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Farmer, M. G.

Publications and source records attributed to Farmer, M. G..

At least 19 records
All sourcesNASA NTRSDOE OSTI
NASA NTRS1984

Model mount system for testing flutter

A wind tunnel model mount system is disclosed for effectively and accurately determining the effects of attack and airstream velocity on a model airfoil or aircraft. The model mount system includes a rigid model attached to a splitter plate which is supported away from the wind tunnel wall several of flexible rods. Conventional instrumentation is employed to effect model rotation through a turntable and to record model flutter data as a function of the angle of attack versus dynamic pressure.

Farmer, M. G.↗
NASA NTRS1983

Two-Degree-of-Freedom Mount System for Flutter Models

Flexible rods replace conventional bearing supports to minimize structural damping. Aerodynamic damping not masked by effects of mount system, making more accurate studies possible of how aerodynamic damping varies as flow over model changed. New system called PAPA.

Farmer, M. G.↗
NASA NTRS1983

Adaptive flutter suppression, analysis and test

Methods of adaptive control have been applied to suppress a potentially violent flutter condition of a half-span model of a lightweight figher aircraft. This marked the confluence of several technologies with active flutter suppression, digital control and adaptive control theory the primary contributors. The control algorithm was required to adapt both to slowly varying changes, corresponding to changes in the flight condition or fuel loading and to rapid changes, corresponding to a store release or the transition from a stable to an unstable flight condition. The development of the adaptive control methods was followed by a simulation and checkout of the complete system and a wind tunnel demonstration. As part of the test, a store was released from the model wing tip, transforming the model abruptly from a stable configuration to a violent flutter condition. The adaptive algorithm recognized the unstable nature of the resulting configuration and implemented a stabilizing control law in a fraction of a second. The algorithm was also shown to provide system stability over a range of wind tunnel Mach numbers and dynamic pressures.

Johnson, E. H.↗
NASA NTRS1982

Flutter clearance of the horizontal tail of the Bellanca Skyrocket II airplane

The Skyrocket II is an all composite constructed experimental prototype airplane. A flutter clearance program was conducted on the horizontal tail so that the airplane could be safely flown to acquire natural laminar flow aerodynamic data. Ground vibration test data were used in a lifting surface flutter analysis to predict symmetric and antisymmetric flutter boundaries. Subcritical response data which were acquired during flight tests are compared with the analytical results. The final flutter clearance placard speed was based on flight test data.

Ricketts, R. H.↗
NASA NTRS1982

A two-degree-of-freedom flutter mount system with low damping for testing rigid wings at different angles of attack

A wind tunnel model mount system for conducting flutter research using a rigid wing was developed. The wing is attached to a splitter plate so that the two move as one rigid body. The splitter plate is supported away from the tunnel wall by a system of rods with fixed fixed and conditions. The rods flex in such a way that only pitch and plunge oscillations are permitted. At the tunnel wall the rods are attached to a remotely controlled turntable so that angle of attack can be varied. Wind tunnel data obtained by using the mount system are presented for a supercritical and a conventional airfoil. Both classical flutter and stall flutter data are presented.

Farmer, M. G.↗
NASA NTRS1982

Measured and calculated effects of angle of attack on the transonic flutter of a supercritical wing

The effects of angle of attack between 0 and 4 degrees were studied. The results indicate that increasing angle of attack from zero can produce substantial changes in the transonic flutter characteristics that are favorable or unfavorable depending on Mach number and angle of attack. The bottom of the transonic flutter-boundary 'bucket' is shown to occur at lower Mach number as angle of attack increases. These flutter results correlate well with the effects of Mach number and angle of attack on aerodynamic behavior, especially on the development of transonic flow phenomena. The calculated flutter characteristics are in good agreement with the experimental data at zero angle of attack, but at nonzero angles of attack the experiments show sharply declining and backward-turning transonic flutter boundaries that are not indicated by the calculations. These may be caused by variations in static aeroelastic deformation.

Yates, E. C., Jr.↗
NASA NTRS1981

Prediction of transonic flutter for a supercritical wing by modified strip analysis and comparison with experiment

Use of a supercritical airfoil can adversely affect wing flutter speeds in the transonic range. As adequate theories for three dimensional unsteady transonic flow are not yet available, the modified strip analysis was used to predict the transonic flutter boundary for the supercritical wing. The steady state spanwise distributions of section lift curve slope and aerodynamic center, required as input for the flutter calculations, were obtained from pressure distributions. The calculated flutter boundary is in agreement with experiment in the subsonic range. In the transonic range, a transonic bucket is calculated which closely resembles the experimental one with regard to both shape and depth, but it occurs at about 0.04 Mach number lower than the experimental one.

Yates, E. C., Jr.↗
NASA NTRS1981

Prediction of transonic flutter for a supercritical wing by modified strip analysis and comparison with experiment

The experiments of Farmer, Hanson, and Wynne (1976) showed that use of a supercritical airfoil can adversely affect wing flutter speeds in the transonic range. Inasmuch as adequate theories for three-dimensional unsteady transonic flow are not yet available, the modified strip analysis published by Yates in 1958 has been used to predict the transonic flutter boundary for the supercritical wing tested by Farmer, Hanson, and Wynne. The steady-state spanwise distributions of section lift-curve slope and aerodynamic center, required as input for the flutter calculations, were obtained from pressure distributions measured by Harris (1972). The calculated flutter boundary is in excellent agreement with experiment in the subsonic range. In the transonic range a 'transonic bucket' is calculated which closely resembles the experimental one with regard to both shape and depth, but it occurs at about 0.04 Mach number lower than the experimental one.

Yates, E. C., Jr.↗
NASA NTRS1980

Wind tunnel test of a fighter aircraft wing/store flutter suppression system: An international effort

A 30% scale, half span model of a lightweight fighter aircraft with an active wing/store flutter suppression system was tested in the NASA Langley Research Center sixteen foot transonic dynamics tunnel. The test featured a store configuration that was intentionally designed to exhibit a violent flutter condition. In addition to Northrop organized control laws, three European countries also contributed control laws to stabilize this condition. After the control laws were mechanized by Northrop, they were tested at the Langley facility. The model was tested up to 170% of the open loop flutter dynamic pressure in a number of cases, with the indication that a substantially greater improvement was achievable. Some special features of the test model are discussed and the design and implementation of the control laws as well as the test monitoring techniques and results are presented.

Hwang, C.↗
NASA NTRS1979

Demonstration of aircraft wing/store flutter suppression systems

A wind tunnel test program was conducted to demonstrate the active wing/store flutter suppression systems on a lightweight fighter aircraft. The program, completed in mid-1978, included the design, analysis, fabrication, and testing of a scale model. The tests were conducted at the NASA Langley 16-ft Transonic Dynamics Tunnel. Three store configurations were selected for testing. Two of these configurations were deliberately designed to exhibit low flutter speeds with rapid reductions in damping at the incipient flutter condition. After initial tunnel entries, which showed the need for certain improvements in the model and the control system design, substantial increases in the flutter speeds were achieved using both leading- and trailing-edge control surfaces, separately. For the most critical configuration, a demonstrated improvement of 18% and a projected improvement of 29% in the dynamic pressure were achieved.

Hwang, C.↗
NASA NTRS1977

A wind tunnel technique for determining stability derivatives from cable mounted aeroelastic models

System identification techniques in common use for extracting stability derivatives from flight test data have been adapted for application to data obtained from aeroelastically-scaled flutter models flown in a wind tunnel on a cable mount system. The concept has been applied with reasonable success to data from rigid models of a Space Shuttle Orbiter and a fighter tested in the NASA Langley transonic dynamics tunnel. Further application of this technique should permit extraction of derivatives that include scaled flexibility effects, thereby obtaining additional information from the testing of expensive flutter models.

Bennett, R. M.↗
NASA NTRS1976

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.↗
NASA NTRS1976

Comparison of supercritical and conventional wing flutter characteristics

A wind-tunnel study was undertaken to directly compare the measured flutter boundaries of two dynamically similar aeroelastic models which had the same planform, maximum thickness-to-chord ratio, and as nearly identical stiffness and mass distributions as possible, with one wing having a supercritical airfoil and the other a conventional airfoil. The considerations and problems associated with flutter testing supercritical wing models at or near design lift coefficients are discussed, and the measured transonic boundaries of the two wings are compared with boundaries calculated with a subsonic lifting surface theory.

Farmer, M. G.↗
NASA NTRS1976

Comparison of supercritical and conventional wing flutter characteristics

A wind-tunnel study is described in which it was attempted to compare the measured flutter boundaries of two dynamically similar aeroelastic models with identical planform, maximum thickness-to-chord ratio, and as nearly identical stiffness and mass distributions as possible, but with one wing having a supercritical airfoil and the other a conventional one. At subsonic Mach numbers, the flutter boundary for the supercritical wing was above that of the conventional wing, as predicted by flutter calculations using subsonic lifting theory. In the transonic region, however, the supercritical wing boundary decreases more rapidly and the minimum flutter point occurs at a dynamic pressure below the conventional wing boundary. Airfoil shape effects may account for some of the difference in the flutter boundaries of the two airfoils.

Farmer, M. G.↗
NASA NTRS1975

A preliminary study of the effects of vortex diffusers (winglets) on wing flutter

Some experimental flutter results are presented for a simple, flat-plate wing model and for the same wing model equipped with two different upper surface vortex diffusers over the Mach number range from about 0.70 to 0.95. Both vortex diffusers had the same planform, but one weighed about 0.3 percent of the basic wing weight, whereas the other weighed about 1.8 percent of the wing weight. The addition of the lighter vortex diffuser reduced the flutter dynamic pressure by about 3 percent; the heavier vortex diffuser reduced the 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.↗