7. Rubbing and Maintaining Clearances
In this text, rubbing is taken to mean contact between surfaces, usually sealing surfaces with radial, tangential, and/or axial relative movements. This occurs when the clearance between a rotor and static engine parts is bridged (Fig. "Rub tolerant blade tip systems").
In the following, two types of rubbing (tribo-) systems are differentiated by their respective functions (Fig. "Properties and tests of rub systems"). A rub-tolerant system is created when the wear of friction occurs on the blade tip or labyrinth tips. These systems are present when hard rub-tolerant coatings are used across from compressor stator vane tips (e.g. on a rotor intermediate ring/spacer).
An abradable system is created when wear occurs on a soft abradable coating (e.g. in the compressor housing), which is worn down without noticeable wear on the opposing surface (e.g. rotor blade tip). Abradable systems are advantageous in those area, where repairing parts (e.g. depositing compressor rotor blade tips) is too cumbersome or prone to damage. Rubbing systems are used in those areas where the coating must be sufficiently hard for resisting high mechanical operating loads (e.g. centrifugal force in an intermediate ring) or where high erosion stress would wear down a soft coating too quickly.
In order to minimize leakages, clearances must be as small as possible. This can result in contact between the sealing surfaces, even in systems where such (abrasive) contact is not intended. In abradable systems, the bridging of clearance during operation is consciously designed to occur in order to minimize clearance, at least for the present operating conditions.
Influence of Seals on Engine Performance:
The performance of an engine and its components is to a large degree dependent on the air- and air/oil sealing systems between the rotor and static components (
Engines on commercial aircraft show a slower decrease in performance over their typically long run times than engines on military aircraft with typically high maneuvering loads and more frequent start-up cycles (Ref. 7.0-4)
Apart from the concomitants of performance decreases, there are many potentially damaging effects:
- Temporary insufficient engine output for certain maneuvers can result in accidents (thrust gap).
- Overheating of disks and shafts due to hot gas incursion (e.g. around seal rings in the turbine).
- Decreased effectiveness of hot part cooling (e.g. cooling of high-pressure turbine blades) shortens the life spans of parts and/or causes them to fail due to overheating.
- Main bearing damage due to altered thrust loads (to high or unstressed) resulting from pressure changes around the rotor disks.
- Oil escaping and, in extreme cases, oil fires if air seals in bearing chambers fail.
- Ignition of an oil fire in a bearing chamber or a titanium fire through sparks and local overheating.
- Fatigue fractures in the blading after oscillations following blade rubbing (Example "Blade rubbing in a turboprop engine").
- Crack initiation in labyrinth- and blade tips due to thermal fatigue and/or dynamic fatigue resulting from high friction temperatures that have damaged the part zones (e.g. embrittled, de-hardened).
- Erosion of the compressor and turbine due to metal removed during rubbing.
- Metal removed during rubbing blocking cooling-air ducts and thus overheating cooled hot parts.
- FOD to the blading due to large coating fragments breaking out.
- Shaft separation and/or detachment of a rotor due to wear and high friction temperatures.
Rub-tolerant systems are primarily used as seal systems for blade tips and labyrinths. A special function is also served by rub-tolerant surfaces that decelerate the rotor if a main rotor shaft fails due to axial offset. This minimizes the risk of titanium fires igniting. In low-pressure turbines, the decelerating effect is increased by blade contact (intermeshing principle, see Volume 1), in which case the blading is destroyed. Rub-tolerant systems are used primarily in situations where the coating is on the rotating part. A typical example is a spacer ring in a high pressure compressor rotor. These are usually hard thermal-spray layers on a ceramic base with special adhesive layers that can handle the forces (centrifugal and friction) and cyclical elastic strain (under centrifugal force, heat strain).
Hot parts (e.g. ring segments in the high pressure turbine housing with a zirconium oxide spray layer) use rub-tolerant systems with hard ceramic coatings. Their heat resistance and oxidation resistance give them an acceptably long operational life.
A special type of rub-tolerant system is washer bushing seals in bearing chambers for oil/air mixtures. In this case, graphite or ceramic rings run on an oil film against steel rings.
Abradable systems are used to minimize the clearance between the tips of compressor rotor blades and the housing, and also in labyrinth seals. These coatings are easily removable and minimize damage to blade and labyrinth tips. In order to make this material removal possible, these coatings are usually porous or are made from a phase mixture with low inner hardness (e.g. Ni/C spray coatings, filled silicon rubber, filled synthetic resin). The hardness of the coating can be increased, depending on the hardness of the blade tips that dig into it. Even in abradable systems, a sufficiently high hardness is required in order to prevent long-term damages from erosion.
If the blade tips of the turbine rotor are coated with special wear-resistant armor (e.g. soldered CBN or SiC particles), then these can form an abradable system with relatively hard ceramic layers.
Soft abradable coatings are usually applied to static part surfaces such as inner housing walls. However, they can also be used in labyrinth seals on the inside of rotating hollow shafts and rings.
Figure "Seals and engine performance" (Refs. 7.0-1 to 7.0-4):
In civilian aircraft (top right diagram) deterioration of engine performance in shape of a SFC increase usually occurs during the first 1000 hours of operation. This can be plausibly explained by the relatively small tolerances in new engines that are milled out in the first few engine runs under typical operating conditions. This effect is reinforced by seal coatings opposite blade tips and labyrinths with honeycomb structure, casing treatment, or porosity. In these types of seals, the sealing effect of the small clearances decreases significantly at the start of the enlarging process. Once clearance has reached a certain size, the amount of leakage from clearance increases does not increase nearly as much as it did initially (Fig. "Seal gaps and fuel consumption").
Tactical aircraft exhibit a considerably faster performance loss than civilian aircraft. The top left diagram shows data from two different groups of military aircraft engines at a constant turbine intake temperature.
In general, the greater loss of thrust in military engines is natural, since thrust changes and maneuvers are considerably more strenuous than in civilian aircraft. It is difficult to quantify the individual factors that influence thrust loss.
In cases where the same engine type was used in aircraft with different maneuvering loads, the thrust loss in aircraft with greater maneuvering loads was about twice that of the other aircraft. The best evidence, however, can be taken from the behavior of engines in testing rigs. Engines in testing rigs have roughly half the thrust loss of engines in flight after 400 hours in a accelerated mission cycle test.
The lion`s share of this difference is evidently the maneuvering loads (g-forces) of the engine in flight. Similar effects can be observed in the life span of main bearings. The thrust loss is primarily caused by rubbing wear between rotating and static parts and also FOD (erosion, damage to blade edges).
Apart from a pronounced break-in period, gradual performance deterioration occurs throughout the entire operational life, and cannot be completely restored even by repairing the engine.
The bottom diagrams (Ref. 7.0-5) show the comparatively small thrust loss of a civilian three-shaft engine (indicated by turbine intake temperature or SFC) compared to two-shaft engines of the same performance class. Low turbine intake temperatures lengthen the life span of hot parts considerably. The manufacturer explains the especially advantageous behavior of the three-shaft engine as being due to the relatively short and stiff shafts, i.e. short main bearing intervals and good housing design (Fig. "Sealing advantages of 3-shaft engines"). Structural housings that absorb external forces are separated from the inner, flow-carrying housings.
Figure "Seal gaps and fuel consumption" (Refs. 7.0-8, 7.0-14): The top diagram shows performance loss in the form of an SFC increase during cruising flight of the individual components of an older large engine type with a high bypass ratio. It is clear that a 0.5% drop in engine performance is caused by clearance increases in the low- and high-pressure compressors (compressor and turbine). In the high-pressure turbine, clearance is most likely additionally affected by noticeable heat strain.
The bottom diagram (Refs. 7.0-9 and 7.0-10) shows, that the increased fuel consumption is due primarily to blade tips and labyrinth seals rubbing during the first few cycles of material removal. The loss in performance then slows considerably over the relatively long life of this civilian engine type, corresponding to erosion and heat strain.
This behavior illustrates the importance of suitable material removal runs, which are problematic due to the fact that they are often not easily comparable with regard to the influence they have on clearance during serial operation.
Clearance increases due to stressing factors during operation (e.g. vis inertia, gyroscopic force, or aerodynamic loads affecting the large engine nacelle) usually only have a minor influence. Surprisingly, the SFC-increase is so small that it was not included in the diagrams.
Figure "Engine behaviour and tip gaps" (Ref. 7.0-6 and 7.0-7): The top left diagram shows that clearance increases in the compressor not only increase specific fuel consumption (SFC), but also decrease engine efficiency and the mass flow rate. This decreases the surge margin (see Fig. "Coatings at tip seals"), i.e. flow instabilities are more likely to occur (top right diagram). In order to counter this effect, the gas-intake temperature at the high-pressure turbine must be increased so as to reach initial performance levels.
The clearance between the rotor blade tips and housing has a considerably greater impact on engine performance than the clearance between the guide vane tips and the rotor hub (middle right diagram).
Rotors with labyrinth intermediate rings and guide vane tips outfitted with a shroud are clearly less sensitive to clearance changes than designs with smooth spacers and shroudless guide vanes (bottom right diagram).
Clearance changes in the high-pressure areas (compressor and turbine; Fig. "Seal gaps and fuel consumption") have the strongest impact on SFC. However, a decrease in efficiency in the low-pressure areas (fan and low-pressure turbine) increases fuel consumption (middle left diagram) more than an equal efficiency loss in the high-pressure areas. If the SFC-increase is nevertheless affected more strongly by the high-pressure areas, then this indicates that the changes during operation are more extreme in the high-pressure area than in the low-pressure area. It is therefore understandable if extra measures are also taken to minimize alterations in the high-pressure area during operation.
In radial compressors (bottom left diagram), the axial clearance “s” between the blade tips and the housing at the compressor exit is especially important. The low blade height “h” in this region means that even small gaps are damaging. Therefore, minimizing the axial clearance at the exit of radial compressors is an important design task.
Example "Surges in the high-pressure compressor due to increased clearance" (Ref. 7.0-11):
“Following a rash of single - and dual engine surge events …the FAA is requiring airlines to perform special tests to determine stability of the high pressure compressor section…
The HPC surge issue surfaced…when a number of high-time…engines began to experience surges under take off power, chiefly due to excessive compressor blade clearances in the HPC caused by different thermal growth rates of the compressor rotor and stator assemblies. If the clearance between the compressor blade tip and the stator assembly is too large, under high power conditions the airflow can become distorted and reduce compressor stability.
(Ref. 7.0-12): The FAA is ordering on-wing tests…of series engines to prevent high-pressure compressor (HPC) surges, which can occur during takeoff and climb.
(Ref. 7.0-13): …Although the engine has been prone to surge events since it was introduced, the double surge is believed to have led to calls for urgent action…the surge margin, which appears to have deteriorated faster than expected, particularly on older engines.
Comment: The crossing of the surge margin in the high-pressure compressor during takeoff and climb is explained by a thermally induced clearance increase between the rotor and housing. Engines with longer run times and correspondingly larger deterioration are naturally more sensitive. The time this effect begins can be plausibly explained by the high engine output. Additionally, elastic deformations of the high-pressure compressor housing (backbone bending) can promote the surge process (see Chapter 7.1.4).
Figure "Coatings at tip seals" (Ref. 7.0-6): Clearance loss can dangerously decrease the margin between the surge limit and the operating levels (surge margin) in two ways (top diagram). Smooth housing walls lead to surge margin decreases and increases in the operating levels after clearance losses. Molded and/or open coatings (honeycomb) can increase the surge margin. The surface structure and possible open hollow spaces in the anti-rub coating influence the behavior of compressor after clearance changes. Although the leak rate may be increased by molded surfaces when compared to smooth ones, the surge limit can be increased significantly.
The bottom diagrams schematically illustrate these effects in a fan. The dashed line represents the surge limit with a clear housing wall. If the housing wall has an open-celled honeycomb structure, then the surge limit is increased (dash-dotted line). Certain molded patterns, such as optimized circumferential rings (bottom right diagram), have the same effect as an open honeycomb pattern. This is referred to as “casing treatment”.
Openly porous coatings, such as metal felts, can have similar effects.
This design detail explains why fan housings in Russian engines perform so well. They are equipped with a circumferential open canal with axially diagonal open slits to the flow canal (Ivanov configuration).
7.0-1 W.R. Beverly, J.G. Sweeny, “Life Cycle Fuel Consumption of Commercial Turbofan Engines”, AIAA Paper No. 76-645 of the AIAA/SAE 12th Propulsion Conference, Palo Alto, California, 26-29 July 1976.
7.0-2 G.P Sallee, H.D. Ruckenberg, E.H. Tooney, “Analysis of Turbofan Engine Performance Deterioration and Proposed Follow-on Tests”, NASA CR-134769.
7.0-3 J.F. Jr Dugan, J.E. Mc Aulay, T.W. Reynolds, W.C.Storck, “Fuel Conservative Engine Technology”, NASA-SP-381
7.0-4 G.M. Mulenburg, J.G. Mitchell, “Simulation of Turbine Engine Operational Loads”, Paper 77-912 of the AIAA/SAE 13th Propulsion Conference, Orlando, Florida/July 11-13, 1977. (Sec. HT0108)
7.0-5 Technical manual of the R.R. Co. for engine RB211.
7.0-6 L.P.Ludwig, “Gas Path Sealing in Turbine Engines”.
7.0-7 C.W. Elrod, “Advances in Gas Turbine Engine Sealing”, Paper G5 of the Proceedings of the “9th International Conference on Fluid Sealing” S. 922 - 312.
7.0-8 D.L. Nored, J.F. Dugan jr, N.T. Saunders, J.A. Zioemanski, “Aircraft Energy Efficiency (ACEE) Statur Report”, Paper 2092 of the Proceedings of the NASA Conference “Aeropropulsion 1979”, Cleveland, Ohio.
7.0-9 R.H. Wulf, “CF6-6D Engine Performance Deterioration”, Paper NASA -CR -15978G, 1978.
7.0-10 R.L. Martin, W.J. Olsson, “Operating Flight Loads and Their Effect on Engine Performance”, SAE-Paper 811071
of the “Aerospace Congress& Exposition”, Anaheim, California, October 5-8, 1981.
7.0-11 E.H.Phillips, “FAA Targets PW4000 Engines”. periodical “Aviation Week & Space Technology”, May 3, 1999, page 54.
7.0-12 G. Norris, “PW4000 operators face surge inspection” periodical “Flight International” 10 - 16 March 1999, page 5.
7.0-13 periodical “Aviation Week & Space Technology, April 26, 1999, page 24.
7.0-14 G.P. Sallee, “Performance Deterioration Based on In-service Engine Data”, JT9D Jet Engine Diagnostics Program, NASA-CR-159525, pages 91-108.