DE4240600C1 - Structural defect detection system for aircraft - uses modal analysis via detected oscillation of aircraft structure and comparison with aircraft model - Google Patents

Structural defect detection system for aircraft - uses modal analysis via detected oscillation of aircraft structure and comparison with aircraft model

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Publication number
DE4240600C1
DE4240600C1 DE4240600A DE4240600A DE4240600C1 DE 4240600 C1 DE4240600 C1 DE 4240600C1 DE 4240600 A DE4240600 A DE 4240600A DE 4240600 A DE4240600 A DE 4240600A DE 4240600 C1 DE4240600 C1 DE 4240600C1
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Prior art keywords
aircraft
signals
finite element
element model
sensors
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Expired - Fee Related
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DE4240600A
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German (de)
Inventor
Otto Sensburg
Heinz Dr Hoenlinger
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Airbus Defence and Space GmbH
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Deutsche Aerospace AG
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    • G—PHYSICS
    • G01—MEASURING; TESTING
    • G01M—TESTING STATIC OR DYNAMIC BALANCE OF MACHINES OR STRUCTURES; TESTING OF STRUCTURES OR APPARATUS, NOT OTHERWISE PROVIDED FOR
    • G01M5/00—Investigating the elasticity of structures, e.g. deflection of bridges or air-craft wings
    • G01M5/0033—Investigating the elasticity of structures, e.g. deflection of bridges or air-craft wings by determining damage, crack or wear
    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B64—AIRCRAFT; AVIATION; COSMONAUTICS
    • B64F—GROUND OR AIRCRAFT-CARRIER-DECK INSTALLATIONS SPECIALLY ADAPTED FOR USE IN CONNECTION WITH AIRCRAFT; DESIGNING, MANUFACTURING, ASSEMBLING, CLEANING, MAINTAINING OR REPAIRING AIRCRAFT, NOT OTHERWISE PROVIDED FOR; HANDLING, TRANSPORTING, TESTING OR INSPECTING AIRCRAFT COMPONENTS, NOT OTHERWISE PROVIDED FOR
    • B64F5/00—Designing, manufacturing, assembling, cleaning, maintaining or repairing aircraft, not otherwise provided for; Handling, transporting, testing or inspecting aircraft components, not otherwise provided for
    • B64F5/60—Testing or inspecting aircraft components or systems
    • G—PHYSICS
    • G01—MEASURING; TESTING
    • G01M—TESTING STATIC OR DYNAMIC BALANCE OF MACHINES OR STRUCTURES; TESTING OF STRUCTURES OR APPARATUS, NOT OTHERWISE PROVIDED FOR
    • G01M5/00—Investigating the elasticity of structures, e.g. deflection of bridges or air-craft wings
    • G01M5/0016—Investigating the elasticity of structures, e.g. deflection of bridges or air-craft wings of aircraft wings or blades
    • G—PHYSICS
    • G01—MEASURING; TESTING
    • G01M—TESTING STATIC OR DYNAMIC BALANCE OF MACHINES OR STRUCTURES; TESTING OF STRUCTURES OR APPARATUS, NOT OTHERWISE PROVIDED FOR
    • G01M5/00—Investigating the elasticity of structures, e.g. deflection of bridges or air-craft wings
    • G01M5/0066—Investigating the elasticity of structures, e.g. deflection of bridges or air-craft wings by exciting or detecting vibration or acceleration
    • G—PHYSICS
    • G01—MEASURING; TESTING
    • G01M—TESTING STATIC OR DYNAMIC BALANCE OF MACHINES OR STRUCTURES; TESTING OF STRUCTURES OR APPARATUS, NOT OTHERWISE PROVIDED FOR
    • G01M7/00—Vibration-testing of structures; Shock-testing of structures

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  • Engineering & Computer Science (AREA)
  • Aviation & Aerospace Engineering (AREA)
  • Physics & Mathematics (AREA)
  • General Physics & Mathematics (AREA)
  • Manufacturing & Machinery (AREA)
  • Transportation (AREA)
  • Measurement Of Mechanical Vibrations Or Ultrasonic Waves (AREA)

Abstract

The defect detection system uses a generator providing sinusoidal or noise signals, with sensors measuring the resulting oscillation of different parts of the aircraft. A processor provides a transfer function using the quotient of the sensor signals and the generator signals, for modal analysis of the aircraft structure. The measured transfer function is approximated by mathematical functions to obtain the characteristic oscillation of the aircraft structure compared with that for a model of the aircraft structure to allow defects to be identified. ADVANTAGE - Allows defect to be detected and localised.

Description

Die Erfindung betrifft ein Verfahren zum Erkennen und Bewerten von Struk­ turschwächen von Flugzeugen entsprechend dem Oberbegriff des Anspruchs 1.The invention relates to a method for recognizing and evaluating structure weaknesses of aircraft according to the preamble of claim 1.

Ein solches Verfahren ist für den Fall "Sinussignale" aus der DE-OS 26 52 361 bekannt. Dabei wird das Flugzeug einer vorgegebenen räumlichen Verteilung von Kräften unter Verwendung von mit Rüttlern gekoppelten Er­ regersignalen ausgesetzt. Zur Messung der Antwortsignale, die durch Si­ nussignale erzeugt werden, sind Beschleunigungsmesser als Sensoren vor­ handen. Die Schwingungserzeugung erfolgt durch einen Generator. Es ist auch bekannt, die angelegten Schwingungssignale in einem Analog/Digital-Konverter zu digitalisieren. Die Aufgabe dieser bekannten Einrichtung besteht darin, ein Strukturanalysensystem anzugeben und die Transferfunktion eines Flugzeugs zu bestimmen. Eine Weiterverar­ beitung der Schwingungscharakteristik zum Erkennen von Strukturschwächen ist insofern vorgesehen, als daran gedacht wird, instabile Vibrationsschwingungsformen zu dämpfen oder durch Änderung der Konfiguration des Flugzeugs zu vermeiden.Such a method is for the case of "sine signals" from DE-OS 26 52 361 known. The aircraft becomes a predetermined spatial Distribution of forces using Er coupled with vibrators exposed to rain signals. To measure the response signals by Si nut signals are generated, accelerometers are provided as sensors act. The vibration is generated by a generator. It is also known, the applied vibration signals in one Digitize analog / digital converter. The task of this known Setup consists of specifying a structural analysis system and the To determine the transfer function of an aircraft. A further processing Processing the vibration characteristics for Detection of structural weaknesses is provided in that is thought to dampen unstable forms of vibration or by changing the configuration of the aircraft.

Aus dem Firmenprospekt JN 6506-0105/86 der Fa. Gen Rad, Milpitas, CA, USA "VIBRATION TESTING AND ANALYSIS" 1986 ist bekannt, die Schwingungs­ prüfung von Flugzeugen mit Gleitsinus vorzunehmen. Aus D. J. Inman, VIBRATION with Control, Measurement and Stability, Englewood Cliffs, USA, 1989, S. 12, 13, 184, 185, 204, 205, 294, ist bekannt, bei der Schwingungs­ prüfung eine Modal-Analyse vorzunehmen und eine Finite-Elemente-Modell aufzustellen. Aus "SHOCK AND VIBRATION HANDBOOK", C. Harris and E. Crede, eds 2. Auflage, McGraw-Hill Book Co./New York, 1976, S. 18-30, 18-31, 27-1, 27-2, 27-3, sind Finite-Elemente-Programme bekannt. Die Bestim­ mung einer Transferfunktion im Zusammenhang mit der Modal-Analyse ist aus der DE-PS 25 32 801 bekannt. Aus Materialprüfung 12 (1970), Nr. 11, November, Seiten 369 bis 372, ist die Steuerung von Schwingungsprüfma­ schinen durch Prozeßrechner und die Verwendung von Rauschgeneratoren be­ kannt. Aus "The Journal of the Acoustical Society of America", Vol. 41, Nr. 4, 1967, S. 840-848, ist die Verwendung von Gleitsinus- und Rauschgeneratoren bei der Schwingungsprüfung bekannt. Aus der DE 38 17 777 A1 ist bekannt, bei der Schwingungsprüfung von einer Welle mit Rotor (Prüfling) ein Modell des Prüflings zu verwenden und daran Defekte anzubringen.From the company brochure JN 6506-0105 / 86 from Gen Rad, Milpitas, CA, USA "VIBRATION TESTING AND ANALYSIS" 1986 is known the vibration check aircraft with sliding sine. From D. J. Inman, VIBRATION with Control, Measurement and Stability, Englewood Cliffs, USA, 1989, pp. 12, 13, 184, 185, 204, 205, 294, is known in the vibration to conduct a modal analysis and set up a finite element model. From "SHOCK AND VIBRATION HANDBOOK ", C. Harris and E. Crede, eds 2nd edition, McGraw-Hill Book Co./New York, 1976, pp. 18-30, 18-31, 27-1, 27-2, 27-3 Finite element programs known. The determinations a transfer function in connection with the modal analysis known from DE-PS 25 32 801. From material test 12 (1970), No. 11, November, pages 369 to 372, is the control of vibration test be through process computers and the use of noise generators knows. From "The Journal of the Acoustical Society of America", Vol. 41, No. 4, 1967, pp. 840-848, is the use of sliding sine and noise generators  known the vibration test. From DE 38 17 777 A1 it is known at the vibration test of a shaft with rotor (test object) a model of the device under test and attach defects to it.

Ausgehend von einem Verfahren gemäß dem Oberbegriff des Anspruchs 1 ist es die Aufgabe der Erfindung, dieses Verfahren derart weiterzubilden, daß damit Schäden am Flugzeug erkannt und lokalisiert werden können und die Restfestigkeit ermittelt werden kann.Starting from a method according to the preamble of claim 1 it is the object of the invention to develop this method in such a way that damage to the aircraft can be identified and localized and the residual strength can be determined.

Ausgehend von einem Verfahren zum Erkennen und Bewerten von Struktur­ schwächen von Flugzeugen der eingangs angegebenen Art, wird diese Aufga­ be erfindungsgemäß dadurch gelöst, daß die Strukturschwächen Struktur­ schäden sind, daß Gleitsinussignale konstanter Amplitude bzw. Rauschsi­ gnale verwendet werden, daß ein Finite-Elemente-Modell der Flugzeug­ struktur aufgestellt wird, dessen Schwingungsformen berechnet werden und als Referenz-Schwingungsformen dienen, daß die aus der Modal-Analyse ge­ wonnenen Schwingungsformen mit den zugehörigen Referenz-Schwingungsfor­ men verglichen und Abweichungen ermittelt werden, daß beim neuen Flug­ zeug bei Abweichungen das Finite-Elemente-Modell angepaßt wird, aber beim gebrauchten Flugzeug aus den Abweichungen gegenüber dem Finite-Ele­ mente-Modell Schäden lokalisiert werden, daß die lokalisierten Schä­ den ins Finite-Elemente-Modell übernommen werden, die Festigkeit des schadenbehafteten Finite-Elemente-Modells berechnet und als Restfestig­ keit des Flugzeugs angesetzt wird.Based on a process for recognizing and evaluating structure weaken aircraft of the type specified above, this task be solved according to the invention in that the structural weaknesses structure damage are that sliding sinusoidal signals of constant amplitude or noise gnale used to be a finite element model of the aircraft structure is set up, the waveforms are calculated and serve as reference waveforms that the ge from the modal analysis won waveforms with the associated reference vibration form men compared and deviations are determined that the new flight in the event of deviations, the finite element model is adapted, however in the used aircraft from the deviations from the finite el mente model damage can be localized that the localized Schä which are adopted in the finite element model, the strength of the damaged finite element model calculated and as residual strength speed of the aircraft.

Der große Vorteil besteht darin, daß jederzeit eine Struk­ turdiagnose am Flugzeug durchgeführt werden kann, wobei eine Bestimmung von Strukturschäden und der Restfestigkeit der Schadensstelle vorgenom­ men wird. Somit kann das Verfahren nach jedem Flug durchgeführt werden, wodurch der Wartungsaufwand deutlich reduziert werden kann. Weiterhin kann das Verfahren auch bereits bei der Entwicklung von Flugzeugen an Prototypen angewendet werden, um Strukturen mit zu hoher oder zu niedri­ ger Festigkeit zu erkennen und damit an Strukturgewicht zu sparen. The big advantage is that a structure is always available Door diagnosis can be carried out on the aircraft, with one determination of structural damage and the residual strength of the damage site men will. Thus, the procedure can be carried out after every flight, whereby the maintenance effort can be significantly reduced. Farther The method can also be used in the development of aircraft Prototypes are used to design structures with too high or too low recognizable strength and thus save on structural weight.  

Als Sensoren zum Messen der erregten Schwingungen werden gemäß Anspruch 3 zweckmäßig Beschleunigungsmesser oder Dehnungsmeßstreifen verwendet, von denen gemäß Anspruch 4 minde­ stens zwei in jede Tragfläche oder gemäß Anspruch 5 in den Rumpf eingebaut werden. Der Prozessor ist zum Weiterverarbeiten der erzeugten Schwingun­ gen und zum Vergleich von im Prozessor umgewandelten Eigenschwingungs­ formen mit Referenz-Schwingungsformen eines Finite-Elemente-Modells der Flugzeugstruktur zum Feststellen und Erkennen der Schäden und von zu er­ mittelnder Restfestigkeit erforderlich.As sensors for measuring the excited vibrations are useful according to claim 3 Accelerometers or strain gauges used, of which at least 4 according to claim least two are installed in each wing or according to claim 5 in the fuselage. The processor is for further processing the generated vibration conditions and for comparison of natural vibrations converted in the processor shape with reference waveforms of a finite element model Aircraft structure for the detection and detection of the damage and to it Average residual strength required.

Die Erfindung wird nachstehend anhand der Zeichnung in einem Ausführungsbeispiel näher erläutert. Es zeigtThe invention is illustrated below in a drawing Embodiment explained in more detail. It shows

Fig. 1 den Ablauf eines Verfahrens zum Erkennen und Bewerten von Struk­ turschäden an einem Flugzeug; Fig. 1 shows the sequence of a method for detecting and evaluating structural turschäden on an aircraft;

Fig. 2 Erläuterung des Prinzips des Verfahrens an einem Biegebalken und Fig. 2 Explanation of the principle of the method on a bending beam and

Fig. 3 Darstellung der Transferfunktionen des Biegebalkens von Fig. 2. Fig. 3 showing the transfer functions of the bending beam of FIG. 2.

Ein Flugzeug 1 hat einen Rumpf 2 und zwei Tragflächen 3 und 4. An jeder der Tragflächen 3 und 4 sind vier Beschleunigungsmesser 5 adaptiert. Die Tragflächen 3 und 4 werden in nicht dargestellter Weise über Ruderstell­ motoren von Querruderflächen 6 und 7 mit einem Gleitsinussignal konstan­ ter Amplitude P symmetrisch erregt. Das Gleitsinussignal konstanter Am­ plitude hat eine variable Frequenz. In dem Blockdiagramm der Fig. 1 ist schematisch der Verfahrensablauf dargestellt, der programmiert in einem im Rumpf 2 eingebauten Prozessor 8 abläuft. Infolge Erregung der Trag­ flächen 3 und 4 durch das Gleitsinussignal konstanter Amplitude P werden in den Sensoren 5 Antwortsignale A erregt. Im Prozessor 8 werden die Sensorsignale digitalisiert und es wird zu jedem Sensorsignal A eine Transferfunktion F(jω) als Quotient von Sensorsignal A zum Gleitsinus­ signal konstanter Amplitude P gebildet. Die Transferfunktionen F(jω) werden mit Modal-Analyse weiterverarbeitet, wobei die gemessenen Trans­ ferfunktionen F(jω) durch mathematische Funktionen aproximiert wer­ den. Die mathematischen Funktionen werden verwendet zur Berechnung von Eigenschwingungsformen. Die berechneten Eigenschwingungsformen werden mit einer Referenz-Schwingungsform eines dynamischen Finite-Elemente- Modells der Flugzeugstruktur verglichen, das zur Entwicklung des Flug­ zeugs erstellt und bei der Flugversuchserprobung verwendet wurde. Bei Abweichungen werden die Eigenschwingungsformen am Finite-Elemente-Mo­ dell angepaßt und es werden aus den Abweichungen gegenüber dem Finite- Elemente-Modell die Schäden lokalisiert sowie durch Vergleich mit dem Sollwert der Struktur die Restfestigkeit an der Schadensstelle ermittelt.An aircraft 1 has a fuselage 2 and two wings 3 and 4 . Four accelerometers 5 are adapted to each of the wings 3 and 4 . The wings 3 and 4 are excited in a manner not shown via rudder actuators of ailerons 6 and 7 with a sliding sinusoidal constant amplitude P symmetrically excited. The sliding sinusoidal signal at constant amplitude has a variable frequency. In the block diagram of FIG. 1, the process sequence is shown schematically, which is programmed in a processor 8 installed in the fuselage 2 . As a result of excitation of the wings 3 and 4 by the sliding sine signal constant amplitude P 5 response signals A are excited in the sensors. The sensor signals are digitized in processor 8 and a transfer function F (jω) is formed for each sensor signal A as a quotient of sensor signal A to the sliding sine signal of constant amplitude P. The transfer functions F (jω) are further processed with modal analysis, the measured transfer functions F (jω) being approximated by mathematical functions. The mathematical functions are used to calculate natural vibration forms. The calculated natural waveforms are compared with a reference waveform of a dynamic finite element model of the aircraft structure, which was created for the development of the aircraft and used in the flight test. In the event of deviations, the natural vibration forms are adapted to the finite element model and the damage is localized from the deviations from the finite element model, and the residual strength at the point of damage is determined by comparison with the nominal value of the structure.

Angenommen, es sei am Tragflügel 3 eine beschädigte Stelle 9 ermittelt worden, bei der noch eine Restfestigkeit von 70% vorhanden ist. Als Konsequenz daraus darf vor der Ausbesserung der Stelle 9 die Tragfläche nicht mehr so hoch belastet werden. Der Pilot darf z. B. bei Flugmanövern das Flugzeug an Stelle von 9 g nur noch mit 6,3 g belasten.Assume that a damaged area 9 was found on the wing 3 , in which there is still a residual strength of 70%. As a consequence of this, the wing may no longer be subjected to such high loads before the repair of point 9 . The pilot may e.g. B. in flight maneuvers the aircraft instead of 9 g only with 6.3 g.

In Fig. 2 ist das Prinzip des in Fig. 1 erläuterten Verfahrens an einem Biegebalken 11 dargestellt, dessen Masse M durch die dargestellten Kur­ ven symbolisiert ist. Bei einer Belastung schwingt der unbeschädigte Biegebalken 11 in die dargestellten Endstellungen 11a nach oben und un­ ten aus, dabei ist der Ausschlag nach unten die Grundbiegung und nach oben die Oberbiegung des Balkens 11. Die auftretende gedämpfte Schwin­ gung beträgt f(t) = mx + dx + cx. Dabei ist mx die Masse, dx die Dämp­ fung und cx die Steifigkeit. Ein beschädigter Balken hat eine bei der gleichen Belastung entsprechend den Kurven 11b nach oben und unten um den Betrag ΔX größere Schwingungsweite. Die auftretende gedämpfte Schwingung ist in dem Diagramm der Fig. 3 gezeigt, in der in Abhängig­ keit von der Frequenz ω die jeweiligen Amplituden A der Transferfunk­ tionen F(ωa) und F(ωb) dargestellt sind.In Fig. 2, the principle of the method explained in Fig. 1 is shown on a bending beam 11 , the mass M is symbolized by the ven shown ven. When there is a load, the undamaged bending beam 11 swings up and down from the end positions 11 a shown, the deflection downward is the basic bend and upward the upper bend of the beam 11 . The damped vibration occurring is f (t) = mx + dx + cx. Here mx is the mass, dx the damping and cx the stiffness. A damaged bar has a B at the same load corresponding to the curves 11 up and larger by the amount .DELTA.X down vibration width. The occurring damped oscillation is shown in the diagram of FIG. 3, in which the respective amplitudes A of the transfer functions F (ω a ) and F (ω b ) are shown as a function of the frequency ω.

Claims (6)

1. Verfahren zum Erkennen und Bewerten von Strukturschwächen von Flugzeugen, bei dem ein Generator für Erregersignale, nämlich Sinus- oder Rauschsignale verwendet wird, bei dem Schwingungen messende Sen­ soren am Flugzeug angebracht werden, bei dem mit den genannten Erreger­ signalen Schwingungen in der Flugzeugstruktur erregt und die Schwingun­ gen durch die Sensoren gemessen werden, bei dem die genannten Erregersi­ gnale und die von den Sensoren gemessenen Signale digitalisiert und in einen Prozessor eingespeist werden, bei dem im Prozessor Transferfunktio­ nen als Quotient aus Sensorsignalen und Erregersignalen gebildet werden und bei dem eine Modal-Analyse der Flugzeugstruktur vorgenommen wird, wobei die gemessenen Transferfunktionen durch mathematische Funktionen aproximiert werden und damit die Eigenschwingungsformen der Flugzeug­ struktur berechnet werden, dadurch gekennzeichnet, daß die Struktur­ schwächen Strukturschäden sind, daß Gleitsinussignale konstanter Ampli­ tude bzw. Rauschsignale verwendet werden, daß ein Finite-Elemente-Mo­ dell der Flugzeugstruktur aufgestellt wird, dessen Schwingungsformen be­ rechnet werden und als Referenz-Schwingungsformen dienen, daß die aus der Modal-Analyse gewonnenen Schwingungsformen mit den zugehörigen Refe­ renz-Schwingungsformen verglichen und Abweichungen ermittelt werden, daß beim neuen Flugzeug bei Abweichungen das Finite-Elemente-Modell ange­ paßt wird, aber beim gebrauchten Flugzeug aus den Abweichungen gegenüber dem Finite-Elemente-Modell Schäden lokalisiert werden, daß die lokali­ sierten Schäden ins Finite-Elemente-Modell übernommen werden, die Fe­ stigkeit des schadenbehafteten Finite-Elemente-Modells berechnet und als Restfestigkeit des Flugzeugs angesetzt wird. 1. A method for recognizing and evaluating structural weaknesses of aircraft, in which a generator for excitation signals, namely sine or noise signals is used, in which vibration-measuring sensors are attached to the aircraft, in which signals excite vibrations in the aircraft structure with the said exciters and the vibrations are measured by the sensors, in which the said excitation signals and the signals measured by the sensors are digitized and fed into a processor, in which transfer functions are formed in the processor as a quotient of sensor signals and excitation signals and in which a modal -Analysis of the aircraft structure is carried out, the measured transfer functions being approximated by mathematical functions and thus the natural vibrations of the aircraft structure being calculated, characterized in that the structure is weak structural damage, that sliding signals of constant amplitude or R Exchange signals are used that a finite element model of the aircraft structure is set up, the waveforms are calculated and serve as reference waveforms, that the waveforms obtained from the modal analysis are compared with the associated reference waveforms and deviations are determined that the finite element model is adjusted in the new aircraft in the event of deviations, but damage is localized in the used aircraft from the deviations from the finite element model, that the localized damage is transferred to the finite element model, which The strength of the damaged finite element model is calculated and used as the residual strength of the aircraft. 2. Verfahren nach Anspruch 1, dadurch gekennzeichnet, daß die Er­ regung mit den Gleitsinussignalen konstanter Amplitude bzw. Rauschsigna­ len mit flugzeugeigenen hydraulisch oder elektrisch angetriebenen Ruder­ stellmotoren von Steuerflächen vorgenommen wird.2. The method according to claim 1, characterized in that the Er excitation with the sliding sinusoidal signals of constant amplitude or noise signal len with the aircraft's own hydraulic or electric rudder Actuators of control surfaces is made. 3. Verfahren nach Anspruch 1, dadurch gekennzeichnet, daß als Sen­ soren Beschleunigungsmesser oder Dehnungsmeßstreifen verwendet werden.3. The method according to claim 1, characterized in that as Sen accelerometers or strain gauges are used. 4. Verfahren nach Anspruch 3, dadurch gekennzeichnet, daß minde­ stens zwei Sensoren in jede der Tragflächen eingebaut werden.4. The method according to claim 3, characterized in that mind at least two sensors are installed in each of the wings. 5. Verfahren nach Anspruch 3, dadurch gekennzeichnet, daß minde­ stens zwei Sensoren in den Flugzeugrumpf eingebaut werden.5. The method according to claim 3, characterized in that mind at least two sensors are installed in the fuselage. 6. Verfahren nach Anspruch 1, dadurch gekennzeichnet, daß die Ab­ weichungen beim Vergleich der Eigenschwingungsformen mit den Referenz-Schwingungsformen und damit die Strukturschäden aus unterschiedlichen Amplitudenhöhen ermittelt werden.6. The method according to claim 1, characterized in that the Ab softening when comparing the natural vibration forms with the Reference waveforms and thus the structural damage different amplitude levels can be determined.
DE4240600A 1992-12-03 1992-12-03 Structural defect detection system for aircraft - uses modal analysis via detected oscillation of aircraft structure and comparison with aircraft model Expired - Fee Related DE4240600C1 (en)

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US7506549B2 (en) 2003-08-27 2009-03-24 Airbus Uk Limited Method and apparatus suitable for measuring the displacement or load on an aircraft component
DE102007055090A1 (en) * 2007-11-16 2009-05-28 Eads Deutschland Gmbh Monitoring device for monitoring the outer skin of an aircraft
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