Patent Application
20100054914
This application relates to a gas turbine engine component wherein an internal cooling air passage is divided into two separate paths, and wherein the two separate paths are formed from a single core in a lost core molding process.
Gas turbine engines are known and typically include a compressor section compressing air and delivering it into a combustion section. The air is mixed with fuel in the combustion section and ignited. Products of the combustion pass downstream over turbine rotors, driving the turbine rotors.
There is a good deal of design that goes into the structure of the turbine rotors, and a number of components that are utilized to control the flow of the products of combustion such that they are directed along desired flow paths. One such component is called a blade outer air seal. A blade outer air seal sits slightly radially outwardly of an outer tip of a turbine blade in a turbine rotor, which is driven to rotate by the products of combustion. By having the blade outer air seal closely spaced from the rotor, leakage of the products of combustion around the turbine rotor is reduced.
The blade outer air seals are subject to very high temperature. Thus, it is known to provide cooling air through the blade outer air seal.
Cooling air from a source of air cooler than the product of combustion is circulated through channels in the blade outer air seal. Recently, these channels have become thinner in a radial dimension. It is known that as the channels become thinner relative to an axial width of the channel, the flow characteristics of the cooling air may degrade. That is, when an aspect ratio of a circumferentially-flowing channel (where the aspect ratio is the radial dimension divided by the axial dimension), is relatively high, then there is good circulation of air and desirable heat transfer characteristics. On the other hand, as the aspect ratio drops, which occurs as the (radial) height of the channel becomes smaller, the heat transfer effectiveness may decrease and/or friction losses may increase. Having a thinner radial dimension is desirable to enable higher cooling effectiveness for the same amount of air flow, or achieving the same cooling effectiveness with reduced air flow. The usage of bleed air for cooling parts rather than producing thrust causes a reduction in turbine efficiency.
Thus, it is known in the prior art to form two separate channels where there was one when there is a relatively radially thin cooling air passage.
Such components are typically formed by lost core molding processes. In a lost core molding process, a core is created for all hollow spaces that are to be formed in the blade outer air seal. Thus, a core would be formed to form the cooling air passages within the blade outer air seal. The core is inserted into a mold, and metal is molded around the core. The core may then be leached away leaving a hollow within the blade outer air seal. The prior art solution of providing two separate channels requires two separate cores, and is thus somewhat undesirable.
In addition, with two separate cores there must be two separate inlet and exit holes. The use of two separate inlet and exit holes can result in a reduced total cross-sectional area due to the two allowable tolerances. With the reduced cross-sectional areas, frictional losses can increase. The frictional losses associated with each hole can add undesirably large pressure drops, especially when the radial height is small and there are significant frictional losses along the passage itself.
Also, existing gas turbine engines already have locations for the inlets and the exit holes that must be maintained. Thus, it would not always be possible to add additional inlet and exit holes.
In a disclosed embodiment of this invention, a gas turbine engine component has a cooling air passage with two distinct flow paths formed by a single core, a single inlet hole, and a single exit hole. The invention also extends to a core and method for forming the component.
In addition, an improved method and an inventive core are also claimed.
These and other features of the present invention can be best understood from the following specification and drawings, the following of which is a brief description.
A gas turbine engine rotor blade 20 is illustrated in
Since the slot 44 extends through the entire width, then the dividing wall 30 will extend entirely between upper and lower walls of the passage 26. This can be appreciated from
The core 40 provides both passages 27 of the channels 26.
The use of the single core to form both passages 27 results in maintaining a single inlet and a single exit hole. Thus, the problem mentioned above of increased frictional losses will not occur. In addition, the method allows the redesign of existing components to achieve smaller radial cross-sections while at the same time maintaining the location of inlet and exit holes.
Although an embodiment of this invention has been disclosed, a worker of ordinary skill in this art would recognize that certain modifications would come within the scope of this invention. For that reason, the following claims should be studied to determine the true scope and content of this invention.
This invention was made with government support under Contract No. N00019-02-C-3003 awarded by the United States Navy. The Government may therefore have certain rights in this invention.
