Figure "Stall at compressor blades" (Ref. 11.2.1.1-4): The function of a compressor blade can be compared with that of an airplane wing. Unlike wings, which can be adjusted, changing the angle at which the flow strikes the compressor blade is done by decreasing the mass flow during deceleration (Fig. "Technical compressor testing"). The top diagram shows the ascending force in relation to the angle of the blade. The ascending force initially increases linearly with the blade angle. At a profile-specific maximum blade angle, the flow suddenly breaks off in the rear section of the suction side of the blade (top of wing) and the ascending force decreases sharply. The same forward energy that creates ascending force in wings is transferred to the flow as a speed increase (kinetic energy) in compressors by the rotating motion of the blading and the redirection of the flow. In the larger blade ducts of the stators, the speed of the flow is decelerated and transformed into pressure.
On the suction side of the blades, the flow accelerates in the forward section and is decelerated to the exit speed in the rear section. The resulting allowable pressure increase is dependent on the boundary layer and limited by the stall of the flow.
The bottom left diagram explains rotating stalls (mild stalls, cold stalls) with the especially pronounced stall type known as a deep stall (Fig. "Axial compressors operating characteristics"):
While it is commonly assumed that a stall forms almost at once around the entire circumference of the blading, the process actually involves the flow stalling first at the blades of a limited circumferential area (cell) and then spreading to the rest of the circumference (Ref. 11.2.1.1-8). For example, stalls tend to occur at blades with geometric flaws, areas with greater tip clearances, or areas with local flow disturbances. The force necessary for a stall-causing influence decreases the closer the blading is to the maximum flow redirection. The local flow stall acts as a barrier and redirects the flow in such a way that the flow strikes the following blade in the same blading at a poor angle, but the blade that is running through the disturbance is again struck by a normally-behaving flow. To an outside observer, the cell of stalled flow seems to be moving at 10-50% of the circumferential speed in the direction of rotation of the rotor (i.e. 50-90% against the rotor movement). This is why it is referred to as a rotating stall. Several of these stall cells can occur arround the circumference with different sizes, radial lengths, shapes, and locations (Fig. "Axial compressors operating characteristics").
Some damage symptoms (crack initiation, crack distribution, crack shape) can be explained with the following hypotheses regarding unstable flows. However, specialists are still debating the correctness of these explanations.

Rotating instability (RI): This can be described as a source mechanism that rotates relative to the disk and differs from a rotating stall by periodic pressure changes in the source (bottom right diagrams). The turbulence stall of the RI can incite high-order circumference modes, while the speed of rotation is roughly half of that of the rotor. If the circumference modes resonate with the number of blades, it results in greater sonic radiation that can incite dangerous blade vibrations (Ref. 11.2.1.1-18). The creation of an RI can be explained as follows: there must be a clearance gap between the blade tips and housing that is large enough for an axial reverse flow to occur at the blade tips. This flow then causes periodic turbulent stalls on the suction sides of the blade tips.

Figure "Influences at loads by flow disturbing influences": The flow redirection caused by the blade creates a reactionary force that corresponds to the aerodynamic ascending force and the resistance of the profile. The rotor blade transfers these forces as a decelerating moment onto the shaft and as an axial force directed against the direction of flow (forward). The ascending force at the blade profile is created by the difference in flow speed on the top surface (suction side) and the bottom surface (pressure side). The aerodynamic loads correspond to the relative speed difference between the deceleration of the flow on the suction side Wmax and the exit speed W2 based on Wmax. If the angle of the blade profile is increased (Fig. "Stall at compressor blades"), i.e. the direction of the flow relative to the blade becomes steeper, then a flow stall will occur on the suction side and the ascending force will collapse. This means that the pressure differences between the suction and pressure sides were too great for the dominant flow conditions. When the ascending force collapses, the blade ceases to move any air.
The angle below which the air leaves these “stall ends” has changed relative to the blade with normal flow. The blade of the next stage that follows in the direction of flow is also subjected to a poor flow angle. If a stall occurs here, as well, then the local stall can rapidly spread throughout the compressor. The bottom diagram explains the important phenomenon known as the “blade passing frequency.” It is a pressure fluctuation at the blade tips with a frequency that corresponds to that of the passing blades at a point on the housing. This equals a very powerful supersonic load.
The total amplitude of the pressure fluctuation corresponds to the pressure difference between the pressure and suction sides of the blade profile near the tips. This pressure difference increases along with the aerodynamic loads on the blade, since it is related to the difference in flow speeds on the pressure and suction sides.
The greater the pressure levels in a section of the compressor, the greater the pressure amplitude (bottom left diagram). In the rear area of a compressor with high aerodynamic loads, the total pressure amplitudes can reach several bar. This puts extremely high dynamic loads on the soft abradable coatings on the inside of the housings, as well as on the typically thin-walled titanium alloy housings themselves. Corresponding dynamic fatigue damages are frequently observed on the coatings (Volume 2 Chapter 7.1) and housings.

Illustration 11.2.1-3 (Lit 11.2.1.1-6): Dangerous blade vibrations can be incited by a partially or completely stalled air flow (stall flutter). This will quickly result in blade failures (Fig. "Flutter problems at a fan"). The danger of this occurring increases with the thin blade profiles under high aerodynamic loads (Fig. "Influences at loads by flow disturbing influences") that are common in modern compressors. However, flutter can also occur without stalls. This occurs at supersonic speeds and is referred to as “ unstalled supersonic flutter” or “ choke flutter”.
Flutter is a self-increasing condition that is initiated by small vibrations of the blade, which can be caused by many different factors, such as flow disturbances ahead of and behind the blades (Ill. 12.6.3.3-20). Auto-incitement occurs when vibrations influence the gas flow and create pulsating gas forces that in turn intensify the blade vibrations. This is an interaction between mechanical and aerodynamic forces. It is very difficult to arrest flutter. Decreasing pressure is usually not sufficient; the speed of the gas flow must generally be changed to arrest flutter.
The top diagram depicts the process by which flutter is incited. In order for flutter to occur, the blade must first be sufficiently deflected by vibrations. These vibrations can be caused by various factors (see Chapter 12.6.3). Typical causes include flow disturbances or changes in the clearance gaps at the circumference. When this occurs, the blade not only twists during torsional vibrations (especially dangerous), but also during flexural vibrations, since the angle of the blade chord has changed. This means that the angle at which the flow encounters the blade changes. If the twisting of the blade sufficiently changes the angle at which the compressor blade encounters the flow and deflects it against the direction of rotation, it will result in a stall (1). If this causes the deflecting gas forces to weaken, the blade will swing back in the direction of rotation. This again minimizes the angle and the flow resumes (2), allowing gas forces to build up again against the direction of rotation.
The deflection of the blade also depends on the stiffness of the entire vibrating system, the element of which is the blade. In rotor blades, the element is the blade, in stators it is the shroud and the housing.
There are other types of flutter besides stall flutter (Fig. "Types of flutter excitement"). For example, supersonic gas flow in bladings can cause a type of flutter that is based on the reciprocal effect of the blade profiles with the shockwaves (Ill. 12.6.3.3-20).
Flutter occurs when a flutter limit is reached (Fig. "Types of flutter excitement"). This type of vibration is not limited only to the blades of axial compressors. The blades of radial compressors and turbines can also flutter. With flutter, the reciprocal influence of two neighboring blades in a blading on one another is important. This means that the behavior of a blading is considerably different from that of an airplane wing. The fluttering of a flag simply depends on a rhythmic stalling of flow eddies, which makes comparisons with blades problematic.

Figure "Types of flutter excitement" (Ref. 11.2.1.1-5): This diagram shows the position of different types of flutter incitement in a characteristic diagram for compressors.
Stall flutter (Fig. "Overloading blades by flutter") usually occurs in the subsonic range, but can also take place if the flow has local supersonic areas (transonic stall flutter).
Supersonic bending stall flutter occurs near the surge limit (stall limit) at high RPM and a high total pressure ratio.
Supersonic unstalled flutter occurs at high RPM below the surge limit (Fig. "Flutter problems at a fan"). This type of flutter can be further divided into “ high incidence supersonic flutter” slightly below the surge limit and “ low incidence supersonic flutter” near the work line (Fig. "Overloading blades by flutter").
Choke flutter (negative incidence flutter) occurs near the choke limit (Fig. "Diagram of compressor operating behavior") with negative angles of flow encounter.

Figure "Technical compressor testing" (Ref. 11.2.1.1-7): The compressor characteristic diagram (Fig. "Diagram of compressor operating behavior") describes the operating behavior. It requires knowledge of the RPM characteristic lines. To this end, the following data are determined at a certain specific RPM:

• total pressure at compressor inlet
• total pressure at compressor exit
• throughput (air volume over time, corrected for pressure and temperature of the ingested air)
• efficiency

The top diagram shows the schematic configuration of a suitable testing rig. The demands of a compressor characteristic diagram for an entire engine only permit the calculation of small parts of the RPM characteristic lines.
The exhaust gas flow can be restricted with a valve by changing the exit surface.
The test begins with the maximum exit surface, which is incrementally decreased (restricted). Sufficient time is left between the individual steps to ensure that the compressor reaches steady operation. The restriction causes the exit pressure at the compressor to increase and decreases the air flow (bottom left diagram). The points of measurement gained through this process are connected by the RPM characteristic line. If the gas flow is restricted further, the pressure increases become less pronounced (curve levels off) until the pressure finally begins to decrease. During restriction, the mass flow and the flow speed from the stator vanes (from CA to CB, bottom right diagram) decrease. The angle of attack of the flow on the profile becomes correspondingly steeper (from WA to WB). This causes a stall (surge) combined with powerful pressure waves (also see Fig. "Stall at compressor blades"). The surge limit is crossed. If possible, pronounced surges should be avoided since they can damage the compressor.
This type of measurement line is then repeated for other RPM values. Several measurement lines can be used to determine the surge line/limit (Fig. "Diagram of compressor operating behavior"). Lines with equal performance can be entered into the compressor characteristic diagram. If the mass throughput of the airflow is reduced with temperature and pressure, the resulting characteristic diagram is dependent on the meteorological values on the specific day.
The operating line is used in engines in which exit surface is fixed, i.e. the thrust jet is not adjustable. In this case, each RPM line will only have one work point. The operating line connects these work points.

The operating line should be located in the characteristic diagram in such a way, that the compressor/engine can be operated at maximum efficiency. In accordance with the depicted testing procedure, it is only strictly valid for steady operating states. If the specific values change quickly (e.g. startup, shutdown, output changes) then the operating state is unsteady and the position of the operating line in the characteristic diagram changes. At the same time, the engine regulators must ensure that the surge limit is not crossed.

Figure "Axial compressors operating characteristics" (Refs. 11.2.1.1-4 and 11.2.1.1-8): The two top diagrams are compressor characteristic diagrams with RPM characteristic lines. These are determined by the different patterns of the distribution and expansion of the stall zones in the compressor sections. Changes to individual blades (e.g. FOD or manufacturing tolerances) can cause local stalls even well below the surge limit (Fig. "Stall at compressor blades").
The undisturbed characteristic line in the characteristic diagram with no flow stalls is referred to as the primary characteristic. Characteristic lines that are disturbed by stall cells are referred to as secondary or tertiary, depending on their temporal order.
A seconday characteristic (rotating stall, Fig. "Stall at compressor blades") can have up to eight evenly distributed stall cells (2) and occur in a hysteresis range below the surge limit in compressors with multiple stages. Situations with two symmetrical cells are referred to as work lines with double characteristics, i.e. a special type of secondary characteristic. Diagram 3 shows that rotating stalls cause a relatively low decrease in performance. Only after the surge limit has been crossed does a deep stall (tertiary characteristic) occur with serious decreases in performance (1). In the deep partial load range (low RPM), permanent rotating stalls occur (4) even during “ normal” operating states. Because this can dynamically overstress the blades, extended operation at such low RPM is not recommended.
Every RPM value has a tertiary characteristic, which is a particularly serious type of rotating stall (deep stall). It occurs at very low RPM and is related to serious decreases in pressure and performance.
A surge is the cyclical stalling of the entire compressor flow with extreme vibrations between 5 and 30 Hz.
These processes, as well as certain types of surge (e.g. stall stagnation Fig. "After burner triggered compressor surge"), are described in more detail in Chapter 11.2.1.2.

Figure "Diagram of compressor operating behavior": The mass throughflow (reduced) is recorded on the abscissa of the compressor characteristic diagram. The RPM and mass throughflow are usually given as reduced values in order to make them independent of the meteorological conditions of the day on which the test occurred. This corrected form results from the inlet temperature and the suction pressure. This is used to determine the universal diagram (similarity diagram). If the values are compared with the design point, then RPM and throughflow can be calculated as percentages. The ordinate contains the total pressure ratio.
The diagram is split into two regions by the surge limit. Compressor operation is stable below this line (stable zone, light grey area). In the area above the curve (unstable area, dark grey), stalls lead to surges (Fig. "Stall at compressor blades", Fig. "Technical compressor testing", Fig. "Axial compressors operating characteristics" and Chapter 11.2.1.2). The compressor is normally operated on its work line (operating line). For the most common, steady operation, this work line should run through the area of the characteristic diagram with the best degrees of performance (Fig. "Technical compressor testing"). During flight operation, especially military flight operation, frequent unsteady operating states with rapid RPM changes are unavoidable. The corresponding operating point on the operating line must be reached as quickly as possible.
It is unavoidable that temporary deviations from the steady operating line will occur, but they must be below the surge limit. Problems can occur especially during the acceleration phase. Increasing output is done by increasing the fuel flow to raise the gas temperature and, therefore, the gas volume. The restricting function of the components that are farther along in the gas flow (combustion chamber, turbine thrust nozzle) causes the pressure to increase at the compressor exit. In order to avoid reaching the surge limit, the steady operating line must be sufficiently far from the surge limit (about 10-20% of the pressure ratio; surge margin). During deceleration, the operating line moves further away from the surge limit. The RPM characteristic lines in the diagram show the compressor behavior at a constant RPM rate and the changes in mass throughflow due to restriction (Fig. "Technical compressor testing"). The RPM rate is given in relation to the design value (=100 %). Steep RPM characteristic lines are, at first, a sign of stable compressor behavior. However, the steeper the RPM lines become and the less they curve off towards the abscissa, the narrower the areas with high efficiency and the adjustable range of the work line become.

Fig. "Factors worsening the compressor behavior"

Figure "Factors worsening the compressor behavior": This illustration shows influences that reduce stability. These influences can promote stalls by reducing the surge margin, and must therefore be prevented if possible.
These influences can raise the operating line and/or lower the surge limit. Effects like the increase of radial clearances (Volume 2 Chapter 7.0) can have both of these undesirable results. The operating line is generally raised by factors outside of the compressor. These, in turn, usually demand measures such as higher compressor exit pressures or increased mass through-flow to prevent the amount of power available to the compressor from decreasing.
The lowering of the surge limit is primarily caused by factors in the compressor itself.

Raising of the work line due to:

• increased cooling air volume
• power takeoff
• acceleration
• poor compressor and turbine efficiency
• increased pressure losses in the combustion chamber
• smaller turbine cross-sections (for example, due to poor blade tolerances)

Lowering of the surge limit due to:

• manufacturing tolerances of the blading
  
• damaged or rough blading
• increased clearances (increased radial blade play, ovalized housings)
• uneven inflow (pressure, angle, temperature, see Chapter 11.2.1.2)

These factors are dealt with separately later on.

Reduced turbine cross-sections:
The smallest cross-section in an engine is determined by the HPT stator. The smaller this cross-section is, the more the through-flow in the core engine is restricted, increasing the pressure and operating line, which also increases the probability of surges. Even apparently minor changes in the flow cross-section can have major unexpected consequences on the operating behavior of the engine. For example, the effect of manufacturing tolerances (form tolerances of cast parts, clearances, and seals) can result in critical cases in which part combinations must be put together for every single engine. Understandably, this makes the replacement of single stator vanes during repairs on the engine problematic. Even operation-related changes such as bulging or deformation of the blades (see Chapter 12.5 and Fig. "Creep damages at turbine inlet guide vanes "), or deposits in the flow path (Volume 1, Ill. 5.3.2-14) can have the described effect.

Power takeoff and air extraction in the compressor:
When unplanned mechanical (e.g. by the radial transmission path to aggregate components) power takeoff occurs, then the lost power is not available to the compressor for its compression work, putting it under greater aerodynamic stress.
The same effect occurs if there is an increased extraction of compressor air, which necessitates a larger mass flow.

Air extraction in the compressor can occur for various reasons, such as:

• cooling air for the hot parts (turbine blades, etc.). The hot parts, such as combustion chambers, turbine blades, disks, and housings, require a significant portion of the total air flow in order to keep their operating temperatures at sufficiently low levels to guarantee that the desired part life spans are reached. Changes to the hot parts and/or the air distribution system (leakages, etc.) can increase the necessary amount of air and raise the operating line. This can be an important indicator of the condition of an engine.
• cooling air to maintain clearances at the blade tips (ACC=“Active Clearance Control”)
  
• cabin air and pressurized air for the aggregates of the nacelle.
  

Increased clearances:

The radial clearances (tip clearances, Fig. "Factors worsening the compressor behavior" and Volume 2 Chapter 7) of blades, especially rotor blades, have the greatest influence on the surge limit.
Increased radial clearances that are greater than 1% of the blade length lead to a dramatic worsening of parameters such as through-flow, efficiency, and the surge margin (Example "Thermal expansion changing tight clearances "). The surge limit is lowered (flow disturbances at the blade tips) while the operating line is raised (decreased efficiency), decreasing the surge margin from both ends. The high compression ratios in modern engines demand very short blades in the rear (high-pressure) compressor area. The high compression temperatures cause correspondingly large thermal strain, making this area of the compressor especially susceptible to flow disturbances.
Clearance gaps that are not axially symmetrical (Example "Compressor case distortion") can have a worse effect on compressor operating behavior (surge margin) than axially symmetrical clearance gaps (even clearance gap along entire circumference, Ref. 11.2.1.1-15) with the same average clearance gap width. With equal average clearance gap width, the loss of tip pressure in asymmetrical clearance gaps was up to 50% greater than that in axially symmetrical clearance gaps. In comparison with multiple gap maximums, a single gap maximum on the circumference has an especially negative effect as soon as it overlays on a certain sector. Asymmetry has considerably less of an influence on the tip efficiency, which is largely determined by the average gap value.
Labyrinths, both in spacers and especially at the compressor exit, have a considerable influence on the operating behavior of compressors when clearances and leakage rates are increased.

Uneven inflow:

Uneven pressure and temperature distribution around the circumference of the compressor inlet (Fig. "Flutter problems at a fan" and Fig. "Factors increasing intake temperature") have a negative effect on compressor performance. The blading prevents the uneven zones from evening themselves out around the circumference. The following model can explain this phenomenon: the compressor behaves as though it were made up of parallel “pipes” arranged around the circumference. With a set outlet pressure, therefore, the “pipe” with the smallest inlet pressure must have the highest pressure ratio. The surge limit is first reached inside this “pipe”, and acts to destabilize the entire flow system. However, the number of disturbances around the circumference does not necessarily have an effect on the worsening of the surge limit. The determining factor is the sector in which the disturbance is located.
Areas with inlet pressure disturbances and flow stalls are subject to increased energy levels and temperatures. This results in a smaller aerodynamic RPM with the same pressure ratio, which decreases the surge limit. Things to be avoided include unsuitably shaped inlets (Ill.
11.2.1.1-13), missing inlet bellmouths, disturbances due to thrust reverser actuation (Fig. "Dynamic loads by disturbance of the inlet flow"), recirculation of hot exhaust gases, and the ingestion of hot steam (takeoff catapult, Chapter 11.2.1.2). Turbulence in the inlet flow (such as due to a ground vortex; Fig. "Dynamic loads by disturbance of the inlet flow") can cause serious blade vibrations and lead to blade fractures. This type of turbulence can also be caused by ice buildup in the inlet area (for example, due to a de-icing system malfunction; Volume 1 Chapter 5.1.4).

Increased volume of cooling air for the hot parts:

The volume of cooling air can increase if, for example, increasing leakages occur in the cooling air duct to the hot parts due to increases in the seal clearances (labyrinths; see Volume 2 Chapter 7). However, damages such as crack initiation in cooled hot parts (e.g. turbine stator vanes; Chapter 12.6.2 and Fig. "Temperature caused damages at high pressure turbine vanes") can also cause dangerous cooling air losses.

Acceleration:

Acceleration of the engine, i.e. an increase in rotor RPM, requires an increase in the fuel flow. This increases the combustion chamber pressure (Fig. "Diagram of compressor operating behavior") that must be built up by the compressor - this raises the operating line considerably. Only a worsening of compressor efficiency beyond the designed parameters, i.e. the regulator`s capacity, should lead to the surge limit being reached.
Therefore, surges should be seen as an indication that the engine, and especially the compressor, is to be inspected (for example, boroscope inspections of FOD) for possible unallowable deviations.

Worsened compressor and turbine efficiency:

Worsened compressor efficiency necessitates an RPM increase in order to maintain the mass flow, which raises the work line. Typical influences that affect efficiency are covered in the section that deals with the lowering of the surge limit.

If the turbine efficiency worsens, it means that there is less power available for the compressor. This causes the RPM to decrease. In order to bring the RPM back up, the fuel flow must be increased, which in turn raises the required pressure at the compressor exit. Roughness and clearance gap air losses also reduce turbine efficiency and have similar effects.

Increasing pressure losses in the combustion chamber:

Pressure losses in the combustion chamber result in increased resistance for the airflow, which raises the operating line. Pressure losses in the combustion chamber that are greater than the designed values may be caused by foreign objects, blockages, or warping (see Chapter 11.2.2).

Manufacturing tolerances of the compressor blading:

Naturally, the profiles of the compressor blading must be within the allowable tolerances. Because of the large number of blades and the principle of their aerodynamic function, which is especially sensitive to geometric deviations (tolerances!), changes in the statistical distribution can result in unallowable effects. Typical problem zones are the leading edges (radius, transition) and the transition to the root platform. Deviations in the profile and blade angle can also lead to problems.
A special problem is posed by scattered measurements that do not correspond to the required design and were not detected by quality control, and only make themselves known later during operation. Deviations can also occur when special operating demands require coatings such as lacquers or erosion protection coatings.

Damaged or rough blading:

During operation, compressor blading can be deformed by erosion or foreign object strikes or be geometrically altered or roughened by abrasive wear (see Volume 1 Chapter 5.3). Roughness occurs primarily on the leading edge and the pressure side. Micro-deformations play an important role on the leading edge, especially on the transition to the suction side, where the flow tends to dissolve. The higher the pressure levels, the thinner the boundary layer becomes, allowing ever smaller areas of roughness to have damaging effects (Ill. 11.2.1.1-9 and Fig. "Needed smoothness of blades in compressor zones"). For this reason, greater roughness is acceptable in the forward compressor area than in the area nearer the compressor exit.
If a foreign object has caused serious local deformations on a blade, stalls will originate <U>here</U>. Because this type of damage is not easily recognized from the outside (Volume 1 Chapter 5.2), it increases the danger of dynamic fatigue damage.
Fouled blading:

The ingestion of dusts and exhaust gases, combined with sticky blade surfaces (oil vapours, leaked oil) can result in deposits that alter the shape of the blade (Ill. 11.2.1.1-9). The decreased efficiency of the compressor lowers the surge limit. The compressor may require cleaning. This can be done according to prescribed procedures with suitable liquid and/or solid cleaning materials (such as rice husks, ground apricot pits) without disassembling the engine or removing the compressor (Ref. 11.2.1.1-1). It has been observed (Ref. 11.2.1.1-9) that the structure of roughness on the blade is an important determinant of the speed with which deposits build up on the surface. For example, deposits build up more rapidly on surfaces that are cleaned with blasting processes.

Fig. "Blade roughness influencing compressor"

Figure "Blade roughness influencing compressor": Excessive roughness on blade surfaces has several undesirable effects:

• deterioration of compressor efficiency
• can cause flow instabilities leading to rotating stalls
• worsened operating performance (lowered surge limit)

Deterioration of efficiency (Ref. 11.2.1.1-9): Compressor bladings generally make use of turbulent boundary layers, since only these are able to overcome the extreme increase in pressure.
Every turbulent boundary layer has a laminar (or viscuous) underlayer (see lower diagram)near the wall. The thickness of this underlayer depends on the Reynolds number (i.e. the relationship of the flow speed to the kinetic toughness “w/n” of the air) relative to the chord length. This relationship is determined by the flight conditions on one hand, and by the position of the blade in the compressor system on the other. The kinetic toughness contains the air density, which rises continuously towards the compressor exit.
Extensive tests on pipes and flat plates have shown that roughness only increases resistance if the uneven sections jut out of the laminar underlayer and penetrate into the turbulent boundary layer (bottom diagram; also see Ill. 12.2.1.1-10). In other words, large peaks increase resistance. Indication of an average roughness value “Ra” is of minor significance. Therefore, the average roughness value is not as important for friction losses and compressor efficiency deterioration as the largest peaks, which are represented more accurately by the absolute roughness value.
It is important that, with surfaces with oriented roughness, such as those that have been machined or subject to erosion, roughness measurements are done in the direction of the air flow. It has been demonstrated that grooves oriented at an angle of more than 10° from the direction of flow strongly increase resistance and must be covered by a roughness measurement.
In the range of small Reynold`s numbers, greater roughness can have a positive laminar-turbulent effect on the transition and improve the compressor characteristics. However, this is usually accompanied by a decrease in efficiency. Aside from the absolute roughness value, the distribution of the roughness on the blade edges and across the surface is also important. For example, roughness on the suction side of the leading edge (“1”) has a greater influence than it does on the pressure side (“3”) and at the trailing edge (“2”). This is due to a combination of two effects: increased surface friction and relocation of the transition point of the laminar into the turbulent flow to a point farther forward.

Occurrence of rotating stalls (Ref. 11.2.1.1-4): Naturally, larger surface anomalies such as spalling or delaminating coatings (Volume 1, Ill. 5.3.2-6), corrosion scars, and FOD marks, which frequently occur in the inlet area, have an especially strong influence on efficiency and the probability of rotating stalls occurring.
Worsened operating behavior (lowered surge limit): Both of the above effects (efficiency deterioration and premature rotating stalls) can lead to stalls occurring in the compressor under poor operating conditions.

Figure "Surface structure improving flow properties" (Lit 11.2.1.1-21): On aircraft fuselages and wings, the air resistance is reduced with the aid of a special type of roughness known as shark skin. There are also efforts to apply a similar designed roughness to the surface of compressor blades to improve their performance in the air flow. The realistic usable size of the roughness depends on the laminar boundary layer (Fig. "Blade roughness influencing compressor"). This means that the structure of roughness would have to be extremely small in the rear area of a modern compressor. On the other hand, technically designed roughness should be already feasible on the fan blades of large bypass engines. However, this area is affected by operating influences that could cause unallowable damage to the structure of this kind of roughness. These influences include erosion, water drop impacts, fouling, and possibly delamination of the foils that carry the roughness structure. For these reasons, there is no known serial implementation of this type of technology.

Figure "Problems to realize allowable blade roughness" (Ref. 11.2.1.1-20): It is of great practical importance to have the most exact knowledge possible regarding the allowable roughness of compressor blading. If unnecessarily smooth surfaces are required during production, the excess costs can be considerable. On the other hand, roughening of the blading during operation through erosion and/or corrosion (Volume 1, Chapters 5.3 and 5.4) can cause compressor and engine efficiency to deteriorate (Volume 2, Ill. 7.0-3). The effects of this roughening are increased fuel consumption and higher gas temperatures, which shorten the life of the hot parts.

Allowable roughness:

The allowable roughness for the highest efficiency decreases towards the compressor exit and is dependent on the pressure ratios and flight condition. The greater the density of the air, the smoother the blades must be. The diagram shows roughness as a function of the determining parameter w/n = Re/c (see Fig. "Technical compressor testing" and Fig. "Needed smoothness of blades in compressor zones").
An important difference is whether the guarantee points of an engine are at high altitudes (civilian) or near the ground (usually military). In these cases the determining factor is whether the area concerned is the inlet stages or the rear section of the high-pressure compressor. The allowable roughness in the inlet stages is several times greater than in the exit stages.

Definition of a roughness k_p that is relevant for compressor blades:

Fig. "Blade roughness influencing compressor" already indicated the problems with a practical blade roughness. The critical roughness value of technical blade roughness was determined by analyzing test data from the 1970`s. In this process, the allowable roughness value k_p was defined as the difference of the arithmatical average values of the ten highest tips and ten deepest troughs over a measuring area of 1 mm. A perthometer can be used to determine these data.
There is currently no correlation between the given analysis and the similar definition “Rz”, and this task would require reworking on the basis of the analyzing methods and quality standards that are common today.

Figure "Needed smoothness of blades in compressor zones" (Ref. 11.2.1.1-4): The diagrams each depict two zones:

• turbulent boundary layer with a hydraulically smooth surface; the roughness peaks do not penetrate the laminar underlayer
  

• turbulent boundary layer with a hydraulically rough surface; the roughness tips penetrate the turbulent boundary layer
  

Experiments (by Nikuradse) determined that hydraulically smooth surfaces merely change the friction resistance by the Reynold`s number (Re), regardless of the roughness (as long as it is sufficiently minor). Increasing Re numbers in this range cause the friction resistance to decrease. The Re number increases along with the flow speed and pressure (the kinematic toughness decreases with air density), and decreases as the temperature rises (kinematic toughness increases as temperature increases). However, in hydraulically rough surfaces, in which the roughness peaks penetrate the laminar underlayer of the turbulent boundary layer, the friction resistance increases as roughness increases, independent of the Re number. The relationship between friction, flow, and roughness of the blade surface, dependent on the Re number, is shown in the top diagram. One can determine a limit specific value, Re_ks,all= w.k_s/v = 100, below which the roughness has no influence. The roughness value k_s is comparable to “sand roughness”. Roughness created by machining in the production process has a different profile from the non-directional and jagged “sand roughness”. Surfaces produced by machining have roughness cross-sections with soft waves across the direction in which they were machined. Roughness measurements of surfaces with oriented profiles, which are done to determine aerodynamic effects, must be done in the direction of flow and include grooves that are angled more than 10° from the direction of flow. For these surfaces, Re_k,all is between 75-260. For forged and etched/electrochemically manufactured blades, this value is about 90. In these cases, Re was determined by the allowable roughness value k_p that was defined in Fig. "Surface structure improving flow properties". The average roughness value “Ra” is not suitable for determining Re_all because, as explained earlier, the determining factor is whether or not the highest peaks penetrate the laminar layer. In general, for the above technically manufactured surfaces, k can be determined by the relationship k=9 Ra.
This allows calculation of the maximum allowable roughness of blade surfaces for a specific compressor stage. This is important not only for the manufacture of blades, but also for considering the effects of potential damage that may occur in later operation (erosion, corrosion).
The bottom diagram depicts the polytrope efficiency of the compressor over the Re number. In this case, the hydraulically smooth area is below the curve. If the intake pressure and Re number increase in the hydraulically smooth area, then friction decreases and the efficiency increases (efficiency deterioration is reduced). This is true up to a specific point, where the roughness peaks penetrate the laminar underlayer of the boundary layer. The curve then flattens off. This means that efficiency is independent of the Re number and does not improve any further. This transition occurs earlier at greater roughnesses.
For the critical roughness Re_k,all , the only determining factor is the quotient of the flow speed “w” and kinematic toughness “v”. This quotient corresponds the Re number relative to the chord length “l”. Towards the compressor exit, “Re” increases along with pressure and speed, while “l” decreases, meaning that Re/l increases sharply. For this reason, the rear stages are the first to exceed the limit roughness value, and the roughness requirements also increase towards the compressor exit. During high-speed flight, the absolute roughness (Rt) of a blade must be below 0.001 to be hydraulically smooth.
Attempts to improve aerodynamic performance through designed surface structures (shark skin) must also take the described relationships into account. This means that these structures would be too small for technological realization on blades in the rear compressor area.

The roughness of compressor blading can increase during operation in several ways:

• erosion
• corrosion
• abradable coating particles building up after rubbing
• adhering dust particles (fouling)
• spalling or delaminating coatings (lacquers)

The largest portion of surface friction is created by the front half of the suction side of the blade (Ill. 11.2.1.1-9). The boundary layer is extremely thin here, especially at high speeds. Fortunately, erosion occurs primarily on the suction side of blades and has a relatively minor influence on their air resistance. Even after long run times under extreme operating conditions (e.g. erosion tests), acceptable compressor performance is to be ensured within the framework of engine development.
Experience has shown that increased roughness, as long as it is below 0.025 mm (Rt=25), will affect the efficiency and performance of an engine, but will have very little direct influence on the throughflow and surge limit of a compressor. This is provided that erosion has not noticeably changed the profile of the blades. However, the worsened efficiency raises the operating line of the compressor (Fig. "Diagram of compressor operating behavior" and Fig. "Factors worsening the compressor behavior") and therefore indirectly reduces the surge margin.

Figure "Causes of stalls and or surges in compressors": This diagram summarizes factors that promote stalls in compressors, and categorizes them according to the components to which they are causally related.
In addition to influences that lower the surge limit and/or raise the operating line (Fig. "Factors worsening the compressor behavior") during “normal” operation, the diagram also includes those that cause stalls due to a sudden damaging effect, such as ingested large foreign objects that alter the shape inlet cross-section, blade fractures, blade damage due to bird or ice strikes, and damage that changes the shape of flow paths in the hot part area.
Many influences can accumulate over long operating times. Typical examples include the deterioration of seals (e.g. increasing tip clearances due to erosion and heavy rubbing) and roughening of blades (Ill. 11.2.1.1-9).
One can see that these influences can also have their source far from the compressor, such as in the afterburner (see Chapter 11.2.1.2).

Figure "Uneven pressure and temperature in compressor inlet" (Lit 11.3.1.1-4): Uneven inlet pressure and temperature distribution around the circumference (Chapter 11.2.1.2) is especially damaging for compressors (top left diagram) because pressure anomalies around the circumference only even out to a very limited degree. Instead, the compressor operates as though it were comprised of two or more parallel compressor sectors that channel air into a common container. This means that the sector with the lowest inlet pressure must create the greatest pressure ratios, and this sector will also be the first to reach its stalling point and destabilize the gas flow (Fig. "Factors worsening the compressor behavior"). Experience has shown that a destabilized sector of about 80° - 120° results in a maximum lowering of the stalling point (top right diagram). The stalling point does not sink any further beyond this sector size. It must also be mentioned that several narrow destabilized sectors distributed around the circumference do not lower the stalling point any more than a single large destabilized sector.
Uneven pressure distribution with serious effects on the stalling point can be caused by, for example, sharp-edged supersonic inlets with large angles relative to the flow (Fig. "Surge by flow disturbances at the compressor inlet") that, during supersonic flight, cause the inlet flow to stall.
Even more serious disturbances occur during supersonic flight with only minor negative angles, since the flow stalls and creates intense local fluctuations due to the interference of the boundary layer and the compression surges emanating from the central body (nose cone).
In compressors with many stages or multiple shafts, inlet pressure disturbances are transformed into sector-like temperature disturbances (increases) by increased energy flow into the heavily disturbed areas, which occurs in the front stages or, in multi-shaft engines, in the front compressors. This uneven temperature distribution causes the disturbed sector to run at a lower Mach number or aerodynamic RPM N/T^-1/2~. However, it must still create the same pressure ratio as the unaffected zones. This reduces the surge limit (bottom diagram). This damaging effect is comparable to a corresponding pressure disturbance at the engine inlet. Temperature disturbances can occur in several different ways: through operation of the thrust reverser, due to the formation of a ground vortex (Fig. "Compressor problems by ground vortex") caused by recirculation of exhaust gases (e.g. in VTOLs), through the ingestion of hot steam (catapult takeoff), or through firing weapons (Fig. "Surge by flow disturbances at the compressor inlet").

Figure "Surge by flow disturbances at the compressor inlet": Firing weapons such as cannons (A) or rockets creates “temperature ramps” inside the entire compressor inlet. During tests, the gas temperatures increased at rates between 2000 and 8000 K/s. Such a sudden, brief temperature increase can seriously decrease the efficiency and aerodynamic stability of an engine (Ref. 11.2.1.1-10). If the hot gases of the on-board cannon enter the engine with the inlet air flow, it can alter the operating behavior of the compressor to the extent that it causes the combustion chamber to flame out (Ref. 11.2.1.1-3). In order to prevent this from occurring, the cannon is located above the strake so that the hot gases are directed over the wing and cannot enter the inlet duct. In addition to the hot gases, pressure waves from cannon mouths and rockets can disrupt the flow. As shown in Fig. "Uneven pressure and temperature in compressor inlet", their effects include temperature increases in the air flow. Another similar type of pressure wave is caused by lightning strikes on the aircraft fuselage, which can also cause engines to flame out (Volume 1 Chapter 5.1.3). The extent of the influence a temperature increase has depends on the pressure levels, the sector angle of the disturbed area, the compressor RPM, the absolute temperature increase, and the speed of the temperature increase at the compressor inlet.
In helicopters, VTOLs, and activated thrust reversers, there is a danger of hot gases being recirculated and/or the creation of a ground vortex (Fig. "Ground vortex caused by thrust reverser", Chapter 11.2.1.2). Both effects can have an unallowable influence on the flow at the engine inlet, and have evidently caused several aircraft accidents.
If a fighter aircraft with engine inlets mounted on the side of the fuselage (B) yaws, it will result in a serious performance decrease if the disrupted boundary layer of the nacelle, which is situated at an angle relative to the flow, affects the engine inlet (Ref. 11.2.1.1-7).
(C) The fuselage flow creates a thick boundary layer. In order to prevent this layer from entering the engine inlet, the inlet is located on the fuselage in a way that allows the boundary layer to flow past through a gap between the inlet and fuselage. The edge of the inlet duct, which is in a forward position parallel to the fuselage, acts as a splitter plate. It prevents compression surges from entering the fuselage boundary layer and increasing resistance by causing flow disruptions. In order to prevent disruptions at the splitter edge due to steep angles, the edge is rounded off. The boundary layer on the splitter plate is drawn through a field of bores ahead of the inlet in order to improve the quality of the inlet flow.
If an overly strong side wind acts on a nacelle engine (D) at low forward speeds (such as takeoff), the already high speed of the inlet flow on the side of the nacelle lip exposed to the wind increases to the speed of sound. When this flow is decelerated in the diffusor section of the inlet duct, it can cause flow stalls. If these flow stalls enter the compressor, they can incite unallowably powerful blade vibrations. Therefore, the manufacturer of aircraft types in which this problem occurs usually requires that the engine is only brought up to full power when the ground roll speed is sufficiently high. This flattens the angle of the side wind relative to the engine axis, corresponding to the speed triangle consisting of the side wind and the head wind created by the aircraft`s motion. With one large military transport aircraft type, for example, a rolling takeoff is required if the side wind exceeds 45 km/h.
According to Ref. 11.2.1.1-7, another anomaly occurs with supersonic airflow. An unstable compression surge with a tendency to vibrate is created in the inlet duct. The surge is swallowed with the frequency of a low buzz and pushed out again (diffusor growl). This sound indictates a dangerous operating state (Example "Sensitivity of supersonic inlet").

Figure "Secondary flow in the intake duct" (Ref. 11.2.1.1-12): Swirl in an engine inlet influences engine efficiency to various degrees, depending on the size, type, and vulnerability of the affected engine. Depending on the direction of rotation of the engines of two-engine fighter aircraft, only the engine with poorer inflow conditions may tend towards inflow instability.
The symmetry of the inlet ducts combined with the flight envelope (see Fig. "Reasons of aircraft flight envelope limits" and Fig. "Application specific weak points of engines") create swirls that are caused by two different swirl types overlaying in the supersonic inlets of all modern fighter aircraft.

• Twin swirl
• Bulk swirl

Moving the inlet opening of the intake duct and engine, which leads to a S-shaped diffusor duct (Fig. "Dynamic loads by disturbance of the inlet flow"), has a decisive influence on swirl formation.

Twin swirl: This type of swirl is created by the simple curving of the air flow caused by the boundary layer in the duct. S-shaped double bends in ducts can also cause this type of swirl. These swirls occur outside of the central axis at the engine inlets of two-engine fighter aircraft with separated inlet ducts (diagrams at bottom and in Fig. "Dynamic loads by disturbance of the inlet flow"). In single-engine fighter aircraft with two symmetrical inlet ducts on the fuselage, two symmetrical swirl zones occur (Fig. "Dynamic loads by disturbance of the inlet flow"). These are caused by high-energy flow zones (higher pressure, greater density) that are pressed outward in the curves of the duct, while low-energy zones are located at the inside of the curves.
Twin swirls are very stable and it is extremely difficult to suppress them by installing specialized components in the inlet duct.

Bulk swirl: If a stall occurs at the duct entrance (Fig. "Surge by flow disturbances at the compressor inlet"), it creates an additional type of swirl that can twist the main air flow. The resulting flow twist is an average value of the flow angle on a fixed radius. In two-engine fighter aircraft with separate inlet ducts, the creation of bulk swirl can be explained by using a model representing half of a twin swirl (Fig. "Dynamic loads by disturbance of the inlet flow"). Depending on whether the flight speed is subsonic or supersonic, and the position of the ramp, the high-energy flow zones form at the top or bottom of the duct. Therefore, the deflection in the duct occurs in different directions of rotation, and enters the engine either with or against the direction of fan rotation (top diagrams). Unlike twin swirls, bulk swirls are sensitive to external flow conditions, which influence the location and intensity of the low-energy flow zones (low total pressure, low density).
In pronounced cases, swirl can cause engine surges and flutter vibrations in fan blades (Fig. "Overloading blades by flutter").

A combination of twin swirl and bulk swirl occurs as a typical flow form in supersonic inlets.

Figure "Dynamic loads by disturbance of the inlet flow" (Refs. 11.2.1.1-12 and 11.2.1.1-7): The top diagrams depict typical pressure distributions that form in the inlet ducts in front of the engines of single- and two-engine fighter aircraft due to swirl (Fig. "Secondary flow in the intake duct").
One must expect the flow in the plane at the end of the inlet duct ahead of the engine to have uneven pressure distribution. This pressure distribution is largely determined by the operating conditions and can change suddenly. Depending on the sensitivity of a specific engine to this type of uneven pressure distribution (distortion), manufacturers limit it with a distortion index (DC parameter). If this limit value is exceeded, it can result in compressor surges or vibrations of fan blades that travel through the uneven pressure flow.
The DC60 parameter shows the unevenness of the pressure distribution in a 60° sector of the flow cross-section ahead of the engine that has the lowest pressure values (i.e. the greatest flow losses; bottom diagrams). P_m is the average value of all total pressures in the inlet. P_60min is the corresponding average value in the 60° sector. The average impact pressure “q” over the inlet cross-section serves as a reference point. Compressor surges are safely prevented as long as DC60 is smaller than the value determined by the manufacturer throughout the entire inlet duct. In fan engines, this only applies to the flow portion of the core engine. The limit values are determined in wind tunnels and/or during flight testing of an aircraft type.

Figure "Ground vortex caused by thrust reverser" (Ref. 11.2.1.1-13): It has long been observed that there is evidently a relationship between the use of thrust reversers and compressor surges in the affected engines, and that this relationship should be seen in connection with the ingestion of recirculated hot gases (Fig. "Compressor problems by ground vortex"). This evidently causes temperature distortion in the air inflow. Compressors react to these temperature changes, which occur in the very short time period of a single rotor rotation, by causing surges.
In order to prevent this type of inflow disturbance, there are special operating guidelines for the use of thrust reversers. In a large fan engine on a testing rig, the blowing off of compressor air and/or the positioning of guide vanes minimized the risk of stalls, but this was not true for actual operation. This indicates that stalls are not caused by the uneven temperature distribution alone. An investigation showed that serious pressure disturbances occurred in the inflow immediately after the thrust reversers were activated. At first it was assumed that these flow disturbances were in a 60° sector (Fig. "Dynamic loads by disturbance of the inlet flow"), but subsequent tests showed that the main cause for the stalls was to be found in the formation of a ground vortex (inlet vortex, sketch 1; also see Volume 1, Chapter 5.2.1.2). The speed and pressure distribution in this type of vortex is shown in sketch 4.
Vortex formation is aided by tailwind in the same degree as the foot of the vortex moves forward (Ref. 11.2.1.1-19). This causes the vortex to occur nearer to the engine axis (2), and allows it to enter the high-pressure compressor (core engine). These conditions are created during thrust reverser operation. The vortex creates pressure and temperature disturbances in the inflow of the core engine (5). In the case in question, the cool fan air flow and hot gases from the turbine-side thrust reverser (3) were carried deep into the compressor of the core engine (high-pressure compressor) by the vortex.

Figure "Pressure shock by suddenly blocked inlet flow" (Ref. 11.2.1.1-22): If an air flow in a duct is suddenly decelerated, its kinetic energy becomes a pressure pulse. This effect is known from water lines (water hammer, left diagram). In extreme cases, this can cause serious damage to pressure lines and fittings. A comparable effect also occurs in ducts with an air flow, such as the inlet duct of a fighter aircraft engine. Engine surges can suddenly totally or partially block the inlet duct like a valve (bottom diagram). The disruption then travels forward from the engine inlet against the flow towards the duct inlet. At the same time, pressure increases in a few tenths of a second in the affected section of the duct (top diagram). A hard engine stall in an engine can cause a shock-like pressure increase in the inlet duct (hammer shock). In this way, compressor surges can even cause damage to peripheral components outside of the engine. Recorded incidents involved rivets and metal sheets coming
loose, especially the rods of pop rivets in prototype assemblies. This can cause extensive foreign object damage in the engine.

These destructive powerful pressure pulses dictate the structure and strength configuration of the inlet ducts of modern fighter aircraft. In these aircraft, the adjustable ramps (Fig. "Hammer Shock consequences") and the seals of the flaps (Refs. 11.2.1.1-24 and 11.2.1.1-25) are especially threatened.
The strength and duration of a surge pulse determines the power of a hammer shock and depends on various factors:

• The type of mechanism that causes the compressor surge. The location of the stall is important (in the fan or hub region). Other important factors are whether the stall occurred in the compressor and spread to the fan, and whether the stall only affects a small section at the circumference of the fan.
• The Mach number at the engine inlet.
• The total pressure ratio.
• The bypass ratio; the stress on fan engines is lower than that on simple single-shaft engines. However, there are also reported cases of damage in front of fan engines (Fig. "Hammer Shock consequences").
• Size of the gap between the fan and compressor (i.e. splitter). The greater the gap, the greater the intake cross-section into the compressor, and the less sensitive the fan is to the influence of flow disturbances from the inlet and the compressor. A portion of the pressure surge from the compressor can more easily escape into the bypass duct, where it can minimize the pulse intensity at the fan, which causes hammer shocks and determines their strength.

In the depicted case of a test aircraft from the 1950`s, a buzz vibration in the afterburner (see Fig. "Excitement of combustion vibrations") caused a rotating stall and about 15 % increase in pressure (Fig. "Indicating values for combustuion vibrations") at the wall of the inlet duct in the plane of the engine inlet. This caused a hammer shock. Evidently, fixed ramps in the inlet duct were plastically deformed and seals were damaged. The pulse was especially hard because the engine was a single-shaft engine, i.e. there was no damping bypass duct. The most powerful pressure shocks can be expected when the inlet flow is completely blocked. Relative pressure peaks, i.e. the peak pressure in the hammer shock relative to the total pressure in the undisturbed airflow, of 1.6 were measured. This means that the stress levels of the walls of the inlet duct were increased by 60% due to the internal pressure. The size and temporal progression of the pressure pulse depends on several factors:

• Volume of the inlet duct.
• Distribution and weakening of the engine surge pulse
• The greater the engine`s bypass ratio, the lower the expected stress on the inlet duct.
• Presence of a bypass at the duct inlet.
• Measures to control the boundary layer at the duct entrance (blow-off).
• Size of the separation gap between the fan and compressor inlet (splitter; bottom diagram).
• Flight speed. The maximum of the relative hammer shock pressure peak increases from flight speed 0. At about Mach 1, it reaches a maximum and decreases slightly at higher flight speeds.

Figure "Hammer Shock consequences" (Ref. 11.2.1.1-25): In this case an auxiliary air-intake flap (middle diagram) that was mounted on the front side of the inlet duct (top diagram) was damaged by a hammer shock (Fig. "Pressure shock by suddenly blocked inlet flow") following heavy engine surges. This is notable because the bypass duct reduces the hammer shock stress relative to a single-shaft engine (Fig. "Pressure shock by suddenly blocked inlet flow"). The surges were caused during a test flight of a prototype by bird strikes or too much fuel in the afterburner (see Fig. "After burner triggered compressor surge" and Example "Unfavourable afterburner conditions at high altitude"). The flaps each consists of two sections that are connected by joints and attached to the wall of the inlet duct by a hinge at one end. When the flap was partially or completely opened, the hammer shock was able to enter the inner space of the flaps and force the flap parts apart (bottom left diagram). This caused a joint to fail (bottom right diagram). The damage required a redesigning in which the flaps and joints were strengthened and the joint location was optimized.