Patent Description:
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 small 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.

<CIT> and <CIT> disclose adaptive machining methods for correcting the form of an as-cast airfoil including generating a number of measured data points along the surface of the trailing edge of the as-cast airfoil, comparing these measured data points to nominal data points, and machining the as-cast airfoil along a maching path generated from the differences in the measured and nominal data points.

According to an aspect of the disclosure, a method of manufacturing an aerodynamic element with an edge is provided as described in claim <NUM>. The aerodynamic element may include a turbine airfoil having a root and a tip, pressure and suction surfaces extending from the root to the tip.

The aerodynamic element may include a ceramic core.

The cutting machine may include one or more of a CNC machine, a ball endmill, an electro-dynamic machining (EDM) electrode and an electro-chemical machining (ECM) electrode.

The method may further include feeding cutting fluid through the aerodynamic element during the driving.

The predefined number of data points may be three.

The cutting toolpaths adapted toward correcting the as cast condition may be defined along one or more of radial, axial and circumferential dimensions.

Each of the cutting toolpaths adapted toward correcting the as cast condition may include one or more passes on each side of the trailing edge such that the trailing edge has a curvature at each side thereof.

The curvature at each side may be 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 is provided as described in claim <NUM>.

The following descriptions should not be considered limiting but are provided by way of example only.

The fan section <NUM> drives air along a bypass flow path B in a bypass duct, while the compressor section <NUM> drives air along a core flow path C for compression and communication into the combustor section <NUM> and then expansion through the turbine section <NUM>.

The exemplary gas turbine engine <NUM> generally includes a low speed spool <NUM> and a high speed spool <NUM> mounted for rotation about an engine central longitudinal axis A relative to an engine static structure <NUM> via several bearing systems <NUM>.

A combustor <NUM> is arranged in the gas turbine engine <NUM> between the high pressure compressor <NUM> and the high pressure turbine <NUM>. The engine static structure <NUM> is arranged generally between the high pressure turbine <NUM> and the low pressure turbine <NUM>. The engine static structure <NUM> further supports the bearing systems <NUM> in the turbine section <NUM>.

The core airflow is compressed by the low pressure compressor <NUM> and then the high pressure compressor <NUM>, is mixed and burned with fuel in the combustor <NUM> and is then expanded over the high pressure turbine <NUM> and the low pressure turbine <NUM>. The high and low pressure turbines <NUM> and <NUM> rotationally drive the low speed spool <NUM> and the high speed spool <NUM>, respectively, in response to the expansion. For example, geared architecture <NUM> may be located aft of the combustor section <NUM> or even aft of the turbine section <NUM>, and the fan section <NUM> may be positioned forward or aft of the location of geared architecture <NUM>.

The gas turbine engine <NUM> in one example is a high-bypass geared aircraft engine. In a further example, the gas turbine engine <NUM> bypass ratio is greater than about six (<NUM>), with an example embodiment being greater than about ten (<NUM>), the geared architecture <NUM> is an epicyclic gear train, such as a planetary gear system or other gear system, with a gear reduction ratio of greater than about <NUM> and the low pressure turbine <NUM> has a pressure ratio that is greater than about five. In one disclosed embodiment, the gas turbine engine <NUM> bypass ratio is greater than about ten (<NUM>:<NUM>), the fan diameter is significantly larger than that of the low pressure compressor <NUM>, and the low pressure turbine <NUM> has a pressure ratio that is greater than about five (<NUM>:<NUM>).

The geared architecture <NUM> may be an epicyclic gear train, such as a planetary gear system or other gear system, with a gear reduction ratio of greater than about <NUM>:<NUM>. 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.

The fan section <NUM> of the gas turbine engine <NUM> is designed for a particular flight condition--typically cruise at about <NUM> Mach (<NUM>/s) and about <NUM>,<NUM> feet (<NUM>,<NUM> meters). The flight condition of <NUM> Mach (<NUM>/s) and <NUM>,<NUM> ft (<NUM>,<NUM> 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 lbm of fuel being burned divided by lbf of thrust the engine produces at that minimum point.

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 <FIG>, a method of manufacturing an aerodynamic element with a trailing edge utilizing autonomous and/or adaptive machining is provided. As shown in <FIG>, the method includes producing the aerodynamic element with an initial or as-cast condition from an investment casting process (<NUM>) and cooling the aerodynamic element produced from the investment casting process (<NUM>). The method further includes generating a predefined number of data points sufficient to characterize contours of the trailing edge (<NUM>) and comparing the data points to a nominal condition to derive transformation parameters that are applicable to cutting toolpaths to thereby adapt the cutting toolpaths to the initial or as-cast condition (<NUM>). In additioon, the method includes driving a cutting machine in accordance with the cutting toolpaths adapted to the initial or as-cast condition (<NUM>) and, optionally, feeding cutting fluid through the aerodynamic element during the driving (<NUM>). In accordance with embodiments, the cutting machine can include or be provided as one or more of a CNC machine, a ball endmill, an electro-dynamic machining (EDM) electrode and an electro-chemical machining (ECM) electrode.

With reference to <FIG> and in accordance with embodiments, the aerodynamic element with the initial or as-cast condition can include or be provided as a turbine airfoil <NUM> for use in, for example, the gas turbine engine <NUM> of <FIG>. The turbine airfoil <NUM> has a root <NUM> and a tip <NUM> opposite the root <NUM>, a pressure surface <NUM> and a suction surface <NUM> opposite the pressure surface <NUM> where the pressure surface <NUM> and the suction surface <NUM> extend from the root <NUM> to the tip <NUM>, a leading edge <NUM> and a trailing edge <NUM> at leading and trailing sides of the pressure surface <NUM> and the suction surface <NUM>, respectively.

While the aerodynamic element has been described above as a turbine airfoil <NUM>, it is to be understood that other embodiments are possible. For example, with reference to <FIG>, the aerodynamic element could also be provided as one or more of turbine blade or vane, a fan, propeller or rotor blade, a ceramic core <NUM> used in the casting process of any of the above, etc. The following description will relate to the case in which the aerodynamic element is provided as the turbine airfoil <NUM>. This is being done for clarity and brevity and should not be interpreted as limiting the disclosure in any way.

With continued reference to <FIG> and with additional reference back to <FIG>, the generating of the predefined number of data points of operation <NUM> includes one or more of scanning, probing and measuring the turbine airfoil <NUM> with the initial or as-cast condition, the predefined number of data points are sufficient to characterize a position, size and shape of the turbine airfoil <NUM> with the initial or as-cast condition and the predefined number of data points are sufficient to characterize the contours of the trailing edge <NUM> relative to the position, the size and the shape of the turbine airfoil with the initial or as-cast condition.

With continued reference to <FIG> and with additional reference to <FIG>, the initial or as-cast condition of the turbine airfoil <NUM> is characterized in that the turbine airfoil <NUM> has an offset discharge. In such cases, the turbine airfoil <NUM> can be formed to define a discharge cavity <NUM>, through which coolant can be discharged from the turbine airfoil <NUM> during operations thereof, and this discharge cavity <NUM> is not in its correct or nominal position. That is, as shown in <FIG>, an initial shape of the trailing edge <NUM> of the turbine airfoil <NUM> with the initial or as-cast condition is generally squared-off with the expectation that the squared-off portion will be machined into a final edge-shape.

Where the turbine airfoil <NUM> has a nominal condition, as shown in <FIG>, the discharge cavity <NUM> should be aligned with the expected position of the trailing edge <NUM> in the final edge-shape (i.e., the discharge cavity <NUM> should be defined along the camber line of the turbine airfoil <NUM> proximate to the trailing edge <NUM>). However, where the turbine airfoil <NUM> has the initial or as-cast condition characterized in that the turbine airfoil <NUM> has the offset discharge, the discharge cavity <NUM> is at least initially mis-aligned with the expected position of the trailing edge <NUM> in the final edge-shape (i.e., the discharge cavity <NUM> is not defined along the camber line proximate to the trailing edge <NUM>).

Where the turbine airfoil <NUM> has the initial or as-cast condition characterized in that the turbine airfoil <NUM> has the offset discharge, the operation of generating the predefined number of data points of operation <NUM> (see <FIG>) includes one or more of scanning, probing and measuring the turbine airfoil <NUM> with the offset discharge, wherein the predefined number of data points are sufficient to characterize a position, size and shape of the turbine airfoil <NUM> with the offset discharge and the predefined number of data points are sufficient to characterize the contours of the trailing edge <NUM> relative to the position, size and shape of the turbine airfoil with the offset discharge.

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 <NUM> (see <FIG>), the data points are comparable with corresponding data points of a turbine airfoil with the nominal condition so that transformation matrices for the data points can be derived and these transformation matrices are applied to the cutting toolpaths resulting in the cutting toolpaths being adapted toward correcting the offset discharge. Thus, the driving of the cutting machine of operation <NUM> (see <FIG>) can be in accordance with the cutting toolpaths having been adapted toward correcting the offset discharge.

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 <FIG>) and, as shown in <FIG>, each of the cutting toolpaths adapted toward correcting the offset discharge can include one or more passes on each side of the trailing edge <NUM> such that the trailing edge <NUM> has a curvature <NUM> and <NUM> at each side thereof (i.e., a pressure-side curvature <NUM> at the pressure surface <NUM> with a predefined radius of curvature and a suction-side curvature <NUM> at the suction surface <NUM> with a predefined radius of curvature).

In an exemplary case, as shown in <FIG>, point A can be determined from an intersection of the chord length M as defined by the nominal condition and the camber line, radius R is determined by the section thickness at point A, point D is determined by the radius R and the camber line and vectors DB and DC are determined by the radius R and the intersection with the airfoil surface. For more complex forms, the radius R can be defined to vary between the pressure surface <NUM> and the suction surface <NUM> (with different point Ds), one can define the final shape to include a "flat" occupying some distance surrounding point A, in which case, the vector DA will be less than radius R. Also, instead of a circular radius, the rounding can be elliptical. Thus, in accordance with further embodiments, the pressure-side curvature <NUM> can be one or more of spherical, elliptical and complex and/or variable along one or more of the radial, axial and circumferential dimensions (see <FIG>) and the suction-side curvature <NUM> can be one or more of spherical, elliptical and complex and/or variable along one or more of the radial, axial and circumferential dimensions (see <FIG>).

With reference to <FIG>, a manufacturing machine <NUM> is provided for executing a method of manufacturing an aerodynamic element. The aerodynamic element can be any aerodynamic element including, but not limited to, the turbine airfoil <NUM> described above. The following description of the manufacturing machine <NUM> will relate to the case where the manufacturing machine is provided to manufacture the turbine airfoil <NUM> although it is to be understood that this is done for purposes of clarity and brevity.

The manufacturing machine <NUM> includes a casting unit <NUM>, a cooling element <NUM>, a cutting machine <NUM> and a processing system <NUM>. The casting unit <NUM> is configured to execute an investment casting process to produce the turbine airfoil <NUM> with an as-cast condition. As described above, the as-cast condition is characterized in that the turbine airfoil <NUM> has an offset discharge. The cooling element <NUM> is configured to cool the turbine airfoil <NUM> and the cutting machine <NUM> is configured to machine the turbine airfoil <NUM> following the cooling by the cooling element <NUM>. The processing system <NUM> is coupled to and disposed in signal communication with at least the cutting machine <NUM> and includes a processor, a memory unit, a servo control unit by which the processor can control operations of the cutting machine <NUM> 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 <NUM> as a whole to generate a predefined number of data points sufficient to characterize contours of the turbine airfoil <NUM> (i.e., the contours of the trailing edge <NUM> where the as-cast condition is characterized in that the turbine airfoil <NUM> 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 <NUM> 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. Additioinal 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 <NUM>", 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.

Claim 1:
A method of manufacturing an aerodynamic element (<NUM>) with a trailing edge (<NUM>), the method comprising:
casting (<NUM>) the aerodynamic element with an initial, as-cast condition;
wherein an initial shape of the trailing edge with the as-cast condition is generally squared-off with an expectation that the squared-off portion will be machined into a final edge-shape, and
wherein the aerodynamic element with the as-cast condition is formed to define a discharge cavity (<NUM>) through which coolant is discharged from the aerodynamic element during operations thereof, and the discharge cavity is at least initially mis-aligned with an expected position of the trailing edge in the final edge-shape;
cooling (<NUM>) the aerodynamic element;
generating (<NUM>) a predefined number of data points,
wherein the generating (<NUM>) of the predefined number of data points comprises one or more of scanning, probing and measuring the aerodynamic
element with the as-cast condition,
wherein the predefined number of data points are sufficient to characterize a position, size and shape of the aerodynamic element with the as-cast condition, and
wherein the predefined number of data points are sufficient to characterize contours of the trailing edge relative to the position, the size and the shape of the aerodynamic element with the as-cast condition;
comparing (<NUM>) the data points to a nominal condition of the aerodynamic element to derive transformation parameters applicable to cutting toolpaths to adapt the cutting toolpaths toward correcting the as-cast condition,
wherein when the aerodynamic element has the nominal condition, the discharge cavity is aligned with the expected position of the trailing edge in the final edge-shape; and
driving a cutting machine in accordance with the adapted cutting toolpaths.