Fluid reaction surfaces (i.e. – impellers) – With heating – cooling or thermal insulation means – Changing state mass within or fluid flow through working...
Reexamination Certificate
1999-02-12
2001-03-27
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...
C416S09600A, C415S115000
Reexamination Certificate
active
06206638
ABSTRACT:
BACKGROUND OF THE INVENTION
1. Field of the Invention
The present invention relates generally to cooling air circuits of turbine rotor blades and stator vanes in gas turbine engines and, more specifically, to serpentine cooling circuits feeding cooling air to sidewall impingement cooling chambers and which blades are castable with a single pull core.
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. The temperatures within gas turbines may exceed 2500 degrees Fahrenheit, and cooling of turbine blades is very important in terms of blade longevity. Without cooling, turbine blades would rapidly deteriorate. Improved cooling for turbine blades is very desirable, and much effort has been devoted by those skilled in the blade cooling arts to devise improved geometries for the internal cavities within turbine blades in order to enhance cooling. 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 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.
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 and suction sides. 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.
U.S. Pat. No. 5,660,524, entitled “Airfoil Blade Having A Serpentine Cooling Circuit And Impingement Cooling”, discloses an airfoil blade, such as a jet engine turbine rotor blade with an internal serpentine coolant circuit that has a last downstream passageway bounded by four monolithic inner walls which are monolithic with at least a portion of the outer walls. Two of the inner walls are spaced from the outer walls and contain air impingement orifices creating two film cooling chambers. Some coolant in the serpentine circuit exits the airfoil blade through a coolant exit in the blade tip. The remaining coolant in the circuit passes through the impingement orifices and exits the blade through film cooling holes in the outer walls.
U.S. Pat. No. 5,813,836, entitled “Turbine Blade”, discloses an airfoil having a double-wall construction for side-wall impingement cooling on the pressure side and a forward flowing multi-pass serpentine cooling air circuit along the suction side of the blade which flows cooling air forward with respect to the aft flowing hot gases through the turbine. The airfoil also includes a leading edge cavity having a plurality of radial film cooling holes supplied by the three pass serpentine cooling circuit. As cooling air flows along the passageways, it convectively cools the portions of the turbine blade adjacent these passageways. The airfoil further includes a trailing edge cavity to cool the trailing edge flow region of the airfoil. A plurality of impingement cavities are located on the pressure sidewall section and impingement holes provide cooling air from the serpentine passageways of the inner cavity and the impingement cavities. Multi-row, compound angle film holes extend from the impingement cavities so that cooling air from the impingement cavities can be discharged from the airfoil.
Known turbine airfoil cooling techniques include the use of internal passages forming a serpentine cooling circuit. Particularly, serpentine passages, leading edge impingement bridges, film holes, pin fins, and trailing edge holes or pressure side bleed slots are utilized for blade cooling. It would be desirable to provide improved blade cooling. In providing even better blade cooling, it is also desirable to avoid significantly increasing the blade fabrication costs. Casting of the blades involves the use of a ceramic core around which the blade is cast. This core is then leached out leaving behind an air flow passage and the internal surface configuration. The core itself is formed by injecting slurry into a mold formed by dies. The dies must then be opened to obtain the core. Some shapes require multiple pull dies because of the complex configuration. It is preferable to minimize the number of dies used and with as few pulls as possible. It is highly desirable to have only two dies surrounding the airfoil with a single pull require to release the core. It is also highly desirable to have a single piece core to improve the quality and accuracy of the casting process and internal passages and cavities of the hollow airfoil.
SUMMARY OF THE INVENTION
A gas turbine engine airfoil for a vane or a blade includes an airfoil outer wall having widthwise spaced apart pressure and suction sidewall sections extending chordally between leading and trailing edges of the airfoil and extending longitudinally from a base to a tip. inside the airfoil is at least one internal cooling circuit having a plurality of longitudinally extending circuit channels between longitudinally extending internal ribs extending widthwise between the pressure and suction sidewall sections and a longitudinally extending first sidewall film cooling chamber positioned between one of the sidewall sections and a first inner wall bounding the cooling circuit. A first plurality of sidewall film cooling holes extend through the pressure sidewall section from the first sidewall film cooling chamber. The internal ribs have corresponding rib angles with respect to a centerline, the first inner wall has a corresponding first wall angle with respect to the centerline, and each of the rib angles and the first wall angle are constant in a longitudinal direction from the base to the tip.
A first exemplary embodiment includes a plurality of first impingement cooling apertures extending from at least a first one of the circuit channels through the first inner wall to the internal cooling circuit. In the preferred embodiment, all the
Carlson Clay K.
Glynn Christopher C.
Martus James A.
Wilson Paul S.
General Electric Company
Hess Andrew C.
Look Edward K.
Woo Richard
Young Rodney M.
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