Patent Application
20210004636
Exemplary embodiments of the present disclosure relate generally to airfoils and, in one embodiment, to methods of manufacturing tairfoils with rounded trailing edges.
Airfoils are present in many aerodynamic applications including, but not limited to, turbines of gas turbine engines. These turbine airfoils each have a root, a tip, pressure and suction surfaces that extend from root to tip and leading and trailing edges at leading and trailing sides of the pressure and suction surfaces. In a turbine, the turbine airfoils can aerodynamically interact with high temperature and high pressure fluids to cause a rotor to rotate.
In cascade testing, it has been shown that turbine airfoils having rounded trailing edges reduce unsteady mixing effects and increase thermodynamic efficiency as compared to turbine airfoils that have squared trailing edges. The turbine airfoils with the rounded trailing edges achieve this effect by creating a wake effect. Even if these turbine airfoils have relatively large trailing edge diameters, the wake effect is similar to what would be produced by a turbine airfoil having a trailing edge with a relatively large trailing edge diameter.
Rounded profiles on trailing edges can be difficult to produce, however, and typically have only been producible on uncooled airfoils due to the need for a core printout from the trailing edge of a cooled airfoil resulting from investment casting processes. As such, a center-discharge airfoil thus often has an extended length that must be trimmed back, and a pressure side-discharge airfoil thus also has an encapsulation that must also be removed. This trimming is typically done manually to a witness line with belt grinders and hand-held rotary grinders, leaving sharp corners with only de-burring applied.
While certain machining processes, such as CNC, would be an approach to automate the process of trimming back the extended length of an airfoil, rigidly-programmed toolpaths (even with offsets) are insufficiently capable of accounting for variabilities in part-to-part shapes that are inherent in investment casting processes and it quickly becomes cost-prohibitive to inspect and program bespoke CNC code for each casting. Likewise, pre-fab electro-dynamic machining (EDM) and electro-chemical machining (ECM) electrodes covering the entire trailing edge are often unable to account for the casting variabilities. Pressure-sensitive robotic deburring has been attempted, but it does not have the necessary cutting power required to perform trailing edge finishing from a rough cast state, and the multiple degrees of freedom (DOF) in robotic arm articulation introduces more variation than desired.
According to an aspect of the disclosure, a method of manufacturing an aerodynamic element with an edge is provided. The method includes producing the aerodynamic element with an initial condition, cooling the aerodynamic element, generating a predefined number of data points sufficient to characterize contours of the edge and comparing the data points to a nominal condition to derive transformation parameters applicable to cutting toolpaths to adapt the cutting toolpaths to the initial condition.
In accordance with additional or alternative embodiments, the aerodynamic element includes a turbine airfoil having a root and a tip, pressure and suction surfaces extending from the root to the tip and the edge is one of a leading edge and a trailing edge at leading and trailing sides of the pressure and suction surfaces, respectively.
In accordance with additional or alternative embodiments, the aerodynamic element includes a ceramic core.
In accordance with additional or alternative embodiments, the generating of the predefined number of data points includes one or more of scanning, probing and measuring the aerodynamic element with the initial condition, the predefined number of data points are sufficient to characterize a position, size and shape of the aerodynamic element with the initial condition and the predefined number of data points are sufficient to characterize the contours of the edge relative to the position, the size and the shape of the aerodynamic element with the initial condition.
In accordance with additional or alternative embodiments, the initial condition is an as-cast condition and the as-cast condition is characterized as an offset discharge, the cutting toolpaths are adapted toward correcting the as-cast condition and the method further comprises driving a cutting machine in accordance with the cutting toolpaths adapted toward correcting the as-cast condition.
In accordance with additional or alternative embodiments, the cutting machine includes one or more of a CNC machine, a ball endmill, an electro-dynamic machining (EDM) electrode and an electro-chemical machining (ECM) electrode.
In accordance with additional or alternative embodiments, the method further includes feeding cutting fluid through the aerodynamic element during the driving.
In accordance with additional or alternative embodiments, the cutting toolpaths adapted toward correcting the as-cast condition are defined along one or more of radial, axial and circumferential dimensions.
In accordance with additional or alternative embodiments, each of the cutting toolpaths adapted toward correcting the as-cast condition includes one or more passes on each side of the trailing edge such that the trailing edge has a curvature at each side thereof.
In accordance with additional or alternative embodiments, the curvature at each side is one or more of one or more of spherical, elliptical and complex and variable along one or more of radial, axial and circumferential dimensions.
According to another aspect of the disclosure, a method of manufacturing a turbine airfoil having a root and a tip, pressure and suction surfaces extending from the root to the tip, and leading and trailing edges at leading and trailing sides of the pressure and suction surfaces, respectively, is provided. The method includes producing the turbine airfoil with an as-cast condition from an investment casting process, cooling the turbine airfoil, generating a predefined number of data points sufficient to characterize contours of the trailing edge and comparing the data points to a nominal condition to derive transformation parameters applicable to cutting toolpaths to adapt the cutting toolpaths to the as-cast condition.
In accordance with additional or alternative embodiments, the generating of the predefined number of data points comprises one or more of scanning, probing and measuring the turbine airfoil with the as-cast condition, the predefined number of data points are sufficient to characterize a position, size and shape of the turbine airfoil with the as-cast condition and the predefined number of data points are sufficient to characterize the contours of the trailing edge relative to the position, the size and the shape of the turbine airfoil with the as-cast condition.
In accordance with additional or alternative embodiments, the as-cast condition is characterized as an offset discharge and the cutting toolpaths are adapted toward correcting the as-cast condition.
In accordance with additional or alternative embodiments, the method further includes driving a cutting machine in accordance with the cutting toolpaths adapted toward correcting the as-cast condition.
In accordance with additional or alternative embodiments, the cutting machine comprises one or more of a CNC machine, a ball endmill, an electro-dynamic machining (EDM) electrode and an electro-chemical machining (ECM) electrode.
In accordance with additional or alternative embodiments, the method further includes feeding cutting fluid through the turbine airfoil during the driving.
In accordance with additional or alternative embodiments, the cutting toolpaths adapted toward correcting the as-cast condition are defined along one or more of radial, axial and circumferential dimensions.
In accordance with additional or alternative embodiments, each of the cutting toolpaths adapted toward correcting the as-cast condition includes one or more passes on each side of the trailing edge such that the trailing edge has a curvature at each side thereof.
In accordance with additional or alternative embodiments, the curvature at each side is one or more of one or more of spherical, elliptical and complex and variable along one or more of radial, axial and circumferential dimensions.
According to another aspect of the disclosure, a manufacturing machine for manufacturing an aerodynamic element. The manufacturing machine includes a casting unit configured to execute a casting process to produce the aerodynamic element with an initial condition, a cooling element configured to cool the aerodynamic element, a cutting machine configured to machine the aerodynamic element following cooling by the cooling element and a processing system. The processing system is configured to generate a predefined number of data points sufficient to characterize contours of the aerodynamic element, compare the data points to a nominal condition to derive transformation parameters applicable to cutting toolpaths to adapt the cutting toolpaths toward correcting the initial condition, and drive the cutting machine in accordance with the cutting toolpaths adapted toward correcting the initial condition.
These and other advantages and features will become more apparent from the following description taken in conjunction with the drawings.
The following descriptions should not be considered limiting in any way. With reference to the accompanying drawings, like elements are numbered alike:
These and other advantages and features will become more apparent from the following description taken in conjunction with the drawings.
A detailed description of one or more embodiments of the disclosed apparatus and method are presented herein by way of exemplification and not limitation with reference to the Figures.
The exemplary gas turbine engine 20 generally includes a low speed spool 30 and a high speed spool 32 mounted for rotation about an engine central longitudinal axis A relative to an engine static structure 36 via several bearing systems 38. It should be understood that various bearing systems 38 at various locations may alternatively or additionally be provided, and the location of bearing systems 38 may be varied as appropriate to the application.
The low speed spool 30 generally includes an inner shaft 40 that interconnects a fan 42, a low pressure compressor 44 and a low pressure turbine 46. The inner shaft 40 is connected to the fan 42 through a speed change mechanism, which in exemplary gas turbine engine 20 is illustrated as a geared architecture 48 to drive the fan 42 at a lower speed than the low speed spool 30. The high speed spool 32 includes an outer shaft 50 that interconnects a high pressure compressor 52 and high pressure turbine 54. A combustor 56 is arranged in the gas turbine engine 20 between the high pressure compressor 52 and the high pressure turbine 54. The engine static structure 36 is arranged generally between the high pressure turbine 54 and the low pressure turbine 46. The engine static structure 36 further supports the bearing systems 38 in the turbine section 28. The inner shaft 40 and the outer shaft 50 are concentric and rotate via bearing systems 38 about the engine central longitudinal axis A which is collinear with their longitudinal axes.
The core airflow is compressed by the low pressure compressor 44 and then the high pressure compressor 52, is mixed and burned with fuel in the combustor 56 and is then expanded over the high pressure turbine 54 and the low pressure turbine 46. The high and low pressure turbines 54 and 46 rotationally drive the low speed spool 30 and the high speed spool 32, respectively, in response to the expansion. It will be appreciated that each of the positions of the fan section 22, compressor section 24, combustor section 26, turbine section 28, and fan drive gear system 48 may be varied. For example, geared architecture 48 may be located aft of the combustor section 26 or even aft of the turbine section 28, and the fan section 22 may be positioned forward or aft of the location of geared architecture 48.
The gas turbine engine 20 in one example is a high-bypass geared aircraft engine. In a further example, the gas turbine engine 20 bypass ratio is greater than about six (6), with an example embodiment being greater than about ten (10), the geared architecture 48 is an epicyclic gear train, such as a planetary gear system or other gear system, with a gear reduction ratio of greater than about 2.3 and the low pressure turbine 46 has a pressure ratio that is greater than about five. In one disclosed embodiment, the gas turbine engine 20 bypass ratio is greater than about ten (10:1), the fan diameter is significantly larger than that of the low pressure compressor 44, and the low pressure turbine 46 has a pressure ratio that is greater than about five 5:1. Low pressure turbine 46 pressure ratio is pressure measured prior to inlet of low pressure turbine 46 as related to the pressure at the outlet of the low pressure turbine 46 prior to an exhaust nozzle. The geared architecture 48 may be an epicycle gear train, such as a planetary gear system or other gear system, with a gear reduction ratio of greater than about 2.3:1. It should be understood, however, that the above parameters are only exemplary of one embodiment of a geared architecture engine and that the present disclosure is applicable to other gas turbine engines including direct drive turbofans.
A significant amount of thrust is provided by the bypass flow B due to the high bypass ratio. The fan section 22 of the gas turbine engine 20 is designed for a particular flight condition—typically cruise at about 0.8 Mach and about 35,000 feet (10,688 meters). The flight condition of 0.8 Mach and 35,000 ft (10,688 meters), with the engine at its best fuel consumption—also known as “bucket cruise Thrust Specific Fuel Consumption (‘TSFC’)”—is the industry standard parameter of 1 bm of fuel being burned divided by 1 bf of thrust the engine produces at that minimum point. “Low fan pressure ratio” is the pressure ratio across the fan blade alone, without a Fan Exit Guide Vane (“FEGV”) system. The low fan pressure ratio as disclosed herein according to one non-limiting embodiment is less than about 1.45. “Low corrected fan tip speed” is the actual fan tip speed in ft/sec divided by an industry standard temperature correction of [(Tram ° R)/(518.7° R)]0.5. The “Low corrected fan tip speed” as disclosed herein according to one non-limiting embodiment is less than about 1150 ft/second (350.5 m/sec).
As will be described below, a method of manufacturing is provided and utilizes autonomous adaptive machining to accomplish trailing edge rounding of an aerodynamic element, such as a turbine airfoil, with the necessary tolerances for reliable aerodynamic benefit and process capability for producibility and affordability.
With reference to
With reference to
While the aerodynamic element has been described above as a turbine airfoil 301, it is to be understood that other embodiments are possible. For example, with reference to
With continued reference to
With continued reference to
Where the turbine airfoil 301 has a nominal condition, as shown in
Where the turbine airfoil 301 has the initial or as-cast condition characterized in that the turbine airfoil 301 has the offset discharge, the operation of generating the predefined number of data points of operation 203 (see
In accordance with embodiments, the number of the data points can be as little as three and up to a number which is substantially larger than three assuming sufficient computing resources are available. For operation 204 (see
In accordance with embodiments, the cutting toolpaths adapted toward correcting the initial or as-cast condition can be defined along one or more of radial, axial and circumferential dimensions (see
In an exemplary case, as shown in
With reference to
The manufacturing machine 601 includes a casting unit 610, a cooling element 620, a cutting machine 630 and a processing system 640. The casting unit 610 is configured to execute an investment casting process to produce the turbine airfoil 301 with an as-cast condition. As described above, the as-cast condition can be characterized in that the turbine airfoil 301 has an offset discharge. The cooling element 620 is configured to cool the turbine airfoil 301 and the cutting machine 630 is configured to machine the turbine airfoil 301 following the cooling by the cooling element 620. The processing system 640 is coupled to and disposed in signal communication with at least the cutting machine 630 and includes a processor, a memory unit, a servo control unit by which the processor can control operations of the cutting machine 630 and an input/output (I/O) bus by which the processor can communicate with the memory unit and the servo control unit. The memory unit has executable instructions stored thereon, which are readable and executable by the processor. When the executable instructions are read and executed by the processor, the executable instructions effectively cause the processor to operate as described herein.
For example, when the executable instructions are read and executed by the processor, the executable instructions effectively cause the processor and the processing system 640 as a whole to generate a predefined number of data points sufficient to characterize contours of the turbine airfoil 301 (i.e., the contours of the trailing edge 308 where the as-cast condition is characterized in that the turbine airfoil 301 has an offset discharge), to compare the data points to a nominal condition to derive transformation parameters applicable to cutting toolpaths to adapt the cutting toolpaths toward correcting the as-cast condition and to drive the cutting machine 630 in accordance with the cutting toolpaths adapted toward correcting the as-cast condition.
Benefits of the features described herein are the provision of turbine airfoils with rounded trailing edges that are produced when the turbine airfoils are cooled airfoils, with the associated benefits to performance and incidental shop part damage prevention. Additional benefits are that variations from investment casting processes (e.g., airfoil bow, lean, twist, wall thickness, etc.) are autonomously adjusted, the rounded profiles are controllable in three dimensions to tolerances of roughly 0.001″, coat-down effects can be fed back into computer-aided modeling (CAM) routines for correction at the casting level and cost avoidance from manual production of rounded trailing edges.
The term “about” is intended to include the degree of error associated with measurement of the particular quantity based upon the equipment available at the time of filing the application.
The terminology used herein is for the purpose of describing particular embodiments only and is not intended to be limiting of the present disclosure. As used herein, the singular forms “a”, “an” and “the” are intended to include the plural forms as well, unless the context clearly indicates otherwise. It will be further understood that the terms “comprises” and/or “comprising,” when used in this specification, specify the presence of stated features, integers, steps, operations, elements, and/or components, but do not preclude the presence or addition of one or more other features, integers, steps, operations, element components, and/or groups thereof.
While the present disclosure has been described with reference to an exemplary embodiment or embodiments, it will be understood by those skilled in the art that various changes may be made and equivalents may be substituted for elements thereof without departing from the scope of the present disclosure. In addition, many modifications may be made to adapt a particular situation or material to the teachings of the present disclosure without departing from the essential scope thereof. Therefore, it is intended that the present disclosure not be limited to the particular embodiment disclosed as the best mode contemplated for carrying out this present disclosure, but that the present disclosure will include all embodiments falling within the scope of the claims.
