Document
--- ._.- ------------------ FLIGHT FLUTTER TESTING OF SUPERSONIC INTERCEPTORS M. DubLin, R. PeLLer - Convair, S cm Diego, CtlLij ornicl Abstract results in a considerable expense of time, money and material to obtain a fix; it delays getting the vehicle This paper presents a summary of experiences into operational use ; and it can sometimes result in in connection with flight flutter t'esting of supersonic permanent restrictions on the airborne vehicle which interceptors. It contains a description of the planning limits its operational capability.
and operational aspects involved, comments on the difficulties encountered, and shows correlation be - From an analytical point of view the deter- tween measurement and theory . The paper con - mination of the flutter stability boundaries is difficult cludes with recommendations for fu ture research and because of lack of precise knowledge of all the para - development to advance .the science of flight flutter meters used in the equations of motion; Reference (1) testin g. outlines these difficulties in more detail and also considers difficulties encountered in flutter model INTRODUCTION testing. Flight flutter tests are therefore made to insure freedom of the vehicle from flutter over its During the last ten years, as noted in Reference operating envelope and environment, and to assist (1), more than fift y different cases of flutter have been the flutter analyst in improving his ability to make encountered on United States piloted military air - analytical predictions.
craft. In addition a certain number of cases of flutter have also been encountered on United States commer - cial and private aircraft. Further, a numberof cases PLANNING ASPECTS of flutter have been encountered on United States military missiles. Although detail statistics are not The steps which must be taken in planning a available, it is known that a number of cases of flutter flight flutter test program are as follows : have occu'rred on foreign aircraft and missiles. Thus, over the last ten year period, it is estimated that at a. Establish desired data and measurements.
least several hundred cases of flutter have been encountered in airborne vehicles of the world. b. Selection of test equipment and installation.
These cases of flutter have had various con - c. Establish test procedure and execute test.
sequences. In some cases mild structural damage occurred and the aircraft was landed safely. In some d. Data analysis and interpretation.
cases very severe structural damage occurred and the aircraft had to be abandoned. With regard to the flutter Although there are a variety of approaches for cases encountered in theUnited States over the afore - each of the above steps, this paper will only consider mentioned time period, insofar as the authors know, the approaches used by Convair in flight flutter testing no loss of life was encountered; whether the same of supersonic interceptors. Figure 1 shows a photo- applies to flutter cases encountered on forei gn air- grap h of one of the configurations tested. Practical craft is not known to the authors. Other aspects of difficulties encountered during the flight flutter test encountering flutter which are important are that it program will also be discussed .
fuel tank configurations. Three fuel tank configur- ations were selected, namely external tanks with full fuel, external to-nks with half fuel and forward center of gravity, and external tanks with half fuel and aft center of gravity. Special compartmented tanks were used for these tests.
TEST EQUIPMENT AND INSTALLATION The equipment used for the tests consists of: a. Excitation system b. · Pickups · -_ O 'f' c. Recording system d. Data analysis system Description of this equipment is discussed hereunder.
EXCITATION SYSTEM Figure 1. USAF F-102A Supersonic Interceptor Based on an examination of the theoretical vibration modes, it was established that the shakers Desired Data and Measurements should be located near t he wing tips in order to obtain satisfactory airplane response for all desired The method chosen for establishing flutte r exciting frequencies. The wing depth available at the stability was to obtain . plots of the damping co- selected location was 4.5 inches for the shaker and its effic ie nts at selected locations on the airplane versus mounting. Since no commercially available shaker airspeed for selected resonant frequencies (Le., both existed which met this space requirement and at the symmetric and anti symmetric) , for selected altitudes same time provided desired force output for satis- and fo r selected airplane mass configurations. Re- factory airplane response, it was necessary to design quired measurements using this method were the air- and develop a shaker system specifically tailored for plane responses (at the selected locations due to an this airplane. Convair developed such a shaker sys- excitation of the airplane), the airspeed, and the tem which is essentially a closed loop servo system altitude.
combining hydraulics and electronics to command and control the movementoftwo reCiprocating masses.
Two excitation methods were employed, namely A functional block diagram of the system is shown in sinusoidal excitation by two inertia shakers, and Figure 2; detail description of the system is contained pilot control excitation. To establish that the shakers in Referenc e 2. The essential elements of this system were functioning properly, it was also necessary to consists of the following: measure the frequency and the displacement of each shak er mass and the phase of one shaker mass a. Pilot's stick swit ch. This is a spring loaded displacement with respect to the other shaker mass on-off switch which when actuated causes the displacement. Pilot control forces or displacements shaker to perform the functions selected on were not measured since movable control surface the pilots control panel.
resp o nses were adequate to establish initiation of pilot control excitation .
b. Pilot's control panel. Three two pOSition toggle switches are located on the pilots A problem area arose in selecting the airplane control panel which permit him to select mass configurations. For the airplane configuration either a manual or an automatic mode of without external wing fuel tanks, the fuel weight is operation. If the manual mode of operation appr ox imately 25 % of the airplane takeoff gross is selected this causes the shakers to sweep weight. Since fuel is expended at a fairly rapid through a specified frequency range at a rate , it was not practical to specify a mass configur- programmed rate of sweep, and at a pro- ation for which measurements should be taken at grammed shaker force ; in this case the spec if ic speeds and altitudes ; it was necessary to take pilot must also select the phasing of the measurements at points on the flight envelope at the shakers (Le., symmetric or antisymmetric), mass configuration which existed. This, of course, and he must also select the sweep cycle leads to one of the difficulties in correlating mea- (Le., ascending frequency or descending surements with theory since in practice analytical frequency) . If the automatic mode of oper- investigations are usually made for a limited number ation is selected this permits obtaining of we ight configurations. However, for the airplane decay responses ; in this mode of operation configuration with external wing fuel tanks, it was six frequencies (either symmetric or anti- poss ible to t ake measurements for various external Force Output Port Shaker Pilot's Pilot Pilot's Stick Control In put Switch Panel Starboard Force Output 't-----l Shaker Hydraulic Electrical Power Power Supply Supply Figure 2. F u nctional Block D iagram of Shaker System symmetric) can be pre s elect~d and the shak - e. Two hydraulically actuated shakers. These su pply the force input to the airplane. Figure ers will exc i te the air p lane for a specif i ed 4 shows a photograph of an assembled time at a given frequency; stop the shakers shaker, Figure 5 shows a photograph of the for a specified time and automatically step to the next frequency -- this process is shaker partially disassembled; the cylinder repeated as long as the pilot stick switch is in the photograph is the shaker mass.
engaged. A programmer is used to accom ~ plish these functions in the automatic mode f. Electrical power supply, This consisted of of operation. Figure 3 shows a photograph the airplane 400 cycle A. C. and 28 volt of the programmer. D. C. power supplies.
c. Function generator. This is used to gener- g. Hydraulic power supply. A separate 3,000 ate the desired sine wave shape . psi hydraulic power supply was installed in the airplane for the shaker system.
d. Two servos. These are used to control the force output of the shakers.
Figure 3. Shaker Programmer Figure 4. Assembled View of Shaker _ __ I b . Sweep rate. In sweeping from 5 cps to 50 cps the sweep rate could be made variable from 55 seconds to 90 seconds.
c. Stopping time of shaker. To obtain decay curves the shaker could be stopped in one- half of a cycle.
d. Synchronization of shakers. Excellent syn- chronization of the shakers was achieved.
Phase desired between one shaker force out- put and the other was within the accuracy of reading the traces .
e. In the automatic mode of operation, the excitation time could be varied from 2 seconds to 7 seconds; the time for decay, after stop - ping excitation, could also be varied from 2 seconds to 7 seconds independent of the ex- Fi gure 5. Exploded View of Shaker citation time.
Other pertinent design characteristics of the f. The shaker mass weight was 8.5 lbs. and its shaker system are : travel was ± 1.0" .
a. Force vs frequency. A linear variation of Trouble encountered with the shaker system force versus frequency was desired. How- were : ever , due to valve characteristics the force- frequency curve actually obtained was as a. Hydraulic leaks .
shown in Figure 6. It is noted that identical force outputs for both starboard and port b. D ete rioration and failure of tubes in the shake r s were not obta ined.
electronic control system.
c. Shorts in programmer stepping switches.
/ d. Potentiometers in pro g rammer were sen- , sitive to temperature.
"
I I e. Human errors in operating and maintaining I I the shaker system.
I I 2 50 / Starboard Shaker Th e shakers were ins ta ll ed in the wing on rigid structure as shown in Fi gu re 7 . Shaker force was ui Port Shaker established from measurements of the shaker mass ;3 / 2 00 I displacement and frequency.
$ / / r.:I / Pilot cont rol excitation simply consisted of the () p:; 150 / pilot "banging" the control stick (Le., lo ngitudinally 0 I and laterally) with his hand or the rudder pedal ~ ~ with his foot. This method would only excite the lowest symmetric and the lowest antisymmetric vi - bration modes.
PICKUPS
/
Fixed surface resp ons es were obtained by seven MB-124 linear ve locit y pickups located as shown 40 50 in Figu,re 7. Movable surface responses were obtained 10 20 30 by three MB-124 lin ear velocity pickups which were FREQUENCY (cps) modified (i.e., by counterbalancing the movable arma- ture) to sense angular velocity; these were located Figure 6. Sh aker For ce v s. Frequency as shown in Figure 7.
Legend: + Linear ve locity pickups • Angular velocity pickups • Shakers Figure 7. Sketch of F-102A Showing Location of Instrumentation The displacements of the shaker masses were plane airspeed, altitude, and outside air temperature were recorded on a photopanel by means of a movie obtained by a variable reluctance pickup excited by a 3,000 cps voltage source. camera.
Difficulties encountered with these pickups were : Correlation between the photopanel and the telemetered signal was maintained by a data correlator a. Linear velocity pickups bottomed at in- which recorded a counter number on the photopanel cremental airplane c.g . normal load fac- as a series of lights and as an electrical pulse on the tors of approximately ±0.4 g's. telemetered signal. This number was changed every two seconds throughout the flight.
b. Sensitivity of angular velocity pickups was The telemetered signals were received at a not as high as desired for easy reading of ground station where they were: traces.
The airplane speed was obtained by a Kollsman a. Recorded as an electrical signal on magnetic airspeed indicator, Type 739 DX-6-059. With this tape.
instrument, as with any otp.er available instrument , it b. Recorded on an oscillograph (to check instru- was difficult to predict exact speeds in the transonic speed regime due to pOSition errors existing in the mentation in the field).
system.
c. Put through appropriate discriminators and The altitude was obtained by a Kollsman alti- recorded on Sanborn recorders.
meter, Type 1846 X, -4-01. In the transonic speed Communication between the g round st ation and regime it was difficult to predict exact altitudes due to position errors existing in the system. Addition- the aircraft was maint a ined by radio at all time .
ally, during dives at high rates of descent it was dif- ficult to predict exact altitudes due to lags in the Difficulties encountered with the recording sys- tem were: altitude measuring system.
a. Loss of telemetering signal due to airplane RECORDING SYSTEM position or distance from the ground station.
Outputs of the veloCity pickups and the shaker pickups were fed into an FM / FM telemetering trans- b. Loss of telemetering signal due to electrical mitter for transmittal to a ground station. The air- failure in the airplane.
c. Necessity of changing tape during the flight slow paper speed of approximately 0.5 inches per when only one tape recorder was available. second.
d. Failure of recording pens on Sanborn equip - The data correlati on trace was recorded on ment.
the oscillograph records along with the airplane responses . This allowed complete correlation with e . Radio failure (eit her airplane or ground the speed information obtained from the photopanel.
radi o).
Photopanel records were developed by standard procedures and read by means of projection eqUipment.
DATA ANALYSIS SYSTEM The data station proved to be a very reliable The data which was stored in the form of an piece of equipment. Such difficulties as were en- elect r ical signal on magnetic tape was processed in a countered could be attributed to human errors.
data s tation. The signals were put through appropriate disc r i minators and oscillograph traces obtained . A TEST PROCEDURE AND EXECUTION OF TEST stan da rd proce dur e which recorded all pickups with An initial plan was made outlining the desired 60 cycle low-pass filters was run off first. If these rec or ds proved unr eadab le because of excessive res- speed-altitude pOints which were required. This plan was flexible in that speed increments could be in- pons e due to atmosphe ric turbule nce, a band-pass creased or decreased depending on the results obtained filte r from 25 to 50 cps was used to eliminate the low fr eq uency responses. If the higher frequencies from each flight. Figure 8 shows a typical speed- made decays in the fundamental modes unreadable , altitude test plan. Tests were initiated at subsonic a 5 to 25 cps band pass filter was used. speeds at the hi g hest altitude chosen . Tests at lower altitudes were always made in such a manner The os cillograph records were, in general, that the equivalent speed obtained at high altitude reco r ded at a paper speed of 4 inches per second. was not exceeded . Frequencies at which decays were obtained were established from sweep records For s pecial conditions , (i. e. , obtaining an overall view of a sweep) the records were recorded at a taken in flig ht at selected intervals.
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M ac h Numbe r Legend: o Symmetr ic & anti -sy mmetric sw ee ps • Sy mmetric & anti- sym me tri c de ca ys x Pilot control pulses (e l eva tor & a il er on or rudd er ) Ve Equiva lent airspe ed, k nots Vc Indicated airspe ed c orrected for instrum ent a nd po sition error, knots Figure 8. Typi c al Sp e ed-Altitude T e st Pl an ~--- - - - ----------------------------~ Prior to flight a ground checkout procedure was b. Meteorological problems. Examples are ex- established which accomplished the following tasks: cessive winds preventing take-off, excessive turbulence and gusty air which would mask the response due to shaker or pilot excitation, a. Insured proper functioning of shaker system.
and excessive outside air temperature which prevented achieving some of the desired b. Insured proper calibration of instrumenta- speeds.
tion.
c. Operational problems. Examples are un- c. Insured proper operation of telemetering availability of chase airplane or chase air- equipment.
plane mechanical problems, limited fuel sup- ply, necessity of going off-base for low alti- d. Set programmer parameters in accordance with desired measurements. tude testing, conflicts with higher priority testing, short time for taking measurements during dives at high descent velOCities, and During flight, ground monitoring was used to: location of data reduction equipment away a. Check proper functioning of shakers, instru- from test base.
mentation and telemetering.
The above difficulties either contributed to lengthen- b. Check proper positioning of pilot's shaker ing the duration of the flight flutter test program or controls. decreased the reliability and accuracy of the mea- surements.
c. Notify pilot if data is unsatisfactory (Le., due to turbulence or gusts). ~equest repeat DATA ANALYSIS AND INTERPRETATION measurements or flying an alternative flight plan.
The two conventional methods were used to ob- tain the experimental damping coeffiCients, namely d. Inform pilot of satisfactory completion of from the response of velocity versus frequency plot, frequency sweep (Le., to reduce test time).
and from the response of velocity versus time (Le., decay) plots.
e. Estimate damping coefficients from decay records, and inform pilot either to continue A typical section of a sweep record is shown testing at higher speeds or to discontinue in Figure 9. Figure 10 shows a sample decay testing until records can be analyzed in record.
detail.
A typical plot of the experimental damping Following the analysis of data for each flight, versus Mach number curve is shown in Figure 11 at it is necessary to re-examine the test plan and deter- a 35,000 foot altitude for the second coupled anti- mine what modifications, if any, need to be made.
symmetric vibration mode. Points on this curve which Typical changes in the plan are: are dotted were simply demonstrated and no mea- a. Decrease speed increments due to a large surements with excitation were taken. A negative decrease in the damping coeffiCient, or damping coefficient denotes a stable system. Similar alternatively, increase speed increments due plots were obtained for all other significant vibration to a steady increase in damping coefficients. modes at various altitudes to demonstrate that the airplane is free from flutter over its design envelope.
b. Repeat test points due to failure of photopanel camera, which results in no speed and altitude The experimental results shown in Figure 11 data. were also compared with theoretical results. The following explanatory comments are made in con- c. Repeat test points to check scatter in data. nection with these theoretical calculations. Theor- etical anti-symmetric mode flutter calculations at a The main difficulties encountered in executing 35,000 foot altitude were made using the lowest five the flight test program were: anti-symmetric coupled vibration modes and for a gross weight corresponding to a 60% full fuel con- a. Development problems with the airplane. dition; two dimensional OSCillatory aerodynamic co- Some examples are electrical power failure, efficients were used in the analysis. Reference 3 malfunction of cabin pressurization system, contains details of these calculations. The minimum compressor stalls, failure of afterburner to damping occurred in the second anti-symmetric light, malfunctioq of fuel quantity indicator, coupled vibration mode with a natural frequency of malfunction of fire warning indicator, and a 13.4 cps. Figure 11 shows this theoretical damping supersonic noise problem. These resulted plotted versus Mach number -- in the transonic speed in either aborted flights or temporary re- region the curve is shown dotted since no calculations strictions on the airplane.
were made - - at zero airspeed the structural damping
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.. ,:', ".J .':' . ~ ,.. "t'" • • .". ," "~I "", .. "I't.".' ~ t •• Ti me NOTES: 1. Record is for right wing pickup 2, Anti-symmetric resonant frequencies , 36 cps and 43 cps 3, M = 1.1 5, h = 15,000 ft.
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Left wing pickup
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NOTES:
1. Symmetric resonant frequency = 43 cps
2. M = 1.3, h = 35,000 ft.
F i gure 10 . Sample Decay Rec o rd h 35,000 Ft.
o. .1 .2 .3 .4 .5 .6 .7 .8 .9 1.1
1.0 1.2 1.3 1.4 1.5 M
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g = Damping coefficient Theoretical
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M = True Mach number
0---0 Experimental
h = Altitude
Figure 11. Comparison of Theoretical and Experimental Damping Coefficients Second Anti-Symmetric Coupled Vibration Mode coefficient was obtained from response curves ob- d. Because excitational forces could not be tained during a ground vibration test. It is noted made exactly equal, unsymmetric responses that the general shape of the experimental and were obtained.
theoretical plots are similar; however the actual magnitude of the damping coefficients differ by a e. Because of the temperature sensitive p()- noticeable amount, tentiometers in the shaker programmer unit, difficulty was encountered in setting the desired frequencies for decays. Thus, less The following problems arose in connection with than the maximum possible response was analysis of the data and its interpretation: obtained.
a. Considerable scatter was found in the damping f. Indication of " false resonances" were ob- coefficients obtained. This lead to difficulty tained because of necessity of using sweep in extrapolating the speed-damping curve times less than theoretically desirable .
to the next intended speed. Possible reasons for these apparent discrepancies were: g. Masking of the lowest coupled vibration modes responses by a gust response gave an erron- 1. Use of two different methods for ob- eous indication of damping .
taining damping factors.
RECOMMENDATIONS FOR FUTURE RESEARCH AND DEVELOPMENT 2. Transfer of energy of vibration between various portions of the airplane.
Flight flutter testing is an ever changing type of testing in which no technique may be considered 3. Difference in mass configurations of the perfect . . As with most any type of testing, hindsight airplane during test.
is a wonderful thing and many changes in technique and different avenues of approach present themselves b. Measured damping factors varied between as testing progresses.
different pickups in the same vibrational mode.
Experiences with the testing discussed in this paper lead to the following recommendations for future c. Altitude trends were difficult to establish.
-
rese a rch and development in the field of flight flutter bility character i stics in lieu of damping testing for manned aircraft flying at moderate super- coefficient versus speed. This approach sonic speeds: is analagous to the method outlined in Refer - ence 4. Purpose of this is to determine a. Make the excitation system completely auto- flutter stability from a single output (L e ., matic. This . system should have a pro- work) instead of multiple outputs (Le ., damp- grammer where the desired excitation and ing coefficients at a number of locations on duration can be pre-set on the ground for a the vehicle).
given flight plan. It should have an auto- matic force and phase synchronizer when Extrapolating current experience to very high two or more shakers are used to accurately supersonic or hypersonic manned and unmanned ve- contrdl the force inputs. Further, it should hicles, a number of new factors enter into the have an automatic vibration mode seeker problem of flight flutter testing. The most important which would determine the peak responses of these are high temperatures , flights at very high in flight.
angles of attack, very high rates of climb or de- scent, and very high longitudinal, lateral and vertical b. Make the data recording and data reduction accelerations. For some configurations in this systems completely automatic. An auto- category it will not be possible to stabilize the matic plot of ihe data is desired in the form vehicle for specified parameters long enough to obtain which is used to interpret the stability measurements concerning flutter stability. In this characteristics of the vehicle.
case , it appears that we will have to revert to a "go-no-go" type of testing.
c . Consider the possibility of using plots of wor k v.ersus speed for interpreting the sta- REFERENCES 1. NACA RM 56 I 12, "A Survey and Evaluation of 3. Convair Report ZU -8-032, "Theoretical Flutter F l utter Research and Engineering," (CONFIDEN- Invesitgation of the F-102A Airplane", (C O NFI- DENTIAL) TIAL) 2. Co nvair Report 56-51, "In-Flight Vibration System, 4. N. O. Myklestad, "Vibra ti on AnalYSiS", McGraw- En gineering and Operations Manual" Hill Book Co.