Patent Application
20140064976
This disclosure relates to a rotor for a gas turbine engine. More particularly, the disclosure relates to fillet geometry for the rotor.
A typical gas turbine engine includes multiple of compressor stages upstream from a combustor section. A turbine section is arranged downstream from the combustor. In one example configuration, at least one end of a compressor rotor is secured to a shaft by a hub. The hub engages a hub engagement feature on the rotor to secure the rotor relative to the shaft. Typically, a nut is received on the correspondingly threaded portion of the shaft and applies a clamping load to the rotor via the hub.
The hub engagement feature on the rotor is provided by first and second annular surfaces that are at a right angle to one another. A fillet joins the first and second surfaces, which are arranged tangentially relative to the fillet.
In one exemplary embodiment, a rotor for a gas turbine engine includes an annular structure having a blade slot. A hub engagement feature is provided on the annular structure. The hub engagement feature includes first and second surfaces transverse to one another and joined by a fillet that is recessed with respect to the first and second surfaces.
In a further embodiment of any of the above, the first and second surfaces are normal to one another.
In a further embodiment of any of the above, the fillet includes a peened surface.
In a further embodiment of any of the above, the annular structure is constructed from a nickel alloy.
In a further embodiment of any of the above, the first and second surfaces are provided by ground surfaces.
In a further embodiment of any of the above, the first and second surfaces are non-tangential to the fillet.
In another exemplary embodiment, a rotor assembly for a gas turbine engine includes a shaft. The rotor assembly includes a rotor that supports a blade and includes a hub engagement feature. The hub engagement feature includes first and second rotor surfaces transverse to one another and is joined by a fillet that is recessed with respect to the first and second rotor surfaces. A hub is supported on the shaft and engages the hub engagement feature.
In a further embodiment of any of the above, the rotor assembly includes a nut secured to the shaft and applies a clamping load to the hub engagement feature via the hub.
In a further embodiment of any of the above, the hub includes first and second hub surfaces respectively engaging the first and second rotor surfaces under the clamping load. The first and second hub surfaces are spaced from the fillet.
In a further embodiment of any of the above, the first and second rotor surfaces are normal to one another.
In a further embodiment of any of the above, the fillet includes a peened surface.
In a further embodiment of any of the above, the rotor is constructed from a nickel alloy.
In a further embodiment of any of the above, the first and second rotor surfaces are provided by ground surfaces.
In a further embodiment of any of the above, the first and second rotor surfaces are non-tangential to the fillet.
In another exemplary embodiment, a method of manufacturing a rotor includes the steps of machining an annular hub engagement feature into a rotor. The hub engagement feature includes first and second surfaces transverse to one another and is joined by a fillet that is recessed with respect to the first and second surfaces. The method includes the step of peening the fillet, and grinding the first and second surfaces.
In a further embodiment of any of the above, the grinding step is performed after the peening step.
In a further embodiment of any of the above, the first and second surfaces are non-tangential to the fillet.
The disclosure can be further understood by reference to the following detailed description when considered in connection with the accompanying drawings wherein:
Although the disclosed non-limiting embodiment depicts a turbofan gas turbine engine, it should be understood that the concepts described herein are not limited to use with turbofans as the teachings may be applied to other types of turbine engines; for example a turbine engine including a three-spool architecture in which three spools concentrically rotate about a common axis and where a low spool enables a low pressure turbine to drive a fan via a gearbox, an intermediate spool that enables an intermediate pressure turbine to drive a first compressor of the compressor section, and a high spool that enables a high pressure turbine to drive a high pressure compressor of the compressor section.
The example 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.
The low speed spool 30 generally includes an inner shaft 40 that connects a fan 42 and a low pressure (or first) compressor section 44 to a low pressure (or first) turbine section 46. The inner shaft 40 drives the fan 42 through a speed change device, such 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 (or second) compressor section 52 and a high pressure (or second) turbine section 54. The inner shaft 40 and the outer shaft 50 are concentric and rotate via the bearing systems 38 about the engine central longitudinal axis A.
A combustor 56 is arranged between the high pressure compressor 52 and the high pressure turbine 54. In one example, the high pressure turbine 54 includes at least two stages to provide a double stage high pressure turbine 54. In another example, the high pressure turbine 54 includes only a single stage. As used herein, a “high pressure” compressor or turbine experiences a higher pressure than a corresponding “low pressure” compressor or turbine.
The core airflow C 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 mid-turbine frame 57 includes airfoils 59 which are in the core airflow path. The turbines 46, 54 rotationally drive the respective low speed spool 30 and high speed spool 32 in response to the expansion.
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 ten (10), the geared architecture 48 is an epicyclic gear train, such as a star 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 5. 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 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. It should be understood, however, that the above parameters are only exemplary of one embodiment of a geared architecture engine and that the present invention 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. The flight condition of 0.8 Mach and 35,000 ft, 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 per hour divided by lbf of thrust the engine produces at that minimum point. “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 [(Tambient deg R)/518.7)̂0.5]. The “Low corrected fan tip speed” as disclosed herein according to one non-limiting embodiment is less than about 1150 ft/second.
Referring to
The rotor 60, which is constructed from a nickel alloy, includes one or more slots 65 that support multiple circumferentially spaced blades 64. It should be understood, however, that the blades 64 may be integrated with the rotor 60.
A first air seal 66 is supported by the rotor 60, in the example, which cooperates with a second air seal 68 supported by the engine static structure 36. The first air seal 66 may integral with or separate from the rotor 62.
The hub engagement feature 71 is shown in more detail in
Referring to
Although an example embodiment has been disclosed, a worker of ordinary skill in this art would recognize that certain modifications would come within the scope of the claims. For that reason, the following claims should be studied to determine their true scope and content.
