Patent Application
20190309629
Exemplary embodiments of the present disclosure relate generally to a method of forming a blade or a vane for a gas turbine engine and, in one embodiment, to a method of forming a blade or a vane for a gas turbine engine using a positive witness band.
A gas turbine engine typically includes a fan section, a compressor section, a combustor section and a turbine section. Air entering the compressor section is compressed and delivered into the combustion section where it is mixed with fuel and ignited to generate a high-energy exhaust gas flow. The high-energy exhaust gas flow expands through the turbine section to drive the compressor and the fan section. The compressor section typically includes low and high pressure compressors, and the turbine section includes low and high pressure turbines.
Both the compressor and turbine sections include rotating blades alternating between stationary vanes. The vanes and rotating blades in the turbine section extend into the flow path of the high-energy exhaust gas flow. All structures within the exhaust gas flow path are exposed to extreme temperatures and need to be formed to certain, particular dimensions.
Accordingly, it is desirable to provide for a method of forming at least vanes that provides for greater dimensional control.
According to an aspect of the disclosure, a method of forming a turbine blade or vane is provided and includes forming a wax pattern in a turbine blade or vane shape and additively forming a positive witness band around a tip portion of the wax pattern.
In accordance with additional or alternative embodiments, the additive forming includes at least one of fused deposition modeling (FDM), selective laser sintering (SLS), direct metal laser sintering (DMLS), laser beam melting (LBM), electron beam melting (EBM), inkjet 3D printing and stereolithography.
In accordance with additional or alternative embodiments, the additive forming of the positive witness band is executed such that the positive witness band is about 0.030″ wide and protrudes from the tip portion of the wax pattern by about 0.005-0.010″.
In accordance with additional or alternative embodiments, the additive forming of the positive witness band is executed such that the positive witness band comprises a fillet.
In accordance with additional or alternative embodiments, the additive forming of the positive witness band is executed such that the positive witness band traverses sacrificial portions of the wax pattern.
In accordance with additional or alternative embodiments, the method further includes forming a mold of the wax pattern and the positive witness band, casting the turbine blade or vane with a positive witness band formation corresponding to the positive witness band in the mold and cutting the turbine blade or vane along the positive witness band formation.
In accordance with additional or alternative embodiments, the turbine blade or vane with the positive witness band formation has a substantially uniform wall thickness proximate to the positive witness band formation.
In accordance with additional or alternative embodiments, the cutting includes driving a cutting tool to execute the cutting along the positive witness band formation.
According to another aspect of the disclosure, a method of forming a turbine blade or vane is provided and includes forming a wax pattern in a turbine blade or vane shape, additively forming a positive witness band around a tip portion of the wax pattern, performing an investment casting using the wax pattern to form the turbine blade or vane with a positive witness band formation corresponding to the positive witness band in the mold and cutting the turbine blade or vane along the positive witness band formation.
In accordance with additional or alternative embodiments, the additive forming includes at least one of fused deposition modeling (FDM), selective laser sintering (SLS), direct metal laser sintering (DMLS), laser beam melting (LBM), electron beam melting (EBM), inkjet 3D printing and stereolithography.
In accordance with additional or alternative embodiments, the additive forming of the positive witness band is executed such that the positive witness band is about 0.030″ wide and protrudes from the tip portion of the wax pattern by about 0.005-0.010″.
In accordance with additional or alternative embodiments, the additive forming of the positive witness band is executed such that the positive witness band comprises a fillet.
In accordance with additional or alternative embodiments, the additive forming of the positive witness band is executed such that the positive witness band traverses sacrificial portions of the wax pattern.
In accordance with additional or alternative embodiments, the turbine blade or vane with the positive witness band formation has a substantially uniform wall thickness proximate to the positive witness band formation.
In accordance with additional or alternative embodiments, the cutting includes driving a cutting tool to execute the cutting along the positive witness band formation.
According to yet another aspect of the disclosure, a wax pattern for a casting of a turbine blade or vane is provided. The wax pattern includes a root, a tip portion opposite the root, opposed pressure and suction surfaces extending between leading and trailing edges from the root to the tip portion and an additively manufactured positive witness band extending around the tip portion.
In accordance with additional or alternative embodiments, the additively manufactured positive witness band includes a material which is similar to or different from that of the tip portion.
In accordance with additional or alternative embodiments, the wax pattern further includes sacrificial portions disposed at least along the trailing edge and cooling circuit elements in an interior of the wax pattern, wherein the additively manufactured witness band traverses the sacrificial portions.
In accordance with additional or alternative embodiments, the positive witness band is about 0.030″ wide and protrudes from the tip portion by about 0.005-0.010″.
In accordance with additional or alternative embodiments, the positive witness band includes a fillet.
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 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 exemplary gas turbine 20 between the high pressure compressor 52 and the high pressure turbine 54. An 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 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 then the high pressure compressor 52, mixed and burned with fuel in the combustor 56, then expanded over the high pressure turbine 54 and low pressure turbine 46. The turbines 46, 54 rotationally drive the respective low speed spool 30 and high speed spool 32 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, gear system 48 may be located aft of combustor section 26 or even aft of turbine section 28, and fan section 22 may be positioned forward or aft of the location of gear system 48.
The engine 20 in one example is a high-bypass geared aircraft engine. In a further example, the 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 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 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 lbm of fuel being burned divided by lbf 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).
Although a two-stage high pressure turbine is illustrated, other high pressure turbines are considered to be within the scope of various embodiments of the present disclosure.
With reference to
With reference to
As shown in
The additively manufactured positive witness band 430 can be formed from various additive manufacturing processes including, but not limited to, at least one or more of fused deposition modeling (FDM), selective laser sintering (SLS), direct metal laser sintering (DMLS), laser beam melting (LBM), electron beam melting (EBM), inkjet 3D printing, stereolithography, or any other suitable additive layer manufacturing (ALM) process. The principle behind additive manufacturing processes involves the selective melting of atomized precursor material and producing the lithographic build-up of the workpiece. In some ALM processes, powder beds are melted by a directed energy source. The melting of the powder occurs in a small localized region of the energy beam, producing small volumes of melting, called melt pools, followed by rapid solidification, allowing for very precise control of the solidification process in the layer-by-layer fabrication of the workpiece. An example of a particular type of system is a PBF-L (powder bed fusion-laser) additive system where the energy beam is a laser. Any of the above devices may be directed by three-dimensional geometry solid models developed in Computer Aided Design (CAD) software systems.
In addition, the additively manufactured positive witness band 430 may include or be formed of a material that is similar to or different from that of the tip portion 403 as long as the material is compatible with the additive manufacturing process being employed for its formation.
In accordance with further embodiments and, as shown in
In accordance with further embodiments and, with reference to
With continued reference to
As used herein, the terms “turbine blade”, “turbine vane” or “turbine blade or vane” can be used interchangeably with “turbine aerodynamic element,” which is disposable within a flow of high pressure and high temperature fluids in a turbine to aerodynamically interact with such high pressure and high temperature fluids.
Benefits of the features described herein are the provision of a method of forming a turbine blade that includes a positive witness band formation which can be used to improve cutting processes. Since the positive witness band formation results from the presence of an additively manufactured positive witness band, the positive witness band formation protrudes outwardly from the turbine blade surface and facilitates the cutting.
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.
This application claims the benefit of U.S. Provisional Application 62/653,327, which was filed on Apr. 5, 2018. The entire contents of U.S. Provisional Application 62/653,327 are incorporated herein by reference.
| Number | Date | Country | |
|---|---|---|---|
| 62653327 | Apr 2018 | US |
