User:SidewinderX/Sandbox/Turbine blade

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Turbine blade from a Turbo-Union RB199 jet engine.

A turbine blade is a shaped vane, typically of metal, which converts directional flow of a fluid (whether liquid or gas) to rotational motion of a turbine. Fluids range from water in water turbines commonly used in hydroelectric plants, to steam in steam turbines, to combustion products in a gas turbine engine. These mechanisms may contain hundreds or thousands of turbine blades, and a failure of any one can lead to catastophic failure of the turbine.

Such failures may be caused by poor tolerances in the dimensions of the blades or their casing, pitting or other erosion caused by cavitation or foreign materials in the fluid, corrosion, poor balance and vibration, fatigue and stress, and thermal deformation. In the most demanding applications, such as jet engines, the turbine blades are often the limiting component .[1] To survive in this difficult enviroment, turbine blades often use exotic materials like superalloys and many different methods of cooling, such as internal air channels, boundary layer cooling, and thermal barrier coatings.

Introduction[edit]

Diagram of a two spool jet engine. The high pressure turbine is connected by a single shaft to the high pressure compressor (purple), and the low pressure turbine is connected to the low pressure compressor by a second shaft (green).

In a gas turbine engine, a single turbine section is made up of a disk or hub that holds many turbine blades. That turbine section is connected to a compressor section via a shaft (or "spool"), and that compressor section can either be axial or centrifugal. Air is compressed, raising the pressure and temperature, through the compressor stages of the engine. The pressure and temperature are then greatly increased by combustion of fuel inside the combustor, which sits between the compressor stages and the turbine stages. That high temperature and high pressure fuel then passes through the turbine stages. The turbine stages extract energy from this flow, lowering the pressure and temperature of the air, and transfer that energy to the compressor stages along the shaft. This is process is very similar to how an axial compressor works, only in reverse.[2]

The number of turbine stages varies in different types of engines, with high thrust, high bypass ratio, engines tending to have the most turbine stages.[citation needed] The number of turbine stages can have a great effect on how the turbine blades are designed for each stage. Many gas turbine engines are two shaft designs, meaning that there is a high pressure shaft and a low pressure shaft. Other gas turbines used three shafts, adding an intermediate pressure shaft between the high and low pressure shafts. The high pressure turbine is exposed the the hottest, highest pressure, air, and the low pressure turbine is subjected to cooler, lower pressure air. That difference in conditions leads the design of high pressure and low pressure turbine blades to be significantly different in material and cooling choices even though the aerodynamic and thermodynamic principles are the same.[3]

Enviroment and failure modes[edit]

Turbine blades are subjected to very strenuous enviroments inside a gas turbine. They face high temperatures, high stresses, and a potentially high vibration enviroment. All three of these factors can lead to blade failures, which can destroy the engine, and turbine blades are carefully designed to resist those conditions.[4]

Turbine blades are subjected to stress from centrifugal force (turbine stages can rotate at tens of thousands of revolutions per minute (RPM)) and fluid forces that can cause fracture, yielding, or creep[nb 1] failures. Additionally, the first stage (the stage directly following the combustor) of a modern turbine faces temperatures around 2500 °F (1400 °C),[5] up from temperatures around 1500 °F (800 °C) in early gas turbines.[6] Modern military jet engines, like the Snecma M88, can see turbine temperatures of up to 2900 °F (1600 °C).[7] Those high temperatures weaken the blades and make them more susceptible to creep failures. The high temperatures can also make the blades susceptible to corrosion failures. Finally, vibrations from the engine and the turbine itself (see blade pass frequency) can cause fatigue failures.[4]

Materials[edit]

A key limiting factor in early jet engines was the performance of the materials available for the hot section (combustor and turbine) of the engine. The need for better materials spurred much research in the field of alloys and manufacturing techniques, and that research resulted in a long list of new materials and methods that make modern gas turbines possible.[6]

The development of superalloys in the 1940s and new processing methods such as vacuum induction melting in the 1950s greatly increased the temperature capability of turbine blades. Further processing methods like hot isostatic pressing improved the alloys used for turbine blades and increased turbine blade performance.[6] Modern turbine blades often use nickel-based superalloys that incorporate chromium, cobalt, and rhenium.[4][8]

Aside for alloy improvements, a major breakthrough was the development of directional solidification (DS) and single crystal (SC) production methods. These methods help greatly increase strength against fatigue and creep by aligning grain boundaries in one direction (DS) or by eliminating grain boundaries all together (SC).[6]

A turbine blade with thermal barrier coating.

Another major improvement to turbine blade material technology was the development of thermal barrier coatings (TBC). Where DS and SC developments improved creep and fatigue resistance, TBCs improved corrosion and oxidation resistance, both of which become greater concerns as temperatures increased. The first TBCs, applied in the 1970s, were aluminide coatings. Improved ceramic coatings became available in the 1980s. These coatings increased turbine blade capability by about 200 °F (90 °C).[6] The coatings also improve blade life, almost doubling the life of turbine blades in some cases.[9]

Most turbine blades are manufactured by investment casting (or lost-wax processing). This process involves making a precise negative die of the blade shape that is filled with wax to form the blade shape. If the blade is hollow (i.e., it has internal cooling passages), a ceramic core in the shape of the passage is inserted into the middle. The wax blade is coated with a heat resistant material to make a shell, and then that shell is filled with the blade alloy. This step can be more complicated for DS or SC materials, but the process is similar. If there is a ceramic core in the middle of the blade, it is dissolved in a solution that leaves the blade hollow. The blades are coated with an TBC they will have, and then cooling holes are machined as needed, creating a complete turbine blade.[10]

List of turbine blade materials[edit]

Note: This list is not inclusive of all alloys used in turbine blades.[11][12]

  • U-500 This material was used as a first stage (the most demanding stage) material in the 1960s, and is now used in later, less demanding, stages.[12]
  • Rene 77[12]
  • Rene N5[13]
  • Rene N6[13]
  • PWA1484[13]
  • CMSX-10[13]
  • Inconel
    • IN-738 - GE used IN-738 as a first stage blade material from 1971 until 1984, when it was replaced by GTD-111. It is now used as a second stage material. It was specifically designed for land-based turbines rather than aircraft gas turbines.[12]
  • GTD-111 Blades made from directionally solidified GTD-111 are being using in many GE Aviation gas turbines in the first stage. Blades made from equiaxed GTD-111 are being used in later stages.[12]
  • EPM-102 (MX4 (GE), PWA 1497 (P&W)) is a single crystal superalloy jointly developed by NASA, GE Aviation, and Pratt & Whitney for the High Speed Civil Transport (HCST). While the HCST program was canceled, the alloy is still being considered for use by GE and P&W.[14]

Cooling[edit]

Another strategy to improving turbine blades and increasing their operating temperature, aside from better materials, is to cool the blades. There are three main types of cooling used in gas turbine blades; convection, film, and transpiration cooling. While all three methods have their differences, they all work by using cooler air (often bleed from the compressor) to remove heat from the turbine blades.[15]

Convection cooling works by passing cooling air through passages internal to the blade. Heat is transferred by conduction through the blade, and then by convection into the air flowing inside of the blade. A large internal surface area is desirable for this method, so the cooling paths tend to be serpentine and full of small fins.[15][16]

Rendering of a turbine blade with cooling holes for film cooling.

A variation of convection cooling, impingement cooling, works by hitting the inner surface of the blade with high velocity air. This allows more heat to be transfered by convection that regular convection cooling does. Impingement cooling is often used on certian areas of a turbine blade, like the leading edge, with standard convection cooling used in the rest of the blade.[16]

The second major type of cooling is film cooling (also called thin film cooling). This type of cooling works by pumping cool air out of the blade through small holes in the blade. This air creates a thin layer (the film) of cool air on the surface of the blade, protecting it from the high temperature air. The air holes can be in many different blade locations, but they are most often along the leading edge.[15] A United State Air Force program in the early 1970s funded the development of a turbine blade that was both film and convection cooled, and that method has become common in modern turbine blades.[6]

One consideration with film cooling is that injecting the cooler bleed into the flow reduces turbine efficiency. That drop in efficiency also increases as the amount of cooling flow increases. The drop in efficiency, however, is usually mitigated by the increase in overall performance produced by the higher turbine temperature.[17]

Transpiration cooling, the third major type of cooling, is similar to film cooling in that it creates a thin film of cooling air on the blade, but it is different in that that air is "leaked" through a porous shell rather than injected through holes. This type of cooling is effective at high temperatures as it uniformly covers the entire blade with cool air.[16][18] Transpiration-cooled blades generally consist of a rigid strut with a porous shell. Air flows through internal channels of the strut and then passes through the porous shell to cool the blade.[19] As with film cooling, increased cooling air decreases turbine efficiency, so that decrease has to be balanced with improved temperature performance.[17]

See also[edit]

Notes[edit]

  1. ^ Creep is the tendency of a solid material to slowly move or deform permanently under the influence of stresses. It occurs as a result of long term exposure to high levels of stress that are below the yield strength of the material. Creep is more severe in materials that are subjected to heat for long periods, and near the melting point. Creep always increases with temperature. From Creep (deformation).

References[edit]

  1. ^ Boyce, p. 368.
  2. ^ Flack, p. 406
  3. ^ Flack, p. 407
  4. ^ a b c Flack, p. 429.
  5. ^ Flack, p. 410
  6. ^ a b c d e f Koff, Bernard L. (2003). "Gas Turbine Technology Overview - A Designer's Perspective". AIAA/ICAS International Air and Space Symposium and Exposition: The Next 100 Years. 14–17 July 2003, Dayton, Ohio. AIAA 2003-2722.
  7. ^ Dexclaux, Jacques and Serre, Jacque (2003). "M88-2 E4: Advanced New Generation Engine for Rafale Multirole Fighter". AIAA/ICAS International Air and Space Symposium and Exposition: The Next 100 Years. 14–17 July 2003, Dayton, Ohio. AIAA 2003-2610
  8. ^ Magyar, Michael J. "Mineral Yearbook: Rhenium" (PDF). United States Geological Survey.
  9. ^ Boyce, p. 449
  10. ^ Flack, p. 430-3
  11. ^ Boyce, p. 440-2
  12. ^ a b c d e Schilke, P. W. (2004). Advanced Gas Turbine Materials and Coatings. GE Energy. August 2004. Retrieved: 16 June 2010.
  13. ^ a b c d MacKay, Rebecca A., et. al. (2007). Low-Density, Creep-Resistant Superalloys Developed for Turbine Blades. NASA Glenn's Research & Technology. Updated: 7 November 2007. Retrieved: 16 June 2010.
  14. ^ S. Walston, A. Cetel, R. MacKay, K. O’Hara, D. Duhl, and R. Dreshfield (2004). Joint Development of a Fourth Generation Single Crystal Superalloy. NASA TM—2004-213062. December 2004. Retrieved: 16 June 2010.
  15. ^ a b c Flack, p.428.
  16. ^ a b c Boyce, p. 370.
  17. ^ a b Boyce, p. 379-80
  18. ^ Flack, p. 428-9
  19. ^ Boyce, p. 375
Bibliography
  • Flack, Ronald D. (2005). "Chapter 8: Axial Flow Turbines". Fundamentals of Jet Propulsion with Applications. Cambridge Aerospace Series. New York, NY: Cambridge University Press. ISBN 9780521819831.
  • Boyce, Meherwan P. (2006). "Chapter 9: Axial Flow Turbines and Chapter 11: Materials". Gas Turbine Engineering Handbook (3rd ed.). Oxford: Elsevier. ISBN 9780750678469.
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