FLUTTER SUBSTANTIATION TESTS ON A TRANSAVIA … · The Transavia Airtruk has been in production for...

7 Al'O 895 AERONAUTICAL RESEARCH LABS MELBOURNE (AUSTRALIA) F/B 1/3

FLUTTER SUBSTANTIATION TESTS ON A TRANSAVIA PL 12/T-300 AIRTRUK--ETC(U)

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DEPARTMENT OF DEFENCE SUPPORT

q DEFENCE SCIENCE AND TECHNOLOGY ORGANISATION

1 i AERONAUTICAL RESEARCH LABORATORIES

mMELBOURNE, VICTORIA

Stntures Tachnical Mmorandum 341

FLUTTER SUBSTANTIATION TESTS ON A TRANSAVIA

FL-12/T-30O AIRTRUK

A. GOLDMN

ELECTE

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DEPARTMENT OF DEFENCE SUPPORT

DEFENCE SCIENCE AND TECHNOLOGY OGAXISATCIGN

AERONAUTICAL RESEARCH LABORATORIES

Structures Technical Mtmorandum 341

FLUTTER SUBSTANTIATION TESTS ON A TRANSAVIA

PL-12/T--3(0 AIPTRUK

by

A. 3OLDMAN

SU'JMMRY

A resonance test and subsequent flight tests havebeen conducted on a Transavia Airtruk. The natural modesand frequencies of vibration were measured in the groundtests, and attempts made to induce flutter during flighttests. The results of these tests are presented.

Q COMMONWEALTH OF AUSTRALIA

POSTAL ADDRESS: Chief Superintendent, Aeronautical Research Laboratorien,P.O. Box 4331, Melbourne, Victoria, 3001, Australia.

"-- .... -' " _8 I

CONTENTS

PAGE NO.

1. INTRODUCTION 1

2. TEST EQUIPMENT AND PROCEDURE 1

2.1 Ground Resonance Test 1

2.2 Flight Tests 4

3. TEST RESULTS 5

3.1 General 5

3.2 Details of Modes 5

3.3 Flight Tests 7

4. DISCUSSION 8

5. CONCLUSIONS 9

REFERENCES

TABLES

FIGURES

DISTRIBUTION

DOCUMENT CONTROL DATA

Lccossion For

[TS t r l

Av ii11 d/or

Dist Special

1. INTRODUCTION

The Transavia Airtruk has been in production for 10 years,more than 100 having been sold, and there is no record of loss dueto structural failure in this time. In order to expand overseasmarkets, particularly in North America, full F.A.A. certification wasrequired and the Aeronautical Research Laboratories were requestedto assist with the flutter substantiation tests.

The aircraft has been frequently modified since the firstproduction model and the aircraft tested was as detailed in Table 1,with major dimensions shown in Fig. l(a). For the ground resonancetests, the aircraft was set up on partially inflated w-in wheelswith the nose suspended, on a soft spring system, from the factoryroof structure. With fuel tanks 50 per cent full and the hopperempty, the rigid body modes of pitch, roll, and heave were 1.4 hertz,4.2 hertz and 4 hertz respectively.

The aircraft was complete and serviceable, with all componentspresent in their correct locations.

The control surfaces were clamped to the main airframe forall the airframe modes except for the mode at 39.4 hertz, and freedfor all the control surface measurements except tab rotation.

The undercarriaqe had the hydraulic dampers replaced withfixed struts, but the dampers were attached to the wheels to retainthe correct mass distribution.

The same aircraft was used in the flight tests. All controlcables were at the lower service limits of tension. No attempt wasmade to simulate in-service backlash increases.

2. TEST EQUIPMENT AND PROCEDURES

2.1 Ground Resonance Tests

A maximum of ten electro-magnetic vlbrators was used forthese tests. These vibrators were nominally identical having a maximumthrust of 138 newtons and diaphragm stiffness of approximately 5kiloneytons/metre. The vibrators were initially located at theextremities of the wing front and rear spars, the port front sparand starboard rear spar of each tailplane, and in a horizontaldirection at the lowest point of each fin.

The vibrators were mounted on adjustable height stands andattached to the aircraft structure through light-weiqht, telescopicrods using ball-jointed connectors bonded to the skin with a quick-setting adhesive. Relocation of vibrators was found to be necessary

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for the wing torsion modes where the rear spar was very close to thenoda2 line, and for the control surface modes. For the wing torsionmodes the two vibrators on the rear spars were relocated towards thewing trailing edges, and for the control surface modes vibrators weregenerally located at the trailing edge of the control surfaces.

The vibrators were driven by high output-impedance power-amplifier3 to ensure that the applied forces were always in phase,irrespective of the motion of the structure. Frequency control wasprovided by a variable oscillator.

Accelerometers were located on the structure directly overeach vibrator attachment and at six other locations, on tLie tailplanesand fins, to provide an indication of the phase of the response relativeto the oscillator output signal. The sixteen accelerometer signals,and a signal from the oscillator in quadrature with the driving signal,were passed through a bank of phase-matched low-pass filters havinga roll-off frequency of 45 hertz. This was to remove distortions inthe signals caused by local panel resonances and other noise. Thesignals were then fed into a Resolved Component Ratiometer (RCR) anda multi-channel oscilloscope. The RCR operates by resolving eachaccelerometer signal into its components in phase and quadrature witha reference signal, summing the similar components, displaying theaverage value of in-phase and quadrature components, and displayingthe ratio of the quadrature to the in-phase components of the response.Using a reference signal in quadrature with the forcing signal makesthis ratio small at resonance. A pure resonant mode woull provide aratio of zero, whereas, if a reference was used in phase with theforce, the ratio would approach infinity and overload the displayof the ratiometer.

Likewise, the use of the quadrature reference signal providesLissajous figures on the oscilloscope which are straight lines atresonance. Use of a signal in phase with the force would providecircles at resonance, which would make tuning rather difficult.

As an aid to determination of the resonant frequencies ofthe airframe, a random noise signal, of bandwidth 0 to 50 hertz, wasfed into a vibrator at three consecutive locations - wing tip, tail-plane tip, and fin tip. The response of the accelurometer directlyover the vibrator was analysed, together with the random signal, ona dual-channel real-time spectrum analyser. In Fig. 2 the transferfunction of response, with reference to random input, is displayedand the peaks annotated with the modes that were subsequently foundclose to the frequencies indicated. The only modes found which didnot show up on these responses were wing torsion, and some of theboom bending and torsion modes. At each of the resonant frequenciesindicated by the spectrum analyser, the force levels in all the

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vibrators were manually adjusted until the acceleration responses atall the monitoring stations were as close as possible to being inquadrature with the exciting force. This was determined by observationof the Lissajous figures, and the RCR display, to obtain a conditionwith all figures as close as possible to straight lines, and the ratiodisplayed being a minimum. with a complex structure it is generallynot possible, with a limited number of force inputs, to obtain a pureresonant mode.

Hodes were found at all the frequencies indicated on thespectrum analyser. Other modes were found at frequencies not indicatedinitially. Modal measurements were made with a single movableaccelerometer attached to the structure, by a suction cup, successivelyat each point of the array of measuring stations. The layout of themeasurement points on the aircraft is indicated in Fig. 1(b) in whichhalf the aircraft is shown. The distribution of measurement pointswas symmetrical about the centre line. The actual locations are listedand dimensioned in Table 2.

The accelerometer signal was fed to a digital transferfunction analyser, which displayed he in-phase and quadrature componentsof the response with reference to the forcing signal. Modal dampingmeasurements were obtained from records of the decay rates of thevibration when the exciting signal from the oscillator was disconnected.

On completion of the airframe vibration mode measurements,the vibrators were disconn~ected from the structure, and the controlsurfaces freed by removing all clamps. The airframe was then rigidlysupported at the wings, booms and tailplanes usinq contour boards onstands with the airframe held onto them with bags of sand. This wasdone to permit measurement of the uncoupled control surface modes.Accelerometers and vibrator s were relocated to suit the individualcontrol surfaces, and resonant modes excited as before, using randominput and the spectrum analyser to detc - resonances. The oscillatorwas then used to fine-tune each resonant vibration mode. The aileron,elevator and rudder resonances were obtained by attaching one ormore vibrators to the trailing edges of the respective control surface.The tab rotational frequencies were obtained by clamping the elevatorJ1 to the tailplane and attaching the vibrator to the elevator. The onlyattachment to the tab was an accelerometer on the trailing edge. Arandom signal was fed into the vibrator and the transfer function ofthe tab was determined on the spectrum analyser for the three positionsof each tab. The tabs are of the spring loaded type.

The elevator control is connected to a downspring which holdsthxe control column in a full forward position when released. To holdthe column central required several restrainin7 rubber cords. The teston the elevators was repeated with a single rubaber cord on the controlcolumn to try and simulate a hand held situation, and also with thedownspring removed completely.

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2.2 Flight Tests

Lightweight accelerometers, havirv a mass of 8 grams each,were fixed to the aircraft structure at thirteen locations as listedin Table 3. Wherever possible, the accelerometers were fixed insidethe structure. The accelerometers on the elevator and rudder weremounted externally. In all locations the accelerometers were fixedto the structure with a quicksettin' epoxy resin adhesive. On theexternal surfaces they were also faired with tape to improve airflow.Cables were run, under covers and along wing trailing edges, toamplifiers located in the passenger compartment. This compartmentalso housed the power supplies, DC to AC invertor, and a 14-trackinstrumentation tape-recorder. A remote control station and microphonewere installed in the pilot's compartment.

Two test flights were made, to cover a total of 12 testpoints. This involved making several dives from approximately 3400metres to 1200 metres altitude and holdin each required airspeedsteady while the control stick was sharply struck in both fore andaft, and sideways directions, to induce a sudden disturbance on thesurfaces. The test points, and air speeds are listed in Table 4.

The records were replayed into a multi-channel chart recorderto observe the actual time histories at each test point. In orderthat the traces were not swamped by noise signals outside the rangeof interest, the signals were passed through a bank of low-pass filtershaving a roll-off at 45 hertz. Samples of this time history arepresented in Fig.15.

To investigate the frequency response of the structure tothe general turbulence present at different speeds the recorded datawere analysed on a Fourier analyser, two channels at a time. Thespectra obtained for the speeds of 145 and 175 knots are presentedin Figs. 16 to 18. In order to obtain the impulse response of thestructure, the Random Decrement procedure was used. This procedureis based on methods described in Ref. 1. Two techniques were usedto produce Figs 19 to 22 which illustrate the random decrementsignatures at different parts of the structure. Initially, the datawere passed through a low-pass filter set successively at 10, 15, 20,25 and 30 hertz to examine responses in each of these ranges, thehighest frequency present beinq dominant in each case. The signalwas then digitised at a rate of 1000 samples per second and passedthrough a microprocessor to provide an acceleration decrement sign-ature of 0.5 spv-odr diration. rig. 22 is the only figure presentedwhich used this technique. The results at the lower frequencies weredistorted due, it is thought, to the poor cut-off characteristics ofthe analogue filters used, and the closeness of various vibrationfrequencies. To improve matters, digital filters were used after theanalogue data had been digitised at the rate of 1280 samples per

second using a time history 1.6 seconds long. At each test pointand location three such time histories were used to provide the datawhich were then analysed using a random decrement programme on aDEC PDP-10 computer to produce Figs. 19 to 21.

3. TEST RESULTS

3.1 General

The vibration mode shapes for all the measured flexiblemodes are listed in Table 5 and plotted in Figs. 3 to 14. In allinstances only the response of the structure in quadrature with theexciting force is plotted. For clarity of presentation, not everymeasured point is plotted, and in cases where the vibration amplitudeson certain components of the aircraft structure were negligible,measurements were not recorded.

3.2 Details of Mode Shapes

Mode at 5.14 Hz. (Fig. 3)

This mode is the antisymmetric lateral bending of the booms. Becauseof the high tailplane some torsion is also introduced. There is alsoslight antisymmetric bending of the wings.


This is antisymmetric vertical bending of the booms. The motion ofthe wings is essentially rigid body motion.

Mode at 9,7 Hz. (not illustrated)

This is the rotation frequency of the elevators. The control springhad been removed and a rubber cord was used to hold the controlcolumn central. It was considered that this best simulated actualflying conditions. The response of each elevator was independantof the other and could be made to oscillate in-phase or in opposition.


This is symmetric torsion of the booms. The two booms are slightlydifferent, and with the close proximity of the antisymmetric modeat 10.5 Hz some difficulty was experienced initially in separatingthe two modes.


This is the antisymmetric torsion mode of the booms. This mode wasexcited mainly by forces on the starboard tail with a slight restrain

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force on the front tail whereas the mode at 9.9 Hz. was excitedsolely from forces on the port tail. The main wing is indicatingsymmetric bending in spite of the antisymmetric nature of the reactiontorque applied at the boom connection. This is no doubt due to theclose proximity of the fundamental wing bending mode at 11.4 Hz.


This is fundamental wing bending coupled with symmetric boom torsion.The boom torsion is the main cause of the sharp change in bending atthe intersection of boom and wing.

Mode at 14.2 Hz. (not illustrated)

This is elevator rotation with the control column fixed. The column,with its control spring, was lashed in two directions using rubbercords. The damping measured in this mode is quite high; in excessof 5 percent of critical damping. The elevator rotation could bein-phase or anti-phase depending upon direction of force input.This mode appears very similar to the mode at 9.7 hertz.


This is the wing antisymmetric bending mode. Antisymmetric bendingof the booms occurs with rotation of tailplanes.


This is anti-phase rotation of the two rudders with the controlcolumn free. This mode was excited with a force on the port rudderonly.

Hiode at 23.4 Hz. (not illustrated)

This is symmetric rotation of the ailerons with the stick free.

Mode at 30.8 Hz. (riot illustrated)

This is the sy.nnetric bending of the stub wing, and was excited viaforce inputs to the main wings.


This is the symmetric wing torsion mode coupled with spanwise aileron

bending.

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Mode at 38 Hz. (Fig. 10)

This mode is essentially rudder bending about its hinge line, coupledwith antisymmetric tailplane torsion. The rudder was clamped tothe fin at the hinge line with a single clamp. The tailplane torsionis present because of the close proximity of the mode at 38.4 Hz.measured with the control surfaces free.


This is antisy-nmetric torsion of the elevator.


This mode is antisymmetric wing torsion and antisymmetric ailerontorsion coupled with aileron bending. This mode was measured withthe ailerons free to rotate and with the airframe on its soft support.


This is rudder bending coupled with some boom torsion which causesthe tailplane to roll.

Mode at 46.5 Hz. (Fia. 14)

This mode is symmetric tailplane bending.

Mode at 48 Hz. (not illustrated)

This is the rotation frequency of the tailplane tab when in themaximum down position.


This is the rotation frequency of the tailplane tab when in themean position.


This is the rotation frequency of the tailplane tab when in themaximum up position.

3.3 Flight Tests

On the samples of time history produced in Fig. 15 the impulsesapplied to the ailerons are well defined, as are the responses on thewings and boom to the impulses. The impulses applied to the elevatorsare not obvious no doubt due to the damping effect of the long control

cables, which pass over several pulleys and two 90 degree changes indirection before reaching the elevator. The general level of randomacceleration at the elevator, as measured by the accelerometers andshown in Figs. 15(b) and 15(d), is almost equal in amplitude to theimpulses applied to the ailerons, and should have been sufficient toinitiate flutter if the aircraft had crossed a flutter boundary. Itshould be noted that the vertical scale used in Figs. 15(c) and 15(d)is 10 times that used in Figs. 15(a) and 15(b). The accelerometeramplifier sensitivity was also changed follotving the first flight,as saturation of the amplifiers wias suspected. The top record onFig. 15(d) indicates that saturation was a problem with that particular

amplifier even at the lower sensitivity. Figs. 16 to 18 show thefrequency spectra of the accelerations at six locations and two airspeeds. The most significant feature appears on Figs 16(a) and16(b) where a sharp peak can be seen at approximately 21 hertz onboth wing and aileron.

Detailed investigation of the records at other speeds andaltitudes indicates that the vibration frequency and amplitude doesnot vary between speeds from 100 to 175 knots, or altitudes from1230 to 2770 metres, and generally does not exceed 1 millimetre peak-to-peak at the aileron trailing edge. The oscillation disappearedat various times during the flights, particularly at times when thepilot may have been undertaking manouvres using the ailerons.

In Figs. 19 to 21 (produced by Random Decrement Analysis),which indicate the impulse response of the structure at six locationsat a speed of 175 knots, it can be seen that vibrations around thetail region are damped. In Fig. 20(a) the wing and aileron show arelatively low damped oscillation at 10 hertz. This is due to leakageof the response of the wing bending mode of about 12 hertz, (seeFig. 16(a)), through the filter at 10 hertz. Fig. 20(b) shows thismode correctly at about 12 hertz with higher damping.

Fig. 20(d) shows the aileron having signs of a lightlydamped oscillation in the 20 to 2S hertz range. To illustrate theaileron vibration further, Fig. 22 is presented to show the rapiddecay from. the initial peak response to the continuing limitedamplitude oscillation. The two curves indicate that the frequencyand damping of the aileron oscillation is independent of air speedin the range covered by the test.

4. DISCUSSION,

The approach to determining the resonant frequencies of thestructure by use of a spectrum analyser appears to have sove short-comings. The method adopted in this test, of using a random inputand a transfer function of the response with reference to the input,

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generally showed the reqions in which resonances occurred, but wasunable to separate close resonances, such as the symmetric andantisymmetric boom torsion modes, even when using the band expansioncapabilities of the analyser. However, the frequency spectra shown inFig. 2 indicate all the reqions in which there were airframe vibrationmodes except the wing torsion modes. This omission was caused byplacing the accelerometer and shaker close to the torsional nodalline. This would have been avoided if the winq-tip leading-edge hadbeen used.

With the flight tests, the fact that the pilots cabin andtheinstrumentation compartment were well separated introduced theproblem of monitoring the operation of equipment in flight. It isapparent that a telemetry system, with remote control of amplifiergains, would be an advantage in this type of work. This would alsoobviate the need to carry the tape recorder on the flight.

5. CONCLUSIONS

The aircraft has been ground tested, in relation to its

airframe vibration characteristics, and all the expected modes ofvibration measured. The flight tests have been carried out andevery reasonable effort made to induce flutter at speeds up to 175knots. The aircraft, in the configuration tested, has been foundto be clear of flutter at air speeds up to and including 175 knots.

REFERENCES

1. HA. Cole. On-line failure detection and damping measurementof aerospace structures by Random Decrementsignatures.NASA CR-2205, 1973.

TABLE I(a)

DETAILS OF AIRCRAFT FOR GROUND RESONANCE TEST

Aircraft type: Transavia PL-12/T-700 Airtruk

Serial No.: G996

Reqistration: VHoUOT

Enqine: Lycominq Type IO-540-KIAS

Serial Vo: L-18871-48A

Propeller. Hartzell Type HC-C3YR-lRF/F8468A-2!<

Serial ilo: DY903

Hopper confiquration: Viet/Dry Doors

Flap Control: Standard manual

Modifications incorporated:

No. 140 Horn balance elevators

No. 141 Bevel Trailinq Ldqe Ailerons

I]o. 143 Elevator Dotm Sprinq

Total mass approx. 1000 kiloqrams

Centre of Gravity approx. - 420 millimetres aft of main wingleadinq edqe.

TABLE l(b)

AIRCRAFT CONFIGURATION FOR FLIGHT TESTS

'ass of aircraft plus oil, fuel, pilot, 1408 kilograms

parachute, battery pack, tape recorder,amplifiers, etc.

Maxim.um permitted mass of aircraft 1742 kilograms

Centre of Gravity of aircraft at 0.532 metres aft of datumtake-off.

Limits of Centre of Gravity 0.417 to 0.589 metres aft

of datum

Moment about Centre of Gravity 7350 Newton-metres

Control surface balances:-

Trim tab 0.3 kilograms T.E. heavy

Elevator 0.14 kilograms L.E. heavy

Aileron 1.59 kilograms T.E. heavy

Rudder 0.75 kiloqrams T.E. heavy

TABLE 2(a)

LOCATION OF WING MEASURING STATIONS

0.015 0.202 0.635 0.765 0.995

+0.995 - WIB WIC WlD -

-0.995 - W12B 12C W12D -

+0.844 W2A W2B W2C W2D W2E

-0.844 WiA W11B W11C WID WIlE

+0.665 W3A PJ3B W3C W3D W3E

-0.665 W1OA W1OB WIOC WIOD W1OE

+0.470 W4A W4B W4C W4E

-0.470 W9A W9B W9C I - W9Ett

+0.318 W5A W5B W5C I W5D W5E

-0.318 W8A W8B W8C W8D W8E

+0.131 W6A W6B W6C W6D W6E

-0.131 w7A W7B W7C W7D W7E

is proportion of wing chord aft of wing leading edge

n is proportion of wing semi-span from aircraft centreline

TABLE 2(b)

LOCATION OF TAILPLAME MEASURING STATIONS

0.042 0.375 0.625 0.986

0.875 PTIA, STIA PTE, STIB PT1C, STiC PT1D, STiD

- 0.875 PT4A, ST4A PT4B, ST4B PT4C, ST4C PT4D, ST4D

+ 0.484 PT2A, ST2A PT2B, ST2B i PT2C, ST2C PT2D, ST2D

0.484 PT3A, ST3A PT3B, ST3B PT3C, ST3C PT3D, ST3D

is proportion of tailplane chord aft of tailplane leading edge.

n is proportion of tailplane semi-span from boom centreline.

TABLE 2-(c)

LOCATION OF FI, RUDDER AND BOOM MEASURING STATIOUiS

2640 3810 4220 4440 4950w.L. _ _ ___ _

305 PF4A PF4B PF4CI- ISF4A SF4B SF4C

75 PF3B PF3C

SF3B SF3C

0 PBIH, SBIH PB2H, SD2H

PS1V, SBIV t PB2V, SB2V

+ 355 PF2 P2 PF2C

SF2A SF2B SF2C

+510 I PFlA PF1B PF1C

SFIA SF1B SFlC

B.S. - Body Station in millimetres from main wing root leading edge.

W.L. - Water line in millimetres from boom centreline.

TABLE 3

LOCATION OF ACCELEROMETERS FOR FLIGHT TESTS

Port aileron - measuring station W1E

Port-wing fron spar -'measuring station W11B

Port-wing rear spar - measuring station W11C

Starboard-wing front spar - measuring station W2B

Starboard-wing rear spar - measuring station 32C

Port fin - measuring station PF2B

Port fin - measuring station PF4B

Port boom - 0.5 metres aft of PB2V

Port tailplane - inboard of PTIB

Port elevator - measuring station PT4D

Port elevator - measuring station PT1D

Port rudder - measuring station PF2C

Port rudder - measuring station PF4C

(Refer Fig. l(b) for locations of measuring stations)

TABLE 4

FLIGHT TEST POINTS

Test point I.A.S. Altitude Date flownKnots metres(feet)

1 106 2770 (9000) 10th July 1981

2 117 2000 (6500) 10th July 1981

3 128 1230 (4000) 10th July 1981

4 128 2770 (9000) 10th July 1981

5 138 2000 (6500) 10th July 1981

6 148 1230 (4000) 10th July 1981

7 145 2770 (9000' 12th July 1981

8 152 2000 (6500) 12th July 1981

9 160 1230 (4000) 12th July 1981

10 160 2770 (9000) 12th July 1981

11 166 2000 (6500) 12th July 1981

12 175 " 1230 (4000) 12th July 1981

TABLE 5

SUMMARY OF MEASURED MODES

Description of Mode Natural Damping Ratiometer Fig.Frequency % Rdg. No.

Hz Critical Q/R

Antisymmetric lateral bending 5.14 2.2 0.26 3of booms

Antisymnetric vertical 6.07 2 0.05 4bending of booms

Elevator rotation - stick 9.7 - - N.I.

free and downspring removed

Symmetric torsion of booms 9.9 - 0.07 5

Antisymmetric torsion of 10.5 1.7 0.08 6booms

Symmetric wing bending with 11.4 1 0.08 7symmetric boom torsion

Elevator rotation - stick 14.2 >5 - N.I.fixed

Antisynmetric wing bendinc 14.4 2.9 0.05 8with antisymmetric boomtorsion

Rudder anti-phase rotation - 21.6 - - N.I.

stick free

Aileron symmetric rotation 23.4 2.4 - N.I.

Stub wing symmetric bending 30.8 - - N.I.

Symmetric w.ing torsion with 35.6 2.5 0.2 9aileron bending

Rudder bending with tailplane 38 2.2 0.15 10antisymmetric torsion

Elevator antisymmetric torsion 38.4 1.4 - 11

Antisymmetric wing torsion and 39.4 2.4 - 12antisymmetric aileron torsionwith aileron bending

Rudder bending 43.2 1.5 0.11 13

Symmetric tailplane bendinq 46.5 1.8 0.1 14

Tab rotation - max. down 46 - - N.I.position

Tab rotation - mean position 53.C - - N.I.

Tab rotation - max. up 55.2 - N.I.position

rr - m

m-----3.18 m

1.76 m

2.3 rrn

0.9 m

-- 6.6m

FIG.1(a) GENER~IAL CAMTANE 01- AIRCRAFT WITH MAJOR DIMMSIC4S

+PF1

+ A + + +

W12 + + + +

*F2V

PB2VPT4 PT3 PT2 PTI

+ + AB

+ + c+ + 0

FIG.1(b) LOCATIcMS OF M~EASUJRIN~G STATICNS (only half showan here)

zp

E-z

Q

E1

00

?u-cpuaq x

2uipeq. 2uiM zHi LL- -

0U 2UTpuoq UOOR ZHj9 ~ UimhIV)0l

9TdTlSI - 211 9V-.-U

2u-Ep-.Gq .Iappria

PU-C uoTSjcoq 9umTdTp31 -H 9i

0E-zo

0

z

U2

CLa

Cr)

NzIH333

F-

CLLz

zH

-

-LJ

F-

UO T .1-0 !

LUJ iooq -l 611

z0CL0fLU

NOIIYY131333V

FIN/RUD LATERA, FIN/RUD LATERAL

STBO TAILPLANE VERT PORT TAILPLANE VERT

WING VER71CAL

* LEADING EDGE* TRAILING EDGE

STUB WING VERTICAL

FIG.3(a) MODE AT 5.14 Hz

WING BOOM TAIL

STBD BOOM-.- - c x- -.c VERTICAL

PORT BOOMVERTICAL

STBD BOOM

K LATERAL-

PORT BOOM

LATERALK=

-- - PTI

PT2W3

W3_____PT 3

ws PT 4

W5- PT4 '

W5 } PT&Ix

WO ST2

0ST31

WI- ST, i .T..

I G r.ID PT r-14. . H - "I. .. .

el INIRUD LATERALFIN/RUD LATERAL 0

STEID TAILPLANvERT PORT TAILPLANE VERT

WING VERTfCAL

x LEADING EDGE0 TRAILING EDGE

STUB WINC VEPTICAL

F:G.4(a) MODE AT 6.07 Hz

WING .BOOM "AIL

STBD BOOM•. . VERTICAL

PORT BOOM

VERTICAL

STBD BOOMLATERAL

PORT BOOMLATERAL

W - PTI

I 1 ri Pr2X-,W3W3 X

PT3

W5 . x PT i

w~oST3

X-K I

WO M AT 6.07ST 3

W12 N

S, 14

FIN/RUG LATERAL F!N 'PL:D LATERAL

STSD TAILPLANE VERT PORT TAILPLANE VERT

WING VERT;CAL

x LEADING EDGE0 TRAILNG EDGE

STUB WING VERTICAL

FIC.5(a) MODE AT 9.9 }z

STBD BOOM

-I VERTICAL

PORT BOOM

VERTICAL

1- STBOD BOOMLATERAL

PORT BOOM

LATERAL

xx

PT 2

PT3

Wrl - x

W5 PT4.

W8 L

W10 STi

WlO0 ST3

W12

FIG.5(b) MODE AT 9.9 Hz

FIN~RIJDLATEAL FN//PWLILATERAL

STRO TAILPLANE VER~T PORT TAILPLANE VER T

WC VER CAL

x LEADING EDGEo TRAILING EDGE

STUB WiNrj VERTICAL

FIO.6(e-) MOMl AT' 10.5 Hiz

STBO BOOM

VER TICAL

PORT BOOMVERTICAL

____ ___ ___ ___ ___ ___ ___ ____ ___ ___ ___ ___ ___ ___ STBD BOOM

LATERAL

____ ___ ___ ____ ___ ___ ____ ___ ___ ___ PORT BOOM

-'-. x LATERAL

Wi PTI

W5 PT14I

Wa I

W3o PT2

W3 ST3

STIS

F1c6b MODE AT 10.5I Hz

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FIG.15(O) TEST POINT 12 -175 KNOTS I.A.S.

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DISTRIBUTION

AUSTRALIA

Partment of Defence-and Department of Defence SuPort

Central Office

Chief Defence ScientistDeputy Chief Defence Scientist copySuperintendent, Science and Technology ProgrammesController, Programmes and Analytical StudiesDefence Science Representative (U.K.) (Doc Data sheet only)Counsellor, Defence Science (U.S.A.) (Doc Data sheet only)Defence Central LibraryDocument Exchange Centre, D.I.S.B. (17 copies)Joint Intelligence Organisation

Aeronautical Research Laboratories

Chief SuperintendentLibrarySuperintendent StructuresDivisional File - StructuresAuthor: A. GoldmanG. LongP. FarrellT. RyallB. Emslie

Haterials Research Laboratories

Library

Defence Research Centre

Library

Central Office

Directo. ('.n l - Army Development (NSO) (4 copies)Defence Industry & Materiel Policy, PAS

Navy Office

Navy Scienti.*c Adviser

Directorate of 14aval Aircraft Engineering

ArEMOff ice

Army Scientific Adviser

.... /cont.

DISTRIBUTION (CONT.)

Air Force Office

Aircraft Research & Development UnitScientific Flight GroupLibrary

Air Force Scientific AdviserTechnical Division LibraryRAAF Academy, Point Cook

Government Aircraft Factories

ManagerLibrary

Department of Aviation

LibraryFlying Operations and Airworthiness Division

First Assistant SecretaryMr. M. ChandivertVr. M. Aubury

Statutory & State Authorities and Industry

Transavia Pty. Ltd., General Manager (2 copies)

Universities and Colleges

Adelaide Barr Smith Library

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SPARES (4 copies)

TOTAL (65 copies)

D1prl-terr l ot Defence

DOCUMENT CONTROL DATA

[Ta .AR NO IW " sb,0l~srt I , 2. Docune,'t Date 3 TwIk No

At-002-p-0 ARL-STRUC-TECH-ME.40-341 June, 1982 89/118

-- 65 Security b. No Pagesa. document

Fw>U"T.7R SUiBSTANTIATION TESTS ON A UNCLASSIFIED 15

i AN'SAVIA PL-i2/T-300 AIRTRUK b title c. Justrdct 7 rof' U U1

9 Dow nralinq ltnsliuctioni

A. -MAN -

10. Cwoporate Author and Address 11. AuthOrity (as approprAirtel

Aeronautical Rese-Z cI; Lab.rator-05 t $iuicr b.Souity r.Dmwat* d.Aia

*'.P.O. Box 433i, mI.BOURNL, VIC. '1001

12. SecondarV Distribution tot this uocum:ntI

Appoved for Public Release.

Overseas enquirers outside utated l,m;tatitins should be refer red through ASOIS. Defence Information Services Brarnch,oepartment of Defenme. Campbell Park. CANBERRA ACT 2601

;3. a This document mvy be ANNOUNCED in caalogs and awareness r rvices available to

No limitations

13. b. Cirntiwn tor other -puraoses 'is tisuil anno n e ,ratj nay be 1i rJ unrestricted6MXl jwj

14 Descriptors 115. COSATI Group

Fultter Utility aircraft 2004Flutter analysis Airframes 0103Aeroelasticity Transavia AirtrukVibration tests I16 Abst,.Kt

A resonance test and subsequent flight tests have beenconducted on a Transavia Airtruk. The natural modes and frequenciesot vibration were measured in the ground tests, and attempts made toinduce flutter during flight test.. The results of these tests axepresented.

PF 65

FLUTTER SUBSTANTIATION TESTS ON A TRANSAVIA … · The Transavia Airtruk has been in production for...

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