Fluid reaction surfaces (i.e. – impellers) – With heating – cooling or thermal insulation means – Changing state mass within or fluid flow through working...
Reexamination Certificate
2001-02-23
2002-12-10
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
06491496
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 metering plates at the base of roots used to meter flow to the cooling circuits within airfoils of the blades and vanes.
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 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 cooling air from 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. In pursuit of higher cooling effectiveness, modem blades have led to multi-pass cooling circuits.
It is also known to pass the cooling air through serpentine cooling air circuits and other passages in the interior of the blade which warms up the cooling air as it travels through the passages before being impinged on the leading edge of the blade. The temperature difference across the leading edge is lower than directing cooling air through the root of the blade for impingement resulting in lower thermal stresses in the blade leading edge and the life of the blade is enhanced. This makes efficient use of cooling flow since the flow is able to internally cool the blade over much of the blade mid-span before flowing out radial leading edge cooling holes to film cool the blade airfoil externally.
However, this technique also adversely effects “backflow margin”. As air flow travels through the internal passages of the blade, pressure losses due to turns and turbulence promoters cause the cooling flow pressure to drop to a level, such that under certain operating conditions, hot gas ingestion into the blade leading edge may occur. This undesired condition is referred to as backflow. One approach for providing more backflow margin is to increase the inlet pressure of the cooling air which is supplied to the blade. This approach is not always feasible because the increase in supply pressure can increase cooling flow leakages to an undesired level. To overcome this problem, leading edge flow passages supplying cooling air for impingement were supplied with refresher passageways. The refresher passageway is connected to a portion of a first channel of the cooling circuit in the root of the blade. This circuit supplies impingement cooling air to the leading edge cavity from its last channel referred to herein as a leading edge supply channel. The refresher passageway refreshes the airflow in this last channel after the airflow has flowed through the rest of the circuit and has become warmed. See U.S. Pat. Nos. 5,387,086 and 5,813,826.
The refresher passageway to the leading edge supply channel is connected to the serpentine cooling circuit inlet channel which passes through the root of the blade and, thus, is coupled to the flow rate through the serpentine cooling circuit. It is desirable to be able to tune the cooling flow through the refresher passageway independent of the flow through the serpentine cooling circuit in order to adjust the flow and pressure to the leading edge supply channel. This would be particularly useful if a blade casting is produced which has higher than desired pressure drops through the serpentine passages or impingement holes. This is also desirable for adjustments to the blade cooling circuit that are useful or necessary due to blade hole and serpentine circuit deterioration such as may be caused by blockages and wear.
Known turbine airfoil cooling techniques include the use of internal cavities forming a serpentine cooling circuit. Particularly, serpentine passages, leading edge impingement bridges, turbulence promoters and turbulators, film holes, pin fins, and trailing edge holes or pressure side bleed slots are utilized for blade cooling. It is desirable to provide improved blade cooling. In providing even better blade cooling, it is also desirable to avoid significantly increasing the blade fabrication costs.
SUMMARY OF THE INVENTION
A gas turbine engine hollow airfoil has an airfoil outer wall with widthwise spaced apart pressure and suction side walls joined together at chordally spaced apart leading and trailing edges of the airfoil and extending radially from a radially inner base to a radially outer airfoil tip. A cooling circuit within the airfoil has radially extending first, middle, and last channels arranged respectively in series with the first channel in fluid communication with a source of cooling air from outside the airfoil. The last channel is in fluid communication with one of the edges. A refresher passageway extends through a radially inner wall bounding a radially inner portion of the last channel and is in fluid communication with the source of cooling air. The refresher passageway is separate, spaced apart, and independent from the first channel.
One exemplary embodiment of the invention further includes an edge cooling plenum located between the last channel and one of the edges and cooling air discharge apertures disposed through a radially extending rib between the last channel and the edge cooling plenum. The edge cooling plenum may be a leading edge cooling plenum and the cooling air discharge apertures may be impingement cooling holes and may also include leading edge cooling holes leading out of the edge cooling plenum through the outer wall around the leading edge. In another exemplary embodiment, the last channel is bounded by the trailing edge and cooling air discharge apertures are disposed through the trailing edge and may be trailing edge cooling slots.
Exemplary embodiments of the invention may further include a metering plate covering an inlet to the refresher passageway and the metering plate has a metering hole over the inlet and the metering hole is adjustable. Another exemplary embodiment of the invention is a gas turbine engine blade with the hollow airfoil extending radially outwardly from a root. The first channel extends through the root and has an entrance at a bottom surface of the root. The refresher passageway extends through a radially inner wall bounding a radially inner portion of the last channel and through the root. The inlet to the refresher passageway is located at the bottom surface of the root and is separate
General Electric Company
Look Edward K.
Nguyen Ninh
Rosen Steven J.
Young Rodney M.
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