Problems and peculiarities of engine operating behavior are closely related to the historical development of engine technology. Developments are primarily determined by economical (Ill. 11.1-1) and, more recently to a larger degree, environmental demands (Ill. 11.1-2).
The degree of technological development and the problems it causes are indicated by typical engine parameters such as (specific) fuel consumption (Ills. 11.1-3 and 11.1-4), efficiency (Ill. 11.1-4), total pressure ratios (Ill. 11.1-5), turbine inlet temperatures (Ill. 11.1-6), and thrust-related figures (Ill. 11.1-6).
Other important influences that are relevant to problems are related to the mission profile of the aircraft (Ills. 11.1-7 and 11.1-8). These include the takeoff/landing frequency, takeoff conditions (e.g. airfield elevation, surrounding temperatures), frequently used flight altitudes (bird strikes, etc.), and typical flight speeds (especially with combat aircraft). The flight envelope is based on physical limits of safe operation of the aircraft and its engines (Ill. 11.1-9). These are largely determined by the behavior of the compressor and combustion chamber (Ill. 11.1-9) in connection with changing atmospheric conditions, i.e. the properties of the ingested air (Ill. 11.1-11). Thrust and propulsion efficiency are closely related to flight altitudes and speeds (Ill. 11.1-12).
Pressure, temperature, and speed levels of the main gas flow (Ill. 11.1-13) are not only signs of the performance and degree of development of an engine, but also have a decisive influence on the stress placed on the engine, and therefore also the life span of its components, which in turn determines costs.
Individual operating phases such as takeoff (Ills. 11.1-14 to -17), acceleration, operation at constant output, and shutdown (Ill. 11.1-17) play various roles and affect the life spans of different components (Ill. 11.1-17).
Illustration 11.1-1 (Ref. 11.1-1): In accordance with this chart, the propulsion-related costs of a commercial aircraft are more than 50% of the total operating costs for this specific long-distance aircraft type. Unfortunately there is no comparable diagram for military aircraft. However, it can be assumed that the engine and fuel costs also represent a large percentage of operating costs of military aircraft. Of course, these costs depend on the specific mission profiles and their specific related damages (Ills. 11.1-7 and 11.1-8).
Purchasing and maintaining engines, along with fuel costs, are major figures for any aircraft operator. These cost determinants are directly and indirectly related in very complex ways.
For example, if low fuel consumption requires a high total pressure ratio in the core engine, a large bypass ratio, and high gas temperatures with minimal cooling, then the development phase of the engine will be marked by special efforts with new and expensive technologies. The life span of the hot parts, for example, is strongly affected by relatively minor temperature increases (Ill. 12.5.1-4). This increases repair costs. Competitive compressors (manufacturing costs, weight, measurements) with good efficiency that is maintained over long time periods, require a great deal of effort and expensive technology to develop. With increasing degrees of efficiency and pressure ratios, and low numbers of stages (lower manufacturing costs), greater sensitivity to damaging influences such as foreign object strikes (dynamic fatigue fracture risk) and erosion (increase of blade tip clearance, worsening of aerodynamic properties) is to be expected. This results in an increase in all mentioned costs. “Exhausted” compressors require undesirably frequent inspections (e.g. after FOD in the compressor or overheating damage to the hot parts). This results in increased maintenance costs, due to more frequent boroscope inspections, for example.
Illustration 11.1-2: Ever stricter environmental regulations have a direct and indirect influence on engine costs (also see Ill. 11.1-1). These regulations primarily concern the reduction of emissions and noise.
These emissions are directly dependent on the specific fuel consumption and, along with cost reduction, are the reason for the trend towards ever larger bypass ratios and elaborate technologies for controlling high gas temperatures in the hot part zone. The fan, which is large and heavy relative to the core engine, causes powerful aerodynamic forces to act on the nacelle, as well as representing large inertial forces during acceleration (Volume 2, Chapter 188.8.131.52). In addition, thrust must be directed into the engine suspension, causing housing deformations that affect the seal gaps.
Higher gas temperatures must not result in an unallowable decrease in the lives of hot parts. This can be accomplished with the aid of suitable cooling technologies, thermal insulation (ceramic coatings) of the surfaces in the gas flow, or heat-resistant materials (such as single crystals). These technologies have specific problems and damage mechanisms. The use of thermal insulation coatings to ensure sufficiently long engine part life is accompanied by the risk that the failure of said coatings can result in even larger, uncontrollable overheating damage (Ill. 12.1-5). This is contrary to the principle of damage tolerance (Chapter 13).
Increasing the total pressure ratio results in greater mechanical and aerodynamic stress on the compressor, which necessitates additional efforts to ensure that clearances are maintained and that blade surfaces and profiles have the proper qualities. The examples of compressor problems in newer engine types are to be considered with this in mind (see Chapter 11.2).
The slow-rotating fan is important for low fuel consumption and low CO2 emissions. The fan and booster (if present) are powered by a correspondingly “slow” low-pressure turbine. This low-pressure system represents a large amount of the engine`s weight. A further step in development is newer high-speed low-pressure turbines that power the fan via a transmission. The mechanically and thermally highly-stressesd low-pressure turbine and the transmission both represent a large technical challenge and require specific problems to be controlled.
Increasing the size of the fan and adding a transmission increases the engine weight, reducing the expected advantages. The longer and heavier fan blades require a correspondingly sturdy and voluminous housing and containment in order to prevent possible blade fragments from escaping and causing unallowable consequential damages. Realizing this type of containment with an acceptably low weight is an extremely demanding task with considerable development risks (Volume 2, Chapter 8).
The technologies used to improve component efficiency are very cost-intensive in both the production of new parts, as well as repair (e.g. coatings).
With regard to NOx emissions, the combustion chamber becomes the center of development efforts. The central issue is the instability of combustion due to excessive air supply (Ill. 184.108.40.206-4), which is intended to reduce the amount of emitted NOx . These vibrations cause problems such as wear on the combustion chamber`s typical plug connections, as well as inciting dangerous vibrations in the hot parts (Ill. 220.127.116.11-4).
The development of soot-free combustion chambers is supported by the trend towards combustion with excessive air supply, but is plagued by the above problems. In military engines, soot-free combustion is desirable because it reduces the visibility of an aircraft in natural and infrared light.
Noise reduction is another task which focuses on the fan and the gas exhaust system. Reduced noise development affects the operating costs of commercial aircraft in the form of fees levied by airports, and also by influencing the length of the window for takeoff and landing, which is relevant to commercial competition. Slow-running fans of the type common today are a move in the right direction towards meeting noise regulations. Measures to reduce noise increase housing weights and also require special fan blade designs that must be reconciled with aerodynamic optimization. New curved blade geometries can have an influence on factors such as the sensitivity to foreign object damage, containment, and the shape of and stress levels in the blade root.
Illustration 11.1-3 (Refs. 11.1-2 and 11.1-3): The demand for the lowest possible fuel consumption (Ill. 11.1-1) has historically led to an increase in the turbine inlet temperature (civilian engines shown in top diagram). For civilian engines, the curve evidently asymptotically approaches a gas temperature of about 1500 °C. This should be related to the maximum possible material temperatures attainable with the most modern cooling techniques. At the same time, the historical trend is towards increased total pressure ratios. The higher pressure levels in civilian engines is due to the demand for low fuel consumption. However, for fighter aircraft the primary concern is low engine weight. The historical reduction of specific fuel consumption (SFC) in civilian engines is shown in the lower diagram. Well-known engine types are also shown.
As has already been shown in Ill. 11.1-2, an increase in the temperature levels and pressure ratios means greater thermal and mechanical stress on the engine components. However, at the same time, there is a desire to make the overhaul and maintenance intervals longer. This greatly increases the effort necessary to develop and implement suitable new technologies and to prove their reliability.
Illustration 11.1-4 (Ref. 11.1-1): The total efficiency of an engine, calculated by the usable propulsion energy relative to the energy content of the fuel consumed, is a product of three partial efficiencies:
It can be recognized that, in terms of efficiency, the LP system with the fan (abscissa) has improved more than the core engine (ordinate). The cited literature shows that this trend is expected to continue and become more pronounced. This indicates that there is new fan technology that is evidently still in the development stage.
Illustration 11.1-5 (Refs. 11.1-1 and 11.1-4): This diagram depicts the historical increase in the total pressure ratios in civilian engines (top graph) and military engines (bottom graph). The trend from Ill. 11.1-3 is recognizable. The pressure ratio of modern civilian aircraft is not only greater than that in military aircraft engines, but it is also expected to increase more rapidly.
High levels of aerodynamic and mechanical stress make high-performance compressors more sensitive to damaging operating influences (FOD, erosion, maintaining clearances, etc.). Therefore, they require a correspondingly extensive effort to develop and test.
Illustration 11.1-6 (Ref. 11.1-4): The historical trends shown in this diagram apply to fighter aircraft engines. It is possible that the trends may be different for engines used in civilian applications (Ill. 11.1-5).
The single-shaft turbojet engines from a development period up to the mid-1960s reached their limits with regard to specific thrust and thermal efficiency. The top left diagram shows the historical development of specific thrust, i.e. the propulsion power relative to the mass flow (gas flow). The problem with turbojets was the extremely high flow rate at the exhaust jet, which increased inner losses. These old engines normally had uncooled turbine rotor blades, which required low gas temperatures. This also meant that cooling air ducts could not become clogged with ingested dusts. The relatively thick cross-sections of the solid blades make them resistant to oxidation, erosion, and FOD (such as carbon impact).
Blade cooling was already used in older engine types. However, it was limited to stator vanes and was fairly simple.
A relatively low total pressure ratio makes the compressor less susceptible to tip clearance changes, also because of the longer blade lengths in the rear section. This prevents rubbing and damage to the sensitive abradable systems. Due to the relatively flexible compressor rotors with large axial lengths and distances between bearings, large tip clearances were necessary in turbojets. The comparatively low rotor RPM and corresponding low LCF stress meant that there was no life span limit that could be compared with that of modern engine types. In these older engines, the steels used in disks, blades, and housings were not corrosion-resistant under operating conditions (usually 13% Cr steels and low-alloyed mild steels). They required special coatings, which made overhaul difficult.
The next step in development were multiple-shaft engines, in which the low-pressure shaft powered a fan (turbofan). In addition to the air flow to the core engine, a second flow was directed around the outside of it (bypass). These gas flows were mixed at the engine exit, leading to lower exit speeds at the jet and improving the propulsion efficiency.
Unlike civilian aircraft, in military aircraft the thrust-weight ratio is of paramount importance. On the other hand, specific fuel consumption is of secondary importance in military applications. The thrust-weight ratio has been increased several times over through the years (top right diagram).
The historical increase in the turbine inlet temperature (bottom left diagram) is closely related to the introduction of cooled turbine rotor blades. The improved cooling, as well as thermal insulation coatings on the hot part surfaces that came into contact with the gas flow, allowed a further increase in the gas temperatures even though they were already far greater than they had been in the uncooled turbines. The progress of the thrust-weight ratio also reflects this trend. This increase of the gas temperature (turbine inlet temperature, Ills. 11.1-3 and 12.1-4) led to its current levels, which are considerably higher than the melting points of materials used in the high-pressure turbine. For this reason, problems with the blade cooling system can be expected to quickly cause serious damage.
The amount of thrust relative to the engine inlet surface area can be taken as an indicator of the power concentration of an engine. Small inlet surface areas/engine diameters allow fighter aircraft to have smaller fuselage cross-sections, which is very desirable with regard to their flight properties.
The historical increase of the performance parameters of fighter aircraft can be seen in connection with an increase in rotor RPM, and therefore also a considerable increase in the mechanical loads on rotor components. This led to (cycle-) limited life spans (LCF). There are many damage mechanisms and problems related to this. The high centrifugal forces on the blades cause large amounts of stress on the blade roots and increase the risk of cracking, especially in the compressor. Higher pressure ratios with small axial lengths and closely spaced blades increase the probability of dangerous vibrations occurring in the compressor.
In connection with the depicted trends, it should be mentioned that the task of controlling high gas temperatures and total pressure ratios has been made considerably easier by the introduction of electronic regulators (digital regulators). Their use is the only way to ensure acceptable operating safety in engines even in damage-relevant situations such as compressor stalls (Ill. 18.104.22.168-2). Evidently, contrary to widespread concerns (regarding vibration damage, corrosion, thermal strain, etc.), the operating safety of these regulators has been very satisfactory both in military and civilian applications. However, only time will allow an ultimate judgement of their performance to be made.
Example 11.1-1 (Ref. 11.1-7):
Excerpt: …(the engine manufacturer-EM) is planning to incorporate a new, angled igniter fuel tube into all … fighter turbofans following evaluations that show the new design offers' “dramatic” cuts in the number of afterburner no-lights- to slightly more than one in every 1,000 afterburner cycles….(EM)-engineers started the process in the mid 1980 when spotty reports of no-lights in remote corners of the flight envelope were being reported by pilots…
The straight igniter tube used at the time sprayed fuel downward into the afterburner pilot can, where it had to be caught in the swirling air, atomized and ignited. But because the JP8 fuel being used in those aircraft wasn't very volatile and the fuel was being sprayed away from the ignition source, no-lights would occasionally happen. The straight tube worked best in warmer temperatures, lower altitudes and with more volatile JP4 fuel“.
Comments: This is an excellent example of the sensitivity of afterburner ignition at high altitudes to fuel changes in connection with design characteristics.
Illustration 11.1-7: The diagrams show the flight envelope of a supersonic fighter aircraft (i.e. its flight speed with regard to flight altitude). The top diagram (Ref. 11.1-9) shows the limiting influences, which can be seen as being primarily related to the properties of the nacelle. The properties of the engine are important with regard to the thrust-resistance balance and the air-resistance limit.
The bottom diagram (Ref. 11.1-10) shows how the re-lighting of an extinguished combustion chamber depends on the fuel temperature. Unlike cold fuel, warmer fuel does not reduce the flight envelope.
Ill. 11.1-8 shows other potential engine problems and their locations in the flight envelope.
Illustration 11.1-8 (Refs. 11.1-5 to 11.1-8): This diagram shows the typical flight envelope of a fighter aircraft in which the engine must meet the demands of steady and unsteady operating conditions. The term mission defines the time during which the engine is operated under operating conditions determined by the required maneuvers. The duty cycles (relevant to engine life) during a mission determine the overhaul time that is set according to the load cycle changes (Ill. 11.1-9 and Chapter 12.6.1).
One can see that specific problems occur in certain areas, which are primarily in the corners of the flight envelope. The susceptibility to these problems is specific to the airplane and engine type. These problems include:
Compressor stalls in connection with certain states of operation, such as lighting of the afterburner (Ill. 22.214.171.124-14).
Engine re-ignition can become more difficult due to the insufficient oxygen supply during low-speed, high-altitude flight (Chapter 11.2.2).
The risk of oil fires (Volume 2 Chapter 9.2) increases along with the flight speed. In an operating state with a high power output, the temperature level of the oil flow is also generally high, which makes ignition more likely. Titanium fires require a sufficient oxygen supply and are usually ignited as the result of compressor blade fractures. This means that they are more likely to occur during high-speed, low-altitude flight (bird strikes, ignition conditions relatively far forward in the compressor, see Volume 2 Chapter 9.1). Ill. 11.1-9 shows the influence of the mission profile on life span reduction.
Illustration 11.1-9 (Ref. 11.1-8): The term “severity” (of damage: wear, life span reduction) is used here to refer to the relationship between the frequency of damage during a specific mission (operation, environment) and that during a reference mission (top left diagram). This is assessed primarily by inspecting components such as the high-pressure compressor (wear, erosion) and the hot parts (combustion chamber and turbine; see bottom right table). The top right diagram compares cargo aircraft being used in military and civilian applications. One can see that the percentage of damages during steady operation (cruising flight) is about equally low in both the military and civilian aircraft. Understandably, the location of the engines and the greater variety of missions in the military aircraft result in a scattering of the lower values. In cyclical operation, the damage to the engines on the military aircraft is several times greater than that during steady operation or civilian use. This can also be traced back to the greater variety of missions in the military engines. It can therefore be assumed that the same aircraft and engine type will have considerably shorter lives, overhaul intervals, and higher damage rates in military applications than they will in civilian ones. This makes it difficult to estimate operating costs of military aircraft by merely projecting civilian costs for the same aircraft.
The top left diagram shows the damage to a fighter aircraft. The percentage of steady operation at high output is dismissably small. However, cyclical operation is of decisive importance. The severity of engine damage is roughly ten times that of a military cargo aircraft and thirty times that of a commercial aircraft. It is possible to see a good correlation between the overhaul intervals in modern engines in civilian use and those in fighter aircraft. The overhaul intervals for civilian engines are usually between 10,000 and 20,000 operating hours, whereas the engines of fighter aircraft are generally overhauled every 1000 to 2000 hours.
The actual frequency of repair visits for modern fleets of fighter aircraft is considerably greater than the recommended times named above, as can be seen in the bottom left diagram. Depending on the average mission length, the intervals between repairs decrease considerably along with shorter flight times. This is understandably so, since the number of cycles that are primarily responsible for damage (LCF, especially startup/shutdown cycles) increases per hour of flight. The overhaul intervals of long range civilian aircraft are often a factor greater than those of fighter aircraft. Short-range civilian aircraft have damage rates that are several times lower than those of fighter aircraft, but they require work far more frequently than long range aircraft due to their greater number of startup cycles.
Illustration 11.1-10: Different missions have specific potential damage mechanisms. As flight altitude decreases, the risk of bird strikes increases exponentially (Volume 1 Chapter 5.2). The engines of fighter aircraft are most at risk, since they often fly at low altitudes, as is the case with navy squadrons. The danger of icing decreases at higher altitudes (Volume 1, Chapter 5.1) because the moisture (freezing rain, snow) that causes icing is only found at altitudes where weather occurs. The risk of oil fires is generally greatest at high-speed, low altitude flight, due to the high engine output combined with correspondingly high temperatures (Volume 2 Chapter 9.2). The oxygen concentration at low altitudes also makes titanium fires possible (Volume 2 Chapter 9.1). Erosion can primarily be expected near the ground due to ingested dust. It causes the aerodynamic surfaces (blades) to become rough and in some cases changes their shape. This worsens the operating behavior and efficiency of the affected compressor. The life span of hot parts and the LCF life span of some rotor components (Ill. 11.1-8) are closely connected to the frequency of startup/shutdown cycles, which means that the life span reduction of these parts is greatest in fighter aircraft and short-range civilian aircraft. In a similar manner, efficiency deterioration (increase in specific fuel consumption over long operating times) is greatest in these aircraft (Volume 2 Ill. 7.0-2), since unsteady operating conditions lead to clearance losses between static (housings, stators, labyrinth seal surfaces) and rotating parts (blades, rotor disks; Volume 2, Chapter 7.1). The resulting rubbing causes wear and increases the clearance gaps during steady operation. Erosion can also contribute to deterioration by wearing down the seal coatings in the housings opposite the rotor blade tips.
Illustration 11.1-11 (Ref. 11.1-11): The operating behavior of an engine is strongly influenced by the properties of the ingested air. Pressure, temperature, and density of air change along with altitude, and the pressure and temperature can also increase at high flight speeds due to the air building up ahead of the engine. Similarly, pressure and temperature decrease in the inlet area in front of a stationary engine in a testing rig (Ill. 11.1-12).
Temperature influence: Specifications and acceptance values regarding engine performance/thrust are usually based on a temperature of 15°C. However, there are also performance guarantees for higher intake temperatures (flat rate). The lower the intake temperature, the greater the air density/mass flow becomes, increasing engine power. Increased intake temperatures (e.g. in the tropics) cause performance to decrease (top right diagram). The surrounding temperature decreases with altitude up to the tropopause, beyond which it remains constant.
Influence of pressure and density: Increasing air pressure causes greater density and increases performance (middle right diagram). Because air pressure decreases with altitude, at equal performance levels, takeoff from runways at high elevations requires higher gas temperatures than takeoff from low-elevation runways. This effect runs contrary to the decreasing temperatures up to the tropopause. Above the tropopause, however, thrust rapidly decreases with altitude (bottom right diagram).
Illustration 11.1-12 (Ref. 11.1-11): The right diagram depicts the way in which, as the flight speed approaches the speed of the exiting exhaust gases/air, the amount of thrust that is usable for propulsion decreases (speed effect curve).
The air congestion at the engine inlet creates a pressure increase, contributing to an increase in thrust/performance (congestion effect curve). Both effects combine and result in the solid curve.
Because propulsion decreases as the flight speed approaches the exit speed of the gas/air flow, but propulsion efficiency simultaneously increases, the greatest propulsion efficiency of propeller engines is at relatively low flight speeds (right diagram). However, due to the high exit speed of their especially hot exhaust gas flow, turbojets only reach good propulsion efficiency at high speeds. Fan engines lie between these two extremes, the propeller engines and turbojets. Fan engines with small bypass ratios, such as are used in modern fighter aircraft, reach good propulsion efficiency at higher flight speeds than fan engines with larger bypass ratios. The latter are typically used in modern commercial aircraft.
Illustration 11.1-13: The volume, speed, temperature, and pressure of the gas flow through the engine determine many component-specific damages and problems.
Compressor blades give the flow more energy as they build pressure. Therefore, their operating behavior is considerably more sensitive to deviations (such as roughness, Ill. 126.96.36.199-8; tip clearances, Volume 2, Chapter 7.1; profile changes) than that of turbine blades, which absorb the energy from the flow.
Ingested foreign objects and stones cause erosion and impact craters. Due to the pressure-dependent thinner boundary layer in the rear section of the compressor, roughness on the blading has a considerably greater effect on efficiency here than it does farther forward in the engine. The compression temperature affects material selection. In modern compressors, titanium alloys are used up to about 500°C. In areas with sufficient pressure and air flow speed, there is the potential danger of a titanium fire (Volume 2, Ill. 9.1.1-12). Another problem specific to titanium is the damaging influence of fretting wear (Volume 2, Ill. 6.1-19) on dynamic strength. Stages at the compressor end with temperatures of up to 600 °C require the use of nickel alloys.
In the compressor, blade height, thickness, and spacing decrease towards the rear of the engine. This means that blade failures can have catastrophic consequences.
The combustion chamber has a determining influence on the emission of pollutants and also on the thermal stress placed on the hot parts located behind it. An especially important factor is unequal temperature distribution (Ill. 188.8.131.52-8). In the combustion chamber, the temperature of the gas flow greatly increases and causes damage through oxidation and thermal fatigue (Chapters 12.4 and 12.5). Oxidation of the fuel in the injection system can cause extensive consequential damages (erosion of the nozzle, flames escaping sideways). Gas vibrations due to unstable combustion, such as in low-emission combustion chambers, cause dynamic fatigue fractures and friction wear. The high gas pressure stresses the relatively hot combustion chamber housing, with the potential dangers of a pressure cooker with high wall temperatures and cyclical mechanical/thermal loads (Ill. 184.108.40.206-9). The gas speed is decreased in the combustion chamber after the compressor in order to ensure safe ignition and good burning of the fuel.
The expansion of the gas flow in the turbine requires the size of the cross-section to be increased in the direction of the flow, with larger spaces between the blades. This decreases the danger of possible blade fragments causing damage that increases in an avalanche-like manner, such as in the compressor. The operating behavior of the turbine blades, which absorb energy from the gas flow, is far less affected by roughness or damage than that of the compressor blades. The gas temperatures of modern engines exceed the melting points of blade materials in the high-pressure turbine. Therefore, these blades must be intensively cooled. Any compromising of the cooling, which can be caused by partial blocking of the cooling air bores with ingested dust, will lead to a rapid shortening of the life of the hot parts.
If afterburning is used, the gas temperature and speed in the afterburner pipe increase considerably once more relative to the temperature at the turbine exit, while the gas pressure stays fairly constant until the thrust jet. The pressure then decreases when the gas expands in the jet, and the gas speed increases. Vibrations can occur in the gas column in the afterburner pipe and can destroy the pipe, thrust jet, and engine (Ill. 11.2.4-11).
Illustration 11.1-14: Damages can be classified according to the frequency at which they occur in certain operating phases. In the following, this is described using several examples, but is by no means a complete list (Ill. 11.1-12).
“1”, Time between shutdown and restart: When an engine is shut down, the temperatures of the components change, which changes clearance gaps and can lead to temporary jamming of the rotor. When the rotor is cooling unsymmetrically, unallowable imbalances due to rotor bow are to be expected (Volume 2, Chapter 7). If parts of the oil circulation system (e.g. reservoirs, oil lines) overheat due to creeping heat from the hot parts, then there is a danger of coke buildup (line blockage, Ref. 11.2-13; reservoir contamination, Example 11.1-2). After shutdown, an (engine-specific) sufficient amount of time must pass before the engine is restarted, in order to ensure that no unallowable jamming and/or rotor bow occur (Ill. 11.1-17). Very long standing times (days, weeks), which are common with military aircraft, promote corrosion due to condensation water.
“2”, Engine start-up (Ill. 11.1-15): During engine startup, the starter may be damaged (e.g. if the start-up processes are too long or frequent). Inadequate drainage can result in fuel remaining after an unsuccessful start attempt, and restarting the engine with fuel remnants present can cause the hot parts to overheat and/or lead to uncontrollable turbine overspeed.
“3”, Accelerating to idle: If, during engine start-up, the compressor does not reach sufficiently high RPM (for example, due to rotor rubbing or malfunction of the fuel regulator system; Ill. 11.1-16), it can result in rubbing damages, overheating of the hot parts, or compressor stalls, all of which are accompanied by the danger of extensive consequential damages (Chapter 11.2).
“4”, The time at idle determines the temperature distribution in the turbine rotor disks (Ill. 11.1-17, Chapter 13). The induced thermal strain overlays with the centrifugal stress and affects the LCF life span of these parts (Ill. 220.127.116.11-8). With regard to engine part life spans, longer idling phases are desirable, since this allows the temperature distribution to be fairly even in the annulus and hub.
During taxiing, there is an increased risk of foreign objects and erosive particles being sucked up from the runway or from dusty air.
“5 and 6”, Accelerating to takeoff power: During acceleration to takeoff power, the overlaying strains in the rotor disks increase considerably to levels above those during steady operation (Example 11.1-3). This increases the risk of disk damage in situations where there is already sufficiently extensive damage (crack initiation, although uncommon). The temperatures and temperature gradients in the turbine stator vanes often cause crack initiation in typical zones even after short operation in this phase (Ill. 18.104.22.168-7). At first, the crack growth rate will usually decrease as crack length increases, which is due to a relaxing effect (Ill. 12.2-21). This can be taken into account during the design phase and be considered when determining projected life spans.
During takeoff, the low flight altitude means that there is an especially high danger of bird strikes (Volume 1, Chapter 5). Bird strikes in this phase are especially dangerous, since a large amount of power is required from the engines, and even minor damage to the blading can cause considerable loss of performance.
When the aircraft is standing or rolling slowly and the engine is running at high power, a ground vortex can form at the engine inlet and suck large, solid foreign objects up from the ground into the engine and/or cause dangerous vibrations in the forward compressor stages (Volume 1, Chapter 5.2).
“8”, Landing approach: There is also a high risk of bird strikes in this phase. Slow flight at low altitudes promotes ice formation (Volume 1, Chapter 5.1) on the nacelle or in the inlet. Ice that breaks free can pose a serious threat to the engine (ice strikes, Volume 1, Chapter 5.1.4).
“9”, Thrust reverser controls: The air flow directed diagonally forward by the thrust reversers can throw up dust and foreign objects from the runway, which are then ingested by the engine (Volume 1, Chapter 5.3).
“10”, Idling before shutdown: The time of idling before the engine is shut down determines how much the hot parts are cooled by the cooling air. This reduces the risk of coking in the oil system after shutdown, which occurs when the cooling action of the air flow and oil flow are no longer in effect, and the heat from the massive hot parts heats up the oil (heat soaking, Example 11.1-2). This causes blocking of oil lines and leads to contaminated/fatigue-prone roller bearings.
Illustration 11.1-15.1 (Ref. 11.1-14): The captain started the engine, but it did not accelerate. The start was aborted and preparations made for a restart (wet start; Ill. 11.1-16). While conducting the shutdown procedure, the flight crew and ground crew noticed black smoke and flames coming from the engine`s exhaust pipe. The attempted restart was aborted and the aircraft was evacuated.
Inspection of the engine revealed that the isolator of a spark plug on one of the two ignition circuits was cracked (see Ill.22.214.171.124-11). This prevented the ignition spark from reaching the combustion chamber. Spark plugs must be treated as wearing parts and must therefore be replaced in time. This case is a failure to do so.
The spark plug failure, combined with the captain acting against regulations and not completely shutting off the fuel flow, allowed a large amount of fuel to enter into the combustion chamber following the first start attempt. The accumulated fuel was then ignited by the misdirected spark from the faulty spark plug. Despite its alarming external appearance during the start-up, the fire did not cause any significant damage to the hot parts.
Illustration 11.1-15.2 (Ref. 11.1-11): This diagram shows the RPM of the (high-pressure) compressor and the temperature levels in the gas flow at the turbine inlet. Points A, B, and C mark the zones that are relevant for damages. The relatively complex startup procedure should be much less problematic in modern engines with “intelligent” electronic regulators with fast reaction times than it is in older engine types. In older engines, the performance lever (fuel supply) has to be adjusted while observing the RPM and gas temperature. Any mistakes can quickly lead to catastrophic damages. In case of a “hung start” (Ill. 11.1-16), the amount of fuel may be far too great for the small available air supply. This can quickly cause the turbine to overheat catastrophically (Ill. 126.96.36.199-7).
Illustration 11.1-16 (Ref. 11.1-11):
“1” The top diagram depicts a normal start (compare with Ill. 11.1-15).
“2” In the case of a “hung start” (false start) the RPM does not increase to idle after ignition, but gets “hung up”. At the same time, the gas temperature increases to dangerous levels due to insufficient air supply. It is vital that the starting process is aborted in time. Causes include regulator, sensor, or starter problems.
“3” In case of a “wet start” (no start), ignition fails to occur in the specified time (see Ill. 11.1-15.1), which can be seen in the RPM and gas temperatures` failure to increase. The cause is most likely to be a failure of the ignition system or fuel regulator. Following this type of failed start, any fuel remnants must be drained in order to prevent a hot start.
“4” hot start, see also Ill. 188.8.131.52-4): This occurs when the amount of fuel is too great in relation to the air flow. The temperature at the turbine inlet does not come back down to its normal value following overshoot (point B in Ill. 11.1-15). In addition to the causes listed in the diagram, the air flow can be decreased by ice or other large ingested foreign objects (e.g. plastic foils, cardboard) in the compressor.
Improper actions taken after a hung start can also lead to a hot start.
Guidelines given by manufacturers specify time-temperature values that enable one to recognize hot starts. The experience of the personnel is an important factor in recognizing hot starts in time and taking appropriate measures. Electronic regulators and suitable sensors in modern engines should reduce the risk of hot starts considerably.
Manufacturers usually prescribe a procedure with specific steps and rules to be taken following a hot start. If, after dangerously long time periods (seconds) at extremely high temperatures, simple external inspections (of the turbine on the exhaust side), boroscope inspections, vibration controls, and changes in the gas temperatures at specific power levels are not sufficient, disassembly of the hot parts and replacement of damaged parts (signs of overheating, see Chapter 13) may be necessary. Experience has shown that, due to high costs and time constraints, unallowably large measurement tolerances are often used. However, this creates a risk of dangerous consequential damages (including disk failure) that may only occur after longer operating times.
Illustration 11.1-17: Startup and the first few minutes of operation, as well as the shutdown process and the hours following it, have a decisive influence on the life span of and damage to hot parts (Chapter 12.6.1). Therefore, they are depicted again here (compare with Ill. 11.1-14) and related to two typical examples.
The increased risk of disk failure during the startup phase is well known (see Example 11.1-3).
Example 11.1-2 describes coking problems in a bearing chamber in the turbine area. This problem is typical for the shutdown process.
The procedure prescribed by the manufacturer for shutting down a low-output helicopter engine requires two minutes of operation at idling speeds before the engine can be shut down completely. This is intended to ensure that the part temperatures in the area of the rear turbine-side bearing chambers are low enough to prevent dangerous oil coking due to heat soaking.
During the development phase of a helicopter engine, a test run was conducted on the weekend with no abnormal occurrences (vibrations, etc.). The test runs were scheduled to continue after the weekend. Shortly after the engine was accelerated to full power from idle, the turbine disk of the first stage burst. Inspection of the fragments revealed that there had already been a large crack present when the engine was shut down. This crack evidently grew to a dangerous size during the last test run, but the disk did not burst. The high stress levels in the disk during the restart of the engine then caused the crack to become unstable and burst the disk.
11.1-1 B.M.Steinetz, R.C. Hendricks, “Engine Seal Technology Requirements to Meet NASA's Advanced Subsonic Technology Program Goals”. periodical “Journal of Propulsion and Power”, Vol 12, No. 4, July-August 1996, page 786-793.
11.1-2 “Engine oil coking and its effect on performance”, periodical “Aircraft Technology Engineering & Maintenance- Engine Yearbook 1999” page 48-53.
11.1-3 Jane's Aero-Engines, edited by B. Guston, ISBN 0 7106 1405 5, Jane's Information Group ltd, Sentinel House, 1683 Brighton Road, Coulsdon, Surrey CD5YH, UK, 1997, page 50-68.
11.1-4 R.Whitford, “Fundamentals of Fighter Design”, Airlife Publishing Ltd, 101 Longden Road, Shrewsbury, SY3 9EB, England, first published in the UK in 2000, page 80-85.
11.1-5 B.L.Koff, “Designing for durability in fighter engines”, Paper des 29th International ASME-KIVI Meeting in Amsterdam ASME-Paper Nr. 84-GT-164, 1984, periodical “International Journal of Turbo and Jet Engines” 1, 209-222 (1984).
11.1-6 M.Naeem, R.Singh, R.Probert, “Impacts of aero-engine deteriorations on military aircraft mission's effectiveness”, periodical “The Aeronautical Journal”, December 2001, pages 685 to 695.
11.1-7 periodical “Aerospace Propulsion”, “GE Finds Angled Fuel Tubes Solve Afterburner No-Light Problem”, February 1993, page 5.
11.1-8 J.F.Montgomery III, T.R.Sewall, J.J.Batka, “Aircraft Usage and Effects on Engine Life”, ASME Paper Nr. 88-GT-140, page 1-7.
11.1-9 A.Kleivan, A.Roed, Institute of Aviation Safety, Stockholm, Sweden “Flight Safety of Aircraft Structures and Systems”, Aeroroed Consult, Karstorp 159, S-26091 Forslov, Sweden, page 7.
11.1-10 R.A.Rudey, J.S.Grobman, “Impact of Future Fuel Properties on Aircraft Engines and Fuel Systems”, AGARD Lecture Series No96 “Aircraft Engine Future Fuels and Energy Conservation”, 1978, pages 6-1 to 6-29.
11.1-11 I.E.Traeger, “Aircraft Gas Turbine Engine Technology”, Glencoe Macmillan/McGraw-Hill, ISBN 0-07-065158-2, 1979.
11.1-12 G.E.Breitkopf, T.M.Speer, “In-Flight Evaluation of LCF Life Consumption of critical Rotor Components Subjected to High Transient Thermal Stress”, AGARD-CP-368, conference “Engine Cyclic Durability by Analysis and Testing”, pages 1-1 to 1-13.
11.1-13 R. Becker, “Der richtige Dreh zum Abstellen”, periodical “Aerokurier ”, 2/2002, pages 60 and 61.
11.1-14 Transportation Safety Board of Canada, Report Number A95P0138, incident on the 27th of June 1995.