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DOUGLAS EXPERIENCE IN FLIGHT FLUTTER TESTING ]. Philbrick - Douglas, Santa Monica, California Abstract The Douglas Aircraft Company has required extensive flight flutter tests on all aircraft models Douglas Aircraft Company experience in flight and versions which have been produced since 1954.
flutter testing is reviewed briefly, with comments The objectives of these demonstrations have been on state-of-the-art excitation and instrumentation 1) verification of analytical predictions , and 2) de- techniques used up to the present time. The limita- monstration that unpredicted instabilities do not exist.
tions of previous techniques are discussed with em- The responsibility for these demonstrations is shared phasiS on the problem of: jointly by the Design Engineering and Testing Divi- sions. A policy, based on the airplane type, perfor- (a) Establishing a flutter margin of safety for mance capabilities , and the aero-elastic character- predicted marginal flutter modes. istics predicted by theoretical analyses and flutter model tests, has been established for the flight (b) Resolving instances of flutter not predicted conditions, airplane configurations, instrumentation, by theoretical calculations in advance. and the data reduction techniques to be used for these flight demonstrations.
(c) Delaying the airplane demonstration by time consumed in acquisition and reduction of Experience has shown that neither the theore- tical predictions nor the flight test techniques used flutter data.
to date have been infallible. The intent of this paper is to show the shortcomings of earlier techniques Current Douglas philosophy in flight flutter test - ing is presented and a description given of: as revealed by flutter experience obtained from tests of current aircraft.
(a) Steady-state vane excitation system develop - ment. EVOLUTION OF TECHNIQUES (bl' An automatic data handling system. The initial flutter programs were conducted by monitoring the decay of structural motion excited by (c) The potential application of automatic com- manual control surface pulse inputs. Instrumentation puting methods for increasing flutter data consisted of strllin gage type accelerometers installed yield. at the aircraft extremities or at locations having large response amplitudes in the predicted flutter modes.
INTRODUCTION Control surface positions were measured using elec- trical potentiometers to define the character of the The development of high performance aircraft input pulses and to detect coupling of control sur- of various configurations with increased flexibilities faces in the flutter mode. Data were usually obtained and concentrated weight items at structural extrem- on airborne oscillographic recorders ; however , direct ities has made the consideration of flutter not only a writing pen type recorders have occasionally been design criterion but also an important flight demon - used to allow immediate monitoring of the data as stration item.
obtained .
Sharp control surface inputs were made at The next figure (Fig. 2) shows the damping incr emental airspeed and Mach number as the flight trends as indicated from the above aileron input enve lope was extended. The tests were run at a investigation. Data scatter and failure to excite the rela ti vely low altitude to minimize Mach buffet symmet ric flutter mode lower than about 85% of the effe cts during airspeed advances, and, conversely, required demonstration speed did not allow extrapola- at a higher altitude to minimize rough air effects tion to the zero damping speed or instill much con- dur ing Mach number extensions. It was also found fidence in investigating this flutter case further. It adva ntageous to schedule flutter flight tests in the is obvious that a more efficient excitation method earl y morning and/or over the ocean to minimize would be desirable in this case.
atm ospheric turbulence.
Although this approach to flutter testing re- quir ed a minimum of test eqUipment and installation, the quality of the data obtained di~ not always pro- vide consistent stability indications. Data scatter resu lted primarily from 1) the manual pulsing de- pend ed on pilot ability for repeatability of pulse dura tion and magnitude, 2) the pulse energy was not dire ct ed to the desired mode , that is, symmetric wing modes were poorly excited by elevator pulses and not at all by conventional aileron inputs, and 3) the tranducer outputs were often masked by buffet and other extraneous vibration.
Various harmon ic analysis methods were used to extract information from the recorded data. The Fou rier analysis and transfer functions proved useful for s eparating frequency components which could be used to follow flutter trends.
Fi g ure 2.
The results of several flutter programs illus- trate many of the above difficulties. As an example, Figu re 1 shows an oscillograph record obtained during Figure 3 shows the structural response of a sin- an a il eron input while investigating a symmetric wing gle jet airplane following a rudder pulse. The exci- bending-torsion flutter case on a twin jet airplane. tation in this instance was adequate for exciting the The initial asymmetric response degenerates to the aft fuselage torsion-vertical stabilizer bending mode ; desi r ed symmetric mode after apprOXimately four however , the airplane had been previously flown be- cycl e s; however , in view of the background noise, yond the flutter speed where rough air was sufficient it was extremely difficult to obtain accurate structural to precipitate an instability which had not been ex- dam p ing from the decay in the required symmetric cited dUring the initial pulsing program. Fortunately, mM p the flutter, although severe, was non-destructive and Fi g ure 3.
Figure 1.
the pilot had an opportunity to perfect his rudder Sinusoidal excitation from manual elevator in- puts has proved successful for exciting structural pulsing technique by using sharper and harder inputs.
response at frequencies below five (5) cps. The input Subsequent investigation using the perfected rudder pulsing provided consistent stability data which allowed for single frequency and frequency sweeps was co~ trolled by having the pilot synchronize his input rate a definite extrapolation to the flutter speed. This trend to the response of tuned reeds. In one instance, a is shown in Figure 4.
photograph of a rather voluptuous lady encased in a plastic projector had the exact mass required to tune Manual control surface pulse excitation has been a reed for a particular frequency. Airplane and pilot adequate for certain flutter flight testing; however, response to this device was excellent. For some un- in many instances, its use was restricted by pilot known reason, the reed was lost on the last flight of ability, response of the control system, and poor this flutter program.
pulse energy transfer to various parts of airplane (Le., elevator to wing). Except for control systems with extremely slow response rates and the cited Instrumentation for flutter flight testing has al- difficulties, structural modes with frequencies below ways posed a problem. The frequency response and 10 cps can be excited by manual pulsing techniques. output of most commercially available transducers re- quire some compromise to cover the required flutter The shortcomings, as noted above, of manual acceleration and frequency ranges. The strain gage pulsing have led to the investigation of auto-pilot type accelerometer has been an useful device from the inputs, ejection of bombs and stores, and devices standpoint of size, calibration, and maintenance. Strain to pulse flight controls. The low frequency re- gages for load and stress measurement in oscillating sponse of auto-pilots (below 5 cps) and the inadequate components provide cleaner data than the accelerom- energy transfer from control surface inputs have, in eter; but the installation, calibration, and maintenance of gages is much more difficult. Control surface general, negated this method of excitation. Bomb and store ejections have been satisfactory in some in- positions from electrical potentiometers are fairly stances; but, usually, the sharp input, limited bomb reliable, but frequency response and lack of sensitivity carrying capacity, and cost of ejected items have at low amplitudes limit their usage. Greater reso- made this excitation method prohibitive. Devices for lution and frequency response are possible from strain control system pulsing have extended the input capa- gage bending beams operated by a cam On the rotating bilities but are still subject to the limitations as cited member. The output and linearity of these items can for manual pilot inputs. The need for a consistent be adjusted by their physical geometry.
pulse input that could be applied at a discrete structural point has led to development of an impulse Extraneous vibration at frequencies above the generator unit. These units are essentially small flutter range tends to mask the accelerometer outputs.
Several types of electrical filters have been developed.
rocket motors having a specific impulse and burning time dependent on the amount and type of propellant A unit package in a case similar to the standard used. The size of these devices has allowed installa- 350 n galvanometer shunt has proved most useful tion in fairly limited spaces and has provided ex- and provides a 6db/ octave attenuation or can be cellent pulse inputs. The details and usage of the seriesed to give multiples of this attenuation. The impulse generator excitation method were presented units have been designed for roll-off frequencies of 20, in a preceding paper * at this symposium. 30, 40, and 60 cps.
Airborne recorders have been utilized for flutter data recording. The standard 18, 36, and 50 channel CEC oscillogarphs have been used mainly for their frequency response, adaptability to the transducer outputs, and the analog presentation of the record. The photographic developing the oscillograph record has been a delaying factor in someflightflutterprograms.
The currently available direct writing oscillographs and magazines have largely eliminated this problem.
In an effort to increase the airspeed range per flight and to provide simultaneous flight coverage, FM/ FM telemetry has been used during recent flutter testing. Eight (8) standard sub-carrier frequencies from 5.4 to 30 KC combined and transmitted on 230.0 megacycle carrier has been used. The composite signal is received at a ground station where it is tape recorded, discriminated, and displayed as an analog record. One or two flight test engineers can reduce the flutter data from these records and keep a running plot as the flutter test progresses. Portable FM/ FM Figure 4.
---------_. - telemetry stations and relay stations have been used response of auto-pilots and the indefinite cut-off of to extend the receivable test area. rotating inertia devices have made these excitatio n meihods undesirable.
Occas i onally, the manner in which the flutter test is conducted does not reveal the existence of a The Douglas Aircraft Company is presently critical flut t er case. Figure 5 illustrates a flutter evaluating the use of auxiliary airfoils for steady state flutter excitation. The first system was deve- incident of this type. The initial data obtained during 10,000 and 35,000 foot altitude airspeed - Mach number loped by Electrosystems, Inc., Burbank, California, and consists of two vanes to be mounted at the extensions indicated adequate stability in the hori- airplane wing, horizontal stabilizer, and / or vertical zontal stabilizer yaw - aft fuselage roll case. Subse- quen t data obtained at intermediate altitudes showed an stabilizer tips. The vanes are driven in pitch by adve rse Mach-airspeed combination with an instability hydraulic servo valves and actuators which are con- within the required flight envelope. Based on this trolled by an electronic programmer.
resu lt, flutter flight programming. has specified that tests be accomplished at three altitudes. The inter- The vane system is designed to provide sym- med iate alti t ude is chosen at an estimated maximum metric and antisymmetric excitation in the frequency "q" - Mach number combination.
range from 1/2 to 15 cps at a maximum input force of 250 pounds (vector). Individual mode tuning, auto- matic and manual frequency sweeps, and instantaneous cut-off for decay monitoring are possible. The equip- ment will operate with 3 sq u are feet vanes to an air- speed of 300 knots and with 2 square Toot vanes to above 400 knots. The system is schematically shown in Figure 6.
The vanes are hinged and mass balanced forward of the 25 % chord to maintain a stable aerodynamic trail position when inoperative or following an emer - gency shut-off. The emergency shut-off will be accomplished by a by-pass valve in the actuator.
Viscous damping can be introduced for vane stability by varying the restriction in the by-pass valve and line.
Airplane protection is afforded by a force feedback system which maintains the mean vane position at the zero force angle of attack. Automatic shut-off is provided for in the event that the input force or airplane structural response exceed a Figure 5.
The various flutter programs have shown that the excitation methods, data availability and reliability, and the necessity for a complete airspeed-Mach number build-up for each airplane configuration and / or flutter fix have been the primary sources of airplane dem o nstration program delays. The cost of flight test ti me and the hazards involved on current airplanes prov i de sufficient justification for a determined effort to eliminate the items cited above.
FUTURE PLANNING For a number of years steady state excitation has been advocated for flight flutter testing. Auto- pilot cycling of control systems and rotating weight devices have been used; however, the low frequency Figure 6.
J look , playback and editing facilities are included in the pre-selected value . In the event that the automatic co mputer station for scanning and editing flight test shut-off items do not operate , a fracture joint in the vane torque tube is designed to fail and shed the vanes data. The required flight data, transducer calibration at an input load below the airplane structural limit . data, and the analysis program are fed automatically into the digital computer allowing analysis of flight test data in greatly reduced time.
The vane system provides a means for exciting airplane vibration modes in-flight and will allow mode surveys for comparison with the calculated and ground In addition to the above, further savings in the vibration modes . The system also allows excitation time required for flight flutter testing may be possible of the modes deemed flutter critical for monitoring with multiple mode excitation using mixed input signals frequency shifts and damping trends during Mach- with the flutter excitation equipment previously men- airspeed advances .
tioned. The composite response signal is compared to the frequency components of the input signal through It is expected that the vane excitation system an analog-integrator , which rejects the frequency will conserve flight flutter test time, as compared to components different from the selected period of the" previous methods, by providing a more positive ex- integral. The chief advantages of this technique are : citation of the flutter modes, increasing the data 1) various modes can be simultaneously tracked confidence factor , and allowing an evaluation of con- throughout the airplane speed range , 2) modal response figuration changes or flutter fixes from data obtained can be extracted in the presence of noise. The most from a single flight.
serious disadvantage of this approach is the long integration time necessary to establish the response In conjunction with a general effort to improve of a lightly damped mode in the presence of noise overall flight test procedures, the Douglas Aircraft or another mode at nearly the same frequency; Le., Company in conjunction with the Consolidated Electro- a number of integration processes are necessary to dynamiCS Corporation is currently developing an auto- reject the close sideband frequencies. Evaluation matic data handling system (ADHS) to expedite the of this technique and efforts to overcome the cited acquisition, handling, and reduction of flight test data.
disadvantage are being continued.
Although the ADHS was not designed specifically for flutter flight testing, the fiutter data requirements were Separation of the structural response of modes integrated in the design specification.
of small frequency difference may be improved by selecting locations for pickups such that each pickup The ADHS consists of an airborne system, a will discriminate against one or more modes and ground station at the test site, and a computer station.
enhance others. By feeding the selected pickup out- The airborne system will sample the analog voltage puts into an analog-type computer, the read-out will be outputs of the various test data transducers, convert several independent signals, each corresponding to a these outputs to binary digital form , record the single degree of freedom representing an orthogonal digitized information on magnetic tape, telemeter mode of the airplane. A Simplified example of this the digitized information to the ground station over approach would be a sum and difference of the outputs a PCM (pulse code modulated) link, and provide in of pickups located at opposite wing tips of an airplane.
larger airplanes, a " quick-look" facility for a flight Summation of the pickup outputs would magnify sym- test engineer's control information. The sampling metric mode response and minimize anti-symmetric rate and accuracy allow frequency resolution up to response. The selection of pickup locations and the 100 cps and to 1 part in 1000 for 100 data channels.
analog circuitry and constants necessary would be Super and sub-commutation of the input channels al- accomplished either during ground vibration tests or lows either higher frequency resolution or an in- while surveying the in-flight vibration modes. The creased number of input channels, respectively. By combination of pickups and analog to accomplish this modular design, the physical size of the airborne function has been termed a "modal pickup."
unit can be tailored to the aircraft size by restriction of number of data channels. The maximum uncom- mutated high frequency capacity (100 channels) can A combination of the " modal pickup" and multi- be utilized in the larger transport and bomber airplanes frequency excitation techniques may be used to follow the amplitude and phasing of several airplane modes.
and approximately thirty (30) channels in an airplane This could be accomplished by driving a common ex- of the A4D size.
citation system from several OSCillators, each of which is tuned to a different modal frequency, and by cross- The g round station is mobile to permit c overa ge correlation integration of the modal pickup outputs of many test sites and contains the telemeter receiver, with the proper input signal the sine-cosine component a tape recorder, and a "quick-look" analog presenta- tion to allow safety monitoring of the flight test data. and frequency of each mode will be obtained.
The compatibility of the ADHS with a digital computer has been one of the design premises. Although the above equipment and concepts have not been fully flight demonstrated , their preliminary The computer station is somewhat similar to the evaluations appear promiSing.
ground station; however , it will not be mobile. Quick- L ____ ._.
CONCLUS I ONS safe execution of a flight demonstration program .
Similarly, adequate ins t rum e n tat ion, excitation Although the validity of analytical predictions methods, data analyses , and coverage of design flight and fl utter model tests have not been discussed, it is envelopes must be provided to insure valid flight test appar e nt tha t the character of the flutter coupling results.
(cata st rophic or otherwise) must be known for the