Fluid reaction surfaces (i.e. – impellers) – With heating – cooling or thermal insulation means – Changing state mass within or fluid flow through working...
Reexamination Certificate
2001-03-27
2002-10-08
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...
C416S24100B, C029S889100
Reexamination Certificate
active
06461107
ABSTRACT:
FIELD OF THE INVENTION
This invention relates generally to gas turbine engines, and in particular, to a process for cooling a flow path surface region on a turbine airfoil.
BACKGROUND OF THE INVENTION
In gas turbine engines, for example, aircraft engines, air is drawn into the front of the engine, compressed by a shaft-mounted rotary-type compressor, and mixed with fuel. The mixture is burned, and the hot exhaust gases are passed through a turbine mounted on a shaft. The flow of gas turns the turbine, which turns the shaft and drives the compressor and fan. The hot exhaust gases flow from the back of the engine, driving it and the aircraft forward.
During operation of gas turbine engines, the temperatures of combustion gases may exceed 3,000° F., considerably higher than the melting temperatures of the metal parts of the engine aft of the compressor, which are in contact with these hot gases. Operation of these engines at gas temperatures that are above the metal part melting temperatures is a well established art, and depends in part on supplying a cooling fluid to the outer surfaces of the metal parts through various methods. Metal parts of these engines that are particularly subject to high temperatures, and thus require particular attention with respect to cooling, are, for example, combustor liners and the metal parts located aft of the combustor including high pressure turbine airfoils, such as turbine blades and turbine vanes.
The hotter the turbine inlet gases, the more efficient is the operation of the jet engine. There is thus an incentive to raise the turbine inlet gas temperature. However, the maximum temperature of the turbine inlet gases is normally limited by the materials used to fabricate the components downstream of the combustors such as the vanes and the blades of the turbine. In current engines, the turbine vanes and blades are made of nickel-based superalloys, and can operate at temperatures of up to 2100°-2200° F. with appropriate well-known cooling techniques.
The metal temperatures can be maintained below their melting levels with current cooling techniques by using a combination of improved active cooling designs and thermal barrier coatings (TBCs). For example, with regard to the metal blades and vanes employed in aircraft engines, some cooling is achieved through convection by providing passages for flow of cooling air from the compressor internally within the blades so that heat may be removed from the metal structure of the blade by the cooling air. Such blades have intricate serpentine passageways within the structural metal forming the cooling circuits of the blade.
Small internal orifices have also been devised to direct this circulating cooling air directly against certain inner surfaces of the airfoil to obtain cooling of the inner surface by impingement of the cooling air against the surface, a process known as impingement cooling. In addition, an array of small holes extending from a hollow core through the blade shell can provide for bleeding cooling air through the blade shell to the outer surface where a film of such air can protect the blade from direct contact with the hot gases passing through the engines, a process known as film cooling.
In another approach, a TBC is applied to the turbine blade component, which forms an interface between the metallic component and the hot gases of combustion. The TBC includes a ceramic coating that is applied to the external surface of metal parts to impede the transfer of heat from hot combustion gases to the metal parts, thus insulating the component from the hot combustion gas. This permits the combustion gas to be hotter than would otherwise be possible with the particular material and fabrication process of the component.
TBCs include well-known ceramic materials, such as, for example, yttrium-stabilized zirconia (YSZ). Ceramic TBCs usually do not adhere well directly to the superalloys used as substrate materials. Therefore, an additional metallic layer called a bond coat is placed between the substrate and the TBC. The bond coat may be made of an overlay alloy, such as a MCrAlX, or other composition more resistant to environmental damage than the substrate, or alternatively, the bond coat may be a diffusion nickel aluminide or platinum aluminide. The surface of the bond coat oxidizes to form a thin, protective aluminum oxide scale that provides improved adherence to the ceramic top coatings. The bond coat and overlying TBC are frequently referred to as a thermal barrier coating system.
Multi-layer coatings are well known in the art. For example, U.S. Pat. No. 5,846,605 to Rickerby et al. is directed to a coating having a plurality of alternate layers having different structures that produce a plurality of interfaces. The interfaces provide paths of increased resistance to heat transfer to reduce thermal conductivity.
Rickerby et al. teaches a traditional bond coat overlying a metallic substrate bonded to a TBC. The TBC comprises a plurality of layers, each layer having columnar grains, the columnar grains in each layer extending substantially perpendicular to the interface between the bond coating and metallic substrate. The structure is columnar to ensure that the strain tolerance of the ceramic TBC is not impaired. The difference in structure of the layers is the result of variations in the microstructure and/or density/coarseness of the columnar grains of the ceramic. These differences assist in providing to resistance to the transfer of heat across the thermal barrier coating.
Improved environmental resistance to destructive oxidation and hot corrosion is desirable. Additionally, the alloying elements of the bond coat interdiffuse with the substrate alloy at elevated temperatures of operation, changing the composition of the protective outer layer. Over time, as the airfoils are refurbished, walls of the airfoils are consumed, which reduces load carrying capability and limits blade life. Also, this interdiffusion can also reduce environmental resistance of the coating, causing loss of material, as layers of material are lost due to corrosive and oxidative effects. This interdiffusion and its adverse effects can be reduced by controlling the temperature of the component in the region of the bond coat/substrate interface. However, even with the use of advanced cooling designs and thermal barrier coatings, it is also desirable to decrease the requirement for cooling; because reducing the demand for cooling is also well known to improve overall engine operating efficiency.
One efficient cooling technique is film cooling. Film cooling is achieved by passing cooling air through discrete film cooling holes, typically ranging from 0.015″ to about 0.030″ in hole diameters. The film cooling holes are typically drilled with laser or by electro-discharge machining (EDM) or electro-stem (ES) machining. Due to mechanical limitations, each film hole has an angle ranging from 20° to 90° relative to the external surface. Therefore, each film jet exits from the hole with a velocity component perpendicular to the surface. But, because of this vertical velocity component and a complex aerodynamic flow circulation near the tip of a turbine blade, commonly referred to as the “squealer tip”, each film jet will have a tendency to lift or blow off from the external surface and mix with the hot exhaust gases, resulting in poor film cooling effectiveness in the area surrounding the squealer tip.
Thus, there is an ongoing need for an improved thermal barrier coating system surrounding the squealer tip, wherein the environmental resistance and long-term stability of the thermal barrier coating system is improved so that higher engine efficiencies can be achieved. The bond coat temperature limit is critical to the TBC's life and has an upper limit of about 2100° F. Once the bond coat exceeds this temperature, the coating system tends to quickly deteriorate, due to high temperature mechanical deformation and oxidation, as well as from interdiffusion of elements with the substrate alloy and subsequent
Darolia Ramgopal
Lee Ching-Pang
Schafrik Robert Edward
Look Edward K.
McCoy Kimya N
McNees Wallace & Nurick
Narciso David L.
LandOfFree
Turbine blade tip having thermal barrier coating-formed... does not yet have a rating. At this time, there are no reviews or comments for this patent.
If you have personal experience with Turbine blade tip having thermal barrier coating-formed..., we encourage you to share that experience with our LandOfFree.com community. Your opinion is very important and Turbine blade tip having thermal barrier coating-formed... will most certainly appreciate the feedback.
Profile ID: LFUS-PAI-O-2987866