Patent Application
20250092798
Exemplary embodiments of the present disclosure pertain to the art of gas turbine engines, and more particularly to data collection during operation of a gas turbine engine.
Testing of operation of gas turbine engines requires the collection of data regarding various components and locations on the gas turbine engine. This data collection often includes the measurement of clearances between components during operation of the gas turbine engine, or other features. The data is typically gathered via probes or other sensors arrayed about the engine and secured to rotationally fixed components of the engine. Many such sensors are secured in place by a threaded connection to the fixed structure, and it is desired to rotationally secure the sensors in place thus preventing their movement or “backing out” during operation of the engine. Further, the sensors may have brittle or otherwise delicate components, so the rotational securing must be accomplished without damaging the sensor.
One such location for the measurement of clearances is radial proximity between stator segments and a rotating hub.
In one exemplary embodiment, a component assembly of a gas turbine engine includes a component of a gas turbine engine, and an instrumentation probe installed to the component. The instrumentation probe includes a sensor body extending through a component wall of the component, and a threaded fastener installed onto a complimentary thread of the sensor body to retain the sensor body at the component. The threaded fastener is retained to the sensor body via deforming a threaded interface between the threaded fastener and the sensor body.
Additionally or alternatively, in this or other embodiments the deformation is formed via application of a vibra-stake tool to the threaded interface.
Additionally or alternatively, in this or other embodiments the deforming is at one or more of the threaded fastener and the sensor body.
Additionally or alternatively, in this or other embodiments the threaded fastener is a nut.
Additionally or alternatively, in this or other embodiments a probe housing at least partially encloses the sensor body and the threaded fastener.
Additionally or alternatively, in this or other embodiments a probe lead is operably connected to the sensor body and extends through a lead opening in the probe housing.
Additionally or alternatively, in this or other embodiments the deformation is at one or more circumferential locations of the threaded interface.
Additionally or alternatively, in this or other embodiments one or more insulators are positioned radially between the probe body and the component wall.
In another exemplary embodiment, a method of installing an instrumentation probe to a component of a gas turbine engine includes installing a sensor assembly of the instrumentation probe through a component wall of the component, securing a threaded fastener to a complimentary thread of the sensor assembly to retain the sensor body at the component wall, and deforming a threaded interface between the threaded fastener and the sensor assembly to prevent rotation of the threaded fastener relative to the sensor assembly.
Additionally or alternatively, in this or other embodiments the deformation is achieved via application of a vibra-stake tool to the threaded interface.
Additionally or alternatively, in this or other embodiments a housing is installed to at least partially enclose the threaded interface.
Additionally or alternatively, in this or other embodiments a cover is installed onto the housing after deformation of the threaded interface.
Additionally or alternatively, in this or other embodiments a probe lead operably connected to the sensor assembly is routed through a lead opening in the housing.
Additionally or alternatively, in this or other embodiments the threaded fastener is a nut.
Additionally or alternatively, in this or other embodiments the threaded interface is deformed at one or more circumferential locations about the threaded interface.
In yet another exemplary embodiment, a stator assembly of a gas turbine engine includes a stator having one or more stator vanes, and an instrumentation probe installed to the stator. The instrumentation probe includes a sensor assembly extending through a stator wall of the stator, and a threaded fastener installed onto a complimentary thread of the sensor assembly to retain the sensor assembly at the stator. The threaded fastener is retained to the sensor assembly via deforming a threaded interface between the threaded fastener and the sensor assembly.
Additionally or alternatively, in this or other embodiments the deformation is formed via application of a vibra-stake tool to the threaded interface.
Additionally or alternatively, in this or other embodiments the threaded fastener is a nut.
Additionally or alternatively, in this or other embodiments a probe housing at least partially encloses the sensor assembly and the threaded fastener.
Additionally or alternatively, in this or other embodiments a probe lead is operably connected to the sensor body and extends through a lead opening in the probe housing.
The following descriptions should not be considered limiting in any way. With reference to the accompanying drawings, 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.
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).
Referring now to
Referring now to
It is desired to ensure that the clamping force is maintained on the stator 60 so that the position of the sensor body 66 is maintained during assembly and operation of the gas turbine engine 20. To accomplish this, the position of the nut 78 relative to the threaded end 80 is maintained by locking the relative position via a vibra-stake operation applied across the interface between the threaded end 80 and the nut 78 at one or more circumferential locations, for example, four circumferential locations as illustrated in
Another embodiment of instrumentation probe 62 is illustrated in
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.
This invention was made with Government support awarded by the United States. The Government has certain rights in this invention.
