Patent Application
20070050156
(1) Field of the Invention
The invention relates to gas turbine engines, and more particularly to methods for measuring the nozzle flow area between vanes in such engines.
(2) Description of the Related Art
A typical axial flow gas turbine engine operates by compressing ambient air in one or more forward compressors, injecting fuel and burning the mixture in a central combustor, and directing the products of combustion through one or more rearward turbines. The compressors and turbines each comprise alternating stages of rotor blades and stator vanes distributed circumferentially about one or more rotating spools and stationary cases respectively. A common low-pressure spool allows the forward compressor and the rearward turbine to rotate in unison, while a common high-pressure spool allows the rearward compressor and forward turbine to rotate in unison. The turbines convert kinetic energy stored in the combustion gas into mechanical energy for powering the forward compressors. The operation of the turbines directly influences the operation of the compressors, since common spools connect them.
The products of combustion flow rearward through a plurality of individual, semi-annular nozzle areas disposed between adjacent stator vanes. The semi-annular nozzle area is often referred to as the flow area. It is important that the total nozzle flow area of each turbine stage is properly sized to allow the turbines and compressors to operate at their optimum efficiency. It is also important to evenly distribute the individual nozzle flow areas circumferentially to reduce high cycle fatigue on the following blade stage due to combustion gas pulsing.
Because of original part manufacturing tolerances, extended engine operation and subsequent restoration processes, the individual nozzle areas will vary. In order to ensure the total nozzle flow area and individual nozzle flow area distribution are within engine specifications, each individual nozzle flow area must be measured prior to assembly in a turbine. A new or restored vane is typically mounted in a mechanical gage, where a series of mechanical probes contact the vane profile in a few locations on the vane surfaces. The distances between select locations on the vane profile and a nominally sized, adjacent vane is measured by the gage. The individual nozzle flow area is then typically calculated using the Simpson's Rule for calculating the area of a trapezoid. Vanes with nozzle flow areas falling within a pre-determined range of values are assigned a classification number. An algorithm optimizes the circumferential distribution of all the vanes in the stage based on specific engine criteria. Vanes are then assembled circumferentially in a turbine stage so that the difference in classification number between adjacent vanes is no more than one classification number.
While it is possible to calculate the nozzle flow area of new vanes using the Simpson's Rule method as described above, the calculation is only an approximation of the actual nozzle flow area, using only a few measured points. For restored vanes, there are even more obstacles to overcome in order to approximate the nozzle flow area using the Simpson's Rule method. Restored vanes may bow slightly during extended operation in the hot environment of a turbine, and restoration processes may shift the vane's datum locations or thin the nozzle walls from blending. It is difficult to properly locate restored vanes in a mechanical gage in a repeatable manner, and only a few measured locations may not accurately reflect subtle changes in the nozzle wall profile. The resulting classification number represents a range of areas and is not an accurate representation of the actual nozzle flow area of the restored vane.
What are therefore needed, are more accurate methods of measuring the nozzle flow area between vanes in gas turbine engines.
According to a turbine vane nozzle flow area measurement method, a convex surface and a concave surface of a first vane are scanned using a laser scanning system. A laser scanning system comprises a laser spot projector, a laser spot sensor, a multi-axis controller and a computer. A series of scanned points are stored as point clouds with the system. The point clouds are combined and translated into a reference coordinate system. A point cloud representing a nominally sized vane is positioned adjacent to one surface of the combined point cloud. An inlet profile is extracted from the intersection of a nozzle inlet plane located perpendicular to the combustion gases and the leading edge of each point cloud. A nozzle flow area is then accurately calculated using integration techniques from the inlet profile. The process is repeated for the other side of the vane.
According to another turbine vane nozzle flow area measurement method, both a pressure side nozzle wall and a suction side nozzle wall of a first vane are scanned and stored as a first point cloud using the laser scanning system. Then, both a pressure side nozzle wall and a suction side nozzle wall of a second vane are scanned and stored as a second point cloud using the system. A nozzle flow area is then measured between each adjacent point cloud using integration techniques. The scanning and storing steps are repeated for each vane in a turbine stage.
a is a perspective view illustrating a point cloud representing a convex surface of a first-stage vane of
b is a perspective view illustrating a point cloud representing a concave surface of a first-stage vane of
When referring to the above listed drawings, like reference numerals designate identical or corresponding elements throughout the various views.
In a typical axial-flow gas turbine engine, one or more turbine stages 10, as illustrated in
Individual vanes 14 are disposed circumferentially around the turbine 10, and each vane 14 comprises an inner diameter shroud 20, an outer diameter shroud 22 and an airfoil 24 spanning radially there between. The vanes 14 are cantilevered inward by their outer shrouds 22 from an outer case 26, circumscribing the turbine 10. A combination of supports 28 and threaded fasteners 30 secure the outer shrouds 22 to the case 26 radially beneath a number of external flanges 32. Additionally, the inner shrouds 20 may be secured to an inner support 34 or may carry an inter-stage seal 36.
Referring now to
A nozzle 50 cross sectional flow area is measured between a first 114 and a second 214 adjacent vane in a turbine stage. An inlet profile 56 is extracted from the intersection of a plane 58, positioned perpendicular to the mean combustion gas 18 flow vector direction, and the leading edges 42, end walls 46, 48 of adjacent vanes 114, 214.
A laser scanning system 60 for measuring the flow area of a nozzle 50 is illustrated in
A multi-axis controller 68, commonly used throughout various industries for accurate positioning during machining, measurement and other operations, carries the fixture 62 and first vane 114. The controller 68 comprises a servo 70 for driving a cross-slide 72 linearly about each of an X-axis 74 and a Y-axis 76, according to instructions from a computer 78. Since the cross-slides 72 move linearly within an X-Y plane only, movement within a Z-axis 80 is maintained constant. The controller 68 provides access to the airfoil 24 and endwalls 46, 48 without having to remove the first vane 114 from the fixture 62.
A laser spot projector 82 and a laser spot sensor 84 are mounted proximate to one another on one of the cross-slides 72. A small diameter laser beam 86, typically fifty micrometer or less, is directed from the spot projector 82 toward the first vane 114 and the sensor 84 receives a reflected light 88 back from the first vane 114. By measuring where the reflected light 88 contacts the sensor 84, the Z-axis 80 distance from the first vane 114 to the projector 82 may be calculated through triangulation. The z-axis 80 distance varies in response to changes in the topology of the first vane 114. A Keyence, LV series laser spot projector 82 and spot sensor 84 were used in the exemplary system 60.
The computer 78 comprises a memory device 90 and a processor 92, and is connected to the controller 68 via cables or a wireless connection. The computer 78 instructs the controller 68 to position the cross-slides 72 about the X-Y plane by means of the servos 70. Since the projector 82 and sensor 84 are mounted to one of the cross-slides 72, the laser beam 86 scans the first vane 114 as the cross-slides 72 traverse according to the instructions. The scan line density, or distance between constant X-axis 74 and Y-axis 76 scan positions (or scan lines), may be increased or decreased to produce a desired scan resolution. The processor 92 is programmed using C++ or any other suitable programming language.
While scanning the first vane 114, the sensor 84 outputs a calibrated Z-axis 80 distance as an analog voltage to the memory device 90. The corresponding, instantaneous X-axis 74 and Y-axis 76 distances are generated from the servos 70 driving the cross-slides 72. These three data sources: the X-axis 74 and Y-axis 76 distances from the servos 70 and the Z-axis 80 distance calibrated from the spot sensor 84, are captured continuously using a high-speed PC data bus and are stored in the memory device 90 as a series of digital coordinates.
Accordingly,
In a first embodiment of a nozzle flow area measurement method 300, the nozzle 50 of a first vane 114 is measured in relation to a second, nominally sized and positioned vane 214. The second vane 214 is created using a computer aided design (CAD) system and sized according to nominal blueprint dimensions.
The first vane 114 is located in a fixture 62 in step 301 as shown in
Referring now to step 302 and the examples illustrated in
Once the first point cloud 96 representing the first surface 38 or 40 is stored, the datum 64 surfaces are extracted in step 303 and a coordinate system 98 referencing these datum 64 surfaces, is constructed. All the points 94 in the first point cloud 196 are then transformed into the reference coordinate system 98.
The first vane 114 is removed from the fixture 62 and inverted, with the opposite surface 38 or 40 now facing the laser scanning system 60 in step 304. The three spheres 66 that are attached to the vane 114 should remain in the same location in the second scan without being disturbed as these are used for aligning and registering the two scans in subsequent steps. The first vane 114 is now located with a second, convex 38 or concave 40, surface visible by the laser spot projector 82 occultation-free, and within a proper scan depth distance from the projector 82.
The topography of the second surface 38 or 40 is next scanned into digital coordinate points 94 (shown as shading) in step 305, which are filtered to remove all outlying and extraneous points 94 created by stray laser beam 86 reflections and system 60 noise. The points 94 are combined into a second point cloud 296, which is stored in the memory device 90 for further manipulation by the processor 92.
Once the second point cloud 296 of the second surface 38 or 40 is stored, all the points 94 in the second point cloud 296 are then transformed in step 306 into the reference coordinate system 98 created in step 303. The three spherical tooling ball centers serve as gage points for matching the two scans into one scan in the next step. Now, both the first and second surfaces, 38 and 40, are stored as point clouds 196, 296 in the memory 92 and may be manipulated by the processor 92.
The point clouds 196 and 296 representing the first and second surfaces, 38 and 40, are combined in step 307 into a combined point cloud 396 (
A nominal point cloud 496 representing a nominally sized and oriented second vane 214 is positioned adjacent to a first side of the combined point cloud 396 in step 308. The positioning of the nominal point cloud 496 ensures that it is at a proper pitch distance and angular orientation from the combined point cloud 396 representing the first vane 114. The pitch distance is the distance between adjacent airfoils 24 and the angular orientation is measured with respect to the axial and radial planes of the engine.
An inlet plane 58 is located perpendicular to a combustion gas flow vector 18 direction at the nozzle 50 inlet in step 309. An inlet profile 56 is then extracted from the intersection of the inlet plane 55 and the combined point cloud 396 at the leading edge 42 and the outer 46 and inner 48 diameter end walls.
An inlet profile 100 of the second, nominal vane 214 is then extracted in step 310. The inlet profile is extracted from the intersection of the inlet plane 55 and the nominal point cloud 496 at the leading edge 42 and the outer 46 and inner 48 diameter endwalls.
The flow area of the nozzle 50 is calculated in step 311 from the area of the inlet profile 56 extracted from the combined point cloud 396 and the nominal point cloud 496. The area encompassed by the inlet profile 56 is calculated using one or more known numerical integration techniques. The flow area of the nozzle 50 for a second side of the first vane 114 is calculated in step 312 by repeating steps 308 through 311, with the nominal point cloud 496 located adjacent to the second side of the combined point cloud 396.
Steps 301 through 312 may be repeated in step 313 to obtain the nozzle flow areas of all the vanes 14 comprising a new or restored turbine vane stage. Once all the nozzle 50 flow areas are calculated and stored as described in the process steps outlined above, the individual nozzle 50 flow areas are sorted from smallest to largest and the vanes 14 are distributed circumferentially about the vane stage such that the difference in nozzle 50 flow area between any vane 14 and its adjacent neighbors is as small as possible. There are many known optimizers that accomplish this distribution once the nozzle 50 flow areas are measured.
In an alternate embodiment of a nozzle flow area measurement method 300, both first and second vanes 114, 214 are scanned as described in steps 301 through 307. The flow area of the nozzle 50 is then measured and calculated in relation to the first and second vanes 114, 214 and not in relation to a nominal vane. This provides an accurate measurement of the actual nozzle 50 flow area between the two vanes 114, 214.
While specific methods have been described in the context of accurately measuring and calculating nozzle flow area of first-stage, high-pressure turbine vanes, it is to be understood that other stages, low-pressure turbine vanes or even compressor stators would similarly benefit. Accordingly, the present invention is intended to embrace those alternatives, modifications and variations as fall within the broad scope of the appended claims.
