Patent Application
20190039167
This disclosure is related to the field of bladed rotors generally, and more specifically to integrally bladed rotors.
Bladed rotors such as impellers, blisks, etc. are employed in gas turbines and other machines. The design, construction and materials of bladed rotors often dictate operating limits for the turbines in which they are employed. Extensive efforts have been made over the years to develop new alloys, new fabrication techniques, and new component designs which permit operation of these rotors at higher operating temperatures and/or lead to lighter weight, longer lived components, with all their attendant advantages.
The fan, turbine, and compressor sections of gas turbine engines include one or more circumferentially extending rows or stages of airfoils, commonly called rotor blades, which are axially spaced between rows or stages of fixed airfoils (stator vanes). The rotor blades are connected to and extend radially outwardly from a rotor disk. During operation the centrifugal loads generated by the rotational action of the rotor blades must be carried by the rotor disk within acceptable stress limits.
In one type of a conventional bladed rotor assembly, the rotor disk has a plurality of slots around its radially outer periphery. The blades may comprise a root, a platform, and an airfoil. The platform has opposite facing surfaces. The root attaches to the slot in the disk and the airfoil extends radially out from the disk. The slots and the roots have complementary shapes, typically either a dove tail or a fir tree. The root mates with the slot and the blade extends radially outward therefrom. This type of rotor assembly is relatively heavy because the slots are cut through the rim of the disk creating what is called a “dead rim” where the metal between the slot can pull on the disk with well over 10,000 g's and fir tree or dovetail mating structures likewise do not contribute to sustaining the disk's centrifugal load and also pulls with the same 10,000 g load, thereby necessitating that the rotor disk be sufficiently sturdy, and thus heavy, in order to accommodate the stresses resulting from the heavy blade attachment area.
Alternatively, the blades may be secured by bonding or welding, to the rotor disc to thereby form an integrally bladed rotor assembly (IBR). A major advantage of an integrally bladed rotor assembly is that there is often no need for an extended blade root or a blade platform. The airfoil may be secured directly to the radially outer periphery of the rotor disk. The absence of an extended root and a blade platform results in a blade that is lighter than a conventional blade. A lighter blade enables the use of a less rigid and lighter rotor disk, in which case the integrally bladed rotor assembly is overall much lighter than a conventional bladed rotor assembly.
A method of making an integrally bladed rotor is disclosed. The method comprises providing a rotor disk comprising a radially outer rim surface that includes a recessed area thereon. A blade comprising an airfoil and a base is positioned such that a base surface is in contact with the recessed area with a gap between the base surface and the recessed area at a perimeter of the recessed area. Heat, pressure, and motion between the blade and the rotor disk are applied to friction weld the base surface to the recessed area.
In some embodiments, the method comprises ejecting friction welding flash through the gap.
In any one or combination of the foregoing embodiments, positioning of the blade comprises contacting the base surface to the recessed area at a low portion of the recessed area.
In any one or combination of the foregoing embodiments, the gap extends from a contact position between the base surface and the recessed area at the recessed area low portion, to the recessed area perimeter.
In any one or combination of the foregoing embodiments, the gap distance between the base surface and the recessed area increases from the recessed area low portion to the recessed area perimeter.
In any one or combination of the foregoing embodiments, the recessed area comprises a linear extended groove in the outer rim surface corresponding to a chord of the airfoil where the airfoil meets the disk.
In any one or combination of the foregoing embodiments, the base surface includes a linear-extending central apex or a linear extending planar surface, and wherein positioning the blade comprises contacting the linear-extending central apex or linear extending planar surface or the with a linear-extending low portion in the groove.
In any one or combination of the foregoing embodiments, the blade further comprises positioning the base surface at an angle of inclination with respect to the recessed area surface of at least 12°.
In any one or combination of the foregoing embodiments, the base surface includes surfaces extending perpendicular to the linear-extending central apex at an overall angle of inclination of less than 170°.
In any one or combination of the foregoing embodiments, the linear-extending central apex comprises a linear-extending sharp edge, and the linear-extending low portion comprises a linear-extending inverse apex sharp edge.
In any one or combination of the foregoing embodiments, the gap distance between the base surface and the recessed area increases from the linear-extending groove low portion to the linear-extending groove perimeter.
In any one or combination of the foregoing embodiments, the base surface and the recessed area each comprise flat rectilinear surfaces or curved surfaces.
In any one or combination of the foregoing embodiments, the blade base includes a peripheral sink surface perpendicular to the outer rim surface.
In any one or combination of the foregoing embodiments, the airfoil is a solid contiguous structure.
In any one or combination of the foregoing embodiments, each of the blade surface and the recessed area comprises a titanium alloy.
Also disclosed is an integrally bladed rotor made by the method of any one or combination of the foregoing embodiments.
Also disclosed is a gas turbine engine comprising an integrally bladed rotor made by the method of any one or combination of the foregoing embodiments.
Also disclosed is a gas turbine engine comprising a fan that comprises an integrally bladed rotor made by the method of any one or combination of the foregoing embodiments.
The following descriptions should not be considered limiting in any way. With reference to the accompanying drawings, in which like elements are numbered alike:
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.
As mentioned above, a method is disclosed for friction welding a blade onto rotor to form an integrally bladed rotor. One application for such an integrally bladed rotor is on a gas turbine engine, for example as a bladed fan. Other applications include compressor rotors or turbine rotors.
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).
With reference now to
As mentioned above, the blades 104 are welded to the disk 102 by the application of heat (e.g., heat generated by friction), pressure, and motion, such as by linear friction welding (LFW). As further shown in
As shown in
The result of linear friction welding of the blade 104 to the disk 102 is shown in
Various materials can be used for the blade base and disk rim. In some embodiments, both the base and the rim comprise nickel alloys, which can be the same alloy for both parts or different alloys. For example, in some embodiments, the blade can be formed from an alloy such as PWA 1228 or equivalent and the disk can be formed from an alloy such as PWA 1215, PWA 1214, or equivalent. In some embodiments, the blade and the base can be formed from the same material, such as PWA 1215 or equivalent. In some embodiments, the gap configuration described herein can provide a technical effect or promoting ejection or removal of flash from the weld zone. Incomplete or irregular ejection or removal of flash from the weld zone can cause defects in the weld joint. Prior art attempts at managing flash removal from friction welding of integrally bladed rotors have involved the formation of standups on the disk rim for LFW attachment of the blades (see U.S. Pat. No. 7,419,082). However, that approach involves machining away of metal from the disk rim to form the standups. Forming the standups by cutting away the metal between each removes superior metal with fine grain structure properties which tends to exist in disk outer diameter areas, exposing coarser grain material that is less resistant to stress from vibration. The embodiments disclosed herein can provide a technical effect of managing flash ejection or removal of a friction weld attachment of a blade to a rotor while promoting formation of a region of fine grain structure metal in the integrally bladed rotor structure.
It should be pointed out that the above embodiments are provided as examples, which can of course subject to variation and modification by the skilled person. For example,
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. For example, “about” can include a range of ±8% or 5%, or 2% of a given value.
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.
