? 1997 Institute of Noise Control Engineering of the USA, Inc.
AIRCRAFT DAMAGE DETECTION FROM ACOUSTIC AND NOISE
IMPRESSED SIGNALS FOUND BY A COCKPIT VOICE RECORDER*
Authors: Ronald O. Stearman Stuart M. Rohre
Aerospace Eng. and Eng. Mechanics Dept. Applied Research Labs
The University of Texas at Austin The University of Texas at Austin
Austin, Texas 78712 P.O. Box 8029
Austin, Texas 78713
Glen H. Schulze
Data Acquisition Systems email@example.com
5965 W. Morraine Ave. firstname.lastname@example.org
Littleton, CO 80123 email@example.com
Currently, research is being conducted to detect damage through structural acoustics, signal processing, and transducer designs. The present study illustrates that damage detection may be carried out with an existing system acting as a latent signal transducer. One example involved a reliability problem in a commuter aircraft engine mount design where undetected crack growth created a critical whirl flutter condition destroying the aircraft. This reliability problem prompted the need for an in-place damage detection system to identify critical engine mount conditions. Signal analysis of data acquired by a cockpit voice recorder prior to and during the catastrophic aircraft whirl flutter event provided insight into critical signals that indicated the failure onset. Although regularly scheduled inspections failed to detect the problem, cockpit voice recorder signals contained a dynamic signature of this damage feature intermittently throughout the duration of the tape. It is highly probable that this damage signature existed for a much longer period of time, but due to the endless loop configuration of the cockpit voice recorder the earlier data was erased. This study indicated that even in the case of an unused cockpit voice recorder track, careful signal processing can extract surprising details about detecting potential damage along with extracting aircraft breakup signatures.
The motivation for the present study was a 1992 Airline Pilots Association (ALPA) accident report concerning a popular nineteen passenger commuter aircraft . The aircraft had been temporarily removed from its fleet operations for an evening training mission. The ALPA accident report concluded that an inflight right engine separation had occurred during this mission. The free right engine then traveled back striking the tail of the aircraft destroying most of the horizontal surfaces. This caused a complete loss of control of the aircraft resulting in an inflight breakup which was fatal to all occupants. Since this occurrence was prior to recent FAA legislation requiring flight data recorders on all commercial airliners, the cockpit voice recorder was the only on board flight record available to supply clues as to the cause of the accident . This included not only the voice communication at the critical event but structural acoustic as well as other acoustical sounds and noise sources during the breakup phase. In essence, the lack of a flight data recorder prompted an extended study of the CVR tape to determine whether the CVR recorder acted as a latent signal transducer. That is, did the CVR record events other than voice signatures that would be critical to determining if an in flight breakup of the aircraft occurred? The voice signatures on the tape indicated that a final catastrophic event occurred without warning to the pilots. No significant voice stress characteristics were identified up through and including the final event.
Past history has taught us that propeller whirl flutter is a catastrophic event that will remove engines and destroy aircraft, but which does not generally occur unless the supporting engine and propeller structure has been significantly damaged. This consideration first prompted an engine mount reliability study to estimate the characteristic service life of the engine truss and to determine if a history of engine truss service difficulties did in fact exist .
An initial search through the Federal Aviation Agency’s (FAA’s) Service Difficulty Reports (SDR’s) indicated that
at least six engine truss redesign cycles have occurred over the ten year life history of the aircraft. Such a large number of engine mount design changes suggest a significant engine mount service life problem. In view of this, the question was then pursued as to whether or not there was a reliability improvement as the redesign cycles progressed through the ten year period.
Once the service difficulty reports (SDR’s) had been reviewed, it was evident that the mount had been experiencing
frequent cracking, particularly in the vicinity of welded joints near the engine attachment points. The actual truss life was noted to be far below the intended and predicted design life of 20,000 hours. Consistent with extreme value studies, only first-time cracking failures were considered in the reliability study. That is, a repaired truss reentered into service was not considered. Six truss types were identified in the preliminary investigation. Only four of the six were found to have statistical relevance to the study, as the remaining two truss types were only in limited use and the sample data was insufficient. The various truss types were indicative of redesigns that evolved over a ten year period and were developed with the goal of extending the characteristic life of the truss. Modifications included increases in truss tube diameters and tube wall thickness along with the addition of gusset plates to the elements and joints noted to crack most often.
In the parameter identification to characterize the behavior of the failing trusses, the data was analyzed with the versatile Weibull model. The resulting four analyses, one for each of the four trusses included in the study, indicated an alarming trend. Rather than increasing the characteristic life of the truss, the augmented strength in each redesign of the truss resulted in an overall decreasing trend in the characteristic life. A summary of those results are depicted in Figure 1. (Truss types are indicated in chronological order of development from left to right).
In addition, analysis of the shape function, a characterization of the type of failure rate, indicated that this important value was also decreasing and, as of the last design, indicated the possibly of failures in the infant mortality region of the "hazard rate curve”. Figure 2 illustrates the critical transition of the latest truss design toward an infant mortality mode of failure (shape parameter < 1). 88000
Upper LimitShape ParameterLower Limit67000
s-confidence limits of 95%46000