Active data acquisition management system in a gas turbine engine

Description

TECHNICAL FIELD

[0001] The present disclosure relates generally to aircraft sensor data acquisition, and more specifically to a system and method for actively controlling a data acquisition rate of an aircraft system sensor.BACKGROUND

[0002] Gas turbine engines, such as those utilized in commercial and military aircraft, include a compressor section that compresses air, a combustor section in which the compressed air is mixed with a fuel and ignited, and a turbine section across which the resultant combustion products are expanded. The expansion of the combustion products drives the turbine section to rotate. As the turbine section is connected to the compressor section via a shaft, the rotation of the turbine section further drives the compressor section to rotate. In some examples, a fan is also connected to the shaft and is driven to rotate via rotation of the turbine as well.

[0003] During operation of the gas turbine engine, it can be desirable for an engine controller such as a full authority digital engine controller (FADEC) to monitor one or more parameters of the engine operation. The parameters can be monitored using sensors configured to detect the parameters. The data from the sensors is read at an acquisition rate (alternately referred to as a sampling rate). In some instances, such as when certain vibrational events occur, it can be desirable to read data from the sensor more frequently than the base data acquisition rate.

[0004] US 2015 / 330869 A1 and US 2013 / 332025 A1 relate to systems and related methods for determining an aircraft engine's current performance by collecting real-time aircraft engine data and processing the engine data for engine maintenance or other analysis.SUMMARY OF THE INVENTION

[0005] According to an aspect of the invention an aircraft is as defined in claim 1.

[0006] These and other features of the present invention can be best understood from the following specification and drawings, the following of which is a brief description.BRIEF DESCRIPTION OF THE DRAWINGS

[0007] Figure 1 illustrates a schematic view of an exemplary engine. Figure 2 schematically illustrates a notional system configuration for a data acquisition management system. Figure 3 schematically illustrates a data acquisition rate chart during an exemplary aircraft operation. Figure 4 illustrates an exemplary method of operating a data acquisition management system. Figure 5 schematically illustrates an aircraft including a data acquisition management system expanded beyond an aircraft engine. DETAILED DESCRIPTION

[0008] Figure 1 schematically illustrates a gas turbine engine 20. The gas turbine engine 20 is disclosed herein as a two-spool turbofan that generally incorporates a fan section 22, a compressor section 24, a combustor section 26 and a turbine section 28. The fan section 22 drives air along a bypass flow path B in a bypass duct defined within a nacelle 15, and also drives air along a core flow path C for compression and communication into the combustor section 26 then expansion through the turbine section 28. Although depicted as a two-spool turbofan gas turbine engine in the disclosed non-limiting embodiment, it should be understood that the concepts described herein are not limited to use with two-spool turbofans as the teachings may be applied to other types of turbine engines including three-spool architectures.

[0009] 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.

[0010] The low speed spool 30 generally includes an inner shaft 40 that interconnects a fan 42, a first (or low) pressure compressor 44 and a first (or 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 second (or high) pressure compressor 52 and a second (or 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. A mid-turbine frame 57 of the engine static structure 36 is arranged generally between the high pressure turbine 54 and the low pressure turbine 46. The mid-turbine frame 57 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.

[0011] 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 mid-turbine frame 57 includes airfoils 59 which are in the core airflow path C. 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.

[0012] 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 invention is applicable to other gas turbine engines including direct drive turbofans.

[0013] 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 (10,668 meters). The flight condition of 0.8 Mach and 35,000 ft (10,668 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 meters / second).

[0014] The exemplary gas turbine engine 20 includes a vibrational sensor 90 positioned in or on the exterior casing of the engine 20. The vibrational sensor 90 is configured to detect an overall magnitude of vibrations within the engine 20 at any given time. Due to its position, the isolation ability of the vibration sensor 90 is limited to distinct shafts and has limited capability to distinguish between related components. The vibrational sensor 90 is connected to a controller 92, such as a full authority digital engine controller (FADEC). In alternative examples, the controller 92 can be a general aircraft controller, a dedicated sensor controller, or any other type of controller configured to receive and interpret signals from the vibrational sensor 90, and from one or more micro-electromechanical-systems (MEMS).

[0015] Positioned local to each bearing system 38 is a micro-electro-mechanical-system (MEMS) 94. As used herein a component, such as the MEMS 94, is "local to" a system when the component is either positioned on or in the component to which it is local, or is in a position where vibrational forces on the component are dominated by the vibrational forces on the system to which the component is local. Each of the MEMS 94 is connected to the controller 92 either via a wireless connection, or a direct wired connection depending on the particular characteristics of the specific engine 20. While illustrated as positioned local to each of the bearing systems 38 of the engine 20, it is appreciated that the MEMS 94 can be positioned local to any number of different components including, but not limited to, an accessory gearbox, a fuel pump, an oil pump, airframe components, or any other line replaceable component.

[0016] As used herein MEMS 94 refer to microscopic, or similarly small, electromechanical devices configured to detect the occurrence of an event and transmit a signal when the occurrence is detected. In some examples, the MEMS 94 are made up of components ranging from 1 to 100 micrometers in size and the MEMS 94 will typically have an length in its longest dimension ranging from 20 micrometers to 1 millimeter. In some examples the MEMS 94 include a processor and micro sensors configured to sense and interpret the immediately surrounding environment.

[0017] Each of the MEMS 94 is configured to trigger when vibrational forces on the MEMS 94 exceed a predetermined threshold. When triggered, the MEMS 94 provides a signal to the controller 92. The controller 92, in turn, increases a data acquisition rate of the vibrational sensor 90, such that the increased data acquisition rate compensates for a structural response mode of the particular system that the MEMS 94 is local to. The data acquisition rate remains adjusted for a predetermined duration. In some examples the predetermined duration is at least the length of an expected vibrational event. In further examples, the predetermined duration can be within the range of 100-400 ms. Once the predetermined duration has elapsed, the data acquisition rate is decreased to the original rate, and normal data collection from the sensor 90 is resumed.

[0018] While Figure 1 illustrates a single system sensor 90, it is appreciated that multiple system sensors 90 can be connected and operate in the same manner. In such an example, each of the MEMS 94 each correspond to one or more system sensor 90.

[0019] With continued reference to Figure 1, Figure 2 schematically illustrates an exemplary sensor system 100 including an aircraft system sensor 110, and multiple MEMS 120, 122, 124, 126, 128. Each of the MEMS 120, 122, 124, 126, 128 is positioned within an aircraft system, such as a gas turbine engine, and is disposed local to a corresponding component within the aircraft system. Each of the MEMS 120, 122, 124, 126, 128 and the system sensor 110 are connected to a controller 130. The controller 130 includes a MEMS module 132 configured to receive and interpret signals from each of the MEMS 120, 122, 124, 126, 128. The MEMS module 132 can be a software module, stored in a memory of the controller 130, or a hardware portion of the controller 130, depending on the needs and environment of the specific aircraft system, and the controller 130.

[0020] Also included within the controller 130 is a data acquisition module 134. The data acquisition module 134 acquires sensor data from the aircraft system sensor 110 and provides the acquired data to a remainder of the controller 130 systems. The data acquisition module 134 controls a data acquisition rate of the data from the aircraft system sensor 110, and operates at a default base data acquisition rate.

[0021] Stored within the MEMS module 132, or the data acquisition module 134, is a set of data acquisition rates, with at least one data acquisition rate corresponding to each MEMS 120, 122, 124, 126, 128. Each of the data acquisition rates compensates for the structural response modes of the components to which the corresponding MEMS 120, 122, 124, 126, 128 is local, and allows for a sufficient data acquisition rate from the aircraft system sensor 110 when the corresponding MEMS 120, 122, 124, 126, 128 is triggered. In some examples, multiple MEMS 120, 122, 124, 126, 128 can correspond to the same data acquisition rate.

[0022] In some cases, multiple MEMS 120, 122, 124, 126, 128 can be triggered simultaneously, either due to a single event triggering multiple MEMS 120, 122, 124, 126, 128, or due to independent events occurring simultaneously. When multiple MEMS 120, 122, 124, 126, 128 trigger at the same time, the controller 130 causes the data acquisition module 134 to adjust the data acquisition rate to be the highest data acquisition rate corresponding to one of the triggered MEMS 120, 122, 124, 126, 128 while keeping track of all triggered MEMS 120, 122, 124, 126, 128. By adjusting to the highest data acquisition rate, the controller 130 ensures that the data is acquired at a sufficient rate to compensate for the structural vibratory mode of each of the components to which the triggered MEMS 120, 122, 124, 126, 128 is local.

[0023] With continued reference to Figures 1 and 2, Figure 3 illustrates a chart 200 of vibrations at an exemplary MEMS 120, 122, 124, 126, 128. The chart 200 includes axis 210, 212 corresponding to vibration of a component to which the exemplary MEMS is local (axis 210) and corresponding to time (axis 212). The vibration of the component with respect to time is illustrated via a line 230. Each time the vibrations exceed a threshold 232, the MEMS is triggered and a signal is sent to the controller from the MEMS. The signal causes the controller to increase the data acquisition rate for a system sensor to the data acquisition rate corresponding to the triggered MEMS for a predefined time window 240. In some examples, the predefined time window 240 for every MEMS in a given system can be identical. In alternative examples, the predefined time window can be unique to each MEMS, and can be selected based on an expected event type at the component corresponding to the MEMS. In yet further examples, the MEMS can continuously trigger while the threshold is exceeded, and the data acquisition rate is increased for a duration equal to the predefined window plus the length that the threshold is exceeded.

[0024] With continued reference to Figures 1-3, Figure 4 illustrates a method for operating the system 100 of Figure 2. During operation of the aircraft, an initial triggering event occurs local to an engine component at a "Triggering Event Occurs" step 310. In response to the occurrence of the triggering event, a MEMS local to the engine component registers the event in a "Local MEMS Registers Event" step 320. The local MEMS responds to the event by transmitting a signal to a controller, such as a FADEC, in a "MEMS Signals FADEC" step 330. The signal is, in some examples, a simple indication that the MEMS originating the signal registered an event. In other examples, additional data such as a magnitude of the event can be transmitted as well. When the controller receives the signal, the controller adjusts a data acquisition rate (sampling rate) of a corresponding system sensor that senses a parameter of the engine system in a "FADEC Adjusts Sampling Rate" step 340.

[0025] With continued reference to Figures 1-4, Figure 5 illustrates an exemplary aircraft 400 including MEMS 410 disposed at additional sensitive systems outside of the gas turbine engines 402. Each of the MEMS 410 is connected to a corresponding controller 420, and the corresponding controller 420 is configured to increase a sampling rate of a corresponding sensor 430 using a similar process to the process described above with regards to the gas turbine engine. In this way, the system and process of Figures 1-4 can be expanded to additional systems and structures within the aircraft 400. The additional structures can include landing gears 442, wing ailerons 444, tail systems 446, and wing joints 448. Further, while described above with specific relationship to vibrational triggers and a vibrational system sensor, one of skill in the art will appreciate that the system described herein can be utilized in conjunction with any other engine or system parameter sensor and is not limited to vibrational sensors.

[0026] Although an embodiment of this invention has been disclosed, a worker of ordinary skill in this art would recognize that certain modifications would come within the scope of this invention, as defined by the appended claims.

[0027] For that reason, the following claims should be studied to determine the true scope and content of this invention.

Claims

1. An aircraft (400) comprising: at least one overall system sensor (90; 110; 430) communicatively coupled to a controller (92; 130; 420); and a plurality of micro electro-mechanical-systems (MEMS) (120, 122, 124, 126, 128; 410) disposed throughout the aircraft (400), each of the MEMS (120, 122, 124, 126, 128; 410) being local to a corresponding aircraft component, and being communicatively coupled to the controller (92; 130; 420), the controller (92; 130; 420) being configured to define a sampling rate of the at least one overall system sensor (90; 110; 430), wherein: the controller (92; 130; 420) is further configured to increase the sampling rate of the overall system sensor (90; 110; 430) to an increased sampling rate in response to receiving a signal from one of the MEMS in the plurality of MEMS (120, 122, 124, 126, 128; 410); and the increased sampling rate corresponds to a parameter of the component to which each one of the MEMS (120, 122, 124, 126, 128; 410) is local, characterised in that: the corresponding aircraft component comprises a landing gear (442), a wing aileron (444), a tail system (446), and / or a wing joint (448).

2. The aircraft of claim 1, wherein the at least one overall system sensor (90; 110; 430) includes a vibration sensor.

3. The aircraft of claim 2, wherein each of the MEMS (120, 122, 124, 126, 128; 410) is configured to provide said signal to the controller (92; 130; 420) when vibrational forces on the MEMS exceed a predetermined threshold.

4. The aircraft of any preceding claim, wherein the controller (92; 130; 420) is a full authority digital engine controller (FADEC).