Patent Application
20120171045
This invention relates generally to workpiece fixtures, and specifically to fixtures for turbomachinery components. In particular, the invention concerns a workpiece fixture for gas turbine engine components, including rotor blade and stator vane airfoils.
Gas turbine engines include a variety of rotary-type internal combustion engines and combustion turbines, with applications in industrial power generation, aviation and transportation. The core of the gas turbine engine typically comprises a compressor, a combustor and a turbine, which are arranged in flow series with an upstream inlet and downstream exhaust. Incoming air is compressed in the compressor and mixed with fuel in the combustor, then ignited to generate hot combustion gas. The turbine generates rotational energy from the hot combustion gas, and cooler, expanded combustion products are exhausted downstream.
The compressor and turbine sections are usually arranged into one or more differentially rotating spools. The spools are further divided into stages, or alternating rows of blades and vanes. The blades and vanes generally have airfoil-shaped cross sections, which are designed to accelerate, turn and compress the working fluid flow, and to generate lift that is converted to rotational energy in the turbine.
In industrial gas turbines, power is delivered via an output shaft coupled to an electrical generator or other load, typically utilizing an external gearbox. Other configurations include turbofan, turboprop, turbojet and turboshaft engines for fixed-wing aircraft and helicopters, and specialized turbine engines for marine and land-based transportation, including naval vessels, trains and armored vehicles.
In turboprop and turboshaft engines the turbine drives a propeller or rotor, typically using a reduction gearbox to control blade speed. Turbojets generate thrust primarily from the exhaust, while turbofans drive a fan to accelerate flow around the engine core. Commercial turbofans are usually ducted, but unducted designs are also known. Some turbofans also utilize a geared drive to provide greater fan speed control, for example to reduce noise and increase engine efficiency, or to increase or decrease specific thrust.
Aviation turbines generally have two or three-spool configurations, with a corresponding number of coaxially rotating turbine and compressor sections. In two-spool designs the high pressure turbine drives a high pressure compressor, forming the high pressure spool or high spool. The low spool drives the fan or a propeller or rotor shaft, and may include one or more low pressure compressor stages. Aviation turbines also power auxiliary devices including electrical generators, hydraulic pumps and components of the environmental control system, either via an accessory gearbox using bleed air from the compressor.
In high-bypass turbofans, most of the thrust is generated by the fan. Variable-area nozzle surfaces can be deployed to regulate the bypass pressure and improve fan performance, particularly during takeoff and landing. Low-bypass turbofans provide greater specific thrust but are louder and less fuel efficient, and are more common on military jets and other high-performance aircraft. Low-bypass turbofans generally have variable-area nozzle systems to regulate exhaust speed and specific thrust, and military jets typically include afterburner assemblies for short-term thrust augmentation.
In general, gas turbine engine performance is constrained by the need for higher compression ratios and combustion temperatures, which increase efficiency and output, versus the cost of increased wear and tear on turbine components, including blades, struts and vanes, and the associated airfoil, platform and shroud surfaces. These tradeoffs are particularly relevant in the turbine stages downstream of the combustor, where gas path temperatures are elevated and active cooling is employed.
To increase dependability and service life, each step in the manufacturing process should therefore be uniformly controlled and efficiently performed. Fixture design plays a substantial role in this arena, from initial component machining to final part testing and certification.
This invention concerns a fixture system for a turbine component. The fixture comprises first and second end blocks, a load beam, and a compliant mask. The turbine component comprises first and second ends, and first and second regions having different surface or cooling features.
The first and second end blocks are positioned adjacent the first and second ends of the turbine component, and coupled together with the load beam to retain the component therebetween. The compliant mask is positioned to cover the first region of the turbine component, leaving the second region uncovered. A removable coating is applied to coat the surface or cooling features in the second region, leaving the first region uncoated.
As shown in
Turbine component 14 is typically manufactured of a durable, high temperature material such as a nickel or cobalt alloy or superalloy. In the particular embodiment of
Compliant mask 12 is formed of a flexible, pliant, elastic material such as natural or synthetic rubber, or another suitably pliable or compliant composite or polymer material. Compliant mask 12 is positioned to cover a selected region of cooling holes or other features on turbine component 14, leaving a second region uncovered for application of a wax coating or other removable coating material. The coating material provides for airflow testing and other selective processing steps, with compliant mask 12 to improve the speed, reliability and efficiency of these steps as described below.
First and second end blocks 16 and 18 are formed of a rigid, high strength polymer material such as nylon, or another suitably rigid polymer or composite material. End blocks 16 and 18 comprise lands 25 for sealing and retaining ends 23 and 24 of turbine component 14, and coupling structures 26 and 28 for connecting to load beam 20.
Lands 25 have recessed, flush and protruding embodiments, as configured for sealing engagement with first and second ends 23 and 24 of turbine component 14. In particular, lands 25 accommodate blade, vane and strut embodiments of component 14, and shrouded and unshrouded embodiments of first (root) end 23 and second (tip) end 24.
Load beam 20 is formed of a rigid, high strength polymer material such as nylon, or another suitably rigid polymer or composite material. Slots, holes or other coupling structures are provided to receive fasteners 28 in ends 30 of load beam 20, forming mechanical attachments to end blocks 16 and 18. As shown in
Depending on embodiment, turbine component 14 may comprise a rotor blade, stator vane, guide vane or strut component, with or without airfoil surfaces 22A-22D and the platform and shroud structures at first and second ends 23 and 24. One or more protective coatings may also be applied, for example a thermal barrier coating, an abrasive coating, a hard coating to protect against impacts, or a protective sheath. In further embodiments, turbine component 14 may be made from an aluminum or titanium alloy, or a composite material.
End blocks 16, 18 and load beam 20 may also be formed of different materials. Suitable materials include, but are not limited to, plastics and thermoplastics such as nylon, rigid fluoropolymers, acetal polymers, polyacetal plastics and polyformaldehyde plastics, including Teflon® and Delrin®, which are trade names of E.I. du Pont de Nemours & Co. of Wilmington, Del. Composite materials are also used, for example an adhesive matrix impregnated with a fibrous material such as Kevlar®, which is also available from DuPont, or a carbon fiber matt or other fibrous material embedded in a matrix substrate. Some of these materials also have suitably flexible and pliable forms for use in compliant mask 12.
As shown in
First and second end blocks 16 and 18 are attached to load beam 20 by fitting mechanical fasteners 28 into holes or slots 36, and adjusting pivots 26 to retain and seal turbine component 14 between end blocks 16 and 18. Compliant mask 12 is positioned between load beam 20 and turbine component 14, and load beam 20 seals compliant mask 12 against the selected surface of airfoil 22. Scratching, abrasion and other effects on turbine component 14 are reduced by the relatively soft polymer or composite composition of compliant mask 12, end blocks 16 and 18, and load beam 20.
Features 40 comprise film cooling holes and other surface or flow features, which are formed into or onto the different exterior regions and surfaces of airfoil 22, platform 38 and shroud 39. As illustrated in
In cooled embodiments, turbine component 14 may comprise a turbine blade, turbine vane or hot strut component configured for a high temperature or high pressure turbine or compressor section of a gas turbine engine. Alternatively, turbine component 14 comprises a blade, vane or strut component for a lower temperature or lower pressure compressor, or a component for the fan section of a turbofan engine.
Coating 41 comprises a wax coating or other removable material applied to selected surface of turbine component 14, for example by dipping fixture system 10 in a coating reservoir, or by pouring or brushing on the coating material. Coating 41 coats some or all of the uncovered regions of turbine component 14, including uncovered features 40 on unmasked surfaces of airfoil 22. Coating 41 is not applied to the regions of turbine component 14 covered by compliant mask 12, where mask 12 seals off particular features 40 to prevent contact with the coating material, leaving features 40 uncoated in the selected regions.
In the particular embodiment of
Turbine components 14 are generally high value, high precision parts, and individual testing is time intensive. In particular, misapplication of coating 41 reduces efficiency by increasing the error and rework rates, decreasing reliability and throughput, and raising individual part costs.
Fixture system 10 addresses these problems by utilizing compliant mask 12 to increase precision and repeatability of the coating process, improving quality and reliability while reducing manufacturing time and cost. Compliant mask 12 reduces the need for precision brushed-on coating methods by consistently covering the same features 40 in each selected region, even when other (unselected) features 40 are closely spaced. Faster dip and pour coating methods can also be employed, using a recycling reservoir to reduce waste and environmental impact. Dip and pour coating methods can also employ a wider range of coating materials, including natural waxes with different viscosities and lower melting temperatures, and polymer waxes and coatings with better adhesion, reduced thickness, and other desirable properties.
In this particular embodiment, selected features 40 remain uncoated along pressure surface 22A and trailing edge 22D, where they were covered by compliant mask 12. Coating 41 covers exposed features 40 along suction surface 22B and at leading edge 22C, in the regions that were not covered by compliant mask 12.
Coating 41 comprises a polymer material such as a natural or polymer wax, as described above, which is removable by heating turbine component 14. Soluble waxes and polymer coatings are also utilized, where coating 41 is removed by washing in water, or by application of a suitable chemical agent such as tolulene.
Use of removable coating 41 provides for more efficient flow testing and other selective processing of turbine component 14. In particular, some components 14 are subject to complex machining steps and coating processes, which can result in over or under drilling, or produce full or partial flow blockages.
Removable coating 41 allows these faults to be detected by coating and closing off particular features 40, while selected features 40 are covered by compliant mask 12. Mask 12 is then removed to measure the flow through or across selected features 40, and individual turbine components 14 are identified for additional inspection, repair, reprocessing or scrap based on the flow rate.
In particular, individual turbine components 14 are identified based on whether the flow falls within a nominal range, or is atypically high or low. Higher flow rates tend to indicate increased flow through over-drilled or over-machined features 40, while lower flow rates indicate decreased flow through under-drilled, under-machined or partially or completely blocked features 40. Fixture 10 and turbine component 14 are then heated or washed to remove coating 41, and the process is repeated in other regions of interest.
Use of a removable wax or polymer material reduces the melting point or solubility temperature at which coating 41 is removed, allowing for repeated testing with lower processing time and reduced risk of damage to fixture 10 and component 14. In particular, the removal temperature is substantially below the operating temperature of turbine component 14. The removal temperature is also substantially below the temperature at which melting, high-temperature oxidation and other phase transitions occur in particular turbine materials, including, but not limited to, nickel and cobalt alloys, superalloys, thermal barrier coatings, abrasive coatings, and aluminum, titanium and composite turbine blade and sheathing materials.
In some embodiments, the melting or wash removal temperature of coating 41 is less than 100° C., for example about 45° to about 85° C. In additional embodiments, coating 41 is removed at room temperature, for example about 20-30° C. or less, for example via a low-temperature wax melting process or by washing with water or a suitable chemical agent.
In the particular embodiment of
As shown in
While this invention has been described with reference to exemplary embodiments, it will be understood by those skilled in the art that various changes may be made and equivalents may be substituted without departing from the scope of the invention. In addition, modifications may be made to adapt particular situations or materials to the teachings of the invention, without departing from the essential scope thereof. The invention is not limited to the particular embodiments disclosed herein, but includes all embodiments falling within the scope of the appended claims.
