While the causes of and influences on damage in the compressor area are extremely varied, the potential remedies are equally numerous. They can be classified into several groups:
The first requirement, however, is recognizing a dangerous stall. This is fairly simple in the case of surges. Detecting rotating stalls, which can also be very dangerous, is more difficult as they are not easily detectable from the outside. In this case, the problem lies in latent damage to the engine components. This includes, for example, blading damage that can lead to a dynamic fatigue fracture at a later time.
Figure "Detecting and identifying a stall": How can one recognize a dangerous stall?
Figure "Compressor stability control": In order to prevent stalls and compressor surges, or to come out of these conditions as quickly as possible, there are several standard measures and procedures (stall recovery; also see Fig. "Reducing surge events by work line, surge limit"). These measures lower the operating line, thus increasing the margin to the surge limit. They can be used individually or in combination with one another, depending on the specific engine type.
The literature indicates that there are other possibilities that are evidently not in serial use. These include the localized, targeted injection of air into the compressor. This measure seems to be a focal point of current research into early detection and prevention of stalls.
If an engine compressor is in an unstable state (deep stall, surge, Fig. "Axial compressors operating characteristics"), lowering the operating line and/or internal improvement of the flow are the only methods of returning to normal operation. The first measures taken are usually reducing the fuel flow and opening the bleed valves. In single-shaft engines, a similar effect is achieved by opening the thrust nozzle. Because the thrust nozzle only influences the location of the operating line of the low-pressure compressor, in multiple-shaft engines this measure is only effective under certain conditions. These conditions are met when the low-pressure compressor itself stalls, or if the outflow conditions from a highly-stressed low-pressure compressor are poor enough to cause a downstream compressor to stall.
If the instability occurs in a low RPM range, then the problem is a rotating stall. Especially effective countermeasures are intermediate stage bleed and closing variable guide vanes. This allows the compressor stages to better adjust internally. Along with the lowering of the operating line due to the fuel reduction, the above measures additionally improve the stall limit and efficiency of the compressor by further lowering the operating line. All of these combine to increase the surge margin and stabilize the operating behavior of the compressor.
The most feared compressor instability is an unstoppable surge process (lock-in surge; Fig. "After burner triggered compressor surge"). The following are typical examples of causes for this operating condition:
Case 1: Angle of attack of the aircraft is too great (Fig. "Surge by flow disturbances at the compressor inlet"),
Case 2: Compressor damage (FOD, etc.)
Case 3: Flawed interaction between the engine and regulator system.
In Case 1, the flow must be improved by decreasing the angle of the aircraft.
In Case 2, power must be reduced by opening the bleed valves and the adjustable thrust nozzle (if applicable).
In Case 3 (most common) the fluctuating pressures during the surges are not properly valued. This prevents the fuel control system from clearly distinguishing the necessary parameters. It is unable to determine whether the compressor is operating in its primary or tertiary characteristic (Fig. "Axial compressors operating characteristics"). The regulator system then frequently supplies too much fuel, preventing the compressor from making the transition from the tertiary characteristic to the primary one. The compressor is decelerated out of the stable zone. In this situation, the only chance is to completely interrupt the fuel flow for at least 1-2 seconds (fuel blipping,). The compressor is then completely relieved for a moment and returns to its primary characteristic, causing any rotating stall zones to disappear. In especially stubborn cases, the engine must be completely shut down. Only when the RPM rate has dropped to a sufficiently low level can it be restarted with any chance of success.
Experience has shown the digital electronic regulators to be sufficiently fast for implementation of the above corrective measures. This should make the catastrophic overheating in the turbine following a surge largely a problem of the past.
There also seem to be indications that digital regulators could be able to prevent surges sufficiently early. Signals from sensors in the compressor airflow that indicate a preliminary stage of a rotating stall are analysed and initiate appropriate countermeasures (Fig. "Preventing a compressor surge", Ref. 126.96.36.199-2).
The problem with surges lies in the extremely short time intervals between the first detectable problem signals and the start of the compressor surge. The possibility of recognizing rotating stalls sufficiently early is considerably greater.
Figure "Preventing a compressor surge": For years researchers have been striving to develop a serially implementable system for preventing compressor surges and dangerous stalls, which can cause catastrophic overheating in the turbine (see Ref. 188.8.131.52-3). The demands for sufficient safety and reliability in this type of system have evidently not yet been met. Problem areas are the early detection of a developing surge and reliable operation. Sensors and data lines are subjected to influences such as korrosion, oxidation, dynamic fatigue, and erosion. A warning example is the frequent failure of relatively robust fire warning systems. Therefore, the system must be extremely resistant to operating influences.
Simpler attempts to create a functioning system for preventing compressor surges are based on prescribed actions. As soon as the operating conditions approach the surge limit and pass a certain prescribed level, actions such as fuel flow reduction, activation of a bleed valve, or variable guide vane adjustment occur (Ref. 184.108.40.206-2). However, this concept has the drawback that deterioration due to clearance gap increases, roughening of blade surfaces, profile changes, and erosion causes the operating behavior of the compressor to gradually worsen. Either the deterioration is addressed by making the system initially react too early, or the system must be recalibrated later to account for deterioration.
The depicted system, which is suggested in Ref. 220.127.116.11-1, is designed to prevent dangerous surges and functions by measuring the static pressure at the housing walls ahead of the first rotor stage of every compressor. The measured values are digitized, any interfering signals are removed and analyzed for characteristic signal elements that precipitate a rotating stall or surge. They are then continuously evaluated to detect the sudden pressure increases typical of rotating stalls and analyzed with a frequency analysis.
If an instability is detected, the logic module alerts the engine regulator, which should either catch the surge before it begins or at least shorten it considerably.
This type of system must be calibrated for each specific engine to correspond to the dominant, measurable, and analyzable effects.
Example "Electronic regulators" (Ref. 18.104.22.168-4):
Excerpt: “A scheduling problem in the low-pressure stator system of the …(big high bypass engine) was responsible for the single-pulse engine stall which hit the…(engine) during tests…The engine self-recovered after the event, which occurred during stall-margin tests….(the OEM) says that the scheduling problem with the variable stator vanes can be corrected with an adjustment to the electronic control system…“
Comments: This example shows the possibility of catching surges quickly thorugh the use of electronic regulators. Evidently, the solution to the problem in this case was mere reprogramming of the regulator.
Figure "Reducing surge events by work line, surge limit" (Ref. 22.214.171.124-5): In order to prevent stalls (low RPM) and surges (high RPM) before they occur, engines are outfitted with various devices and configurations (Fig. "Compressor stability control" and Fig. "Preventing a compressor surge").
A basic possibility is to set the stage- and total pressure ratios low enough that no stalls can occur throughout the entire flight envelope (Fig. "Application specific weak points of engines"), i.e. realizing extremely surge margins. A prerequisite is that no overly rapid acceleration of the engine (increasing fuel flow too quickly) occurs. This type of engine only requires an acceleration fuel regulator (maximum fuel limiter) that guarantees a sufficient surge margin.
This concept is used in older engine types, but requires considerable overdesign with detrimental effects on engine weight and performance. The performance concentration (thrust/weight ratio) of these engines is unacceptably low for today`s requirements.
In single-shaft engines with thrust nozzles with a variable shape (usually fighter engines with afterburners), the power increase can be faster than in those with thrust nozzles with fixed shapes (fixed exit cross-section). During acceleration, the nozzle can be opened correspondingly. The pressure in the thrust nozzle drops; i.e. it does not rise too quickly. The pressure behind the low-pressure compressor of a two-shaft engine also behaves correspondingly (top left diagram). This relatively low total pressure ratio can be increased again after the critical acceleration process. In this way, even during acceleration of the engine, the operating line of the compressor of a single-shaft engine or the low-pressure compressor of a two-shaft engine can be kept sufficiently far from the surge limit. A high acceleration rate may be necessary especially during touch-and-go starts and landing. The thrust nozzle is then opened farther at low RPM and closed at high RPM.
Dividing the compressor into several smaller compressors powered by their own shaft systems makes it possible to better adjust them to the aerodynamic requirements. This makes it easier to ensure the necessary surge margin. However, in the lower RPM range the end pressure of a compressor can be too high. In this case a countermeasure for the increased surge risk, and also for the more likely occurrence of rotating stall, is the installation of a bleed valve in an appropriate location. This keeps the specific operating line of the compressor low until an RPM range is reached that allows a higher compression ratio (top right diagram).
An alternative solution to the multiple-shaft principle is the installation of variable compressor guide vanes. These are usually installed in several front stages of a compressor. This measure can raise the surge limit (bottom right diagram), but has virtually no influence on the operating line. The bottom left diagram evidently applies to an older model single-shaft engine (e.g. J-79, Ref. 126.96.36.199-17). Modern engines often combine the multiple-shaft principle with variable guide vanes.
Variable inlet guide vanes (variable IGVs) can be used to improve efficiency without greatly affecting the operating line. The inflow angle of the airflow on the first rotor stage is optimized, allowing the amount of inflowing air to be controlled in lower RPM ranges at low engine power levels.
Another important related function of variable IGVs should be mentioned. When firing weapons such as cannons and rockets, hot gases and pressure waves can enter the compressor and cause surges (Fig. "Compressor surge influenced by cannons"). This can be prevented by an increased surge limit if the IGVs are closed immediately before the weapons are fired.
Figure "Blades and houses influencing aerodynamic stability" (Ref. 188.8.131.52-6): Gaps between the tips of the rotor blades and the housings have an especially serious influence on the stability of compressors (right diagram, Fig. "Airflow at compressor blade tips"). Small gaps shift the characteristic line of the compressor towards higher pressure and thus ensure a sufficient surge margin.
Operating influences such as erosion and wear due to rubbing increase the gaps (also see Volume 2 Chapter 7) and cause compressor stabitilty to deteriorate. The mass throughput, efficiency, and surge margin all decrease. Understandably, the relationship of the gap width “s” to the blade length “h” is important. A gap width that is tolerable for the longer blades at the front of the compressor may not be tolerable for the shorter blades in the rear of the compressor. Naturally, for small engines, maintaining small gap sizes is especially important. Appropriate measures ensure acceptable gaps over longer operating times (such as filters in helicopter engines, or robust blade profiles). However, the result is usually a compromise that does not allow a compressor to be completely optimized in both its weight and aerodynamics.
Wider blades, i.e. a smaller length ratio (blade length/chord length), lengthens the characteristic lines towards smaller throughflow (middle diagram) and increases the surge margin by raising the latter.
Casing treatment has a similar effect as changing the yield-strength ratio (right diagram, also see Fig. "Casing treatment minimizing tip clearances"). This can delay the start of rotating stalls.
Figure "Airflow at compressor blade tips" (Ref. 184.108.40.206-7): The tip zone and the gap are subject to special flow conditions. The diagram is intended to give an impression of the complex flow patterns. The flow near the blade tip is extremely important to the aerodynamic behavior of the blades and therefore also to the operating behavior of the entire compressor (Fig. "Blades and houses influencing aerodynamic stability"). In axial compressors, a considerable portion of the losses is related to the leakage flow at the blade tips. Several flow effects combine and alternate in this area:
The leakage flow in the tip clearance gap especially affects the main flow. The sharp, flat leakage flow (location of turbulence) creates vortices (leakage flow vortex) that roll up and dissipate in the direction of the main flow. The greater the gap, the more pronounced the leakage flow vortex and the backflow it creates. In the inlet area of the blading, horseshoe vortices were observed in the main flow. These form at the air buildups at the front blade edges near the housing walls.
Figure "Casing treatment minimizing tip clearances": Because the tip region of the blades is a major influence on the losses in the compressor (Fig. "Airflow at compressor blade tips"), it is a good area to strive for improvements. The primary method is a casing treatment to minimize losses and/or raise the surge limit (to shift it to low throughflow volumes). This can evidently be accomplished in very different ways. The diagram shows typical concepts.
A ring duct that is partially located ahead of the blades` inlet edge and opened inward can be used both in radial compressors (top right diagram, Ref. 220.127.116.11-9) and axial compressors (bottom right diagram, Refs. 18.104.22.168-11 and 22.214.171.124-12). The openings can be in the shape of ring slits or diagonal slits across the axial direction. This configuration evidently creates a vortex in the tip region, which increases the seal effectiveness (minimizes the leakage flow), especially at larger gap widths. The depicted solution is used with great success in the engines of Russian fighter aircraft. It can be assumed that this characteristic contributes to the ability of a certain fighter aircraft type to avoid surges during flights near the ground, where it ingests the air from the top of the wing with considerable flow redirection.
The top left diagram shows a radial compressor in which simple bores at the front of the compressor housing direct the air into the inlet area. This configuration appears to lower the surge limit considerably to lower air throughflow (Ref. 126.96.36.199-8).
Circumferential grooves with varying depth, width, and radial/axial angles are widely reported (Refs. 188.8.131.52-13 and 184.108.40.206-14).
There is no uniform formula with regard to the success and applicability of the individual solution. Evidently, in spite of the available calculating programs and high-performance computers, the selected configuration must be optimized in a testing process for every specific case and application. It is usually impossible to improve all operating characteristics of the compressor to the same degree. This applies, for example, to the surge limit and the efficiency in various operating states at different gap widths. It is more likely for an improvement of one characteristic to cause another to be compromised, or for improvements to be limited to a specific range of operation. This range is usually the lower and middle RPM range, in which rotating stalls occur at the blade tips.
Figure "Compressor stability by blade design" (Ref. 220.127.116.11-16):
Hub treatment (top diagram):
Another procedure evidently under consideration is a process similar to a casing treatment, but at the guide vane tips. This so-called hub treatment involves structuring the spacer surfaces across from the guide vane tips. Ref. 18.104.22.168-15 describes research in rotor spacers that have axial grooves oriented diagonally against the direction of rotation, similar to those found in housings. The tests were evidently successful in the described case (see excerpt):
Excerpt: ”…The experiments with the high stagger blading showed, for the first time, that hub treatment was very effective in delaying stator static pressure rise. For this blading, measurements of the stator exit flow field appear to indicate that the interception of stall was associated with a wall stall. With the treatment, the large blockage in the hub region was greatly decreased and, near the hub, the total pressure at the stator exit was found to be higher than the stator inlet. These results support the hypothesis that endwall treatment is effective when the type of stall that occurs is wall stall.”
However, other publications (Ref. 22.214.171.124-16) claim that the improvement in stability of the surge margin is achieved at the cost of decreased efficiency.
Trenching (middle diagram):
Along with casing treatments and hub treatments, there are several other possibilities for changing the seal effect at the blade tips so it has a positive effect on the flow stability of the compressor. The middle diagram depicts trenching, which is a circumferential groove in the housing roughly the width of the blading. The blade tips move into this groove. An improvement in efficiency in an acceptable operating range was only achieved through a single conical groove (middle right diagram). This had no positive influence on the dependency of efficiency on gap width. Other groove shapes worsen the efficiency and the operating range. The conical design is currently found in many serial engines and has evidently proven itself.
Gap bleed (no diagram):
Use of an air bleed at the blade tips is a possible method of minimizing leakage flow and improving compressor efficiency. However, the cost of blades with suitable air ducts would be considerably higher than that of conventional ones. In addition, the air duct could weaken the blades due to thin walls and the notch effect. This would increase the danger of dynamic fatigue fractures and serious FOD. An important factor for success is the angle and location of the bleed bores that must be spread along the entire mean line. It can be assumed that this configuration would also be useful for coming out of already-stalled flow conditions.
Tip treatment (bottom diagram):
A widely used method is “squealer tips” (aka feather edging; bottom left diagram). This process involves thinning the blade profile from the pressure side. The positive effects of squealer tips include minimization of the gap leakage flow due to increased flow resistance, and also less damaging rubbing behavior (lower thermal and mechanical loads on the blades).
There have also been reports of winglets (bottom right diagram) that worsened compressor efficiency. This behavior can be explained by increased frictional resistance and is the most likely reason for this method not finding a serial application.
Figure "Verification of compressor operation suitavility": The success of a compressor and, therefore, customer satisfaction in serial operation are determined by many factors:
A single test is not sufficient for determining whether or not these properties are satisfactorily achieved (Fig. "What to expect from a testrun"). A combination of different tests and trials is necessary). In the pre-development phase, the influence of various construction characteristics on the operating behavior can be investigated and optimized on a compressor testing rig under laboratory conditions (top diagram). There are at least two different philosophies for this:
The primary concern is the aerodynamic behavior and the verification of the best possible efficiency and good flow stability under reproducible test conditions. If maintaining tight clearances is the main goal, then it is not necessary to make the strength of the housing walls and rotors conform to serial standards. However, this philosophy seems to be a thing of the past. Today, compressor behavior is investigated with the use of the most realistic possible constructions and configurations. As far as the blading is concerned, the process is incremental. The investigation begins with smooth blades (Fig. "Surface structure improving flow properties"). This will determine whether or not their aerodynamic behavior is as expected. In the following steps, the roughness of the blade surfaces can be specifically increased to allow conclusions regarding operating influences such as erosion, corrosion, or fouling. This naturally requires relevant experience with this type of roughness increase. The results of these tests show maximum attainable limit values and reveal minor aerodynamic improvements. However, they do not provide much information about long-term operating behavior under realistic operating conditions (e.g. efficiency and flow stability after longer run times).
Far more accurate indications as to the behavior during serial operation, as well as to the demands of the acceptance regulations (containment, bird strikes, etc.), can be determined on an engine testing rig (middle diagram). This is especially true with elaborate cyclical running programs within the framework of an ETOPS rating process. Shorter runs, such as 150-hour runs, can only determine the functioning of the total system and its components. They are not suitable for determining long-term effects (e.g. wear and dynamic loads).
Flight operation (bottom diagram) including standstill times stresses and tests the engine in several areas that are not covered by the described tests. These include environmental influences (e.g. corrosion, fouling), maintenance (e.g. recognizing damages, practicality of specific maintenance procedures under field conditions), and the effect of acceleration forces due to flight maneuvers.
Figure "What to expect from a testrun": This diagram attempts to estimate which conclusions can be expected from different test runs. The estimations have been subjectively made by the author based on his experience. For this reason, they have no claim to quantitative validity, but only qualitatively indicate trends.
Acceptance and certification runs that serve to verify specific properties (bird strikes, containment) or verify performance and operating behavior do not yield a great deal of information regarding long-term effects in serial operation. 50- and 150-hour runs are better suited to this, but they are also fairly limited in the conclusions that they permit (Fig. "Tests of technologies need precise analysis").
It is clear that the most accurate conclusions regarding serial operation can be expected from ETOPS test runs. These extremely elaborate cyclical runs with long-term effects can be seen as the final step in the development of an engine.
126.96.36.199-1 Th.Herpel, L.Fottner, ““Transient Performance of Turbofan Engines with Respect to Unsteady Compressor Flow Near the Stability Limit”, Proceedings Paper ISABE 95-7124, of the “Twelfth International Symposium on Air Breathing Engines”, September 10-15, 1995 Melbourne, Australia, pages 1345 to 1354.
188.8.131.52-2 W.Ried, U.Blöcker, “Possibilities for Online Surge Suppression by Fast Guide Vane Adjustment in Axial Compressors”, AGARD-CP-421 Proceeding Paper of the AGARD-Konferenz “Advanced Technology for Aero Gas Turbine Components”, pages 31-1 to 31.15.
184.108.40.206-3 I.J.Day, “Active Suppression of Rotating Stall and Surge in Axial Compressors”, Proceeding Paper ASME 91-GT-87 of the “International Gas Turbine and Aeroengine Congeress and Exposition”, Orlando, FL. June 3-6, 1991, pages 1-8.
220.127.116.11-4 “GE90 surge cured”, periodical “Flight International” May 31 - June 6 1995, page 8.
18.104.22.168-5 “Jet Engine Accident Investigation”, General Electric, 1960s.
22.214.171.124-6 A.Schäffler, “Einfluss des Grenzschichtzustandes der Schaufelgitterströmung auf das Betriebsverhalten mehrstufiger Axialverdichter”, Seminar lecture at the TU Munich on January 17, 1978.
126.96.36.199-7 M.Inoue, M.Koroumaru, “Structure of Tip Flow in an Isolated Axial Compressor Rotor”, ASME-Paper 88-GT-251 of the “Gas Turbine Congress and Exposition”, June 5-9, 1988, Amsterdam.
188.8.131.52-8 European patent specification 0 229 519 B1 (Int. Cl.: F04D 27/02) filed 11.4.1990.
184.108.40.206-9 European patent specification 0 545 953 B1 (Int. Cl.: F04D 29/42) filed 30.7.1991.
220.127.116.11-10 M A. Sidorkin, inventor, patent specification RU-2162164-C1 (Int. Cl.: F04D 27/02 and 29/44) published 20.1.2001.
18.104.22.168-11 M A. Sidorkin, patent specification RU-2162165-C1, published 20.1.2001.
22.214.171.124-12 L.M. Wenzel, C.M. Mehalic, “Effect of Casing Treatment on Performance of a Multistage Compressor”, NASA Technical Memorandum NASA TM X-3175, 1975.
126.96.36.199-13 R. Hopkins, “An Experimental Evaluation of Axial Casing Tip Treatment and Diffuser Manifold in a Small Axial-Centrifugal Compressor”,Paper AIAA-83-1350 of the “19th Joint Propulsion Conference”, June 27-29, 1983, Seattle.
188.8.131.52-14 L.H. Smith Jr, “NASA/GE Fan and Compressor Research Accomplishments”,Paper ASME 93-GT-315 of the “International Gas Turbine and Aeroengine Congress and Exposition”, Cincinnati, Ohio, May 24-27, 1983.
184.108.40.206-15 P.Cheng, M.E. Prell, E.M. Greitzer, “Effects of Compressor Hub Treatment on Stator Stall Margin and Performance”,Paper AIAA - 83-1352 of the “19th Joint Propulsion Conference”, June 27-29, 1983.
220.127.116.11-16 J.Hübner, “Randströmung in Verdichtergittern”, FVV report for plan Nr. 574, 1992.
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