Patent Application
20130089421
1. Field of the Invention
This application relates generally to gas turbine engine clearances between relatively rotating airfoil tips and seals and, more particularly, to clearances at airfoil tips of blades and cantilevered vanes.
2. Description of Related Art
Gas turbine engine blades have airfoils have gaps between the rotatable blades and static seals or casings and static cantilevered vanes and rotating seals or rotors in both turbines and the compressors. Often referred to as tip clearances, it is desirable for an efficiency standpoint to reduce the gap between the rotating component and the radially adjacent static part in order to reduce leakage of the gas stream across this gap. The leakage not only reduces efficiency of the compressor or the turbine, but also reduces the life of the turbine blade tips and shroud members because of high temperatures acting on the parts.
Some gas turbine engine designers will set the gap such that the blades will not rub at all. Some designers provide a negative gap in order to produce rub during the initial engine break-in in order to allow for the normal wear from the rub to produce a smooth and close to zero gap as possible. However, this rub can have undesirable results such as high vibratory stress at an airfoil tip of a rotatable blade or static cantilevered vane or at a tip of an impeller. A rub can have undesirable result of a loss of material fatigue strength.
Blade tip clearances are a compromise between avoiding rubs and minimizing leakage for best engine performance. Rubs can lead to loss of strength, cracking and additional maintenance. Currently, clearances are often set more open, especially at LE and TE where blades are thin. This avoids rubs in areas of high stress, maintains material properties, and avoids tip cracking. Thus, it is highly desirable to provide an airfoil tip that has low leakage and avoids rubs in areas of high stress in order to maintain material properties and avoid tip cracking.
A gas turbine engine component includes an airfoil extending from an airfoil base to an airfoil tip at a free end of the airfoil and extending downstream from a leading edge to a trailing edge of the airfoil. At least one recess extends into and circumferentially completely through the airfoil tip and is located inwardly of the leading and trailing edges of the airfoil.
The recess may be located in an area of the airfoil tip subject to high tip vibratory stress and or rubs and may be a circular scallop having a scallop radius. The recess may have a maximum depth of a few mills in a range of about 5-10 mills (0.005-0.01 inches) as measured from a nominal tip edge without the recess.
The airfoil may be on a gas turbine engine radial or impeller compressor blade. The airfoil may be on a gas turbine engine stator vane and extend radially inwardly from a base of the vane airfoil at an outward end of the stator vane to a vane airfoil tip of the vane airfoil at a radial inward end of the stator vane and extend downstream from a leading edge to a trailing edge of the vane airfoil. The recess extends into and circumferentially completely through the vane airfoil tip and the recess is located inwardly of the leading and trailing edges of the vane airfoil.
A gas turbine engine assembly may include a plurality of the airfoils and an airfoil tip clearance between airfoil tips and an annular tip seal surrounding the airfoil tips containing the recesses or scallops.
A method of reducing vibratory stress at an airfoil tip at a free end of an airfoil of a gas turbine engine component includes determining an area or areas of high tip vibratory stress and machining, cutting, or otherwise forming at least one recess or scallop extending into and circumferentially completely through the airfoil tip in the area or areas of high tip vibratory stress.
The airfoil tip and airfoil may be in pluralities of airfoil tips and airfoils from a single stage or circumferential row of blades or vanes surrounded by annular tip seals and the determining an area or areas of high tip vibratory stress includes running or rotating a rotor containing the blades or an annular tip seal surrounding the vanes respectively and observing nicks or scratches in the annular seal or airfoil tips. The recesses or scallops are formed in area or areas of high tip vibratory stresses adjacent the nicks or scratches.
The foregoing aspects and other features of the invention are explained in the following description, taken in connection with the accompanying drawings where:
Illustrated in
The exemplary embodiment of the compressor 14 illustrated herein includes a five stage axial compressor 30 followed by the single stage centrifugal compressor 18 having an annular centrifugal compressor impeller 32. Outlet guide vanes 40 are disposed between the five stage axial compressor 30 and the single stage centrifugal compressor 18. The compressor 14 includes a forward casing 110 and an aft casing 114. The forward casing 110 generally surrounds the axial compressor 30 and the aft casing 114 generally surrounds the centrifugal compressor 18 and supports the diffuser 42 directly downstream of the centrifugal compressor 18.
Referring further to
The stator vanes 204 are cantilevered from and fixed to the forward casing 110 of their radial outward ends 234 and are unsupported at their radial inward ends 236 which are free ends 87. A vane airfoil 225 extends radially between the opposite radial outward and inward ends 234, 236. Each vane airfoil 225 extends radially inwardly from a base 228 of the vane airfoil 225 at the outward end 234 of the stator vane 204 to a vane airfoil tip 138 of the vane airfoil 225 at the radial inward end 236 of the stator vane 204. The vane airfoil 225 extends downstream from a leading edge LE to a trailing edge TE. Each vane airfoil tip 138 is radially spaced apart and outwardly from and adjacent to a rotor seal land 232 on the high pressure rotor 12. An annular vane airfoil tip clearance 66 is defined between the vane airfoil tip 138 and the rotor seal land 232 on the high pressure rotor 12.
Referring to
It is desirable to minimize the rotor airfoil tip clearance 60, the stator airfoil tip clearance 66, and the impeller airfoil tip clearance 80 during the engine operating cycle and avoid or minimize rubs between the associated lands and airfoil and blade tips, particularly, during engine accelerations such as during cold bursts. In more general terms, it is desirable to minimize airfoil tip clearances 140 (illustrated herein as the rotor, vane, and impeller airfoil tip clearances 60, 66, 80) between airfoil tips 142 (illustrated herein as the blade, vane, and impeller airfoil tips 58, 138, 86) and annular tip seals 144 (illustrated herein as the blade rub land 62 on the compressor forward casing 110, the rotor seal land 232 on the high pressure rotor 12, and the centrifugal blade tip shroud 90). To this end, the rotor and vane airfoil tips 58, 138 and the impeller airfoil tips 86, as illustrated in
As exemplified in
The recesses 92 and scallops 94 are formed in or cut into nominal airfoil tip edges 106 of the rotor airfoil tips 58 which is illustrated as a dashed line in
The recesses 92 and scallops 94 are very shallow having a maximum depth D of only a few mills such as in a range of about 5-10 mills (0.005-0.01 inches) as measured from the nominal airfoil and impeller tip edges 106, 108 respectively for the exemplary engine illustrated herein. Larger engines may have larger recesses or scallops. The scallops 94 are circular having a scallop radius SR. The scallop radius SR of the scallops 94 in the impeller tip edges 108 of the impeller airfoil tips 86 is smaller than the first radius of curvature R1 of the impeller blade tip shape 102.
There are known methods of determining areas of high tip vibratory stress in the airfoil tips and impeller blade tips. There are analytical methods such as finite element analysis, empirical methods, and semi-empirical in which a combination of testing and analytical methods are used to determine if stresses will produce cracks or if there is undesirable rubbing. During engine production and overhaul, the engine is often run or rotated and rubbing may occur between airfoil tips and surrounding seals indicating areas of high tip vibratory stress in the airfoil tips and impeller blade tips. Nicks and scratches observed after such engine runs indicate such areas and where the recesses or scallops should be placed on the airfoil tips and impeller blade tips. The recesses and scallops can be machined or cut into the airfoil tips adjacent the nicks and scratches. The size and depth of the recesses and scallops may be tailored for each individual airfoil and, thus, vary from airfoil to airfoil within a stage or annular row of blades, vanes or impeller blades.
The present invention has been described in an illustrative manner. It is to be understood that the terminology which has been used is intended to be in the nature of words of description rather than of limitation. While there have been described herein, what are considered to be preferred and exemplary embodiments of the present invention, other modifications of the invention shall be apparent to those skilled in the art from the teachings herein and, it is, therefore, desired to be secured in the appended claims all such modifications as fall within the true spirit and scope of the invention.
Accordingly, what is desired to be secured by Letters Patent of the United States is the invention as defined and differentiated in the following claims:
