Evaluation of four subcritical response methods for on-line prediction of flutter onset in wind tunnel tests
Previously cited in issue 13, p. 2024, Accession no. A82-30140
Ruhlin, C. L.↗
Engineering topics
Doggett, R. V., Jr.
Publications and source records attributed to Doggett, R. V., Jr..
Evaluation of four subcritical response methods for on-line prediction flutter onset in wind-tunnel tests
The methods were evaluated for use in tests where the flutter model is excited solely by airstream turbulence. The methods were: randomdec, power-spectral-density, peak-hold, and cross-spectrum. The test procedure was to maintain a constant Mach number (M) and increase the dynamic pressure (g) in incremental steps. The test Mach numbers were 0.65, 0.75, 0.82, 0.90, and 1.15. The four methods provided damping trends by which the flutter mode could be tracked and extrapolated to a flutter-onset q. A hard flutter point was obtained at M = 0.82. The peak-hold and cross-spectrum methods gave reliable results and could be most readily used for on-line testing. At M = 0.82, a p-k analysis predicted the same flutter mode as the experiment but a 6-percent lower flutter q. At the subcritical dynamic pressures, calculated damping values were appreciably lower than measured data.
Ruhlin, C. L.↗
Evaluation of four subcritical response methods for on-line prediction of flutter onset in wind-tunnel tests
Four subcritical response methods were evaluated for on-line use in transonic wind-tunnel tests where the flutter model is excited solely by airstream turbulence. The methods were: randomdec, power-spectral-density, peak-hold, and cross-spectrum. Subcritical response data were obtained during tests in the Langley Transonic Dynamics Tunnel of a cantilevered flutter model wing. The test procedure was to maintain a constant Mach number and increase the dynamic pressure (q) in incremental steps. The four methods provided damping trends by which the flutter mode could be tracked and extrapolated to a flutter-onset q. A hard flutter point was obtained at M = 0.82. The peak-mold and cross-spectrum methods gave reliable results and could be most readily used for on-line testing.
Ruhlin, C. L.↗
Measured transonic unsteady pressures on an energy efficient transport wing with oscillating control surfaces
Highlight results are presented from subsonic and transonic pressure measurement studies conducted in the Langley Transonic Dynamics Tunnel on a supercritical wing model representative of an energy efficient transport design. Steady- and unsteady-pressure data were acquired on the upper and lower wing surface at an off-design Mach number of 0.60 and at the design Mach number of 0.78, for a Reynolds number of 2.2 x 10(6) (based on the wing average chord). The model configuration consisted of a sidewall-Mounted half-body fuselage and a semi-span wing with an aspect ratio of 10.76, a leading-edge sweepback angle of 28.8 degrees, and supercritical airfoil sections. The wing is instrumented with 252 static pressure orifices and 164 dynamic pressure gages. Model test variables included wing angle of attack, control-surface mean deflection angle, control-surface oscillating deflection angle and frequency, and phasing between oscillating leading-edge and trailing-edge controls when used together.
Cazier, F. W., Jr.↗
Aeroelastic instability stoppers for wind tunnel models
A mechanism for constraining models or sections thereof, was wind tunnel tested, deployed at the onset of aeroelastic instability, to forestall destructive vibrations in the model is described. The mechanism includes a pair of arms pivoted to the tunnel wall and straddling the model. Rollers on the ends of the arms contact the model, and are pulled together against the model by a spring stretched between the arms. An actuator mechanism swings the arms into place and back as desired.
Doggett, R. V., Jr.↗
Aeroelastic instability stoppers for wind tunnel models
A mechanism for diverting the flow in a wind tunnel from the wing of a tested model is described. The wing is mounted on the wall of a tunnel. A diverter plate is pivotally mounted on the tunnel wall ahead of the model. An actuator fixed to the tunnel is pivotably connected to the diverter plate, by plunger. When the model is about to become unstable during the test the actuator moves the diverter plate from the tunnel wall to divert maintaining stable model conditions. The diverter plate is then retracted to enable normal flow.
Doggett, R. V., Jr.↗
Effects of angle of attack and vertical fin on transonic flutter characteristics of an arrow-wing configuration
Experimental transonic flutter results are presented for a simplified 1/50 size, aspect ratio 1.77, wind tunnel model of an arrow wing design. Flutter results are presented for two configurations; namely, one with and one without a ventral fin mounted at the 0.694 semispan station. Results are presented for both configurations trimmed to zero lift and in a lifting condition at angles of attack up to 4 deg. The results show that the flutter characteristics of both configurations are similar to those usually observed. Increasing angle of attack reduces the flutter dynamic pressure by a small amount (about 13 percent maximum) for both configurations. The addition of the fin to the basic wing increases the flutter dynamic pressure. Calculated results for both configurations in the nonlifting condition obtained by using subsonic doublet lattice unsteady aerodynamic theory correlate reasonably well with the experimental results. Calculated results for the basic wing obtained by using subsonic kernal function unsteady aerodynamic theory did not agree as well with the experimental data.
Doggett, R. V., Jr.↗
Dynamic response of a forward-swept-wing model at angles of attack up to 15 deg at a Mach number of 0.8
Root mean square (rms) bending moments for a dynamically scaled, aeroelastic wing of a proposed forward swept wing, flight demonstrator airplane are presented for angles of attack up to 15 deg at a Mach number of 0.8 The 0.6 size semispan model had a leading edge forward sweep of 44 deg and was constructed of composite material. In addition to broad band responses, individual rms responses and total damping ratios are presented for the first two natural modes. The results show that the rms response increases with angle of attack and has a peak value at an angle of attack near 13 deg. In general, the response was characteristic of buffeting and similar to results often observed for aft swept wings. At an angle of attack near 13 deg, however, the response had characteristics associated with approaching a dynamic instability, although no instability was observed over the range of parameters investigated.
Doggett, R. V., Jr.↗
Wind-tunnel experiments on divergence of forward-swept wings
An experimental study to investigate the aeroelastic behavior of forward-swept wings was conducted in the Langley Transonic Dynamics Tunnel. Seven flat-plate models with varying aspect ratios and wing sweep angles were tested at low speeds in air. Three models having the same planform but different airfoil sections (i.e., flat-plate, conventional, and supercritical) were tested at transonic speeds in Freon 12. Linear analyses were performed to provide predictions to compare with the measured aeroelastic instabilities which include both static divergence and flutter. Six subcritical response testing techniques were formulated and evaluated at transonic speeds for accuracy in predicting static divergence. Two "divergence stoppers" were developed and evaluated for use in protecting the model from structural damage during tests.
Ricketts, R. H.↗
Experimental and analytical transonic flutter characteristics of a geared-elevator configuration
The flutter model represented the aft fuselage and empennage of a proposed supersonic transport airplane and had an all movable horizontal tail with a geared elevator. It was tested mounted from a sting in the transonic dynamics tunnel. Symmetric flutter boundaries were determined experimentally at Mach numbers from 0.7 to 1.14 for a geared elevator configuration (gear ratio of 2.8 to 1.0) and an ungeared elevator configuration (gear ratio of 1.0 to 1.0). Gearing the elevator increased the experimental flutter dynamic pressures about 20 percent. Flutter calculations were made for the geared elevator configuration by using two analytical methods based on subsonic lifting surface theory. Both methods analyzed the stabilizer and elevator as a single, deforming surface, but one method also allowed the elevator to be analyzed as hinged from the stabilizer. All analyses predicted lower flutter dynamic pressures than experiment with best agreement (within 12 percent) for the hinged elevator method. Considering the model as mounted from a flexible rather than rigid sting in the analyses, had only a slight effect on the flutter results but was significant in that a sting related vibration mode was identified as a potentially flutter critical mode.
Ruhlin, C. L.↗
Dynamic wind-tunnel tests of an aeromechanical gust-alleviation system using several different combinations of control surfaces
Some experimental results are presented from wind tunnel studies of a dynamic model equipped with an aeromechanical gust alleviation system for reducing the normal acceleration response of light airplanes. The gust alleviation system consists of two auxiliary aerodynamic surfaces that deflect the wing flaps through mechanical linkages when a gust is encountered to maintain nearly constant airplane lift. The gust alleviation system was implemented on a 1/6-scale, rod mounted, free flying model that is geometrically and dynamically representative of small, four place, high wing, single engine, light airplanes. The effects of flaps with different spans, two size of auxiliary aerodynamic surfaces, plain and double hinged flaps, and a flap elevator interconnection were studied. The model test results are presented in terms of predicted root mean square response of the full scale airplane to atmospheric turbulence. The results show that the gust alleviation system reduces the root mean square normal acceleration response by 30 percent in comparison with the response in the flaps locked condition. Small reductions in pitch-rate response were also obtained. It is believed that substantially larger reductions in normal acceleration can be achieved by reducing the rather high levels of mechanical friction which were extant in the alleviation system of the present model.
Stewart, E. C.↗
Some experimental and theoretical flutter characteristics of an arrow-wing configuration
Experimental and theoretical flutter results are presented for a simplified 1/50-size wind-tunnel model of an arrow-wing configuration. Transonic flutter characteristics are presented for three configurations - wing without engine nacelles, wing with flow-through nacelles, and wing with pencil nacelles (thin, streamline bodies). Experimental results are correlated with analytical results obtained from kernel-function and doublet-lattice unsteady aerodynamic theories. Theoretical results are presented that show the effects on flutter of systematic changes in structural stiffness and mass.
Doggett, R. V., Jr.↗
Preliminary study of effects of winglets on wing flutter
Some experimental flutter results are presented over a Mach number range from about 0.70 to 0.95 for a simple, swept, tapered, flat-plate wing model having a planform representative of subsonic transport airplanes and for the same wing model equipped with two different upper surface winglets. Both winglets had the same planform and area (about 2 percent of the basic-wing area); however, one weighed about 0.3 percent of the basic-wing weight, and the other weighed about 1.8 percent of the wing weight. The addition of the lighter winglet reduced the wing-flutter dynamic pressure by about 3 percent; the heavier winglet reduced the wing-flutter dynamic pressure by about 12 percent. The experimental flutter results are compared at a Mach number of 0.80 with analytical flutter results obtained by using doublet-lattice and lifting-surface (kernel-function) unsteady aerodynamic theories.
Doggett, R. V., Jr.↗
Some applications of the NASTRAN level 16 subsonic flutter analysis capability
The Level 16 flutter analysis capability was applied to an aspect-ratio-6.8 subsonic transport type wing, an aspect-ratio-1.7 arrow wing, and an aspect-ratio-1.3 all movable horizontal tail with a geared elevator. The transport wing and arrow wing results are compared with experimental results obtained in the Langley transonic dynamic tunnel and with other calculated results obtained using subsonic lifting surface (kernel function) unsteady aerodynamic theory.
Doggett, R. V., Jr.↗
Some experiences using wind-tunnel models in active control studies
A status report and review of wind tunnel model experimental techniques that have been developed to study and validate the use of active control technology for the minimization of aeroelastic response are presented. Modeling techniques, test procedures, and data analysis methods used in three model studies are described. The studies include flutter mode suppression on a delta-wing model, flutter mode suppression and ride quality control on a 1/30-size model of the B-52 CCV airplane, and an active lift distribution control system on a 1/22 size C-5A model.
Doggett, R. V., Jr.↗
Geared-elevator flutter study
An experimental and analytical study was made of the transonic flutter characteristics of a supersonic transport tail assembly model having an all-movable, horizontal tail with a geared elevator. Two model configurations, namely, one with a gear-elevator (2.8 to 1.0 gear ratio) and one with locked-elevator (1.0 to 1.0 gear ratio), were flutter tested in the Langley transonic dynamics tunnel with an empennage cantilever-mounted on a sting. The geared-elevator configuration fluttered experimentally at about 20% higher dynamic pressures than the locked-elevator configuration. The experimental flutter dynamic pressure boundaries for both configurations were nearly flat over a Mach number range from 0.9 to 1.1. Flutter calculations (mathematical models) were made for the geared-elevator configuration using three subsonic lifting-surface methods. In one method, the elevator was treated as a discrete surface, and in the other two methods, the stabilizer and elevator were treated as a single warped-surface with the primary difference between these two methods being in the mathematical implementation used. A comparison of the experimental and analytical results shows that the discrete-elevator method predicted best the experimental flutter dynamic pressure level. However, the single warped-surface methods predicts more closely the experimental flutter frequencies and Mach number trends.
Ruhlin, C. L.↗
Geared-elevator flutter study
The paper describes an experimental and analytical study of the transonic flutter characteristics of an empennage flutter model having an all-movable horizontal tail with a geared elevator. Two configurations were flutter tested: one with a geared elevator and one with a locked elevator with the model cantilever-mounted on a sting in the wind tunnel. The geared-elevator configuration fluttered experimentally at about 20% higher dynamic pressures than the locked-elevator configuration. The experimental flutter boundary was nearly flat at transonic speeds for both configurations. It was found that an analysis which treated the elevator as a discrete surface predicted flutter dynamic pressure levels better than analyses which treated the stabilizer and elevator as a warped surface. Warped-surface methods, however, predicted more closely the experimental flutter frequencies and Mach number trends.
Ruhlin, C. L.↗
Flutter suppression by active control and its benefits
A general discussion of the airplane applications of active flutter suppression systems is presented with focus on supersonic cruise aircraft configurations. Topics addressed include a brief historical review; benefits, risks, and concerns; methods of application; and applicable configurations. Results are presented where the direct operating costs and performance benefits of an arrow wing supersonic cruise vehicle equipped with an active flutter suppression system are compared with corresponding costs and performance of the same baseline airplane where the flutter deficiency was corrected by passive methods (increases in structural stiffness). The design, synthesis, and conceptual mechanization of the active flutter suppression system are discussed. The results show that a substantial weight savings can be accomplished by using the active system. For the same payload and range, airplane direct operating costs are reduced by using the active system. The results also indicate that the weight savings translates into increased range or payload.
Doggett, R. V., Jr.↗