Document
EXPERIENCES WITH THE DESIGN AND IMPLEMENTATION OF FLUTTER SUPPRESSION SYSTEMS Jerry R. Newsom and Irving Abel NASA Langley Research Center Hampton, Virginia First Annual NASA Aircraft Controls Workshop NASA Langley Research Center Hampton, Virginia October 25-27, 1983 ABSTRACT A considerable amount of research has been conducted on the application of active Because of its impact on safety of controls to increase aircraft performance.
flight, flutter suppression is probably the active controls concept furthest from practical implementation and, therefore, requires significant attention. This attention spans both analytical and experimental studies. Research efforts at NASA have been directed towards the development of analysis and design methodology and the correlation of experimental results with analytical predictions. The purpose of this paper is to discuss some experiences with the design and testing of several flutter suppression systems. Emphasis will be on the experimental activities.
ACTIVE CONTROLSTECHNOLOGY The application of active controls technology (ACT) to reduce aeroelastic response of aircraft structures offers a potential for significant payoff in terms of aerodynamic efficiency and weight savings. To reduce the technical risk associated with this new technology, research was begun in the early 1970's to advance this concept. The technical.program encompasses three areas: control law synthesis, aeroservoelastic analysis, and experiments aimed at verifying both the analysis and synthesis methodology. In the area of control law synthesis, classical methods are being applied where applicable. The latest "state-of-the-art" optimal methods are being refined and applied to the aeroservoelastic case. Innovative approaches are being developed to take highly theoretical synthesis methods which result in complex (high- order) control systems and modify these methods in the design of simpler (low-order) control systems. Strategies are being developed to investigate the sensitivity of the resulting control systems to uncertainty and to incorporate this knowledge into the design cycle. Analysis methods include a comprehensive program (DYLOFLEX) (ref.
1) for calculating the loads on an aeroelastic vehicle equipped with active controls. The evaluation of vehicle static and dynamic stability is being accom- plished using programs and methods developed in-house at LaRC. The experimental pro- gram is aimed at validating the analysis and synthesis methods by comparison with wind tunnel tests and flight results using a remotely piloted drone. The flight test program, called DAST (Drones for Aerodynamic and Structural Testing) (ref. 2), has become the focal point of the experimental validation.
ACT -7;' CONVENTIONAL fi !I SYNTHESIS
II
u FUNCTION BENEFITS .FLUllER SUPPRESSION OREDUCEDWING LOADS l LOAD ALLEVIATION l LOWERWEIGHT @RIDE QUALITY l INCREASEDSPAN . STABILITY *HIGHER L/D AUGMENTATION 0 STABILITY I I ACTIVE FLUTTER SUPPRESSION Active flutter suppression is a concept to increase the flutter speed of a vehicle through the use of active feedback control. The active control system consists of (1) control surfaces, (2) sensors, and (3) control laws. Through proper selection of the control surfaces and sensors and design of the control laws, the damping of the The aeroelastic system can be augmented and thereby increase the flutter speed.
benefit to be derived from flutter suppression is usually reduced structural weight.
4---z
DAMPING
SPEED
CONTROL LAW DESIGN PROCESS The development of methodology to design flutter suppression systems has been an This development ranges from analytical syn- integral part of this research program.
thesis techniques to the overall control law design process. A flowchart of the The first element approach to the overall control law design process is shown below.
of the process is the selection of design objectives (i.e., gain margin, phase margin, etc.). The second element is the selection of a design point (i.e., Mach number and altitude). Control law synthesis is then performed at the design point.
The next element, analysis, provides information on the performance of the control law at off-design flight conditions. If the design objectives at the off-design then a gain scheduler which may he a function of Mach flight conditions are not met, number and/or dynamic pressure is evaluated. If a gain scheduler will not meet the design objectives, then a path back to control law synthesis is selected.
SELECT
DESIGN OBJECTIVES
I
OPTIMAL CONTROL LAW
.-
SYNTHESIS
[YES
CONTROL LAW SYNTHESIS The major emphasis in the development of design methodology has been in the area of control law synthesis. To accomplish the objectives of control law synthesis, as stated below, the practical problems in control law implementation must be The historical development of control law synthesis methodology has pro- recognized.
ceeded from classical techniques to optimal control theory to optimization techniques Both unconstrained and constrained optimization techniques have been W;,,,“,“’ l Beginning recently, emphasis is being given to the use of constrained optimization techniques since several design objectives can then be satisfied simultaneously.
l OBJECTIVE:
. DESIGN A LOW-ORDERCONTROL LAW FOR A HIGH-ORDER SYSTEM
TO MEET SEVERAL DESlGN OBJECTIVES
. LOW ORDER- SIMPLE IMPLEMENTATION
EH ORDER-CHARACTERISTIC OF AEROELASTICSYSTEMS
l METHODOLOGY DEVELOPMENT
l CLASSICAL TECHNIQUES
. OPTIMAL CONTROL THEORY/ORDER REDUCTION
. OPTIMIZATION TECHNIQUES
l UNCONSTRAINED OPTIMIZATION
. CONSTRAINEDOPTIMIZATION
WIND TUNNEL STUDIES Wind tunnel studies of aeroelastic models have been a cornerstone of the NASA research program. Presented in this chart are a number of models that have been used to demonstrate active control concepts on a variety of configurations. ,The Delta- wing model was an early experimental demonstration of flutter suppression (ref. 9).
The B-52 model was tested in support of a USAF/Boeing flight study on active controls Wing load alleviation was studied in support of a USAF/Lockheed program (ref. 10).
using a C-5A model (ref. 11). The DAST ARW-1 model was used for a variety of flutter suppression studies including an evaluation of a control system that would ultimately be tested on a remotely piloted research flight vehicle. Control laws were synthe- sized and tested on the model using classical, aerodynamic energy, and optimal methods (ref. 12). The F-16 and YF-17 model tests have shown active flutter suppres- sion to be a promising method for preventing wing/external store flutter (refs. 13 and 14). Use of active controls is especially attractive for fighters because of the multitude of possible store configurations. These studies are part of an Air Force Flight Dynamics Laboratory/General Dynamics/Northrop/NASA cooperative effort. A cooperative effort was also conducted with the McDonnell Douglas Corporation on a DC-10 derivative wing. Increases in flutter speeds in excess of 26 percent were demonstrated. This study is reported in reference 15.
DAST ARW-1 DELTA-WING MODEL Experimental studies have made a major contribution to the active control technology The Delta-wing model (whose photograph is in this program developed at NASA.
chart) was the first experimental demonstration of flutter suppression in this At a Mach number of 0.9, increases in the flutter dynamic pressure country (ref. 9).
One ranging from 12.5 percent to 30 percent were demonstrated with active controls.
of the major contributions of this wind tunnel program was the development of minia- These actuators paved the way for future wind tunnel tests ture hydraulic actuators.
To evaluate the performance of an active flutter of aeroelastically scaled models.
suppression (AFS) system, subcritical response techniques must be employed. Three different methods were used to determine subcritical response of the Delta-wing model, Analytical methods were used to predict and the results are described in reference 9.
and the results agreed reasonably well with both open-loop and closed-loop stability, However, for the closed-loop case, it was necessary to use a control the experiment.
surface aerodynamic correction factor that was derived using measured hinge moment data.
,DAST MODEL The aeroelastic model used for this study was originally built to support the DAST flight program (ref. 2). The objective of the wind tunnel study was to demonstrate a 44-percent increase in flutter dynamic pressure. Two control laws were designed (ref. 12). One control law was based on the aerodynamic energy method, and the other was based on the results of optimal control theory. At Mach 0.95, a 44-percent increase in flutter dynamic pressure was achieved with both control laws, thereby validating the two synthesis methodologies. Experimental results indicated, however, that the performance of the systems was not as good as that predicted by analysis.
The results also indicated that wind tunnel turbulence is an important factor in both control law synthesis and experimental demonstration.
SIGNIFICANCE I 0 44 % INCREASE IN FLUTTER DYNAMIC PRESSURE l VALIDATED SYNTHESIS METHODOLOGY a WiND-TUNNEL TURBULENCE EFFECiS ; ’ l HIGH-FREQUENCY CONTROL/ STRUCTURE INSTABILITIES CONTROL LAW PERFORMANCE An illustration of flutter suppression performance for the DAST model is shown below.
On the left, the spectrum of outboard peak accelerations for the wing with system off is compared to that for the system on. The model was being excited by tunnel turbulence. The data were measured at a dynamic pressure just below the system-off flutter boundary at M = 0.90. The decrease in amplitude and shift in the maximum response frequency resulting from the control law is evident. Also presented is a plot of flutter dynamic pressure as a function of Mach number. Data for both system off and system on are shown. The flutter suppression system is most effective (i.e., provides the largest increase in flutter dynamic pressure) at the higher Mach The effectiveness is significantly reduced at lower Mach numbers. A dis- numbers.
cussion of these results is given in reference 12.
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CONTROL SYSTEMOFF
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DC-10 MODEL A cooperative study was conducted with the Douglas Aircraft Company to apply control law design methods developed by NASA to a realistic transport configuration and to provide a rapid transfer of research technology to industry. These studies were an extension of previous wind tunnel tests performed by Douglas (ref. 16). The aeroelastic model (shown in the photograph on this chart) is representative of a wing which has a 4.27-m-span increase over the standard DC-10 wing.
Two control laws were designed at NASA Langley using different design methods (ref..
15). Both control laws resulted in a 59-percent increase in flutter dynamic pressure. The performance of the control laws as a function of gain and phase was also evaluated. Calculations performed prior to wind tunnel testing predicted all experimental trends. both structural damping and phase During the wind tunnel tests, characteristics of the actuator were identified as very important factors related to the effectiveness of the control laws. In addition, a correction factor was used to account for control surface effectiveness and did improve the correlation between measured and predicted characteristics.
!
P .’ ’ :o 58 # Il&tEASE IN FLUTTER 0 <DYNAMIC PRESSURE ‘: ^ , ‘-0 PER;ORMANCE AS A FLI~CTION OF I,“““: .,GAIN AND PHASE ‘, ^ .+ ANALYSis PREDICTED ALL _!,~XPERWIENTAL TRENDS ^I: < II i’, : “9 STRUCTUFjAL DAMPING EFFECTS * L ‘: x, :I & ACTljATOR DYNAMICS ^ ,‘. CONTkL SURFACE AERODVNAMIC .CORhECTION ’ ,,, ‘:A’, 2: MEASUREDAND PREDICTED STABILITY BOUNDARIES AS A FUNCTION OF SYSTEM GAIN AND PHASE FOR DC-10 WIND TUNNEL MODEL Measured and predicted stability boundaries in terms of flutter velocity versus system gain and phase are presented below. Three or four distinct flutter modes are exhibited, depending on phase angle.
For all phase angles analyzed, a decrease in flutter velocity is shown for mode 3 at low values of gain. At negative phase angles, the reduction in flutter velocity is more pronounced. The velocity at which mode 8 goes unstable is nearly independent of system gain and phase.
The mode 4 instability is aggravated by negative phase angles and stays relatively fixed for positive phase angles.
At phase angles of +20" and above, a new flutter mode result- ing from a coupling between the feedback filter mode and the first wing bending mode becomes critical. A detailed discussion of these results can be found in reference 15.
-CONTROLLAW NO. 1 MODE 0 ANALYSIS 0 3 0 8 0 EXPERIMENT 04 ti EXPERIMENT (NO FLUTTER) A FILTER >
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0 .5 1.0 1.5 0 .5 1.0 1.5 Kg 0 -5 do ls5 Kg .5 Kgl.o O 115 DAST: WHAT IS IT?
The concept of the DAST program (ref. 2) is to provide a focus for evaluation and improvement of synthesis and analysis procedures for aerodynamic loads prediction and design of active control systems on wings with significant aeroelastic effects.
Major challenges include applications to wings with supercritical airfoil and tests emphasizing the transonic speed range. The program requires complete solutions to real-world problems since research wings are fabricated and flight tested. Because of the risky nature of the flight testing, especially with regard to flutter, target drone aircraft are modified for use as test bed aircraft.
PRINCIPAL RESEARCH AREAS EMPHASIS
l ACTIVE CONTROLSYSTEMS EVALUATIONS . TRANSONIC REGION
. AEROELASTI C EFFECTS
. AERODYNAMIC LOADS MEASUREMENT
. STRUCTURALINVESTIGATIONS
. STABILITY AND PERFORMANCE STUDIES
-._- ------ DAST: HOW DO WE DO IT?
DAST uses an Air Force version of the Firebee II target drone as the basic test bed.
The standard Firebee wing is removed and replaced with the research wing of as depicted in this chart, involves an air The operational sequence, interest.
launch from beneath the wing of a carrier aircraft; a free-flight test phase of between 20 and 40 minutes (depending on Mach number and altitude); followed by a mid- During the free-flight air retrieval by helicopter via a parachute recovery system.
phase, a test pilot controls the vehicle from a ground cockpit. An F-104 aircraft is used as chase, and the copilot of this aircraft serves as a backup flight controller for the drone in case of a malfunction with the uplink system. Data from the experi- ments are provided in real time to the ground by means of a pulse-code-modulated telemetry system. Experimenters provide real-time assessments of the status of the This assessment is based on research wing and its associated active control systems.
the response of the wing to control surface sweeps and pulses. Flight tests are per- formed at the NASA Dryden Flight Research Facility located at Edwards Air Force Base, California.
, ,: ; .DRONETESTVEkitCLE : DAST: WHAT ARE WE DOING IT WITH?
Two transport-type research wings have been built for flight testing. The first wing, Aeroelastic Research Wing No. 1 (ARW-l), was designed for M = 0.98 cruise and was purposely designed to flutter within the flight envelope. Tests of the first research wing configuration have been terminated due to loss of the aircraft result- ing from vehicle systems problems. However, valuable flutter data and test technique experience were acquired. References 17-20 provide a description of these results.
The wing fabrication and test planning for the second research wing (ARW-2) have been sponsored by the NASA Aircraft Energy Efficiency program. This design involved what is believed to be the first exercise of an iterative procedure integrating aerodynamics, structures, and controls technologies in a design loop resulting in flight hardware. Evaluation of multiple active controls systems operating simultaneously, the operation of which is necessary to preserve structural integrity for various flight conditions, is the primary objective of the flight tests on this fuel-conservative-type research wing.
DAST RESEARCHWINGS ,,-ARW-2 ARW-1 l FLUTTERWITHIN FLIGHT ENVELOPE l ACTIVE FLUTTERSUPPRESSION SYSTEM l SUPERCRITICAL AIRFOIL ARW-2 . FUEL CONSERVATIVEWING DESIGN l HIGH ASPECT RATIO (AR = 10.3) l LOW SWEEP(A = 25O) l ADVANCED SUPERCRITICAL AIRFOIL l MULTIPLE ACTIVE CONTROLSCRITICAL TO FLIGHT OPERATION l FSS l MLA l GLA l RSS CORRELATION OF MEASUREDAND PREDICTED DAMPING AND FREQUENCYVARIATIONS (ARW-1) The frequency and damping of the dominant mode for the symmetric case are shown below. The analysis and flight test data are for a test altitude of 4.56 kilometers. The change in frequency with Mach number is predicted well for both the FSS-off and FSS-on cases. However, analysis overpredicts the damping for both the FSS-off and FSS-on cases.
The experimental flutter speed is extrapolated to be approximately M = 0.80 for the FSS-off case.
An actual flutter point was encountered for the FSS-on case at M = 0.82. Other data comparisons can be found in reference 17.
ALTITUDE =4.57 km: SYMMETRIC
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ARW-2 RIGHT SEMISPAN IN LAB PRIOR TO WIND TUNNEL TESTING The ARW-2 wing panels have been fabricated and are being used to support two ground tests. The left semispan has been used to conduct a hardware-in-the-loop test of the active control system electronics. The right semispan shown in the photograph below has been tested in the Langley Transonic Dynamics Tunnel to obtain unsteady pressure distributions. This is believed to be the first measurement of unsteady pressures on The pressure measurements from the wind tunnel will a flexible supercritical wing.
be compared against those measured during the flight tests. A secondary objective of the wind tunnel test is to investigate possible angle-of-attack effects on the flutter boundary at high transonic speeds.
CONCLUSIONS A large amount of expertise has been acquired through the analytical and experimental studies conducted to date. Many lessons have been learned that help guide the future research directions. A few of these lessons are shown below. The first three les- sons are technical in nature and have been or are presently receiving attention.
However, even though the last lesson is nontechnical in nature, it certainly needs to receive more attention. Several of the future thrusts listed below are being researched at the present. These include the use of transonic time plane unsteady aerodynamics, applying flutter suppression methodology to other active control functions, and synthesis of multiple active control systems. The other thrusts are not presently being emphasized but are still on the list of future work.
FUTURETHRUSTS
LESSONSLEARNED
l UNSTEADYAERO THEORYNEEDS l TRANSONIC TIME PLANE UNSTEADY
AERO
l CONTROLSURFACE
.ARBITRARY MOTION 0 SYSTEMATIC METHODSFOR
LOCATING CONTROLSURFACES
l ACCURATEDEFINITION OF AND SENSORS
ACTUATORDYNAMICS
l APPLY FLUllER SUPPRESSION
l ACCURATETURBULENCE MODEL METHODOLOGY TO OTHERACTIVE
CONTROLFUNCTIONS
l CLOSERCOOPERATIONBETWEEN
0 SYNTHESIS OF MULTIPLE ACTIVE
AEROELASTICIANAND CONTROLS
CONTROLSYSTEMS
ANALYST
0 CONTROLCONFIGURED VEHICLES
REFERENCES Perry, B. III; Kroll, R. I., Miller, R. D.; and Goetz, R. C.: DYLOFLEX: A 1.
Computer Program for Flexible Aircraft Flight Dynamic Loads Analyses With Active Controls. J. Aircraft, Vol. 17, No. 4, April 1980, pp. 275-282.
Drones for Aerodynamic and Structural 2. Murrow, H. N.; and Eckstrom, C. V.: Testing (DAST)--A Status Report. J. Aircraft, Vol. 16, No. 8, August 1979, pp.
521-526.
3. Newsom, J. R.: A Method for Obtaining Practical Flutter Suppression Control Laws Using Results of Optimal Control Theory. NASA TP 1471, Aug. 1979.
4. Gangsaas, Dagfinn; and Ly, Vy-Loi: Application of a Modified Linear Quadratic Gaussian Design to Active Control of a Transport Airplane. AIAA Paper No.
79-1746, August 1979.
5. Mahesh, J. K.; Stone, C. R.; Garrard, W. L.; and Hausman, P. D.: Active Flutter Control for Flexible Vehicles. NASA CR-159160, Nov. 1979.
6. Nissim, E.; and Abel, I.: Development and Application of an Optimization Procedure for Flutter Suppression Using the Aerodynamic Energy Concept. NASA TP 1137, Feb. 1978.
7. Mukhopadhyay, V.; Newsom, J. R.; and Abel, I.: A Method for Obtaining Reduced Order Control Laws for High Order Systems Using Optimization Techniques. NASA TP 1876, 1981.
8. Newsom, J. R.; and Mukhopadhyay, V.: Application of Constrained Optimization to Active Control of Aeroelastic Response. NASA TM 83150, June 1981.
9. Sandford, M. C.; Abel, I.; and Gray, D. L.: Development and Demonstration of a Flutter Suppression System Using Active Controls. NASA TR 450, December 1975.
10. Redd, L. T.; Gilman, J., Jr.; Cooley, D. E.; and Sevart, F. D.: A Wind Tunnel Study of B-52 Model Flutter Suppression. AIAA Paper No. 74-401, April 1974.
11. Doggett, R. V., Jr.; Abel, I.; and Ruhlin, C. L.: Some Experiences Using Wind Tunnel Models in Active Control Studies.
NASA TM X-3409, August 1976, pp.
831-892.
12. Newsom, J. R.; Abel, I.; and Dunn, H. J.: Application of Two Design Methods for Active Flutter Suppression and Wind Tunnel Test Results. NASA TP 1653, May 1980.
13. Peloubet, R. P., Jr.; Haller, R. L.; and Bolding, R. M.: F-16 Flutter Suppression System Investigation. AIAA Paper No. 80-0768, May 1980.
14. Hwang, C.; Johnson, E. H.; and Pi, W. S.: Recent Developments of the YF-17 Active Flutter Suppression System. AIAA Paper No. 80-0769, May 1980.
15. Abel, I.; and Newsom, 3. R.: Wind Tunnel Evaluation of NASA-Developed Control Laws for Flutter Suppression on a DC-10 Derivative Wing. AIAA Paper No. 81-0639, April 1981.
16. Winther, B. A.; Shirley, W. A.,; and Heimbough, R. M.: Wind Tunnel Investigation of Active Controls Technology Applied to a DC-10 Derivative. AIAA Paper No. 80-0771, May 1980.
and Pototzky, Anthony S.: Comparison of Analysis and Flight 17. Newsom, Jerry R.; Test Data for a Drone Aircraft with Active Flutter Suppression. AIAA Paper No. 81-0640, April 1981.
18. Edwards, J. W.: Flight Test Results of an Active Flutter Suppression System Installed on a Remotely-Piloted Research Vehicle. AIAA Paper No. 81-0655, April 1981.
Application of a Flight Test and Data Anaysis 19. Bennett, R. M.; and Abel, I.: Technique to Flutter of a Drone Aircraft. AIAA Paper No. 81-0652, April 1981.
20. Newsom, Jerry R.; Pototzky, Anthony S.; and Abel, Irving: Design of the Flutter Suppression System for DAST ARW-1R - A Status Report. NASA TM 84642, March 1983.