure in which the spark is created is considerably lower. Despite the high electric energy, erosion of the electrode is controllable because the ignition duration of engines is usually limited to start-up. Therefore, the ignition should only be used when it is absolutely necessary. There are situations in which constant ignition is required to prevent combustion chamber flame out. For example, during fire fighting missions there may not be sufficient oxygen in the inlet air. If the tolerable application limits are exceeded, the spark plug may overheat (Ref. 18.104.22.168-17).
Every spark causes a small melted crater and evaporation of the electrode material. If this spark erosion progresses far enough, the spark plug will fail. Increasing the diameter of the spark plug disturbs the airflow in the combustion chamber due to the larger hole in the wall. Therefore, it is not possible to increase the diameter of the central electrodes without reducing the distance between the electrodes, which also decreases the spark energy.
The electrodes of engine spark plugs are usually arranged concentrically. There is a rod at the center and a conical electrode around the outside (top right diagram). In order to generate the spark with a relatively low current, a ring of semiconducting ceramic is located on the front between the electrodes (Refs. 22.214.171.124-15 and 126.96.36.199-16). If the spark plug has a configuration that makes the spark leap up in an arc, then the spark plug must not extend above the inner combustion chamber wall into the burning zone. This keeps the temperature levels of the spark plug lower. The spark plug should not experience temperatures greater than 600°C. Above this temperature, erosion, oxidation, and corrosion rapidly decrease spark plug life. The lean fuel/air mixtures in low-emission combustion chambers increase oxidation and increase the stress on the spark plugs. The different thermal strains of metallic electrodes and semiconducting ceramic can create gaps and contact problems. This is especially important for spark plug life in engines with high pressure ratios and correspondingly high heat transfer (Fig. "Wearing of spark plugs and its consequences"). It is understandable that the location of a spark plug in the combustion chamber is important for its functioning and life span. The spark plug must be positioned in the primary zone in a way that the rotating hot gas cloud created by ignition is carried by the recirculating flow and ignites the entire primary zone. Experience shows that this requirement is best met by a spark plug position that is axial to the fuel nozzle.
However, this position creates new problems (bottom diagram). The flow disturbance caused by the opening for the spark plug in the combustion chamber wall creates uneven temperature distribution. Its proximity to the fuel spray cone promotes soot deposits and coking on the front side of the spark plug and compromises spark creation. Under no circumstances may liquid fuel come into contact with the spark plug, either by flowing down the combustion chamber wall or directly from the fuel cone. This must be ensured with a sufficiently large axial distance between the fuel nozzle and the spark plug, proper angling of the latter, and/or the use of a deflector.
Contact with liquid fuel causes various damage mechanisms (top right diagram):
188.8.131.52-1 NTSB Identification: ATL93IA113 with details in the NTSB Imaging System
184.108.40.206-2 NTSB Identification: ATL94IA097 with details in the NTSB Imaging System
220.127.116.11-3 “GE fits new combustors to CF6-80C2 powerplants”, periodical “Flight International”, 13-19 August 1997, page 13.
18.104.22.168-4 R.A. Rudey, J.S. Grobman, “Characteristics and Combustion of Future Hydrocarbon Fuels”, Agard Lecture Series No. 96, “Aircraft Engine Future Fuels and Energy Conservation”, presented 16-17 October, 1978 in Munich, Germany. page 5-1 to 5-23.
22.214.171.124-5 R.A. Rudey, J.S. Grobman, “Impact of Future Fuel Properties on Aircraft Engines and Fuel Systems”, Agard Lecture Series No. 96, “Aircraft Engine Future Fuels and Energy Conservation”, presented 16-17 October, 1978 in Munich, Germany. page 6-1 to 6-29.
126.96.36.199-6 R.N.Hazlett, “Thermal Oxidation Stability of Aviation Turbine Fuels”, ASTM Publication Code Number (PCN) 31-001092-12.
188.8.131.52-7 J.S.Posford, G.K. Widdington, “Cyclic Endurance Testing of the RB211-22B Cast HP Turbine Blade”, AGARD Conference Proceedings No. 368, of the conference on “Engine Cyclic Durability by Analysis and Testing”.Chapter 15.
184.108.40.206-8 M. Brandauer, A.Schulz, S.Wittig, “Mechanisms of Coke Formation in Gas Turbine Combustion Chambers” , Paper No. 95-GT-49 of the “40th International Gas Turbine and Aeroengine Congress and Exhibition”, June 5-8, 1995.
220.127.116.11-9 “Report highlights JT8D problems”, periodical “Flight International”, 18 March 1980, pages 6 and 7.
18.104.22.168-10 “CFMI tests solution to CFM56 cracking problem”, periodical “Flight International”, 18-24 February 1998, page 20.
22.214.171.124-11 J.T. McKenna, “Despite Engine, Wing Problems G5 Exceeds Test Targets”, periodical “Aviation Week & Space Technology”, April 29, 1996, page 31.
126.96.36.199-12 S.W.Kandebo, “USAF Targets Engine Mishaps”, periodical “Aviation Week & Space Technology”, March 29, 1999, page 84.
188.8.131.52-13 F.J.Plumb, “Maintenance Problems in Gas Turbine Components at the Royal Naval Aircraft Yard Fleetlands”, AGARD Conference Proceeding No. 317 of the conference “Maintenance in Service of High Temperature Parts”, pages 3-1 to 3-13.
184.108.40.206-14 F.W. Skidmore, J.M.Bennett, D.E. Glenny, “An Investigation Into Hard Carbon Formation in a Modified Gas Turbine Combustor”, Proceeding Paper ISABE 95-7116, pages 1268 to 1273.
220.127.116.11-15 “The Jet Engine”, Rolls Royce plc, ISBN 0 9022121 2 35, page 131.
18.104.22.168-16 I.E. Traeger, “Aircraft Gas Turbine Engine Technology”, Glencoe, ISBN 0-07-065158-2.
22.214.171.124-17 “All fired-up about engine ignition”, periodical “Aircraft Technology Engineering & Maintenance” Dec/Jan 2001, page 10-17.
126.96.36.199-18 M.Harrison, “The CFM56 in service”, “Engine Yearbook 2002” pages 52 to 63.
188.8.131.52-19 NTSB Identification DCA83IA035, microfiche number 26408A, Incident occurred August 19, 1983.
184.108.40.206-20 NTSB Identification ATL931A113, Incident occurred June 18, 1993.
220.127.116.11-21 “TF30 inspection ordered”, periodical “Flight International”, 31.12.1977, page 1898.
18.104.22.168-22 E. Critchley, P. Sampath, F. Shum, “Cold Weather Ignition Characteristics of Advanced Small Gas Turbine Combustion Systems”, AGARD Proceeding Paper AGARD-CP-480 of the conference “Low Temperature Environment Operations of Turboengines (Design and User's Problems”, Seite 9-1 bis 9-7.
22.214.171.124-23 D.S. Breitman, F.K. Yeung, “Cold Start Development of Modern Small Gas Turbine engines at Pratt & Whitney Canada”, AGARD Proceeding Paper AGARD-CP-480 of the conference “Low Temperature Environment Operations of Turboengines (Design and User's Problems”, pages 13-1 to 13-7.
126.96.36.199-24 S.W. Kandebo “Alliedsignal Commits to LT101 Improvements”, periodical “Aviation Week & Space Technology”, March 11, 1996, page 70.
188.8.131.52-25 TSB Reports-Air1994, Report Number A94C0034, Canada “Uncontained Engine Failure”.