Abstract:
An onboard system for a rotary wing aircraft detects a limit cycle oscillation in the tail mast and provides a timely indication of the limit cycle oscillation to an aircrew before serious damage to the airframe is likely to occur.

Description:
GOVERNMENT INTEREST 
     The invention described herein may be manufactured, licensed, and used by or for the U.S. Government. 
    
    
     BACKGROUND 
     A rotary wing aircraft provides a complex vibratory environment. Vibrations emanate from many different sources including the main rotor, the tail rotor, gearboxes, linkages and engines. Vibrations will vary in intensity and frequency depending on the speeds and relative speeds of rotation of the main and tail rotors, load factors, structural deformation, resonances inherent in the airframe, and aerodynamic forces. 
     Vibrations cause premature failure of mechanical components such as gears and bearings, damage to avionics, flight instruments, fatigue to the airframe and discomfort to passengers and aircrew. Some underlying cause of vibrations include imbalances in rotating parts, uneven friction, meshing of gear teeth, parts that are dragging together, etc. Traditionally, helicopter vibrations have been managed or suppressed by balancing and alignment of rotating parts, reduction of friction, the use of vibration isolation mounts, installation of damping structures, absorption materials, and the like. 
     More recently, electronic systems have been devised to monitor and manage vibrations on rotary wing aircraft. Rotor track and balance systems such as the Rotor Analysis and Diagnostic System (RADS) have focused on providing information that can be used in flight to adjust pitch links, blade weights and trim tabs for smoother operation. Still other systems such as the Active Control of Structural Response system (ACSR) made by Agusta-Westland have been designed to reduce vibrations of the main rotor by active control systems that employ high-frequency force-actuation within the helicopter&#39;s structure. See, for example, U.S. Pat. No. 5,853,144. 
     Unfortunately, not all harmful vibrations have been mitigated by RADS, ACSR, or similar systems. One such vibration that is not mitigated by an ACSR or RADS type systems is referred to a limit cycle oscillation (LCO). In general, an LCO is defined as an oscillation of finite duration and finite amplitude which will return to a steady state value without additional external influences placed upon the system other than those found in the normal system environment. In some helicopters, such as the EH-101/AW-101, manufactured by Agusta-Westland, LCO&#39;s of a significant amplitude and duration have been detected in the vicinity of the tail rotor. These LCO events are unpredictable, occur only rarely and emanate far enough away from the cockpit that they have not been perceived by members of the aircrew until they have reached a magnitude that could cause damage to the aircraft. 
     In an effort to better understand LCO&#39;s in the EH-101/AW-101 and similar helicopters, vibration sensors have been placed in the vicinity of the tail rotor and the signals monitored either on the ground, via telemetry, or by an operator while in the air. Because LCO events have been extremely rare and have taken place under a variety of seemingly unrelated conditions, their causes are not yet well understood. However rare they may be, the consequences of LCOs are potentially catastrophic. Thus, the recommended course of action when a significant LCO is detected is to land the aircraft as soon as practicable, before damage to the airframe can occur. Tasking an aircrew member to monitor for LCO events is simply not practical. Thus, there is an immediate need for a system to detect and alert a helicopter aircrew to unsafe vibration levels from LCOs in the AH-101/AW-101 and similar rotary wing aircraft, before damage from an LCO can take place. 
     SUMMARY 
     In general, in one aspect, an embodiment of a system to detect a limit cycle oscillation event in a rotary wing aircraft according to the present invention includes a vibration sensor that outputs a signal to indicate vibration in the tail mast of the aircraft, a filter to limit the signal output from the sensor to a predetermined frequency range in which the limit cycle oscillation occurs and a switch to provide a first signal when the filtered signal output from the sensor exceeds a first predetermined amplitude for a predetermined duration. 
     In another aspect, an embodiment of a system to detect a limit cycle oscillation event in a rotary wing aircraft according to the present invention includes a second switch to provide a second signal after the first switch has provided the first signal and the filtered signal output from the sensor has exceeded a second predetermined amplitude greater than the first predetermined amplitude. 
     In yet another aspect, an embodiment of a system to detect a limit cycle oscillation event in a rotary wing aircraft according to the present invention includes a third switch to provide a third signal after the second switch has provided the second signal and the filtered signal output from the sensor has exceeded a third predetermined amplitude greater than the second predetermined amplitude. 
    
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
       Embodiments according to the invention are illustrated in the accompanying drawings in which like reference numerals represent like parts throughout and in which: 
         FIG. 1  shows an acceleration-time graph of an exemplary limit cycle oscillation in the context of the present invention; 
         FIG. 2  shows an illustrative simplified side elevation of a helicopter that is equipped with an embodiment of a system for alerting an aircrew to unsafe vibration levels, according to the present invention according to the present invention; 
         FIG. 3  shows a side elevation of the tail portion of the helicopter shown in  FIG. 2 ; 
         FIG. 4  is a simplified block diagram of an embodiment of a system for alerting an aircrew to unsafe vibration levels, according to the present invention; and 
         FIG. 5  shows system output-frequency graph of the overall pass band of a signal conditioning and logic unit in an embodiment of a system for alerting an aircrew to unsafe vibration levels according to the present invention. 
     
    
    
     DETAILED DESCRIPTION 
     In the following detailed description, reference is made to the accompanying drawings which are a part of this patent disclosure, and in which are shown by way of illustration specific embodiments in which the invention, as claimed, may be practiced. This invention may, however, be embodied in many different forms and should not be construed as limited to the embodiments set forth; rather, these embodiments are provided so that this disclosure will be thorough and complete, and will fully convey the scope of the invention to those skilled in the art. 
       FIG. 1  shows an acceleration-time graph that approximates a representative LCO event  100 . An LCO event in the context of the present invention, such as might be experienced during operation of an EH-101/AW-101, is characterized generally by a vibration of the aircraft having a frequency in the range of 10-15 Hz, an acceleration of greater than 0.7 g&#39;s and a duration of more than 0.5 seconds. In this instance, LCO event  100  has occurred in the vicinity of the tail of the aircraft. It begins at approximately 1 second on the time scale (x-axis) and increases rapidly from about 0.2 g&#39;s to a peak acceleration of just under 1 g at second 5. LCO event  100  rapidly diminishes in magnitude so that by second 6.5 the event has ended. Overall, LCO event  100  persists for about 5 seconds. In general, LCO events may be longer or shorter than LCO event  100  but all share the same basic characteristics of having finite duration and finite amplitude and a return to a steady state value without additional external influences. 
       FIG. 2  shows a simplified side elevation view of an EH-101/AW-101 rotary wing aircraft (helicopter)  200  having a fuselage  202 , main rotor  204 , tail rotor  206 , tail fin  302  to which tail rotor  206  is mounted, cockpit  208  and turbines  210 . As shown in  FIG. 3 , tail fin  302  includes a tail rotor gear box  306  from which a tail rotor shaft  305  extends for mounting tail rotor  206 . The angular orientation and forward direction of the long axis of tail fin  302  is indicated by a dashed line directional arrow A. A directional arrow B indicates the long axis and forward direction of airframe  202 . A vibration sensor  304 , such as an accelerometer, is mounted to tail fin  302  aft of tail rotor gear box  306 . Vibration sensor  304  is oriented to detect movement of vertical fin  302  back and forth along the long axis of tail fin  302 , as indicated by directional arrow A. In the prototype, vibration sensor  304  is a 15 VDC model 7290A-10 accelerometer made by Endevco. A wide variety of available vibration sensors based on a number of different technologies, including piezoelectric, capacitance, null-balance, strain gage, resonance beam, piezoresistive and magnetic induction may be employed in alternative embodiments. The analog signal from vibration sensor  304  is carried by a shielded cable (not illustrated), preferably along production aircraft cable runs inside the tail drive shaft covers. The cable then proceeds inside the tail and aft cabin in the overhead and forward to cockpit  208 . In alternative embodiments, a wireless or fiber optic link may be employed and signals from the vibration sensor  304  may be digital. 
       FIG. 4  shows a block diagram of an LCO signal conditioning and logic unit  400  which receives and processes signals from vibration sensor  304  and provides LCO status signals for a display  402  positioned on the helicopter instrument panel in cockpit  208 . Signal conditioning and logic unit  400  includes a first stage signal conditioner  404  to band limit and buffer the analog signal from accelerometer  304  before digitization. First stage signal conditioner  404  includes a voltage follower/buffer amplifier  405 , a low pass analog filter  406  and a high pass analog filter  407 . In the prototype, voltage follower/buffer amplifier  405 , low pass analog filter  406  and high pass analog filter  407  are implemented on a USR-100 manufactured by Teletronics Technology Corporation. Low pass analog filter  406  is preferably a 6-pole Butterworth filter with a 3 dB cutoff frequency F c  of 21 Hz such that frequencies below 15 Hz pass essentially unattenuated (less than −0.09 dB reduction in gain) and higher frequencies, which may cause aliasing and ringing in downstream stages, are attenuated. The output of low pass filter  406  is coupled to the input of high-pass analog filter  407  which has an F c  of 7.14 Hz. High pass analog filter  407  is preferably a 6-pole Butterworth filter and has a passband above 10 Hz with less than −0.09 dB reduction in gain. High pass filter  407  functions to filter out low frequency signals from accelerometer  304  below 10 Hz, including a DC component representing the background gravitational force, low frequency noise from bumps encountered while the aircraft is taxiing as well as airframe modes below 10 Hz. In the prototype, first stage signal conditioner  404  is implemented on a Common Airborne Instrumentation System (CAIS) Data Acquisition Unit (CDAU), TTC M/N CDAU-2016, manufactured by Teletronics Technology Corporation. 
     The filtered output  416  of first stage signal conditioner  404  is coupled into a low pass digital filter unit  408  having multiple channels  408   a - d , which include analog to digital converter stages  409   a - d  and low pass filter stages  410   a - d . Sampling in the prototype is performed by a four channel 12 bit analog to digital (A/D) with f s  of 127 Hz. The inputs of analog to digital converters  409   a - c  are coupled to the output  416  of filter unit  407 . The input of analog to digital converter  408   d  is coupled to the unfiltered output  412  of vibration sensor  304 . Digital filters  410   a - d  are preferably finite impulse response (FIR) filters with 120 taps, 8× oversampling and f s  of 127 Hz. Digital filters  410   a - d  have an F c  at 15.88 Hz and essentially no attenuation at frequencies below 14.82 Hz. Digital filters  410   a - d  are designed with sharp cut off characteristics to attenuate vibratory interference that is close to the frequencies of an LCO, such as main rotor and tail rotor blade passing frequencies, which occur at about 16.2 and 17.8 Hz, respectively. Digital filter unit  408  is implemented in the prototype system on an SCD-608D-2 signal conditioning card manufactured by Teletronics Technology Corporation. The overall passband of the Signal conditioning and logic unit  400  is shown by a curve  500  plotted in  FIG. 5 . 
     The data output from digital filter unit  408  is monitored by a level detector  411  having four channels  411   a - d . Channel  411   a  is configured to provide an indication (i.e., an output signal) whenever the filtered data from vibration sensor  304  indicates that a 0.7 g level has been detected. Since an LCO is known to occur at a frequency of between 10 and 15 Hz, the indication from level detector  408  is configured to persist for at least ½ of the period of the lowest frequency of interest, i.e., 10 Hz. 
     The outputs from level detector  411  are coupled to a logic stage  415  which includes a timer circuit  412 , latches  413   a - c  and light drivers  141   a - d . The output from level detector channel  411   a  is coupled to timer circuit  412  which, in turn, provides an indication when the indication from level detector  411   a  persists for 0.5 seconds or more. Timer  412  may be adjusted to a shorter or longer period in alternative embodiments. The output of timer  412  is coupled to a latch  413   a  which is coupled to light driver  414   a.    
     Level detector channel  411   b  is configured to provide an indication whenever the signal from digital low pass filter channel  408   b  exceeds a magnitude of 2 g&#39;s. The output of latch  413   a  and the output of level detector  411   b  are coupled to the input of latch  413   b . The output of latch  413   b  is coupled to light driver  414   b.    
     Level detector channel  411   c  is configured to provide a indication whenever the signal from digital low pass filter channel  408   c  exceeds a magnitude of 3.5 g&#39;s. The output of latch  413   b  and the output of level detector channel  411   c  are coupled to the input of latch  413   c  which in turn is coupled to light driver  414   c . Level detector  411   d  is configured to provide a indication whenever the signal from digital low pass filter channel  408   d  exceeds 2.5 volts. The output of level detector  411   d  is coupled to light driver  414   d.    
     The outputs from light drivers  414   a - d  are coupled to an LCO annunciator/indicator unit  402  which is positioned in the cockpit where it is easily seen by both pilots. LCO annunciator unit  402  includes four display lights  402   a - d . The first three lights  402   a - c  indicate LCO events and are arranged in a row in order of severity from left to right and include, a green indicator light  402   a  to indicate detection of the least severe LCO event, a yellow indicator light  402   b  to indicate detection of a moderately severe LCO event, and a red indicator light  402   c  to indicate that a severe LCO event has been detected. A blue indicator light  402   d  is positioned below the row of lights to indicate that the data may be relied upon by the aircrew. In alternative embodiments, audible alerts may be provided in addition to or as replacements for one or more indicator lights. 
     Operation of the system will now be described. As noted, signal  412  from vibration sensor  304 , which is not processed by first stage signal conditioner  404 , is monitored by level detector channel  411   d  for the presence of the DC bias voltage from the background gravity field which will always exist when vibration sensor  304  is powered up and working properly. Indicator light  402   d  will be illuminated by light driver  414   d  as long as good data is being received from vibration sensor  304 . The light will not illuminate if there is no power to the instrumentation system, no power to the accelerometer, or the accelerometer fails to return a DC biased signal. 
     Signal  416 , which is processed by first stage signal conditioner  404 , will be monitored by signal conditioning and logic unit  400  to check for three levels of vibration. The first level is currently set in the prototype to indicate whenever an oscillating vibration maximum amplitude reaches ±0.7 G&#39;s. This indication is provided by level detector channel  411   a  and initiates timer  412 . Indicator light  402   a  (green) of annunciator unit  402  will illuminate if the vibration maximum amplitude remains at 0.7 Gs or greater for at least 0.5 seconds and will remain illuminated via latch  413   a . Only if indicator light  402   a  (green) has been illuminated, may second and third levels of detection circuitry be enabled. Indicator light  402   b  (yellow) light will illuminate when the filtered signal reaches ±2.0 G&#39;s. Indicator light  402   c  (red) will illuminate when the signal reaches ±3.5 G&#39;s. 
     Latches  413   a - c  keep indicator lights  402   a ,  402   b  and  402   c  (green, yellow, and red) lit once they have been illuminated until the aircraft is powered down. Both the filtered and unfiltered signals from vibration sensor  304  are preferably recorded continuously. A flag or other indication preferably will be set and recorded in the data stream every time the filtered signal transitions above each of the three data levels to facilitate post-flight analysis. 
     CONCLUSION 
     Although the present invention has been described in considerable detail with reference to certain embodiments hereof, it will be clear to one skilled in the art that the above embodiments may be altered in many ways without departing from the invention. Accordingly, the spirit and scope of the appended claims should not be limited to the description of the embodiments contained herein.