Patent Application
20090238694
(1) Field of the Invention
The present invention relates to a turbine engine component having an improved scheme for cooling an airfoil portion.
(2) Prior Art
The overall cooling effectiveness is a measure used to determine the cooling characteristics of a particular design. The ideal non-achievable goal is unity, which implies that the metal temperature is the same as the coolant temperature inside an airfoil. The opposite can also occur when the cooling effectiveness is zero implying that the metal temperature is the same as the gas temperature. In that case, the blade material will certainly melt and burn away. In general, existing cooling technology allows the cooling effectiveness to be between 0.5 and 0.6. More advanced technology such as supercooling should be between 0.6 and 0.7. Microcircuit cooling as the most advanced cooling technology in existence today can be made to produce cooling effectiveness higher than 0.7.
The Table I below provides the operational parameters used to plot the design point in the durability map.
It should be noted that the overall cooling effectiveness from the table is 0.717 for a film effectiveness of 0.296 and a convective efficiency (or ability to pick-up heat) of 0.573. Also note that the corresponding cooling flow for a turbine blade having this cooling microcircuit is 3.5% engine flow.
There are however field problems that can be addressed efficiently with peripheral microcircuit designs. One such field problem is illustrated in
In accordance with the present invention, a turbine engine component is provided with improved cooling. The turbine engine component broadly comprises an airfoil portion having an airfoil mean line, a pressure side, and a suction side, a first region on the pressure side having a first array of cooling microcircuits embedded in a wall forming the pressure side, a second region on the pressure side having a second array of cooling microcircuits embedded in the wall, and the first region being located on a first side of the mean line and the second region being located on a second side of the mean line.
Other details of the radial split serpentine microcircuits of the present invention, as well as other objects and advantages attendant thereto, are set forth in the following detailed description and the accompanying drawings wherein like reference numerals depict like elements.
The present invention solves several problems associated with the use of serpentine microcircuits in airfoil portions of turbine engine components such as turbine blades. For example, it has been discovered that the heat transfer for a channel used in a peripheral serpentine cooling microcircuit is much superior if the inlet to the channel is at a 90 degree angle with respect to the direction of flow within the channel. When using such an inlet, it is desirable to place the inlet closer to any distress regions wherever possible to address regions requiring enhanced heat transfer. It has also been discovered that it is advantageous to radially place two microcircuit panels with two 90 degree turn inlets instead of using just one panel with a straight inlet. The duplication of the two circuits disposed radially provide large increases in heat transfer when compared with the same region covered by a panel with a straight inlet.
One area of concern regarding traditional microcircuit cooling is the inability to form the microcircuit within positional tolerance embedded in the airfoil walls. It is therefore desirable to take advantage of placement of microcircuits in the airfoil wall to (1) eliminate areas of known distress; (2) alleviate microcircuit positional problems during forming and subsequent casting of the airfoil; and (3) take advantage of pumping (rotational forces) necessary to lead the flow through the microcircuit peripheral cooling solutions.
Referring now to
The airfoil portion 102 has a pressure side 130 and a suction side 132. Embedded within the wall forming the pressure side 130 are a series of peripheral microcircuits in two regions 134 and 136. The region 134 is located above the airfoil mean line 138 at 50% span, while the region 136 is located below the airfoil mean line 138. Within the region 134, there is located a first fluid passageway 140 having a fluid inlet 142 which communicates with one of the feed holes 112. The fluid inlet 142 has a 90 degree bend. Fluid from the passageway 140 flows into a passageway 144 where the fluid proceeds around the tip of the airfoil portion 102, goes around the leading edge 120 via passageway 158 and discharges on the airfoil suction side 132 via outlet (s) 160.
Within the region 134, there is located a fluid inlet 146 which communicates with one of the feed inlets 112 from the leading edge internal circuit 108. The fluid inlet 146 has a 90 degree bend. Fluid from the inlet 146 is supplied to a first fluid passageway 148 and to a second fluid passageway 152. Each of the fluid passageways 148 and 152 has a plurality of film holes 150 for supplying film cooling over the pressure side 130 of the airfoil portion 102.
Further, within the region 134, there is a located a fluid inlet 154. The fluid inlet 154 has a 90 degree bend. The fluid inlet 154 supplies cooling fluid to a fluid passageway 156 so that the cooling fluid flows in a direction perpendicular to the fluid inlet 154. The fluid passageway communicates with a fluid passageway 158 which wraps around the leading edge 120 of the airfoil portion 102. The fluid passageway 158 has one or more outlets 160 for allowing cooling fluid to flow over the suction side 132 of the airfoil portion 102.
Within the region 136, there is located a fluid passageway 162 and a fluid passageway 164. Each of the fluid passageways 162 and 164 receives fluid from an inlet 166 which communicates with one of the inlets 112 in the trailing edge internal circuit 110. The inlet 166 has a 90 degree bend. The fluid passageway 164 has a plurality of film cooling holes 168 for allowing cooling fluid to flow over the pressure side 130. The fluid passageway 162 has a plurality of exit holes 170 for allowing cooling fluid to flow over the trailing edge 126 of the airfoil portion 102.
Also within the region 136, there is a fluid passageway 172 which communicates with a fluid passageway 174 at a right angle to the passageway 172 and a further fluid passageway 176 at a right angle to the fluid passageway 174. The fluid passageway 176 has a plurality of film cooling holes 178 for allowing cooling fluid to flow over the pressure side 130 of the airfoil portion 102. The fluid passageway 172 communicates with an inlet 180 which has a 90 degree bend. The inlet 180 communicates with one of the feed holes 112 in the trailing edge internal circuit 110.
One advantage of the present invention is that the feeds from the inlets 142, 166, and 180 are radially split to increase internal heat transfer. Further, a plurality of ties 182 may be provided to maintain positional tolerance of the cooling microcircuits with the airfoil wall. Still further, each of the inlets 142, 146, 152, 166, and 180 has a 90 degree turn for supplying cooling fluid to each respective cooling microcircuit. The cooling of the leading and trailing edges 120 and 126 of the airfoil portion 102 protects them from external thermal load by the embedded wall microcircuits. It should also be noted that the peripheral microcircuits are tied together around the airfoil portion 102 to facilitate forming onto the airfoil wall; thus improving castability of the part in subsequent casting processes.
It is apparent that there has been provided in accordance with the present invention radial split serpentine microcircuits which fully satisfy the objects, means, and advantages set forth hereinbefore. While the present invention has been described in the context of specific embodiments thereof, other unforeseeable alternatives, modifications, and variations may become apparent to those skilled in the art having read the foregoing description. Accordingly, it is intended to embrace those alternatives, modifications, and variations as fall within the broad scope of the appended claims.
