The text to the illustrations mentioned below in Chapters 22.214.171.124 and 126.96.36.199 already contains suggestions for preventing damage in combustion chambers (Fig. "Typical combustion chamber damages"). For this reason, the following overview is limited to a compilation arranged into categories:
In addition to the actual damages, there are also problems, i.e. flaws and complications, that are more closely related to the function of the combustion chamber but do not affect life spans or safety. These include the requirements for lower emissions (Fig. "Combustion chamber soot problems"). Another issue is damage to other engine components. The turbine blading is especially vulnerable to damage due to poor temperature distribution in the gas flow at the turbine inlet (Fig. "Temperature variation at the combustion chamber outlet") or erosion due to coke and/or thermal barrier particles (Fig. "Measures against coke in combustion chambers").
Figure "Combustion chamber wall cooling technology" (Ref. 188.8.131.52-9): As shown in the diagram, over a span of 15 years the wall temperature of the inner combustion chamber wall was increased by about 600 °C. This evidently permitted a decrease in the cooling air temperature of about 70%. This is an incredible advance. However, it must be mentioned that technology such as carbon fiber-reinforced graphite (oxidation), transpiration cooling (blockages), and effusion cooling are currently unsuited for long-term serial applications.
Figure "Combustion chamber thermal fatigue behavior" (Ref. 184.108.40.206-7 ): The inner combustion chamber liner is a highly stressed part with a limited life due to the cyclical thermal loads. The historical trend of increasing turbine inlet temperatures (TIT, Fig. "Historical trends of of fighter engine problems") requires more and more constructive improvements. The metal sheet constructions that have been used for many years have reached their limits. The areas of the roll seam welds at the openings for the cooling air film are especially highly loaded by thermal strain (also see Volume 2, Chapter 9.3). The cause is large temperature differences between the air-cooled area of the bores (“1”) and the ends of the lips (“2”). The cooling effect of the air film already decreases considerably in this area. This makes these zones have the highest material temperatures in the combustion chamber. Ensuring oxidation- and thermal-fatigue life spans in this critical zone (top right diagram) requires calibration of the
In addition, the air feed for combustion must be taken into consideration to ensure a good distribution of the TIT. The diagram shows how optimizing the construction increases the life of the combustion chamber, if the relevant weak point is in the area of the cooling air slits. Wall rings produced with a chipping process (middle curve) allow more effective cooling, since the cooling air strikes the lip before it creates the cooling film. This cooling feed prevents individual air streams and promotes even cooling air films.
Further life increases can be acheived with dual wall construction. The supporting combustion chamber wall is protected from the hot gases by the inner wall (right). The inner wall is cooled by convection and protected by a cooling air film. Due to the lower mechanical loads, especially oxidation-resistant cast material can be used. This construction has special advantages with regard to thermal fatigue.
So-called shingle designs can fulfil demands for more effective cooling and minimal expansion restriction in modern combustion chambers (schematic diagram at bottom). The various operating loads on the combustion chamber wall are spread across two optimized structures. The relatively cool supporting outer wall (“A”) absorbs the forces from pressure differences and mechanical loads (fastenings, force introduction). It usually carries the inner shingle structure to protect it from thermal overloads (“C”). These shingles are usually attached non-permanently with a threaded connection (“B”), and coated with a thermal insulation coating.
Excerpt: “…After approximately two years of in-service use (in an air force turbo prop engine), inspections revealed that engines incorporating the LSM (low smoke modification) suffered from excessive turbine erosion. An investigation by …(the user) found that the erosion was caused by break-away of hard carbon deposits which formed in a fuel-rich region inside the dome of modified liners. Water tunnel testing, extensive single combustor liner testing and finally an accelerated endurance trial of a modified engine led to the development of an additional modification to the LSM specification to produce the revised low smoke modification (RLSM). The RLSM retained all the desirable properties of the LSM and reduced the hard carbon production to the rate no greater than that measured for standard, unmodified combustors.”
Comments: This is a can-type combustion chamber in an older engine type. The LSM modification of the standard combustion chamber was intended to reduce smoke emission by up to 80%. Evidently, this was a “disimprovement”, as it led to erosive coking. The extremely simple, limited measures depicted in Fig. "Measures against coke in combustion chambers" can evidently combine minimum smoke emission with the erosion-free operation of the standard combustion chamber.
Figure "Combustion chamber deformation" (Ref. 220.127.116.11-6): During the development of a large first-generation fan engine, deformations and buckling occurred in the outer wall of the annular combustion chamber (middle diagram). Evidently the part temperatures were so high, that these creep damages occured in hot zones within operating times typical for that series. The cause was the pressure difference over the combustion chamber wall. Evidently this involved a self-increasing process. The buckling of the lips that create the inner cooling air film (top diagram) led to hot streaks (Fig. "Hot gas streaks as combustion chamber problem"), which in turn promoted the collapse of the combustion chamber wall. The remedy for the damages in this case was to solder reinforcing rings to the outside of the combustion chamber wall. In addition, the lips for the cooling air film were waved (bottom diagram). This type of lip can absorb thermal strain elastically.
Figure "Measures against coke in combustion chambers" (Example "Disimprovement by low smoke modification"; Ref. 18.104.22.168-1): Soot creation in combustion chambers can be minimized with relatively simple measures. This requires knowing the cause of the sooting.
The top diagram shows the dome of a flame tube of an annular combustion chamber. Air enters the dome through several bores that are arrayed radially (“1”). Hook-shaped metal sheets (“2”) direct the flow around the circumference parallel to the dome wall. Evidently, too little air from the resulting swirl reached the radially outward areas of the guide sheets (bottom left diagram). In this zone, this resulted in a rich fuel/air mixture and coke deposits (“4”).
In an improved version (bottom right diagram), the bores (“1”) were made slightly larger. This intensified the air swirl (“5”). The air swirl even reached the outer zones of the combustion chamber dome and the fuel/air mixture in this area became leaner, suppressing coke formation. The bottom diagram shows how successful this was. The amount of coke particles caught in the exhaust gas in an hour is given over the air/fuel ratio.
Figure "Influences on combustion chamber wall heating-up" (Refs. 22.214.171.124-2 and 126.96.36.199-3): Optimizing the temperature of the combustion chamber walls can minimize or eliminate various damage mechanisms and problems. The following damage mechanisms are important for the part`s operating life, pollutant emissions, and repair costs (Fig. "Chamber wall temperature determines overhaul intervals"):
The combustion efficiency and the gaseous emissions are barely affected by the thermal barrier coating (Ref. 188.8.131.52-3).
The thermal conditions in a combustion chamber wall are determined by the amount of energy that is introduced and removed. This is primarily determined by convection (hot gases, cooling air film) and radiation (flame, walls; bottom left diagram; Fig. "Combustion chamber life dependence from flame radiant energy").
The combustion chamber surface facing the wall is especially important for heat transfers. It uses various configurations to create protective cooling air films and insulating coatings (thermal insulation coatings, thermal barriers; also see Volume 1, Chapter 184.108.40.206). This coating usually consists of a thicker surface area of stabilized zirconium oxides and a metallic intermediate layer (bond coating) between it and the base material. Bond coatings must provide a sufficiently strong bond between the thermal insulation coating and the substrate over long operating periods. This task is made more difficult by the tendency of ZrO2 to become an ion conductor at the high typical operating temperatures. This carries oxygen to the bond coating, causing it to oxidize and lose its bond strength. Therefore, the bond coating must have especially high oxidation resistance. Usually, these are spray coatings of the type MCrAlY, with “M” being a heat-resistant metal such as Co or Ni.
There are two types of zirconium oxide thermal insulation coatings currently in serial use.
In the primary zone, heat is transferred away especially by the intense radiation of the hot zirconium oxide coating. At high temperatures, the heat transported by the radiation can be greater than that carried by cooling air, since the radiation energy increases with T4 [°K]. The heat transfer in the combustion chamber wall plays an important role for the transport of the inflowing and outflowing heat.
Many properties of wall coatings influence convection and radiation (top table): coating thickness, coating structure, number of layers, and roughness.
The transparency of the zirconium oxide coating to thermal radiation varies. Transparency depends on whether the coating has a columnar (physical vapor deposition = PVD coating) or a lamelar (thermal spray coatings, see Fig. "Thermal barrier coatings of turbine rotor blades") structure. Important factors include the absorption, reflection, and radiation (emission) of the heat from both the surface of the thermal insulation coating and the bond coating (bottom right diagram). In the diagram, the coating properties are normed with respect to the wavelength of the spectrum of the thermal radiation. This merely allows a comparison between the progress of the lines. Above a wavelength of 2.7 x 10-3 mm, the reflection and radiation of ZrO2 PVD coatings increase considerably. These values drop continuously in the bond coating underneath. These properties can be very different at different wavelengths of the thermal radiation (Ref. 220.127.116.11-2). Because of this temperature dependency, the coatings in the combustion chamber can behave differently in different areas (e.g. primary zone, diffusion zone).
If the surface roughness plays a role, then roughening should be noticeable during the operating period (e.g. due to erosion and thermal fatigue). This changes the thermal conditions of the wall and the combustion chamber behavior. For example, in turbine blades, a frequently observed increase in blade roughness caused by erosion and deposits is assumed to increase the heat transfer by up to 60% (Ref. 18.104.22.168-8).
22.214.171.124-1 F.W. Skidmore, J.M.Bennett, D.E. Glenny, “An Investigation Into Hard Carbon Formation in a Modified Gas Turbine Combustor”, Proceeding Paper ISABE 95-7116, pages 1268 to 1273.
126.96.36.199-2 J.Manara, R.Brandt, J.Kuhn, J.Fricke T.Kell, U. Schultz, M.Peters, W.A. Kayser, “Emittance of Y2O3 stabilised ZrO2 thermal barrier coatings prepared by electron-beam physical-vapour deposition”, periodical “High Temperatures-High Pressures”, 2000, volume 32, pages 361-368.
188.8.131.52-3 H.F. Butze, C.H. Liebert, “Effect of Ceramic Coating of JT8D Combustor Liner on Maximum Liner Temperatures and other Combustor Performance Parameters”, NASA Technical Memorandum, NASA TM X-73581, December 1976, pages 1-11.
184.108.40.206-4 B.Jones, “Requirements in the Development of Gas Turbine Combustors”, Aviation Study Institute, Proceedings of the congress “Combustion and Flow in Engines”, Portugal, 14-25 Sept. 1987, pages 1-13.
220.127.116.11-5 J.J. Faitani, “Smoke Reduction in Jet Engines Through Burner Design”, Proceedings Paper ASE 680348 of the “Air Transportation Meeting”, New York N.Y. April 29 - May 2, 1968, pages 1 to 11.
18.104.22.168-6 B.H.Rowe, “Development Progress on CF6 Engines”, Proceedings Paper ASE 710421 of the “National Air Transportation Meeting”, Atlanta,Ga. May 10-13, 1971, pages 4 and 5.
22.214.171.124-7 B.L. Koff, “Designing for Durability in Fighter Engines”, Proceedings Paper ASME 84-GT-164, 1984, pages 3 and 4.
126.96.36.199-8 J.P.Bons, R.P. Taylor, S.T. McClain, R.B. River, “The Many Faces of Turbine Surface Roughness”, Proceedings Paper No. ASME 2001-GT-163 of the “45th International Gas Turbine and Aeroengine Congress and Exhibition”, New Orleans, Louisiana, June 4-7, 2001, pages 1 to 10.
188.8.131.52-9 C.A. Arana, “The Effect of TB C Utilization in the Design of Robust Aircraft Combustors”, Proceeding No. AGARD-R-823 of the “85th Meeting of the AGARD Structures and Materials Panel”, Aalborg, Denmark 15-16 October 1997, pages 18-1 to 18-9.