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The Effects of Damage and Uncertainty on the Aeroelastic /Aeroservoelastic Behavior and Safety of Composite Aircraft

A Progress Report, JAMS Meeting July 22nd 2009University of Washington – AMTAS

Eli Livne, ProfessorDepartment of Aeronautics and Astronautics

eli@aa.washington.edu

The team:

Department of Aeronautics and Astronautics: Dr. Eli Livne – PI, Professor;

Department of Mechanical Engineering: Francesca Paltera, PhD student and Dr. Mark Tuttle,co-PI, professor and chairman;

at Boeing Commercial, Seattle: Dr. James Gordon, Associate Technical Fellow, FlutterMethods Development, and Dr. Kumar Bhatia, Senior Technical Fellow, Aeroelasticity andMultidisciplinary Optimization;

at the FAA: Technical Monitor - Curtis Davies, Program Manager of JAMS, FAA/Materials &Structures, Dr. Larry Ilcewicz, Chief Scientific and Technical Advisor for Advanced CompositeMaterials, and Carl Niedermeyer, FAA Airframe and Cabin Safety Branch (previously, Boeingflutter manager for the 787 and 747-8 programs)

Motivation and key issues:

Material degradation and damage in composite airplanes operating for many years in diverseoperational and maintenance environments can lead to degradation of the flutter (aeroelasticstability) and gust response (dynamic loads) capability of the airframes. The goals of the FAAfunded R&D effort entitled: “The Effects of Damage and Uncertainty on the Aeroelastic /Aeroservoelastic Behavior and Safety of Composite Aircraft” are:

To develop computational tools (validated by experiments) for automated local/globallinear/nonlinear analysis of integrated structures/ aerodynamics / control systems subject tomultiple local variations/ damage;

To develop aeroservoelastic probabilistic / reliability analysis for composite actively controlledaircraft;

To link with design optimization tools to affect design and repair considerations.

To develop a better understanding of effects of local structural and material variations incomposites on overall Aeroservoelastic integrity;

Progress 2008-2009

Work in 2008-2009 at the University of Washington’s Departments of Aeronautics &Astronautics and Mechanical Engineering focused on the experimental thrust of the overall“Aeroelasticity of Uncertain Composites” R&D program. In this effort:

Modal testing equipment and aeroelastic wind tunnel instrumentation and data acquisition andprocessing equipment were used to test an aeroelastic model of a tail / rudder system, with acomposite-construction rudder.

In a first series of computational simulations and tests the behavior of the system with anumber of composite rudders was studied. The different rudders included different foam coredensities, different damage (delamination) of the rudder itself as well as failure of a number ofhinges attaching the rudder to the tail structure.

In a second series of tests the structural dynamic characteristics of the system were modifiedby introducing new pitch, plunge, and rudder stiffnesses, and then freeplay was introduced inthe rudder rotation attachment to simulate loosening of actuator attachments or the effect oflocal damage in actuator attachment areas of actual tail / rudder systems.

Linear flutter simulations and tests are described in the following short paper presented at theSEM-2009 meeting this year

Flutter Response To Damage Of Composite Aircraft Control Surfaces

(presented at the Society of Experimental Mechanics – SEM - Conference, 2009)

Francesca Paltera, Mark E. Tuttle, Eli LivneDepartment of Mechanical Engineering, University of Washington, Stevens Way, Box 352600

Seattle, WA 98195, U.S.Apaltera@u.washington.edu

ABSTRACT

An experimental model of a typical three degree-of-freedom airfoil section with compositecontrol surface subjected to a two-dimensional, incompressible air-flow has been created. Theflutter response of the system with both, a defect-free and with damaged control surfaces,were measured experimentally. An analytical model was implemented to predict the flutterresponse in the case of a defect-free control surface. Flutter predictions were obtained usingboth the U-G method and the Root Locus Technique. Measurements obtained using thedefect-free rudder compared well with predictions, validating the analytical model and flutterpredictions. Measurements with the damaged rudders show that some forms of damage mayhave little impact on flutter speed, whereas other forms of damage may have a pronouncedeffect. In general, damage to the rudder decreases the flutter speed.

INTRODUCTION

Air flowing over a flexible surface, for example an aircraft wing, results in aerodynamic forcesapplied to the surface. The resulting surface deformations may increase or decrease the initialaerodynamic force. An increase in aerodynamic force will cause still greater surfacedeformations, whereas a decrease will cause an elastic recovery of surface deformations.Hence airflow can cause a feedback process involving aerodynamic, elastic, and inertialforces. Aeroelasticity is the study of this feedback process. Flutter is one of several importantaeroelastic phenomena. Flutter is a structural vibration caused by a constant, steady-state airflow over the surface. In extreme cases flutter can cause serious structural damage. Althoughmost commonly associated with aircraft, flutter can occur for any elastic structure exposed toa steady-state airstream.

This paper describes a laboratory model representing a three degree-of-freedom airfoil sectionwith composite control surface that has been created to study aeroelastic effects observedduring wind-tunnel tests. The focus of current studies is on flutter, although the model hasbeen designed to allow interrogation of additional aeroelastic factors such as limit cycleoscillation, buffeting, or control system reversal. Previous studies [1] compared flutter speedsmeasured experimentally to predictions, obtained using both the U-G method [2] and the RootLocus Technique [3], for the system with defect-free control surfaces. Both these numericalmethods and the laboratoty model developed in the present study are based on the state-space model developed by Edwards et al. [4] and summarized in Figure 1. During the presentwork flutter speeds exhibited by an airfoil with an intentionally damaged composite controlsurface is measured experimentally.

EXPERIMENTAL MODEL

Design and construction of the model is summarized in Figure 2. The model is based on thesymmetric NACA 0012 cross-section and resembles a model developed by Conner et al [5].The model consists of two parts: a main aluminum wing with 40 cm chord (width) and 90 cmspan (length), and a composite control surface with a 13 cm chord and 90 cm span. Thecontrol surface resembles an aileron, flap, or rudder used in actual aircraft wing structures.For clarity it will herein be called the aileron. The aileron is mounted to the trailing edge of themain wing using micro-bearings. The main wing is constructed using an aluminum circularspar that passes through 18 aluminum ribs. The ribs effectively define the NACA 0012 cross-section. The circular spar serves as the pitch axis of the entire structure, and is located at thequarter-chord location from the leading edge. A 0.64 mm thick aluminum sheet was bonded tothe periphery of the 18 internal ribs, defining the external aerodynamic surface. In addition, 2aluminum tubes with 2.54 cm diameter were mounted 5 cm and 38 cm, respectively, from theleading edge. The chordwise center of gravity of the model can be adjusted by adding orsubtracting balance weights from these tubes.

The control surface is constructed using a low-density urethane foam 10 lb/ft3, covered with a[0º/90º/0º]T graphite/epoxy facesheet. An aluminum spar tube (1.5 cm diameter and 0.4 mmwall-thickness) is embedded in the foam and passes through the leading edge of 4 aluminumribs that complete the NACA 0012 trailing edge. The rotational degree of freedom of thecontrol surface relative to the main wing is ± 35°. One end of a relatively stiff 1.6 mm diameter

steel wire is inserted tightly into a slot at one end of the aileron spar beam. The other end ofthe wire is inserted into a support block mounted on the main wing. This wire serves as a leaf-spring and restricts the rotation of the aileron relative to the main wing. The spring stiffnesscan be adjusted by moving the support block so as to increase or decrease the wire lengthbetween supports, or by changing the diameter of the steel wire.

The main wing spar extends through the top and bottom of the 91 cm x 91 cm wind tunnel testsection and is attached to two movable support blocks (see Figure 3). Resistance to plungemotion is provided by two 0.21 cm x 7.5 cm x 51 cm steel leaf-springs. Hence, the two supportblocks move with the main spar, and this motion is restricted to the (transverse) plungedirection. The spar passes through two precisions bearings mounted to the support blocks. Anadjustable pitch rotational stiffness is provided by a steel spring wire, inserted at one end intothe main spar tube and into a bracket on the support block at the other. This design allows forplunge motion that is independent of pitch and also allows for easy modification of plunge andpitch stiffnesses.

The focus of current studies is on the impact that damage within a composite control surfacemay have on flutter speeds. In order to accomplish this purpose, a total of 3 control surfaces,with different kinds of induced damage, were fabricated. Debonding was achieved by using:1) mold release between the shaft tube and the composite;2) Teflon between the foam and the composite;3) combining case 1) and 2).

STRUCTURAL CHARACTERIZATION OF THE AIRFOIL MODEL

Two Uniaxial Accellerometers PCB 352C22 from Pcb Piezotronic were installed in thestructure, on the rudder and on the wing, at the locations showed in Figure 4. Theaccelerometers were used in concert with a Laser Vibrometer (OFV 2600 VibrometerController – OFV 302 Sensor Head) from Polytec, to measure accelerations anddisplacements of the structure during wind tunnel tests. Both the accellerometers and the laservibrometer were monitored using a Jaguar Data Acquisition System from Spectral Dynamic toobtain the time and frequency response of the airfoil.

The experimental model was structurally characterized through a series of static tests in orderto measure stiffnesses and damping ratios corresponding to each of the three-degrees-of-freedom. One degree of freedom at a time was allowed to remain free in the structure, whilethe remaining two degrees of freedom were rigidly fixed. Stiffnesses were measured usinghand-held force gages and displacement sensors.

Uncoupled natural frequencies, coupled natural frequencies and damping ratios were alsomeasured for the system, using a hand-hammer Impact Sensor PCB 086C01 from PcbPiezotronics, for exciting the structure, and the Jaguar System for recording both time andfrequency response of the system. Several sets of springs were used in order to modify thestiffness of the structure, and hence the flutter speed. Uncoupled natural frequencies, couplednatural frequencies and damping ratios were measured again every time a spring waschanged.

EXPERIMENTAL RESULTS

The aluminum airfoil was assembled with each of the five manufactured control surfaces, oneat a time. A multiple series of flutter tests were performed on each particular airfoil/controlsurface combination. Time and frequency response of the system was monitored during eachtest and recorded every time the flutter speed was reached. Experimental data were obtainedthrough the accellerometers installed on the structure and the laser vibrometer pointed on thecontrol surface. The flutter tests were all conducted using the same spring set for the plunge,the pitch and the flap motion, except for one case where different springs where adopted forthe plunge and the pitch motion.

In general, starting of the flutter phenomenon is associated with increasingly rapid oscillationsof the structure as shown in Figure 5 (a) and (b), which rapidly brings the system to oscillate ina simple periodic motion when pure flutter is reached, as can be seen in Figure 5 (c).Frequency of flutter can be read correspondent to the peak that appears in the frequencyresponse plot of the system, as shown in Figure 5 (d).

With linear stability theory the flutter speed is the speed at which the system becomes unstable(and oscillations begin to diverge in time). What it can be seen in tests as fully developedsimple harmonic motion (Figure 5 (c)) is a limit cycle oscillation (LCO) that is made possible,probably, by the nonlinearities in the springs that restrains the amplitude and prevents theoscillation from diverging to infinity (destruction).

Also, coupled natural frequencies were measured for the whole assembled structure. Thefrequency response for the three coupled degrees-of-freedom were analyzed using the JaguarSystem after that the system were excited using the hand-hammer Impact Sensor. A typicalresponse can be seen in Figure 6 where the three higher peaks correspond to the threecoupled natural frequencies.

A total of four different aileron/control surface configurations with induced damage were tested,showing how flutter speed and flutter frequency vary from one case to another. In one of thefour configurations, three bushings were removed along the hinge of the wing in order tosimulate a broken hinge.

CONCLUSIONS

A study into the damage dependency and experimental measurements of the flutter speed of athree-degree-of-freedom airfoil section with composite control surface was conducted under atwo-dimensional, incompressible air- flow. Damage was induced at different locations of thecontrol surface and the effect of several kinds of damages on the flutter response wasinvestigated.

Generally, it has been determined that the presence of damage reduces the stiffness of thestructure, having the effect of reducing the flutter speed of the aircraft. Additional insight intothe effects of damage size, location and form were needed in order to more fully characterizethe response of the wing structure, in regards to the flutter phenomenon.

The conclusions of this study are as follows:1. The flutter speed decreases when the center of gravity of the control surface is moved

closer to its trailing edge.2. The flutter speed decreased due to debonding induced in the control surface at the interface

just aft of the hinge line but forward of the trailing edge. Possible skin flutter was identifiedprior to the system’s flutter based on the interpretation of the time response.

3. The flutter speed decreased due to induced free-play in the control surface at the hinge line,simulating localized hinge bearing damage.

4. The flutter speed substantially decreased when the control surface was free to rotate aboutthe hinge, simulating a broken control surface actuator.

Figure 1 Schematic of the aeroelastic model

Figure 2 Design and manufacturing of the model

Figure 3 Support Blocks Figure 4 Accellerometers Locations

Figure 5 a) time response before to reach flutter; b) time response as approaching flutter; c)time response during flutter; d) frequency response during flutter

Figure 6 Coupled natural frequencies for the system

REFERENCES

[1] F. Paltera, M.E. Tuttle, B. Kuykendall, A Composite Airfoil Section Used to StudyAeroelastic Effects, Proceedings of SEM 2008 Annual Conference and Exposition[2] Evans, W. R., Control-system Dynamics, McGraw-Hill Book Company, pp.117, 1954.

[3] Theodore Theodorsen and I. E. Garrick, Mechanism of flutter a theoretical andexperimental investigation ofthe flutter problem, N.A.C.A. Report No. 685, 1940[4] Edwards, J. W., Ashley, H. & Breakwell, J. V. Unsteady aerodynamic modeling for arbitrarymotions, AIAA

Journal 17, 365-374,1979[5] M. D. Conner, D.M. Tang, E. H. Dowell and L.N. Virgin, Non linear behavior of a typicalairfoil section withcontrol surface freeplay: a numerical and experimental study, Department of MechanicalEngineering andmaterials science, Duke University Durham, NC 27708-0300, U.S.A., 1996

Aeroelastic and Simulations and Wind Tunnel Tests of a Tail / Rudder System withRudder Freeplay

The UW 2D 3dof tail / rudder aeroelastic model is shown in Fig. 7. The figure also showspossible damage scenarios to be studied in the work planned for the 2009-2010 academicyear.

Figure 7: The UW tail / rudder aeroelastic model (following a Duke University model design)

Rudder rotation stiffness is critical in determining the aeroelastic stability and gust response ofthe system, and any degradation or damage may impact the aeroelastic behavior of the wholesystem dramatically. If the resulting stiffness nonlinearity is of the “hardening” type, that is, ifwith larger amplitudes of motion the rudder “sees” higher stiffness than what is encountered atsmall amplitudes, then at speeds well below the flutter speed, the system may exhibitsustained limit cycle oscillations. If, however, damage leads to “softening” nonlinearity, that is,reduced stiffness at higher motion amplitudes, then the result may be destructive flutter wellbelow the flutter speed for which the undamaged structure was designed (Fig. 8).

Figure 8: The effect of different stiffness nonlinearities on the aeroelastic behavior of tail /rudder systems.

Modifications of the UW aeroelastic model to improve pitch stiffness simulation are shown inFig. 9 below:

Figure 9: 2009 UW Aeroelastic model improvement

Predicted flutter speeds as a function of rudder rotational hinge stiffness as well as predictedrudder LCO amplitudes are shown in Fig. 10:

Figure 10: Flutter speed of the UW tail / rudder system with varying rotational stiffness, andpredicted LCO amplitudes of the UW aeroelastic tail / rudder system.

And initial LCO results captured in the wind tunnel at a speed about 81% lower than the flutterspeed of the system without freeplay are shown in Fig. 11:

Figure 11: Preliminary LCO tests of the UW Tail / Rudder system with freeplay.

While varying stiffness in the cases described above is done by changing translational androtational springs, the simulations and tests are aimed at capturing the aeroelastic behavior ofcomposite aeroelastic system subject to degradation and damage. Local effects in suchstructures can translate to overall impact on key stiffness elements, and thus affect the globalaeroelastic behavior of the structures.

Plans for the 2009-2010 year

Bring the UW emerging aeroelastic test capabilities to full productivity and carry outsimulations and tests to study:

Freeplay and other hinge rotation nonlinearities (including dampers actine with no actuatorstiffness, to represent actuator failure);

Rudder models representing realistic composite rudder designs in the presence of skindamage of different sizes, as well as rudder hinge failure.