Patent Application
20160258440
The present disclosure relates generally to gas turbine engines, and more specifically to bladed rotors used in gas turbine engines.
Gas turbine engines are used to power aircraft, watercraft, power generators, pumps, and the like. Gas turbine engines operate by compressing atmospheric air, burning fuel with the compressed air, and then removing work from hot high-pressure air produced by combustion of the fuel in the air. Rows of rotating blades and non-rotating vanes are used to compress the air and then to remove work from the high-pressure air produced by combustion. Each blade and vane has an airfoil that interacts with the gasses as they pass through the engine.
Airfoils have natural vibration modes of increasing frequency and complexity of the mode shape. The simplest and lowest frequency modes are typically referred to as the first bending mode, the second bending mode, the third bending mode, and the first torsion mode. The first bending mode is a motion normal to the working surface of an airfoil in which the entire span of the airfoil moves in the same direction. The second bending mode is similar to the first bending mode, but with a change in the sense of the motion somewhere along the span of the airfoil, so that the upper and lower portions of the airfoil move in opposite directions. The third bending mode is similar to the second bending mode, but with two changes in the sense of the motion somewhere along the span of the airfoil. The first torsion mode is a twisting motion around an elastic axis, which is parallel to the span of the airfoil, in which the entire span of the airfoil, on each side of the elastic axis, moves in the same direction.
Blades are subject to destructive vibrations induced by unsteady interaction of the airfoils of those blades with gasses passing through a gas turbine engine. One type of vibration is known as flutter, which is an aero-elastic instability resulting from the interaction of the flow over the airfoils of the blades and the blades' natural vibration tendencies. The lowest frequency vibration modes, the first bending mode and the first torsion mode, are often the vibration modes that are susceptible to flutter.
When flutter occurs, the unsteady aerodynamic forces on the blade, due to its vibration, add energy to the vibration, causing the vibration amplitude to increase. The vibration amplitude can become large enough to cause damage to a blade. Another type of vibration is known as forced response, which is an aero-elastic response to inlet distortion or wakes from upstream airfoils, struts, or any other flow obstruction. The operable range, in terms of pressure rise and flow rate, of turbomachinery can sometimes be restricted by flutter or forced response phenomena.
Flutter instability can degrade the performance of the rotary compressor and may also lead to fatigue failure or other permanent damage to the compressor. One result of the flutter instability can be blade deformation and/or blade fatigue failure. Thus, it is desirable to avoid rotary compressor blade motion that causes flutter.
The present disclosure may comprise one or more of the following features and combinations thereof.
A rotor for use in a gas turbine engine may include a central wheel, and a plurality of blades having blade tips. The central wheel may be arranged around a central axis. The plurality of fan blades extend outward from the central wheel in a radial direction away from the central axis.
In illustrative embodiments, a flutter control system for a turbomachine fan includes a plurality of optical tip timing sensors located in a fan case of the turbomachine and configured to sense the passing of blade tips of a fan of the turbomachine. A controller is operably connected to the plurality of optical tip timing sensors. A series of nozzles are located in the fan case and directed at the edges of the blades. High pressure air off of the compressor is directly injected into the blade row to alter unsteady pressure. A nozzle actuator is operably connected to the controller, such that the nozzle actuator selectively actuates the nozzles directed at the edges of the blades in response to data from the plurality of optical tip timing sensors indicating flutter or near flutter conditions. Data from the plurality of optical tip timing sensors is compared to a threshold value and the nozzles are actuated based on the comparison to dampen flutter of the plurality of fan blades.
In illustrative embodiments, optical tip timing sensors sense vibrations produced by a rotating blade and generate flutter signals that are a function of the sensed vibration. The flutter signals are transmitted to the processor in the controller. The processor generates a control signal based on the flutter signals and transmits the control signal to a nozzle actuator for controlling the position of the actuator, thereby modulating the discharge of high pressure compressor gases from the nozzles.
In illustrative embodiments, a third aspect of the disclosure is drawn to a method for reducing flutter instability wherein the steps of the method are stored on a computer-readable medium and comprise a method for reducing instability of a fan blade stored on a computer-readable medium comprising, generating a substantially parabolic flutter boundary curve representing flutter parameters of the fan blade, sensing flutter vibrations of the compressor, calculating a differential quantity representative of the difference between the flutter boundary curve and the operating mode, comparing the flutter vibrations to the differential quantity, operating the actuator to permit the flow of high pressure compressor gasses through the nozzles when the magnitude of the flutter vibration is greater than the differential quantity and monitoring the relationship of the magnitude of the flutter vibration and the differential quantity; and discontinuing the flow of high pressure compressor gasses through the nozzles when the flutter vibration is less than the differential quantity.
These and other features of the present disclosure will become more apparent from the following description of the illustrative embodiments.
For the purposes of promoting an understanding of the principles of the disclosure, reference will now be made to a number of illustrative embodiments illustrated in the drawings and specific language will be used to describe the same.
An illustrative aerospace gas turbine engine 100 includes a fan assembly 110 adapted to accelerate/blow air so that the air provides thrust for moving an aircraft as shown in
The fan rotor 10 includes a central fan wheel 12, a plurality of fan blades 14, and a spinner 16 as shown, for example, in
The fan rotor 10 is illustratively mounted to a turbine engine core 120 to be rotated by the engine core 120 as suggested, for example, in
Fan blade 14 illustratively includes an airfoil 30, a root 32, and a platform 34, as shown in
The fan blade 14 has a notional first bend mode node line 36 that extends axially the airfoil 30 from a leading edge 31 to a trailing edge 33 of the airfoil 30 adjacent to the platform 34 as shown in
Fan blade 14 also has notional third bend mode node lines (not shown) that extend axially across the airfoil 30 from the leading edge 31 to the trailing edge 33 of the airfoil 30 and that are spaced apart from the platform 34. Fan blade 14 further has a notional first torsion node line 39 that extends radially along the airfoil 30. Generally, the first, second, and third bend modes along with the first torsion mode make up low order modes that affect fan blade 14.
Fan blades 14 have a high aspect ratio (e.g. tall relative to chord and thickness) and are prone to easily vibrate under certain conditions. Not all fan blades 14 flutter but if a fan blade does flutter, it can cause high vibratory stresses and may result in fatigue failure. There is variability in manufacturing which affects the mode shape and frequency but the bigger unknown is the aerodynamic damping. When designing a fan blade, the aerodynamic damping is either estimated or based on test data from previous designs which may be similar. Blade tip timing (BTT) gives near real time information on fan blade 14 response. An asynchronous response may indicate flutter activity and by altering the phase of the surface unsteady pressure with respect to the airfoil motion, flutter may be inhibited. To control flutter, the twisting and bending motions must be measured along a section of the wing containing the leading and trailing edge control surfaces.
The flutter control system 200 includes a plurality of optical tip timing sensors 212 located in a fan case 112 of a gas turbine engine 100. Optical tip timing sensors 212 are located to observe arrival timing of a plurality of fan blades 14 fixed to a fan rotor 10 as the plurality of fan blades 14 rotate about central axis 11. In one example, optical tip timing sensors 212 are located in the fan case 112 to monitor passing of a leading edge 31, trailing edge 33, and mid-chord 35 of the plurality of fan blades 14. The optical tip timing sensors 212 monitor the leading edge 31 and trailing edge 33 utilized to determine fan blade 14 twist.
Optical tip timing sensors 212 are installed at the leading edge and sometimes at the trailing edge. For blade flutter, a minimum of three optical tip timing sensors 212 would be installed in each plane. If two planes are used, installation may be accomplished on leading edge or trailing edge alone. Optical tip timing sensors 212 would be in close proximity in angular positioning and would not be equally spaced around the circumference of the fan. For example, the optical tip timing sensors 212 may be arranged within 180 degrees or less of the fan circumference, within 90 degrees or less of the fan circumference, within 45 degrees or less of the fan circumference, within 30 degrees or less of the fan circumference, within 15 degrees or less of the fan circumference, or within 10 degrees or less of the fan circumference. This allows for collection of more tip passing data and correlation and/or verification of data.
Together, the information from the optical tip timing sensors 212 is communicated to a digital controller 130. The digital controller 130 compares the passing timing of the fan blades 14 to a threshold, to determine if a fan blade 14 is approaching a flutter condition or is actively fluttering. The digital controller 130 is sensing asynchronous vibration on the fan blade 14. Based on the comparison, the digital controller 130 sends commands to a fan nozzle actuator 132. The fan nozzle actuators 132 direct high pressure air from the compressor to exit through one or more fan nozzles 136 into the fan blade row in order to alter the surface unsteady pressure so that it is out of phase with the blade motion, increase the aerodynamic damping, increase the stall margin and make it more difficult for the fan blades 14 to flutter or if caught early enough, impede flutter potential. Use of compressor gases through fan nozzles 136 ensures that sufficient back pressure is applied to the fan blades 14 to dampen out flutter as measured by the optical tip timing sensors 212.
Fan blades 14 are continually monitored by optical tip timing sensors 212 during the application of high pressure compressor gasses through fan nozzles 136. As fan blade flutter subsides, controller 130 closes actuator 132 to discontinue the flow of compressor gasses to fan nozzles 136. The duration of the application of compressor gasses through fan nozzles 136 is dependent upon the amount of time it takes to dampen fan blade flutter.
The fan nozzle actuator 132 is selectively controlled by the digital controller 130 to control the air pressure through fan nozzle 136. The flutter control system 200 is a closed-loop system and includes one or more optical tip timing sensors 212 and a digital controller 130. The optical tip timing sensors 212 actively and selectively detect the flutter condition of one or more fan blades 14 and communicates with the digital controller 130 to actuate the fan nozzle actuators 132. The illustration provided is highly schematic. In one example, the optical tip timing sensors 212 are a time of arrival type sensor. The optical tip timing sensors 212 time the passage (or arrival time) of one or more fan blades 14 as the fan blades 14 pass a fixed, case-mounted sensor as the air fan blades 14 rotate about the engine longitudinal centerline axis A. The arrival time of the fan blades 14 are timed by the optical tip timing sensors 212. Other fan blades 14 may similarly be timed. The digital controller 130 is programmed to differentiate between which fan blade 14 arrival times correlate to a flutter condition and which fan blade 14 arrival times correlate to non-flutter conditions. According to the present disclosure, rotors for various parts of a gas turbine engine such as compressors and turbines may be provided that are less susceptible to damage as a result of flutter or forced response effects.
According to one method of controlling flutter in a fan included in a gas turbine engine, a system of the present disclosure may perform the steps of: measuring blade tip timing of fan blades included in the fan as they rotate about an axis, determining when a flutter condition or pre-flutter condition exists, and directing high pressure air from a compressor included in the gas turbine engine toward the fan blades to create a disturbance out of phase with unsteady pressures acting upon the fan blades.
The step of measuring blade tip timing of fan blades included in the fan as they rotate about an axis may be performed by optical tip timing sensors positioned along at least one of the leading edge and trailing edge of the fan blades. The optical tip timing sensors may monitor the leading edge and trailing edge of the fan blades and transmit acquired data to the digital controller.
The step of determining when a flutter condition or pre-flutter condition exists may be performed by comparing the blade data measured by optical tip timing sensors to known values to determine whether a flutter condition or a pre-flutter condition is present.
The step of directing high pressure air from a compressor included in the gas turbine engine toward the fan blades to create a disturbance out of phase with unsteady pressures acting upon the fan blades may be performed by a digital controller that operating an actuator to direct high pressure air through one or more nozzles directed at the fan blades.
While the disclosure has been illustrated and described in detail in the foregoing drawings and description, the same is to be considered as exemplary and not restrictive in character, it being understood that only illustrative embodiments thereof have been shown and described and that all changes and modifications that come within the spirit of the disclosure are desired to be protected.
This application claims priority to and the benefit of U.S. Provisional Patent Application Number 62/126,943, filed 2 Mar. 2015, the disclosure of which is now expressly incorporated herein by reference.
| Number | Date | Country | |
|---|---|---|---|
| 62126943 | Mar 2015 | US |
