CA1179529A - Aircraft structural integrity system - Google Patents

Abstract

ABSTRACT OF THE DISCLOSURE

Systems and methods for assessing the effect of at least one of a plurality of forces acting upon a structure are provided herein.
Structural moment detectors, combined with external signal and data-processing components or structural information detectors which are unitary devices which include the signal and data-processing functions, are employed as an integral component of these systems. These systems include basic measure-ment systems, structural integrity measurement systems, applied structural measurement systems, applied load measurement systems, applied communciation-detection systems and additional miscellaneous systems. One preferred such a structural information detector comprises a self-contained, unitary device for collecting and interpreting data reflecting the effect of a force acting on a structure including a housing adapted to be attached at spaced points to the surface of the structure, an optical sensor within the housing for detecting the relative orientation of spaced surface coordinate vectors of the structure, circuitry within the housing for converting signals from the optical sensor to a form usable by signal-processing electronics also located within the housing. The output of this structural information detector embodies useful information which directly indicates the effect of the force acting on the structure.

Description

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This invention relates to systems employing structural moment detectors for collecting and interpreting data reflecting the effect of at least a selected one of a plurality of forces acting on an aircraft or helicopter structure.

Optical sensors, sometimes called structural moment detectors of flexural rigidity sensors, which are basically autocollimators which are insensitive to linear dynamic motion but which respond to angular deflection of one end of the sensor with respect to the other, are known in the art. For example, such sensors are disclosed in the patent to Rossire, U.S.
No. 3,229,511 and in the publication entitled "The Structural Rigidity Sensor:' Applications in Non-Destructive Testing", published by the Air Force Systems Co~nand, United States Air Force (Frank J. Seiler Research Laboratory, Publication SRL-TR-75-0017, October, 1975). See also the U. S. patents to Okubo Nos. 4,159,422 issued June 26, 1979 and 4,164,149 issued August 14, 1979.

Systems which employ structural moment detectors to measure and record certain effects of forces acting on a structure are also disclosed in the publications described above. For example, the Rossire patent discloses an aircraft attitude control system in which a structural moment detector is used to sense wing loading and automatically adjust the attitude of the aircraft to maintain wing loading within safe operational limits. The Air Force publication and the Okubo patents disclose systems which employ structural moment detectors to obtain the ~vibration signatures~ of various structures such as airframes, buildinqs, aerospace vehicles, 11~795~9 rotating machinery bearings, dams and the like.

The present invention relates to improvements in the prior art sensors (SND) systems described above and to novel end-use applications of the.structural information detector (the SMD plus improved raw signal processing systems) which are specially adapted for measurement of the structural integrity of an aircraft.

The details of such systems and applications may vary somewhat depending on the precise objective but, in general, they will comprise a sy~tem for collecting and interpreting data reflecting the effect of at least a selected one of a plurality of forces acting on an airframe structure and will include at least one structural moment detector carried by the structure for generating output cignals in respon~e to the plurality of forces acting on the structure, means for proces~ing the output signals to modify the information content thereof (including rejecting components) of said signals which reflect extraneous forces other than the selected force) and means for manipulating the processed signals to provide secondary signals which are responsive to the condition of the airframe as a result of the application of the selected force.
As used herein, the term "forces acting on a structure~ is intended to include not only primary external forces applied to the structure but also includes secondary external or internal effects which flow from the application of external forces or changes in the environment of the structure, ~uch a~, for example, strain energy released within the structure as a result of cracking, thermal il'7~S~9 stresses, gravity-induced effects, electromagnetic forces and stresses, and the like.
In the accompanying drawings, Figure 1 is a schematic drawing illustrating a typical application of structural moment detectors and related electronics and data processing units to assess the structural integrity of an aircraft;
Fig. 2 shows in greater detail the signal processing the buffering components of the processing electronics of Fig. l;
Fig. 3 depicts a system for assessing altered structural capability of aircraft components; and Fig. 4 depicts an implementation of the systém of the present invention for use in connection with assessing the structural integrity of helicopters.

In accordance with aspects of the present invention, SMD's are employed in systems whlch provide information concerning structural fatigue, active crack detection, overload conditions, load history and vibration for the purpose of assessing structural loads on the aircraft, the remaining life-time of the aircraft, the condition of components of the aircraft, the operational history of the aircraft, the compliance of structural members with known loadings, the current ability of the aircraft to carry its design loadc and to provide more effective and efficient maintenance and test procedures.

11~9529 The ultimate function of the aircraft structural integrity system of aspects of the present invention is to prevent catastrophic structural failure of the aircraft through structural integrity assess-ment.

Airframes are designed to meet desired strength and fatigue life requirements by projecting the average load environment of the alrcraft over the expected airframe lifetime.

1179S;~9 TheJe pro~ectlon~ are, ln turn, b~sed on averaged load hlstory data concerning similar aircraft designed for similar ~-ervice~. Since aircraft receive such diversified use, however, it i~ very difficult to assign confidence limits to th~ ~tatlstics thus assumed. One particular aircraft could experience a load history quite di~ferent from the expected load hi~tory a~sumed $n the original design. ~n term~ of accumul~ted damage, this airframe could actually be much ~older~ than the recorded engine or airframe hour~
. ...
woula suggest. Nevertheless, aircraft are normally inspected for cr~tical flawJ based on the accumulated flying time or engln- operation time. The assumption is that this number of hour- is correlated with the load hi~tory actually exper-lenced by thc aircraft. The pos~ibility for error is obvious7 100 hour- flying along the front range of the Rocky Mountains can repreJent a conslderably différent load en~ironment fro~ 100 hourJ flying over Kansa~ wheatfields. Nevertheless, pre~ent alrcraft ~tructural integrity asses~ment technique-almo~t lnvarlably employ airframe or engine operating time to ~etor~ino when inspections should be made, followed by actual in~pectio~ of the airframe to estimate the remaining fatigue lifetime.

According to an aspect of the present ernbodiment of the invention, aircraft structural integrity assessment is performed using systems which include SMD's which act, in effect, as a damage counter, providing continuously updated real time assessment of the remaining fatigue lifetime of the airframe.

1~795~9 The system achieves this func~ion by one or a combination of monitoring structural cracks as they occur and continuously assessing the remaining fatigue lifetime of the airframe (structural failure monitoring), warning of impending failure of the airframe due to loads which approach or exceed the design loads (load measurement), by warning of impending engine failure (vibration monitoring), and by directly measuring the compliance of structural members subjected to known loads.

Referring to Fig. 1., in the present embodiment of an aspect of the invention, a plurality of SMD's 111 are mounted at selected locations on airframe 112 and are properly oriented to be sensitive to the selected parameter to be measured. The SMD's lll are connected through cabling 113 to the processing electronics 114 (signal processing and bufferlng, A-D conversion, data processing, data recording and editing and labelled MPV on the drawing and the output 115 is suitably tisplayed 116 and/or stored.

Fig. 2 shows in greater detail a system for measuring aircraft structural integrity. The signal processing and buffering components are, as shown in Fig. 2, individual units for each sensor consisting of power supplies 117, voltage followers 118, and signal amplifiers 119.
The power supply provides conditioned power to the sensor and powers the signal processing electronics. The voltage follower converts the sensor output to a buffered analog voltage signal. The signal amplifier amplifies the signal from the voltage follower.

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, The analog to d~g~tal converter either individually or collectively converts the analog information from the ~ensor~ to a dig~tal format compat~ble with that required by the computing, processing, recording, editipg and display unit~. ~a~ically, the A-D converter samples the analog signal 120, holdJ the sampled value a~ a constant voltage 121 and converts the voltage level 122 to binary data.

The data processing computer 123 is a single frame computer capable of accepting the digital data 124 and manipulating it in a predetermined programmed fa~hion to convert the digitized data into a representation 125 related to aircraft or engine structural integrity. The output 125 may include information concerning the existenca of crack~, the ~everlty of tho crack-, the load~ experienced by the aircraft, the load history, the fatigue lieftime of the structure, the integrity of engines or engine mounts, the v~bration signatures and other information involving aircraft or engine ~tructural integrity~

The dat~ recording and editing device 126 prov~de~
for permdnent recording of all or part of the acquired data 125 for u~e at a later time and, additionally, provides the capability to manually edit the acquired data.

The data di~play and keyboard device 126 provide~ a visual display of either the data 125 acquired from the jcomputer 123 or of the recorded edited data 128 from the data I

1~795~

rocordLng and editing unit. Additionally it provides, by means of feedback loop 129, for predetermined alteration of the mean-by which the data processor 123 is transforming acquired data 124 or, by means of feedback loop i3Q, the manner in which the data 125 is recorded, edited, displayed or digitized.
The keyboard 125 is for data/command entry to the ~ystem.

al Structural Pailure Monitoring There are essentially two independent methods of detecting incipient fa$1ure of the structure. The first method involves measuring the frequency response and/or compliance of a structure and its variability with time.
The second method involve~ mea~uring the seismic waves in the structure induced by cracking or other structural failure.

Mo~t aircraft structures can be con~idered to be nearly elastic systems. Even structures formed with composite materials have vibrat~on characteristic3 which are well defined and repeatable. It is known, both analytically and experimentally, that the natural modes of vibration of a structure are distinct, well-defined and a function of the physical d~men~ions of the structure and the internal properties of the ma~erial. Any changes in these factors will change the natural frequencies and the associated vibrational modes of the structure.

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Certaln characteristic~ of the aircraft which change slowly over its lifetime, e.g., elasticity of the various structural members, may ~e init$ally calibrated by obser~ng deflections under ground test conditions and thereafter continuously monitoring the vibratLonal mode frequencie~ and shapes. In accordance with aspects of the invention, impending ~tructural failure can be ob~erved by nitoring the changes in these mode shapes and frequencie~.~ Further, shocX waves frll very 5ma11 crac~c propogation centers can be easily detectod.

Initial calibration of the aerodynamic coefficients may be accomplished by flying aircrrft on specifiable paths, e.g., high-speed taxi runs and take-offs that terminate in a landing on the ~ame runway, optionally in oppo~ite d~rections to eliminate effect~ of win~, after performing such prescribed maneuver~ a~ stall~, turn~ and high-speed pa~ses. Changes ln aerodynamic ¢oefficient~ due to environmental conditions, e.g., rain or ice, are then observed as changes in deflection~ without associated change~ in elasticlty ~as oSservable through structural frequency of the natural modes).
Rapidly changing variables e.g., aircraft mass is initially calibratéd by observing aeflections of the landing gear, suspension points on the ground, and thereafter estimated by counting engine cycles and observing changes in m~de shapes and frequencies associated with fuel depletion.

11'79~3 With th~se techniques, the components structural deflections which may have aerodynamic ~lift, ~ido force ~nd drag), inertial ~g loading), thermal loading, engine vibration ioading, transient aerodynamic loading, operAtion of aircraft subsystems ~i.e., drag brakes, landing ge~r, flaps and engine controls) can be effectively ~eparated.

With information sensed from the SMD sensor system, the changes in the frequencies of thè natural de~ of the structure due to changing physical dimensions and internal materlal properties will be determined by the microproce~sor.

Each structure will have a distinct set of natural frequencle~ and will have a well-dçfine~ impul~e response when the ~tructuro 1- new, unfat~gued ~nd ha~ no exi~ting crack~.
When changes occur, e.g., cracking or fatiguing, the lmpul~o re~ponse wlll change due to changing natural frequencies of the structure. Thus, continuous identification of either the impulse response or equivalently the frequency spectrum of the structure allows for identification of structural change~ of appreciable size.

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To i~plement active cr~ck detection in an air vehicle that ha~ been optimized for ~trength-to-weight rat~o, where m2ny components operate near the structural limits, requires an alternate real time approach. In such aircraft, the materials used, e.g., aluminum, glass fibres and advanced composites, fat~gue can readily reduce the useful lif~ of the aircraft. If such an aircraft i8 subjectea to inadvertent adverse loads or the structure material has unsuspected material defects, the onset of structural failure may occur with a cata~trophic di~continuity. The early signature of such a fa$1ure may be manifested in exceedingly small cracks ~n th~ structure which can result ln ~tructural failure in relat$vely ~hort order. The prior art procedure for crack detection in aircraft involves visual and/or electronic inspection of the aircraft on the ground. Since extremely small cracks are of great significance in these type~ of structure~, inspection procedures are st costly and relatively ineffective.

According to aspects of the invention, these exceedingly small crack~ are detected as they occur in f light by taking advan-tage of the act that the formation of cxacks involve~ a substantial release of strain energy. The release can be detected by the SMD sensor and, when processed by the micro-pxocess~r, the occurrence of the crack can be ~eparated from 11'7~3S"9 the inertlal and Aerodynamic forces simultaneously lnduc~ng the ob~erved structural deflections.
. .

Crack~ occur in a material in an attempt to relieve the stress created by incident force~. This results in a release of ~traln energy. This energy is expended in essen-tlally two forms: 1) the formation of the surface of the crack, and 2) the kinetic energy for crack propagation.

It is well-known that the crack velocity is signi-ficantly less than the velocity of sound in the materlal.
For example, in aluminum, the crack velocity is no more than 3/10 the speed of sound in the materiai. Therefore, if a crack propagate~ only a ~hort distance, the acoustic fault ha~ already been transmitted through the ~tructure and 1-sense~ by the SMD sensor means. This phenomenon i~ similar to an explos~on in the alr where the ~hock wave propagates much faster than the exploslve products.

Materials, e.g., glass at room temperature, undergo brittle fracture. The ~a~e is not trua of metals, for instance, which are capable of deforming by slip and twinning even at very low temperatures. ~t has been observed that even when a metal fail~ by bxittle cleavage a certain amount cf the plas-tlc deformatlon almo~t always occurs prior to fracture. Metal~
therefore do not fracture as a result of pre-existlng crack~
but, in many cases, by cleavage cracks nucleated as a result of the pla~tic-defonmatl~n proceYs. Present theories favor 11795;~

thc concopt of dislocation interactions as inducing cleavage nuclel. D~slocations on different slip planes can combino to form new dislocations on the fracture plane, thereby opening a c~ack. Alternately, slip on a g~ven plane can be impeded by some sort of barrier leading to ~ pile-up of dislocations wh$ch, $n turn, nucleate a crack. An obstacle to slip must be very strong so that lt can stand the high stress at the head of the dislocation pile-up. Deformation twins ana grain boundaries are obstacles with sufficient strength to stand the high stress.
.

Therefore, in materials, e.g., metals, energy is lo~t due to plastic deformation in addition to crack formation.
When slip takes place during the movement of a crack, energy is absorbed ln nucleating and moving dislocations. If the energy r¢quired to ovorcome plastic deformation becomes too large, the crack may decelerate and ~top. Thus, detection of these cracks after they occur is a very difficult task. At the instance of crack occurrence, howevar, energy release will excite lightly driven frequency modes and drive acoustic waves within the material. Because of the broaa ;
freguency response of the SMD, it can easily detect this energy release. Because of the sharpness of the pulse, time of arrival techniques can be used to ascertain the approximate location or origination of the acoustic pulsé.

U

11 7~S~9 4) V~bratlon Monitoring ~ he concept of monitoring the vlbration ~ignature of a mechanical system for an indication of the mechanical health of the system is well establi~hed. Every operating mechan$c~1 system ha~ ~ aistinct vibration signature which 1~ produced when the system i8 operating properly.

When a malfunction occurs the signature changes.
Appropriate observat~on and analysi~ of the vibration signature can therefore provide an early indication of the severity and location of pO8~ ible trouble and can help to prevent costly cata~trophlc failuro.

The actual vibratlon signature of an aircraft cont~ln~
many frequencie~. This 18 a result of different components v$bratinq at varlous discrete frequencies and various mechanical re~onance~ an~ nonllnear combinations of those signals ln the machine. The resultant signal at a measurement point ~ there-~ore a complox vlbration wave form which must be processed to reduce ~t to its dlscrete frequency components or analysis.

~ typical prior art vibration monitoring system con~lsts of sensors (vi~ration transducers), a signal processor tmonitorlng system) and ~uitable displays or alarm generating device~. ~he ~en~or~ commonly ln u~e are the piezoelectric accelerometer and the inductiv~ velocity transducer. Whlle there is limited agreement on the specific cro~sover frequency, there i~ general agreement that vibration severity 1~ propor-tional to velocity at relatively low frequencies and propor-tional to acceleration at high freguencies. Thus, the application~ of the piezoelectric and velocity transaucer~ are 5~9 naturally separated by frequency. In addition, veloc~ty tran~ducers are generally rugged, operate over wide temperature ranges, produce realtively high signal to noise output~, but are limited to 1000 Hz. Piezoelectric accelerometers are more sen~itive to contamination. Both ~ave frequency ranges which are ~ignificantly influenced by the mothod of attachment to the machi~ne.
.
Both the velocity transducer and the piezoelectric accelerometer respond to displacements perpendicular to their unting ~urface. The structural moment detector ~SMD) however, ¦
mea~ure~ th;e difference between plane perpendicular to the ~urface to which it i8 mounted, that is, the measurement motion i8 90- to that of other sensors.

~ ho siginificant point i8 not that suitable mountings will permit direct replacement of velocity transducers, strain gage~ and piezoelectric accelerometers. The s~gni~icant point i8 that the SMD responds to transverse and longitudinal waves ln a body whlch cause the surface to deflect a8 little a~
3.5 x 10 9 raaians across the 1.5 inch length of the sensor ~a surface displacement of 5.3 x 10 3 microinches). This is a new and unigue measurement and it provides new information for vibration nitorins systems, especially suited to aircraft structural integrity systems.

In the measurement of these ultra small deflection~, the SMD has a frequency res~onse which is essentially flat from 0 to 40 k~z. Combinations of velocity and acceleration sensors ~n the best vibration~ ni~oring systems currently use, pn~vide a flat response from 0 tp 20 ~nz only.

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c) Aircraft Load Measurement Another application of the systems of aspects of the present invention i~ the direct measurement o-f the ability of variou~
aircraft ~tructural member~ to carry a specific load. Tho output of the structural moment aetector i~

OFRS ' EI ~f~loading)dx where El is the effective flexural rigidity of a member of tho aircraft ~tructure and f~loading) indicate~ the local bending moment due to the loading on the structural member.
Th ability of th~ structural ~ember to carry its design loaa can ~e a~se~sed by i~posing a known load wh~ch i~ ~mall relatlve to the dcsign load on the structure and using the output o~ the ~D to determine the flexural rigidity of the ~tru¢ture. Changes ~n the ability of the ~tructural member to carry it- design load re~ult in changes in the flexural rigi~ity o~ the member and these changeQ can be translated lnto an ~n~ication of the ~everity of tho change. This mea~urement makes use of the static (IX) output of the SMD.

According to aspects of the present invention, the inverse of the above described operation i~ used to measure the actual load impo~ed to an aircraft structural member. If the flexural ri~dit~

~8 ~nown, then the ~ensor output is directly related to the load.

, With appropriate signal processing it is possible frequently to evaluate the flexural rigidity of the member using autocorrelation techni~ue~ and then use that ~alue of flexural rigid$ty to deter~ine the existing load~. Thus, llt~'~S'~3 knowing both the lo~ carrylng capability of the member and the actual load, it i~ po~sibl~ to predic~ impending ~ailur~.

d) Altered Structural Capability Measurement According to this embodiment of aspects of the invention a ~y~te~ iQ provided for asse~sing the structural capabillty of a~ aircraft- ox o~her structure which has been pu~posely dified or repaired or which has altered capability as a result of the accumulation of ice, w~nd damage or other environmental effect~. This system uses the unique measur2ment of the SMD to infer the exist~ng flexural rigidity of ~tructural member~. This measurement is then oomp~red with aes~gn values and an asses~ment is made.

In one implementation of the system the ~ensor is unted on a beam which ha~ fractured and has bee~ rep~ired by weld$ng. The SMD output 18 processed to provide the flexural rigidity of the welded beam. This value i8 compared with the value of the unfractured beam to determ~n~ if the repair has returned the structure to it~ original ~trength.

In another implementation the sensor is unted to a beam whose strength has been increaffed by the addition of other ~ember~. The SMD output i8 procesffea to provide the flexural rigidity of ths gtrengthened beam. This value ~3 then compared with ths design value to see if the desired ~trength has been achieved.

. . .
Yet a third imple~entation is dep~cted in Figuxe 18.
The SMD 131 i~ mounted ~eas the leading edge of the airfoil 132.

~79S'~9 ~~ Th~ SMD output ~8 processed 134 to determine flexural r~gidi~y.
The value obtainea differ~ from the design value as a re~ult of the accumulation of ice 133. Thi~ difference is used a~
an indication to the pilot that his airplace wing has an accumulation of ice.

e)- Structural Integrity Assessment Systems For Helicopter~

According t~ this embodiment of aspects of the inventio~, a sy~tem i8 provided for assessing the structural integrity of helicopter~. In this embodiment, certain aspects of the aircraft structural integrity systems, described above, and of other load and structural mea urement systems, described belo~, are combined to produce a system tailored especially to the unique requirements of measuring the structural integrity of ~ helicopter structure. Helicopter structures present uni~ue pro~lems because of the relatively low frequency but large amplitude of rotation and vibration of the helicopter rotor system.

Technical failures, involving mater~als or components, which result in helicopter accidents include failure of :
~1) power plant ~2) power train (3) driveshaft (4) rotor blades and (5) rotor hubs. Primary causes of failures include incorrectly fitted parts, high ~tress concentration and cyclic exitation resulting in fatique.

According to this embodiment, a plurality of sensors are mounted throughout the helicopter. S~gnal processing techniques, including adaptive filtering and adaptive noi~e cancellation techniques are used to ~eparate the ~ackground 1179S'~9 no~se from crack init~ation and propogation signals. In addition, v~bration ~ignatures of the major componentæ
mentioned above are nitored. Any significant cha~se~ are used to as~es~ structural $ntegrity.

One mode of implementation of the invention is shown in Figure 4. SMD sensors 135 are ~ounted near rotating mechanical parts and long main structural members. The sensor outputs are fea to the electronics, microprocessing and recording unit 136~ The data is processed according to the preprogrammed instructions and appropriate information i8 then displayea 137 to the p~lot and/or recorded for use by maintenance personnel after landing.

Claims

The embodiments of the invention in which an exclusive property or privilege is claimed are defined as follows:

1. A system for collecting and interpreting data reflecting the effect of at least a selected one of a plurality of forces acting on an aircraft structure, said system comprising, in combination:

(a) at least one structural moment detector carried by said structure for generating output signals in response to said plurality of forces acting on said structures (b) means for processing said output signals to modify the information content thereof, including rejecting components of said signals which reflect the effects of extraneous forces other than said selected one;

(c) means for manipulating said processed signals to provide secondary signals responsive to the condition of said structure as a result of the application of said selected force.