Fluid reaction surfaces (i.e. – impellers) – With heating – cooling or thermal insulation means – Changing state mass within or fluid flow through working...
Reexamination Certificate
1999-09-30
2001-09-18
Verdier, Christopher (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
06290463
ABSTRACT:
FIELD OF THE INVENTION
The present invention relates generally to cooling of turbine rotor blades and stator vanes in gas turbine engine turbines and, more specifically, to impingement cooling of leading edges of airfoils in turbine rotor blades and stator vanes.
DESCRIPTION OF RELATED 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. It is highly desirable 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 or other types of cooling circuits 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.
High pressure turbine blades typically have very high heat loads at the leading edges. In order to cool this leading edge, an impingement cooling technique is often used in the first stage high pressure turbine blade. The impingement cooling is accomplished by directing the cooling air through a row of crossover holes in a wall between a leading edge cavity and a cavity or passage of the cooling circuit. The cooling air is then discharged through shower head holes in the leading edge to provide film cooling on an exterior surface of the leading edge of the airfoil.
Prior art crossover hole configurations are typically circular, ellipse, or race track in cross-section. The crossover holes are typically cast with the entire blade. During a casting process, a parting line between two core die halves is located where a middle of the crossover holes is located to allow the core die halves to be pulled apart in both concave and convex directions. A shift in the core die will result in a mismatch in crossover hole portion of the two halves because the parting line is located where the middle of the crossover holes are to be located. This then requires hand rework on the ceramic core or scrapping of the core. Rework contributes to variations in hole sizes which in turn results in flow variations. Natural core die wear also results in excess core material on the crossover holes requiring additional hand work of cores and increasing the chance of flow variation. Discrete impinging jets through the crossover holes result in local cool spots at the stagnation point of each jet.
Heat transfer coefficients on surfaces between jet stagnation points are less than the heat transfer coefficients at the stagnation points which causes undesirable non-uniform heat transfer distribution. Crossover hole misalignment leads to an even more undesirable and more non-uniform heat transfer distribution. Another problem common to the crossover holes is cracks around the edge of the crossover holes due to the stress concentration created by the discrete holes and the large thermal gradient between blade airfoil surface temperature and the wall in which the crossover holes are formed. Therefore, it is desirable to have an impingement design that requires less or no rework on the impingement holes and/or the core portions for the holes. It is also desirable to have an impingement design with improved heat transfer coefficient distribution and that reduces thermal stress on the wall in which the holes are formed.
SUMMARY OF THE INVENTION
A coolable gas turbine engine airfoil includes an outer airfoil wall with pressure and suction sides extending chordwise between leading and trailing edges of the airfoil, a leading edge cooling plenum formed between a forwardmost span rib and the outer wall along the leading edge of the airfoil, and a cooling air channel within the airfoil bounded in part by the forward most rib. A narrow slotted cooling air impingement means disposed in the span rib for impinging cooling air from the channel on an interior surface of the outer airfoil wall along the leading edge of the airfoil. The narrow slotted cooling air impingement means includes at least one longitudinally extending narrow slot having a slot length that is at least about an order of magnitude larger than a slot width of the narrow slot and the slot width is constant along the slot length. One embodiment of the slotted cooling air impingement means is a single longitudinally extending slot extending along almost an entire length of the forwardmost rib and the longitudinally slot preferably includes longitudinally spaced apart rounded ends. In another embodiment of the airfoil, the slotted cooling air impingement means includes two or more closely spaced apart longitudinally extending slots extending along almost an entire length of the forwardmost rib and each of the longitudinally extending slots preferably has longitudinally spaced apart rounded ends. Yet another embodiment of the invention provides a coolable gas turbine engine blade having the coolable airfoil extending longitudinally outwardly from a platform of the blade to an outer airfoil tip and a root extending longitudinally inwardly from the platform.
Since the number of slots is much less than the number of crossover holes that are typically used, the mismatched area caused by the shift between two core dies used during manufacture of the airfoils is minimized and the need for rework is greatly reduced. Excess core material at parting lines along centers of the crossover holes due to die wear is also reduced by the present invention. This helps reduce variation in impingement flow. The present invention produces a line jet through the slots which in turn results in a more uniform heat transfer distribution in the radial direction as opposed to point jets through the conventional discrete crossover holes. The line jet will also reduce the cross flow effects which exist in the discrete jets. The invention reduces stress created by the thermal gradient between the outer surface and the slot wall because the slot design provides a separation between the slot walls. The impingement slots are more advantageous than crossover holes from a manufacturing standpoint and heat transfer performance and thermal stress standpoint.
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Fargher David G. W.
Lee Ching-Pang
General Electric Company
Hess Andrew C.
Verdier Christopher
Young Rodney M.
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