Fluid reaction surfaces (i.e. – impellers) – With heating – cooling or thermal insulation means – Changing state mass within or fluid flow through working...
Reexamination Certificate
1997-11-17
2001-04-24
Look, Edward K. (Department: 3745)
Fluid reaction surfaces (i.e., impellers)
With heating, cooling or thermal insulation means
Changing state mass within or fluid flow through working...
C415S115000
Reexamination Certificate
active
06220817
ABSTRACT:
BACKGROUND OF THE INVENTION
1. Field of the Invention
The present invention relates generally to cooling of turbine rotor blades and stator vanes in gas turbine engines and, more specifically, to serpentine cooling circuits therein.
2. Discussion of the Background Art
A gas turbine engine includes a compressor that compresses air which is channeled to a combustor wherein it is mixed with fuel and ignited for generating combustion gases. The combustion gases flow downstream through one or more stages of turbines which extract energy therefrom for powering the compressor and producing additional output power for driving a fan for powering an aircraft in flight for example. A turbine stage includes a row of turbine rotor blades secured to the outer perimeter of a rotor disk, with a stationary turbine nozzle having a plurality of stator vanes disposed upstream therefrom. The combustion gases flow between the stator vanes and between the turbine blades for extracting energy to rotate the rotor disk. Since the combustion gases are hot, the turbine vanes and blades are typically cooled with a portion of compressor air bled from the compressor for this purpose. Diverting any portion of the compressor air from use in the combustor necessarily decreases the overall efficiency of the engine. Accordingly, it is desired to cool the vanes and blades with as little compressor bleed air as possible.
Typical turbine vanes and blades include an airfoil over which the combustion gases flow. The airfoil typically includes one or more serpentine cooling passages therein through which the compressor bleed air is channeled for cooling the airfoil. The airfoil may include various turbulators therein for enhancing cooling effectiveness and, the cooling air is discharged from the passages through various film cooling holes disposed around the outer surface of the airfoil.
The temperature profile of the combustion gases channeled over the airfoil is typically center peaked at about 50% to about 80% of the radial height or span of the airfoil. Secondary flow fields between adjacent airfoils may sometimes cause the temperature profile of the combustion gases to shift radially outwardly on the pressure side of the airfoil. Accordingly, the airfoil typically experiences relatively high heat input loading on its pressure side above the airfoil mid-span. Since the serpentine cooling circuits introduce air into the airfoil from its root, the cooling air must be provided with a suitable flow rate to ensure that the outer portions of the airfoil experiencing the greatest heat input are adequately cooled for obtaining a useful life during operation. The inner portions of the airfoil may therefore be over-cooled which is an inefficient use of the valuable compressor bleed air. To overcome this drawback a “Multi-tier turbine airfoil”, disclosed in U.S. Pat. No. 5,591,007, was devised and is incorporated herein by reference. This patent discloses a turbine airfoil having a plurality of internal ribs defining at least two independent serpentine cooling circuits arranged in part in different longitudinal tiers, with an outer tier circuit being disposed in part longitudinally above an inner tier circuit for differentially longitudinally cooling the airfoil. More advanced turbine airfoil designs have been developed that could use a better cooling air distribution.
Typical mid-circuit cooling air, after picking up the heat in the serpentine passage, exits through film cooling holes. One or more rows of film cooling holes are placed on the pressure side and also on the suction side. New highly aerodynamically efficient airfoils in low through flow turbine designs are subject to an external gas path flow along the pressure side that has low velocity. This can result in a very high blowing ratio (mass flux ratio of film cooling air to gas flow) through the film cooling holes and very poor film cooling effectiveness (film blow-off) on the pressure side of the airfoil. Geometrical limitations of at least some of the cavities which supply the film cooling air prevent or make difficult the use of film holes on both pressure and suction sides that have relatively shallow angles from the surfaces of the sides. The use of larger angles would result in significant aerodynamic mixing losses and poor film cooling effectiveness because much of the film cooling air would flow out of the boundary layer. Therefore, it is desirable to have a circuit design which can avoid the use of film cooling in such areas of the airfoil and provide effective and efficient film and convective cooling of the entire airfoil.
SUMMARY OF THE INVENTION
A turbine airfoil includes a plurality of internal ribs defining at least two independent serpentine cooling circuits having outer and inner serpentine portions, respectively, in different longitudinal tiers with the outer serpentine position being disposed longitudinally above the inner tier serpentine position for differentially longitudinally cooling the airfoil. The outer and inner serpentine portions include outer and inner exits and entrances wherein the outer and inner exits are positioned aft of the outer and inner entrances, respectively, so as to have a chordal flow direction aftwards from the leading edge to the trailing edge within the serpentine portions.
The airfoil may include film cooling holes in an outer wall of the airfoil on the suction side of the airfoil and no film cooling holes on a pressure side of the outer wall along a mid-chord portion of the airfoil between the leading and trailing edges.
ADVANTAGES OF THE INVENTION
The present invention provides advantages that include a significant improvement in the cooling of not only, an upper span portion of the turbine airfoil outer wall, but also of a mid-chord portion of the suction and pressure sides of the outer wall. Furthermore, the use of separate leading edge and mid-circuits provides colder cooling air at the upper span portions of the airfoils.
The highly curved or arched contour of the airfoil has span ribs between span channels or cavities nearer to the leading edge that are wider than span ribs between span channels nearer the .trailing edge and, therefore, are also on the average further away from the external hot surfaces of the sides of the outer wall and generally have temperatures closer to the cooling air temperature in the channels. In a downstream wise serpentine circuit design as in the present invention, the cooling air temperature is colder than the cooling air temperature in the same cavities for upstream wise serpentine circuit design. Therefore, a downstreamwise serpentine circuit will have a colder average spanwise rib wall temperature than that of an upstreamwise serpentine circuit and, therefore, have an overall better cooling air temperature distribution in the chordwise direction and a better bulk temperature of the airfoil for better cooling of the entire airfoil.
The two tier circuit design offers additional flexibilities in distributing cooling air in a more efficient manner and also shortens the length of each pass and increases the number of turns which result in a higher heat transfer (cooling) coefficients inside the serpentine passages. The downstreamwise serpentine circuit design also provides an internal cooling air pressure which is more consistent with and tailored to the external gas pressure as the external gas expands in the chordwise or downstream direction through the turbine. This results in a better back flow margin for the blade and a more optimum use of internal cooling potential by trading more pressure consumption for better heat transfer.
Because the outer wall sides closer to the leading edge are cooled by colder fresher air than in those in the prior art, film cooling in this region may not be necessary. This will result in better turbine performance and lower cost in manufacturing. In addition, the film cooling holes closer to the trailing edge can have shallower flow angles from surface than those closer to the leading edge resulting in a better film cooling effectiveness. The ext
Durgin George A.
Lee Ching-Pang
General Electric Company
Hess Andrew C.
Look Edward K.
Woo Richard
Young Rodney M.
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