Experimental transonic flutter characteristics of two 72 deg-sweep delta-wing models

NASA Technical Reports Server (NTRS)

Doggett, Robert V., Jr.; Soistmann, David L.; Spain, Charles V.; Parker, Ellen C.; Silva, Walter A.

1989-01-01

Transonic flutter boundaries are presented for two simple, 72 deg. sweep, low-aspect-ratio wing models. One model was an aspect-ratio 0.65 delta wing; the other model was an aspect-ratio 0.54 clipped-delta wing. Flutter boundaries for the delta wing are presented for the Mach number range of 0.56 to 1.22. Flutter boundaries for the clipped-delta wing are presented for the Mach number range of 0.72 to 0.95. Selected vibration characteristics of the models are also presented.

Transonic flight flutter tests of a control surface utilizing an impedance response technique

NASA Technical Reports Server (NTRS)

Mirowitz, L. I.

1975-01-01

Transonic flight flutter tests of the XF3H-1 Demon Airplane were conducted utilizing a frequency response technique in which the oscillating rudder provides the means of system excitation. These tests were conducted as a result of a rudder flutter incident in the transonic speed range. The technique employed is presented including a brief theoretical development of basic concepts. Test data obtained during the flight are included and the method of interpretation of these data is indicated. This method is based on an impedance matching technique. It is shown that an artificial stabilizing device, such as a damper, may be incorporated in the system for test purposes without complicating the interpretation of the test results of the normal configuration. Data are presented which define the margin of stability introduced to the originally unstable rudder by design changes which involve higher control system stiffness and external damper. It is concluded that this technique of flight flutter testing is a feasible means of obtaining flutter stability information in flight.

Contributions of Transonic Dynamics Tunnel Testing to Airplane Flutter Clearance

NASA Technical Reports Server (NTRS)

Rivera, Jose A.; Florance, James R.

2000-01-01

The Transonic Dynamics Tunnel (TDT) became in operational in 1960, and since that time has achieved the status of the world's premier wind tunnel for testing large in aeroelastically scaled models at transonic speeds. The facility has many features that contribute to its uniqueness for aeroelastic testing. This paper will briefly describe these capabilities and features, and their relevance to aeroelastic testing. Contributions to specific airplane configurations and highlights from the flutter tests performed in the TDT aimed at investigating the aeroelastic characteristics of these configurations are presented.

The effects of rotational flow, viscosity, thickness, and shape on transonic flutter dip phenomena

NASA Technical Reports Server (NTRS)

Reddy, T. S. R.; Srivastava, Rakesh; Kaza, Krishna Rao V.

1988-01-01

The transonic flutter dip phenomena on thin airfoils, which are employed for propfan blades, is investigated using an integrated Euler/Navier-Stokes code and a two degrees of freedom typical section structural model. As a part of the code validation, the flutter characteristics of the NACA 64A010 airfoil are also investigated. In addition, the effects of artificial dissipation models, rotational flow, initial conditions, mean angle of attack, viscosity, airfoil thickness and shape on flutter are investigated. The results obtained with a Euler code for the NACA 64A010 airfoil are in reasonable agreement with published results obtained by using transonic small disturbance and Euler codes. The two artificial dissipation models, one based on the local pressure gradient scaled by a common factor and the other based on the local pressure gradient scaled by a spectral radius, predicted the same flutter speeds except in the recovery region for the case studied. The effects of rotational flow, initial conditions, mean angle of attack, and viscosity for the Reynold's number studied seem to be negligible or small on the minima of the flutter dip.

Transonic flutter study of a wind-tunnel model of a supercritical wing with/without winglet. [conducted in Langley Transonic Dynamics Tunnel

NASA Technical Reports Server (NTRS)

Ruhlin, C. L.; Rauch, F. J., Jr.; Waters, C.

1982-01-01

The model was a 1/6.5-size, semipan version of a wing proposed for an executive-jet-transport airplane. The model was tested with a normal wingtip, a wingtip with winglet, and a normal wingtip ballasted to simulate the winglet mass properties. Flutter and aerodynamic data were acquired at Mach numbers (M) from 0.6 to 0.95. The measured transonic flutter speed boundary for each wingtip configuration had roughly the same shape with a minimum flutter speed near M=0.82. The winglet addition and wingtip mass ballast decreased the wing flutter speed by about 7 and 5 percent, respectively; thus, the winglet effect on flutter was more a mass effect than an aerodynamic effect.

Experimental and analytical transonic flutter characteristics of a geared-elevator configuration

NASA Technical Reports Server (NTRS)

Ruhlin, C. L.; Doggett, R. V., Jr.; Gregory, R. A.

1980-01-01

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.

Stall flutter experiment in a transonic oscillating linear cascade

NASA Technical Reports Server (NTRS)

Boldman, D. R.; Buggele, A. E.; Michalson, G. M.

1981-01-01

Two dimensional biconvex airfoils were oscillated at reduced frequencies up to 0.5 based on semi-chord and a free stream Mach number of 0.80 to simulate transonic stall flutter in rotors. Steady-state periodicity was confirmed through end-wall pressure measurements, exit flow traverses, and flow visualization. The initial flow visualization results from flutter tests indicated that the oscillating shock on the airfoils lagged the airfoil motion by as much as 80 deg. These initial data exhibited an appreciable amount of scatter; however, a linear fit of the results indicated that the greatest shock phase lag occurred at a positive interblade phase angle. Photographs of the steady-state and unsteady flow fields reveal some of the features of the lambda shock wave on the suction surface of the airfoils.

Transonic flutter study of a wind-tunnel model of a supercritical wing with/without winglet

NASA Technical Reports Server (NTRS)

Ruhlin, C. L.; Rauch, F. J., Jr.; Waters, C.

1982-01-01

The scaled flutter model was a 1/6.5-size, semispan version of a supercritical wing (SCW) proposed for an executive-jet-transport airplane. The model was tested cantilever-mounted with a normal wingtip, a wingtip with winglet, and a normal wingtip ballasted to simulate the winglet mass properties. Flutter and aerodynamic data were acquired at Mach numbers from 0.6 to 0.95. The measured transonic flutter speed boundary for each wingtip configuration had roughly the same shape with a minimum flutter speed near M = 0.82. The winglet addition and wingtip mass ballast decreased the wing flutter speed by about 7 and 5%, respectively; thus, the winglet effect on flutter was more a mass effect than an aerodynamic effect. Flutter characteristics calculated using a doublet-lattice analysis (which included interference effects) were in good agreement with the experimental results up to M = 0.82. Comparisons of measured static aerodynamic data with predicted data indicated that the model was aerodynamically representative of the airplane SCW.

Flutter tests (IS4) of the 0.0125-scale shuttle reflection plane model 30-OTS in the Langley Research Center 26-inch transonic blowdown tunnel test no. 547

NASA Technical Reports Server (NTRS)

Kotch, M. A.

1974-01-01

A series of slab wing flutter models with rigid orbiter fuselage, external tank, and SRB models of the space shuttle were tested, in a reflection plane arrangement, in the NASA Langley Research Center's 26-inch Transonic Blowdown Tunnel. Model flutter boundaries were obtained for both a wing-alone configuration and a wing-with-orbiter, tank and SRB configuration. Additional test points were taken of the wing-with-orbiter configuration, as a correlation with the wing-alone condition. A description of the wind tunnel models and test procedures utilized in the experiment are provided.

Application of the CAP-TSD unsteady transonic small disturbance program to wing flutter. [Computational Aeroelasticity Program

NASA Technical Reports Server (NTRS)

Bennett, Robert M.; Batina, John T.

1989-01-01

The application and assessment of a computer program called CAP-TSD (Computational Aeroelasticity Program - Transonic Small Disturbance) for flutter predictions are described. Flutter calculations are presented for two thin swept-and-tapered wing planforms with well-defined modal properties. One planform is a series of 45-degree swept wings and the other planform is a clipped delta wing. Comparisons are made between the results of CAP-TSD using the linear equation and no airfoil thickness and the results obtained from a subsonic kernel function analysis. The calculations cover a Mach number range from low subsonic to low supersonic values, including the transonic range, and are compared with subsonic linear theory and experimental data. It is noted that since both wings have very thin airfoil sections, the effects of thickness are minimal.

Aeroelastic Tailoring of Transport Wings Including Transonic Flutter Constraints

NASA Technical Reports Server (NTRS)

Stanford, Bret K.; Wieseman, Carol D.; Jutte, Christine V.

2015-01-01

Several minimum-mass optimization problems are solved to evaluate the effectiveness of a variety of novel tailoring schemes for subsonic transport wings. Aeroelastic stress and panel buckling constraints are imposed across several trimmed static maneuver loads, in addition to a transonic flutter margin constraint, captured with aerodynamic influence coefficient-based tools. Tailoring with metallic thickness variations, functionally graded materials, balanced or unbalanced composite laminates, curvilinear tow steering, and distributed trailing edge control effectors are all found to provide reductions in structural wing mass with varying degrees of success. The question as to whether this wing mass reduction will offset the increased manufacturing cost is left unresolved for each case.

Flutter parametric studies of cantilevered twin-engine transport type wing with and without winglet. Volume 2: Transonic and density effect investigations

NASA Technical Reports Server (NTRS)

Bhatia, K. G.; Nagaraja, K. S.

1984-01-01

Flutter characteristics of a cantilevered high aspect ratio wing with winglet were investigated. The configuration represented a current technology, twin engine airplane. Compressibility effects through transonic Mach numbers and a wide range of mass-density ratios were evaluated on a low speed and high speed model. Four flutter mechanisms were obtained from test, and analysis from various combinations of configuration parameters. It is shown that the coupling between wing tip vertical and chordwise motions have significant effect under some conditions. It is concluded that for the flutter model configurations studied, the winglet related flutter is amenable to the conventional flutter analysis techniques. The low speed model flutter and the high-speed model flutter results are described.

Uncertainty Quantification of the FUN3D-Predicted NASA CRM Flutter Boundary

NASA Technical Reports Server (NTRS)

Stanford, Bret K.; Massey, Steven J.

2017-01-01

A nonintrusive point collocation method is used to propagate parametric uncertainties of the flexible Common Research Model, a generic transport configuration, through the unsteady aeroelastic CFD solver FUN3D. A range of random input variables are considered, including atmospheric flow variables, structural variables, and inertial (lumped mass) variables. UQ results are explored for a range of output metrics (with a focus on dynamic flutter stability), for both subsonic and transonic Mach numbers, for two different CFD mesh refinements. A particular focus is placed on computing failure probabilities: the probability that the wing will flutter within the flight envelope.

Experimental transonic steady state and unsteady pressure measurements on a supercritical wing during flutter and forced discrete frequency oscillations

NASA Technical Reports Server (NTRS)

Piette, Douglas S.; Cazier, Frank W., Jr.

1989-01-01

Present flutter analysis methods do not accurately predict the flutter speeds in the transonic flow region for wings with supercritical airfoils. Aerodynamic programs using computational fluid dynamic (CFD) methods are being developed, but these programs need to be verified before they can be used with confidence. A wind tunnel test was performed to obtain all types of data necessary for correlating with CFD programs to validate them for use on high aspect ratio wings. The data include steady state and unsteady aerodynamic measurements on a nominal stiffness wing and a wing four times that stiffness. There is data during forced oscillations and during flutter at several angles of attack, Mach numbers, and tunnel densities.

Flutter of a Low-Aspect-Ratio Rectangular Wing

NASA Technical Reports Server (NTRS)

Cole, Stanley R.

1989-01-01

A flutter test of a low-aspect-ratio rectangular wing was conducted in the Langley Transonic Dynamics Tunnel (TDT). The model used in this flutter test consisted of a rigid wing mounted to the wind-tunnel wall by a flexible, rectangular beam. The flexible support shaft was connected to the wing root and was cantilever mounted to the wind-tunnel wall. The wing had an aspect ratio of 1.5 based on the wing semispan and an NACA 64A010 airfoil shape. The flutter boundary of the model was determined for a Mach number range of 0.5 to 0.97. The shape of the transonic flutter boundary was determined. Actual flutter points were obtained on both the subsonic and supersonic sides of the flutter bucket. The model exhibited a deep transonic flutter bucket over a narrow range of Mach number. At some Mach numbers, the flutter conditions were extrapolated using a subcritical response technique. In addition to the basic configuration, modifications were made to the model structure such that the first bending frequency was changed without significantly affecting the first torsion frequency. The experiment showed that increasing the bending stiffness of the model support shaft through these modifications lowered the flutter dynamic pressure. Flutter analysis was conducted for the basic model as a comparison with the experimental results. This flutter analysis was conducted with subsonic lifting-surface (kernel function) aerodynamics using the k method for the flutter solution.

Wing box transonic-flutter suppression using piezoelectric self-sensing actuators attached to skin

NASA Astrophysics Data System (ADS)

Otiefy, R. A. H.; Negm, H. M.

2010-12-01

The main objective of this research is to study the capability of piezoelectric (PZT) self-sensing actuators to suppress the transonic wing box flutter, which is a flow-structure interaction phenomenon. The unsteady general frequency modified transonic small disturbance (TSD) equation is used to model the transonic flow about the wing. The wing box structure and piezoelectric actuators are modeled using the equivalent plate method, which is based on the first order shear deformation plate theory (FSDPT). The piezoelectric actuators are bonded to the skin. The optimal electromechanical coupling conditions between the piezoelectric actuators and the wing are collected from previous work. Three main different control strategies, a linear quadratic Gaussian (LQG) which combines the linear quadratic regulator (LQR) with the Kalman filter estimator (KFE), an optimal static output feedback (SOF), and a classic feedback controller (CFC), are studied and compared. The optimum actuator and sensor locations are determined using the norm of feedback control gains (NFCG) and norm of Kalman filter estimator gains (NKFEG) respectively. A genetic algorithm (GA) optimization technique is used to calculate the controller and estimator parameters to achieve a target response.

Results of flutter test OS6 obtained using the 0.14-scale wing/elevon model (54-0) in the NASA LaRC 16-foot transonic dynamics wind tunnel

NASA Technical Reports Server (NTRS)

Berthold, C. L.

1977-01-01

A 0.14-scale dynamically scaled model of the space shuttle orbiter wing was tested in the Langley Research Center 16-Foot Transonic Dynamics Wind Tunnel to determine flutter, buffet, and elevon buzz boundaries. Mach numbers between 0.3 and 1.1 were investigated. Rockwell shuttle model 54-0 was utilized for this investigation. A description of the test procedure, hardware, and results of this test is presented.

Using transonic small disturbance theory for predicting the aeroelastic stability of a flexible wind-tunnel model

NASA Technical Reports Server (NTRS)

Silva, Walter A.; Bennett, Robert M.

1990-01-01

The CAP-TSD (Computational Aeroelasticity Program - Transonic Small Disturbance) code, developed at the NASA - Langley Research Center, is applied to the Active Flexible Wing (AFW) wind tunnel model for prediction of the model's transonic aeroelastic behavior. Static aeroelastic solutions using CAP-TSD are computed. Dynamic (flutter) analyses are then performed as perturbations about the static aeroelastic deformations of the AFW. The accuracy of the static aeroelastic procedure is investigated by comparing analytical results to those from previous AFW wind tunnel experiments. Dynamic results are presented in the form of root loci at different Mach numbers for a heavy gas and air. The resultant flutter boundaries for both gases are also presented. The effects of viscous damping and angle-of-attack, on the flutter boundary in air, are presented as well.

Digital-flutter-suppression-system investigations for the active flexible wing wind-tunnel model

NASA Technical Reports Server (NTRS)

Perry, Boyd, III; Mukhopadhyay, Vivek; Hoadley, Sherwood Tiffany; Cole, Stanley R.; Buttrill, Carey S.

1990-01-01

Active flutter suppression control laws were designed, implemented, and tested on an aeroelastically-scaled wind-tunnel model in the NASA Langley Transonic Dynamics Tunnel. One of the control laws was successful in stabilizing the model while the dynamic pressure was increased to 24 percent greater than the measured open-loop flutter boundary. Other accomplishments included the design, implementation, and successful operation of a one-of-a-kind digital controller, the design and use of two simulation methods to support the project, and the development and successful use of a methodology for online controller performance evaluation.

Basis Function Approximation of Transonic Aerodynamic Influence Coefficient Matrix

NASA Technical Reports Server (NTRS)

Li, Wesley Waisang; Pak, Chan-Gi

2010-01-01

A technique for approximating the modal aerodynamic influence coefficients [AIC] matrices by using basis functions has been developed and validated. An application of the resulting approximated modal AIC matrix for a flutter analysis in transonic speed regime has been demonstrated. This methodology can be applied to the unsteady subsonic, transonic and supersonic aerodynamics. The method requires the unsteady aerodynamics in frequency-domain. The flutter solution can be found by the classic methods, such as rational function approximation, k, p-k, p, root-locus et cetera. The unsteady aeroelastic analysis for design optimization using unsteady transonic aerodynamic approximation is being demonstrated using the ZAERO(TradeMark) flutter solver (ZONA Technology Incorporated, Scottsdale, Arizona). The technique presented has been shown to offer consistent flutter speed prediction on an aerostructures test wing [ATW] 2 configuration with negligible loss in precision in transonic speed regime. These results may have practical significance in the analysis of aircraft aeroelastic calculation and could lead to a more efficient design optimization cycle

Basis Function Approximation of Transonic Aerodynamic Influence Coefficient Matrix

NASA Technical Reports Server (NTRS)

Li, Wesley W.; Pak, Chan-gi

2011-01-01

A technique for approximating the modal aerodynamic influence coefficients matrices by using basis functions has been developed and validated. An application of the resulting approximated modal aerodynamic influence coefficients matrix for a flutter analysis in transonic speed regime has been demonstrated. This methodology can be applied to the unsteady subsonic, transonic, and supersonic aerodynamics. The method requires the unsteady aerodynamics in frequency-domain. The flutter solution can be found by the classic methods, such as rational function approximation, k, p-k, p, root-locus et cetera. The unsteady aeroelastic analysis for design optimization using unsteady transonic aerodynamic approximation is being demonstrated using the ZAERO flutter solver (ZONA Technology Incorporated, Scottsdale, Arizona). The technique presented has been shown to offer consistent flutter speed prediction on an aerostructures test wing 2 configuration with negligible loss in precision in transonic speed regime. These results may have practical significance in the analysis of aircraft aeroelastic calculation and could lead to a more efficient design optimization cycle.

Buffet test in the National Transonic Facility

NASA Technical Reports Server (NTRS)

Young, Clarence P., Jr.; Hergert, Dennis W.; Butler, Thomas W.; Herring, Fred M.

1992-01-01

A buffet test of a commercial transport model was accomplished in the National Transonic Facility at the NASA Langley Research Center. This aeroelastic test was unprecedented for this wind tunnel and posed a high risk to the facility. This paper presents the test results from a structural dynamics and aeroelastic response point of view and describes the activities required for the safety analysis and risk assessment. The test was conducted in the same manner as a flutter test and employed onboard dynamic instrumentation, real time dynamic data monitoring, automatic, and manual tunnel interlock systems for protecting the model. The procedures and test techniques employed for this test are expected to serve as the basis for future aeroelastic testing in the National Transonic Facility. This test program was a cooperative effort between the Boeing Commercial Airplane Company and the NASA Langley Research Center.

Comparison of the NASA Common Research Model European Transonic Wind Tunnel Test Data to NASA Test Data

NASA Technical Reports Server (NTRS)

Rivers, Melissa; Quest, Juergen; Rudnik, Ralf

2015-01-01

Experimental aerodynamic investigations of the NASA Common Research Model have been conducted in the NASA Langley National Transonic Facility, the NASA Ames 11-ft wind tunnel, and the European Transonic Wind Tunnel. In the NASA Ames 11-ft wind tunnel, data have been obtained at only a chord Reynolds number of 5 million for a wing/body/tail = 0 degree incidence configuration. Data have been obtained at chord Reynolds numbers of 5, 19.8 and 30 million for the same configuration in the National Transonic Facility and in the European Transonic Facility. Force and moment, surface pressure, wing bending and twist, and surface flow visualization data were obtained in all three facilities but only the force and moment and surface pressure data are presented herein.

Analytical and experimental study of the effects of wing-body aerodynamic interaction on space shuttle subsonic flutter

NASA Technical Reports Server (NTRS)

Chipman, R. R.; Rauch, F. J.

1975-01-01

The effects on flutter of the aerodynamic interaction between the space shuttle bodies and wing, 1/80th-scale semispan models of the orbiter wing, the complete shuttle and intermediate component combinations were tested in the NASA Langley Research Center 26-inch Transonic Blowdown Wind Tunnel. Using the double lattice method combined with slender body theory to calculate unsteady aerodynamic forces, subsonic flutter speeds were computed for comparison. Using calculated complete vehicle modes, flutter speed trends were computed for the full scale vehicle at an altitude of 15,200 meters and a Mach number of 0.6. Consistent with findings of the model studies, analysis shows the shuttle to have the same flutter speed as an isolated cantilevered wing.

Results of flutter test OS7 obtained using the 0.14-scale space shuttle orbiter fin/rudder model number 55-0 in the NASA LaRC 16-foot transonic dynamics wind tunnel

NASA Technical Reports Server (NTRS)

Berthold, C. L.

1977-01-01

A 0.14-scale dynamically scaled model of the space shuttle orbiter vertical tail was tested in a 16-foot transonic dynamic wind tunnel to determine flutter, buffet, and rudder buzz boundaries. Mach numbers between .5 and 1.11 were investigated. Rockwell shuttle model 55-0 was used for this investigation. A description of the test procedure, hardware, and results of this test is presented.

Comparison of the NASA Common Research Model European Transonic Wind Tunnel Test Data to NASA Test Data

NASA Technical Reports Server (NTRS)

Rivers, Melissa B.; Quest, Jurgen; Rudnik, Ralf

2015-01-01

Experimental aerodynamic investigations of the NASA Common Research Model have been conducted in the NASA Langley National Transonic Facility, the NASA Ames 11-ft wind tunnel, and the European Transonic Wind Tunnel. In the NASA Ames 11-ft wind tunnel, data have been obtained at only a chord Reynolds number of 5 million for a wing/body/tail = 0 degree incidence configuration. Data have been obtained at chord Reynolds numbers of 5, 19.8 and 30 million for the same configuration in the National Transonic Facility and in the European Transonic Facility. Force and moment, surface pressure, wing bending and twist, and surface flow visualization data were obtained in all three facilities but only the force and moment, surface pressure and wing bending and twist data are presented herein.

Digital-flutter-suppression-system investigations for the active flexible wing wind-tunnel model

NASA Technical Reports Server (NTRS)

Perry, Boyd, III; Mukhopadhyay, Vivek; Hoadley, Sherwood T.; Cole, Stanley R.; Buttrill, Carey S.; Houck, Jacob A.

1990-01-01

Active flutter suppression control laws were designed, implemented, and tested on an aeroelastically-scaled wind tunnel model in the NASA Langley Transonic Dynamics Tunnel. One of the control laws was successful in stabilizing the model while the dynamic pressure was increased to 24 percent greater than the measured open-loop flutter boundary. Other accomplishments included the design, implementation, and successful operation of a one-of-a-kind digital controller, the design and use of two simulation methods to support the project, and the development and successful use of a methodology for on-line controller performance evaluation.

Flutter Analysis of the Thermal Protection Layer on the NASA HIAD

NASA Technical Reports Server (NTRS)

Goldman, Benjamin D.; Dowell, Earl H.; Scott, Robert C.

2013-01-01

A combination of classical plate theory and a supersonic aerodynamic model is used to study the aeroelastic flutter behavior of a proposed thermal protection system (TPS) for the NASA HIAD. The analysis pertains to the rectangular configurations currently being tested in a NASA wind-tunnel facility, and may explain why oscillations of the articles could be observed. An analysis using a linear flat plate model indicated that flutter was possible well within the supersonic flow regime of the wind tunnel tests. A more complex nonlinear analysis of the TPS, taking into account any material curvature present due to the restraint system or substructure, indicated that significantly greater aerodynamic forcing is required for the onset of flutter. Chaotic and periodic limit cycle oscillations (LCOs) of the TPS are possible depending on how the curvature is imposed. When the pressure from the base substructure on the bottom of the TPS is used as the source of curvature, the flutter boundary increases rapidly and chaotic behavior is eliminated.

Selected topics in experimental aeroelasticity at the NASA Langley Research Center

NASA Technical Reports Server (NTRS)

Ricketts, R. H.

1985-01-01

The results of selected studies that have been conducted by the NASA Langley Research Center in the last three years are presented. The topics presented focus primarily on the ever-important transonic flight regime and include the following: body-freedom flutter of a forward-swept-wing configuration with and without relaxed static stability; instabilities associated with a new tilt-rotor vehicle; effects of winglets, supercritical airfoils, and spanwise curvature on wing flutter; wind-tunnel investigation of a flutter-like oscillation on a high-aspect-ratio flight research wing; results of wing-tunnel demonstration of the NASA decoupler pylon concept for passive suppression of wing/store flutter; and, new flutter testing methods which include testing at cryogenic temperatures for full scale Reynolds number simulation, subcritical response techniques for predicting onset of flutter, and a two-degree-of-freedom mount system for testing side-wall-mounted models.

Selected topics in experimental aeroelasticity at the NASA Langley Research Center

NASA Technical Reports Server (NTRS)

Ricketts, R. H.

1985-01-01

The results of selected studies that have been conducted by the NASA Langley Research Center in the last three years are presented. The topics presented focus primarily on the ever-important transonic flight regime and include the following: body-freedom flutter of a forward-swept-wing configuration with and without relaxed static stability; instabilities associated with a new tilt-rotor vehicle; effects of winglets, supercritical airfoils, and spanwise curvature on wing flutter; wind-tunnel investigation of a flutter-like oscillation on a high-aspect-ratio flight research wing; results of wind-tunnel demonstration of the NASA decoupler pylon concept for passive suppression of wing/store flutter; and, new flutter testing methods which include testing at cryogenic temperatures for full scale Reynolds number simulation, subcritical response techniques for predicting onset of flutter, and a two-degree-of-freedom mount system for testing side-wall-mounted models.

Experimental and analytical study on the flutter and gust response characteristics of a torsion-free-wing airplane model. [in the Langley transonic dynamics tunnel

NASA Technical Reports Server (NTRS)

Murphy, A. C.

1981-01-01

Experimental data and correlative analytical results on the flutter and gust response characteristics of a torsion-free-wing (TFW) fighter airplane model are presented. TFW consists of a combined wing/boom/canard surface and was tested with the TFW free to pivot in pitch and with the TFW locked to the fuselage. Flutter and gust response characteristics were measured in the Langley Transonic Dynamics Tunnel with the complete airplane model mounted on a cable mount system that provided a near free flying condition. Although the lowest flutter dynamic pressure was measured for the wing free configuration, it was only about 20 deg less than that for the wing locked configuration. However, no appreciable alleviation of the gust response was measured by freeing the wing.

Unsteady Transonic Flow Past Airfoils in Rigid Body Motion.

DTIC Science & Technology

1981-03-01

coordinate system. Numerical experiments show that the scheme is very stable and is able to resolve the highly non- linear transonic effects for flutter...Numerical experiments show that the scheme is very stable and is able to resolve the highly nonlinear transonic effects for flutter analysis within...of attack, the angle between the flight direction and the airfoil chord. The effect of chanqinthe angle of attack of a conventional symmetric airfoil

Developing, mechanizing and testing of a digital active flutter suppression system for a modified B-52 wind-tunnel model

NASA Technical Reports Server (NTRS)

Matthew, J. R.

1980-01-01

A digital flutter suppression system was developed and mechanized for a significantly modified version of the 1/30-scale B-52E aeroelastic wind tunnel model. A model configuration was identified that produced symmetric and antisymmetric flutter modes that occur at 2873N/sq m (60 psf) dynamic pressure with violent onset. The flutter suppression system, using one trailing edge control surface and the accelerometers on each wing, extended the flutter dynamic pressure of the model beyond the design limit of 4788N/sq m (100 psf). The hardware and software required to implement the flutter suppression system were designed and mechanized using digital computers in a fail-operate configuration. The model equipped with the system was tested in the Transonic Dynamics Tunnel at NASA Langley Research Center and results showed the flutter dynamic pressure of the model was extended beyond 4884N/sq m (102 psf).

Experimental unsteady pressures at flutter on the Supercritical Wing Benchmark Model

NASA Technical Reports Server (NTRS)

Dansberry, Bryan E.; Durham, Michael H.; Bennett, Robert M.; Rivera, Jose A.; Silva, Walter A.; Wieseman, Carol D.; Turnock, David L.

1993-01-01

This paper describes selected results from the flutter testing of the Supercritical Wing (SW) model. This model is a rigid semispan wing having a rectangular planform and a supercritical airfoil shape. The model was flutter tested in the Langley Transonic Dynamics Tunnel (TDT) as part of the Benchmark Models Program, a multi-year wind tunnel activity currently being conducted by the Structural Dynamics Division of NASA Langley Research Center. The primary objective of this program is to assist in the development and evaluation of aeroelastic computational fluid dynamics codes. The SW is the second of a series of three similar models which are designed to be flutter tested in the TDT on a flexible mount known as the Pitch and Plunge Apparatus. Data sets acquired with these models, including simultaneous unsteady surface pressures and model response data, are meant to be used for correlation with analytical codes. Presented in this report are experimental flutter boundaries and corresponding steady and unsteady pressure distribution data acquired over two model chords located at the 60 and 95 percent span stations.

Optical measurement of unducted fan flutter

NASA Technical Reports Server (NTRS)

Kurkov, Anatole P.; Mehmed, Oral

1990-01-01

A nonintrusive optical method is described for flutter vibrations in unducted fan or propeller rotors and provides detailed spectral results for two flutter modes of a scaled unducted fan. The measurements were obtained in a high-speed wind tunnel. A single-rotor and a dual-rotor counterrotating configuration of the model were tested; however, only the forward rotor of the counterrotating configuration fluttered. Conventional strain gages were used to obtain flutter frequency; optical data provided complete phase results and an indication of the flutter mode shape through the ratio of the leading- to trailing-edge flutter amplitudes near the blade tip. In the transonic regime exhibited some features that are usually associated with nonlinear vibrations. Experimental mode shape and frequencies were compared with calculated values that included centrifugal effects.

An Overview of Unsteady Pressure Measurements in the Transonic Dynamics Tunnel

NASA Technical Reports Server (NTRS)

Schuster, David M.; Edwards, John W.; Bennett, Robert M.

2000-01-01

The NASA Langley Transonic Dynamics Tunnel has served as a unique national facility for aeroelastic testing for over forty years. A significant portion of this testing has been to measure unsteady pressures on models undergoing flutter, forced oscillations, or buffet. These tests have ranged from early launch vehicle buffet to flutter of a generic high-speed transport. This paper will highlight some of the test techniques, model design approaches, and the many unsteady pressure tests conducted in the TDT. The objectives and results of the data acquired during these tests will be summarized for each case and a brief discussion of ongoing research involving unsteady pressure measurements and new TDT capabilities will be presented.

A harmonic analysis method for unsteady transonic flow and its application to the flutter of airfoils

NASA Technical Reports Server (NTRS)

Ehlers, F. E.; Weatherill, W. H.

1982-01-01

A finite difference method for solving the unsteady transonic flow about harmonically oscillating wings is investigated. The procedure is based on separating the velocity potential into steady and unsteady parts and linearizing the resulting unsteady differential equation for small disturbances. The differential equation for the unsteady velocity potential is linear with spatially varying coefficients and with the time variable eliminated by assuming harmonic motion. A study is presented of the shock motion associated with an oscillating airfoil and its representation by the harmonic procedure. The effects of the shock motion and the resulting pressure pulse are shown to be included in the harmonic pressure distributions and the corresponding generalized forces. Analytical and experimental pressure distributions for the NACA 64A010 airfoil are compared for Mach numbers of 0.75, 0.80 and 0.842. A typical section, two-degree-of-freedom flutter analysis of a NACA 64A010 airfoil is performed. The results show a sharp transonic bucket in one case and abrupt changes in instability modes.

An experimental and analytical investigation of the effect of spanwise curvature on wing flutter at Mach number of 0.7

NASA Technical Reports Server (NTRS)

Rivera, Jose A., Jr.

1989-01-01

An experimental and analytical study was conducted at Mach 0.7 to investigate the effects of spanwise curvature on flutter. Two series of rectangular planform wings of aspect ration 1.5 and curvature ranging from zero (uncurved) to 1.04/ft were flutter tested in the NASA Langley Transonic Dynamics Tunnel (TDT). One series consisted of models with a NACA 65 A010 airfoil section and the other of flat plate cross section models. Flutter analyses were conducted for correlation with the experimental results by using structural finite element methods to perform vibration analysis and two aerodynamic theories to obtain unsteady aerodynamic load calculations. The experimental results showed that for one series of models the flutter dynamic pressure increased significantly with curvature while for the other series of models the flutter dynamic pressure decreased with curvature. The flutter analyses, which generally predicted the experimental results, indicated that the difference in behavior of the two series of models was primarily due to differences in their structural properties.

Some Effects of Sweep and Aspect Ratio on the Transonic Flutter Characteristics of a Series of Thin Cantilever Wings Having a Taper Ratio of 0.6

NASA Technical Reports Server (NTRS)

Jones, G. W., Jr.; Unangst, J. R.

1963-01-01

An investigation of the flutter characteristics of a series of thin cantilever wings having taper ratios of 0.6 was conducted in the Langley transonic blowdown tunnel at Mach numbers between 0.76 and 1.42. The angle of sweepback was varied from 0 degrees to 60 degrees on wings of aspect ratio 4, and the aspect ratio was varied from 2.4 to 6.4 on wings with 45 degrees of sweepback. The results are presented as ratios between the experimental flutter speeds and the reference flutter speeds calculated on the basis of incompressible two-dimensional flow. These ratios, designated the flutter-speed ratios, are given as functions of Mach number for the various wings. The flutter-speed ratios were characterized, in most cases, by values near 1.0 at subsonic speeds with large increases in the speed ratios in the range of supersonic speeds investigated. Increasing the sweep effected increases in the flutter-speed ratios between 0 degrees and 30 degrees followed by progressive reductions of the speed ratios to nearly 1.0 as the sweep was increased from 30 degrees to 60 degrees. Reducing the aspect ratio from 6.4 to 2.4 resulted in progressively larger values of the flutter-speed ratios throughout the Mach number range investigated.

Buffet test in the National Transonic Facility

NASA Technical Reports Server (NTRS)

Young, Clarence P., Jr.; Hergert, Dennis W.; Butler, Thomas W.; Herring, Fred M.

1992-01-01

A buffet test of a commercial transport model was accomplished in the National Transonic Facility at the NASA Langley Research Center. This aeroelastic test was unprecedented for this wind tunnel and posed a high risk for the facility. Presented here are the test results from a structural dynamics and aeroelastic response point of view. The activities required for the safety analysis and risk assessment are described. The test was conducted in the same manner as a flutter test and employed on-board dynamic instrumentation, real time dynamic data monitoring, and automatic and manual tunnel interlock systems for protecting the model.

Application of Reduced Order Transonic Aerodynamic Influence Coefficient Matrix for Design Optimization

NASA Technical Reports Server (NTRS)

Pak, Chan-gi; Li, Wesley W.

2009-01-01

Supporting the Aeronautics Research Mission Directorate guidelines, the National Aeronautics and Space Administration [NASA] Dryden Flight Research Center is developing a multidisciplinary design, analysis, and optimization [MDAO] tool. This tool will leverage existing tools and practices, and allow the easy integration and adoption of new state-of-the-art software. Today s modern aircraft designs in transonic speed are a challenging task due to the computation time required for the unsteady aeroelastic analysis using a Computational Fluid Dynamics [CFD] code. Design approaches in this speed regime are mainly based on the manual trial and error. Because of the time required for unsteady CFD computations in time-domain, this will considerably slow down the whole design process. These analyses are usually performed repeatedly to optimize the final design. As a result, there is considerable motivation to be able to perform aeroelastic calculations more quickly and inexpensively. This paper will describe the development of unsteady transonic aeroelastic design methodology for design optimization using reduced modeling method and unsteady aerodynamic approximation. The method requires the unsteady transonic aerodynamics be represented in the frequency or Laplace domain. Dynamically linear assumption is used for creating Aerodynamic Influence Coefficient [AIC] matrices in transonic speed regime. Unsteady CFD computations are needed for the important columns of an AIC matrix which corresponded to the primary modes for the flutter. Order reduction techniques, such as Guyan reduction and improved reduction system, are used to reduce the size of problem transonic flutter can be found by the classic methods, such as Rational function approximation, p-k, p, root-locus etc. Such a methodology could be incorporated into MDAO tool for design optimization at a reasonable computational cost. The proposed technique is verified using the Aerostructures Test Wing 2 actually designed

Unsteady transonic flow calculations for two-dimensional canard-wing configurations with aeroelastic applications

NASA Technical Reports Server (NTRS)

Batina, J. T.

1985-01-01

Unsteady transonic flow calculations for aerodynamically interfering airfoil configurations are performed as a first step toward solving the three dimensional canard wing interaction problem. These calculations are performed by extending the XTRAN2L two dimensional unsteady transonic small disturbance code to include an additional airfoil. Unsteady transonic forces due to plunge and pitch motions of a two dimensional canard and wing are presented. Results for a variety of canard wing separation distances reveal the effects of aerodynamic interference on unsteady transonic airloads. Aeroelastic analyses employing these unsteady airloads demonstrate the effects of aerodynamic interference on aeroelastic stability and flutter. For the configurations studied, increases in wing flutter speed result with the inclusion of the aerodynamically interfering canard.

Overview of Recent Flight Flutter Testing Research at NASA Dryden

NASA Technical Reports Server (NTRS)

Brenner, Martin J.; Lind, Richard C.; Voracek, David F.

1997-01-01

In response to the concerns of the aeroelastic community, NASA Dryden Flight Research Center, Edwards, California, is conducting research into improving the flight flutter (including aeroservoelasticity) test process with more accurate and automated techniques for stability boundary prediction. The important elements of this effort so far include the following: (1) excitation mechanisms for enhanced vibration data to reduce uncertainty levels in stability estimates; (2) investigation of a variety of frequency, time, and wavelet analysis techniques for signal processing, stability estimation, and nonlinear identification; and (3) robust flutter boundary prediction to substantially reduce the test matrix for flutter clearance. These are critical research topics addressing the concerns of a recent AGARD Specialists' Meeting on Advanced Aeroservoelastic Testing and Data Analysis. This paper addresses these items using flight test data from the F/A-18 Systems Research Aircraft and the F/A-18 High Alpha Research Vehicle.

Subsonic Transonic Applied Refinements By Using Key Strategies - STARBUKS In the NASA Langley Research Center National Transonic Facility

NASA Technical Reports Server (NTRS)

Paryz, Roman W.

2014-01-01

Several upgrade projects have been completed at the NASA Langley Research Center National Transonic Facility over the last 1.5 years in an effort defined as STARBUKS - Subsonic Transonic Applied Refinements By Using Key Strategies. This multi-year effort was undertaken to improve NTF's overall capabilities by addressing Accuracy and Validation, Productivity, and Reliability areas at the NTF. This presentation will give a brief synopsis of each of these efforts.

A wind-tunnel investigation of a B-52 model flutter suppression system

NASA Technical Reports Server (NTRS)

Redd, L. T.; Gilman, J., Jr.; Cooley, D. E.; Sevart, F. D.

1974-01-01

Flutter modeling techniques have been successfully extended to the difficult case of the active suppression of flutter. The demonstration was conducted in a transonic dynamics tunnel using a 1/30 scale, elastic, dynamic model of a Boeing B-52 control configured vehicle. The results from the study show that with the flutter suppression system operating there is a substantial increase in the damping associated with the critical flutter mode. The results also show good correlation between the damping characteristics of the model and the aircraft.

Winglet effects on the flutter of twin-engine-transport type wing

NASA Technical Reports Server (NTRS)

Bhatia, K. G.; Nagaraja, K. S.; Ruhlin, C. L.

1984-01-01

Flutter characteristics of a cantilevered high aspect ratio wing with winglet were investigated. The configuration represented a current technology, twin-engine airplane. A low-speed and a high-speed model were used to evaluate compressibility effects through transonic Mach numbers and a wide range of mass-density ratios. Four flutter mechanisms were obtained in test, as well as analysis from various combinations of configuration parameters. The coupling between wing tip vertical and chordwise motions was shown to have significant effect under some conditions. It is concluded that, for the flutter model configurations studied, the winglet related flutter was amenable to the conventional flutter analysis techniques.

Test Cases for Flutter of the Benchmark Models Rectangular Wings on the Pitch and Plunge Apparatus

NASA Technical Reports Server (NTRS)

Bennett, Robert M.

2000-01-01

The supercritical airfoil was chosen as a relatively modem airfoil for comparison. The BOO12 model was tested first. Three different types of flutter instability boundaries were encountered, a classical flutter boundary, a transonic stall flutter boundary at angle of attack, and a plunge instability near M = 0.9 and for zero angle of attack. This test was made in air and was Transonic Dynamics Tunnel (TDT) Test 468. The BSCW model (for Benchmark SuperCritical Wing) was tested next as TDT Test 470. It was tested using both with air and a heavy gas, R-12, as a test medium. The effect of a transition strip on flutter was evaluated in air. The B64AOlO model was subsequently tested as TDT Test 493. Some further analysis of the experimental data for the BOO12 wing is presented. Transonic calculations using the parameters for the BOO12 wing in a two-dimensional typical section flutter analysis are given. These data are supplemented with data from the Benchmark Active Controls Technology model (BACT) given and in the next chapter of this document. The BACT model was of the same planform and airfoil as the BOO12 model, but with spoilers and a trailing edge control. It was tested in the heavy gas R-12, and was instrumented mostly at the 60 per cent span. The flutter data obtained on PAPA and the static aerodynamic test cases from BACT serve as additional data for the BOO12 model. All three types of flutter are included in the BACT Test Cases. In this report several test cases are selected to illustrate trends for a variety of different conditions with emphasis on transonic flutter. Cases are selected for classical and stall flutter for the BSCW model, for classical and plunge for the B64AOlO model, and for classical flutter for the BOO12 model. Test Cases are also presented for BSCW for static angles of attack. Only the mean pressures and the real and imaginary parts of the first harmonic of the pressures are included in the data for the test cases, but digitized time

Pressure measurements on a rectangular wing with a NACA0012 airfoil during conventional flutter

NASA Technical Reports Server (NTRS)

Rivera, Jose A., Jr.; Dansberry, Bryan E.; Durham, Michael H.; Bennett, Robert M.; Silva, Walter A.

1992-01-01

The Structural Dynamics Division at NASA LaRC has started a wind tunnel activity referred to as the Benchmark Models Program. The primary objective of the program is to acquire measured dynamic instability and corresponding pressure data that will be useful for developing and evaluating aeroelastic type CFD codes currently in use or under development. The program is a multi-year activity that will involve testing of several different models to investigate various aeroelastic phenomena. The first model consisted of a rigid semispan wing having a rectangular planform and a NACA 0012 airfoil shape which was mounted on a flexible two degree-of-freedom mount system. Two wind-tunnel tests were conducted with the first model. Several dynamic instability boundaries were investigated such as a conventional flutter boundary, a transonic plunge instability region near Mach = 0.90, and stall flutter. In addition, wing surface unsteady pressure data were acquired along two model chords located at the 60 to 95-percent span stations during these instabilities. At this time, only the pressure data for the conventional flutter boundary is presented. The conventional flutter boundary and the wing surface unsteady pressure measurements obtained at the conventional flutter boundary test conditions in pressure coefficient form are presented. Wing surface steady pressure measurements obtained with the model mount system rigidized are also presented. These steady pressure data were acquired at essentially the same dynamic pressure at which conventional flutter had been encountered with the mount system flexible.

Computed and Experimental Flutter/LCO Onset for the Boeing Truss-Braced Wing Wind-Tunnel Model

NASA Technical Reports Server (NTRS)

Bartels, Robert E.; Scott, Robert C.; Funk, Christie J.; Allen, Timothy J.; Sexton, Bradley W.

2014-01-01

This paper presents high fidelity Navier-Stokes simulations of the Boeing Subsonic Ultra Green Aircraft Research truss-braced wing wind-tunnel model and compares the results to linear MSC. Nastran flutter analysis and preliminary data from a recent wind-tunnel test of that model at the NASA Langley Research Center Transonic Dynamics Tunnel. The simulated conditions under consideration are zero angle of attack, so that structural nonlinearity can be neglected. It is found that, for Mach number greater than 0.78, the linear flutter analysis predicts flutter onset dynamic pressure below the wind-tunnel test and that predicted by the Navier-Stokes analysis. Furthermore, the wind-tunnel test revealed that the majority of the high structural dynamics cases were wing limit cycle oscillation (LCO) rather than flutter. Most Navier-Stokes simulated cases were also LCO rather than hard flutter. There is dip in the wind-tunnel test flutter/LCO onset in the Mach 0.76-0.80 range. Conditions tested above that Mach number exhibited no aeroelastic instability at the dynamic pressures reached in the tunnel. The linear flutter analyses do not show a flutter/LCO dip. The Navier-Stokes simulations also do not reveal a dip; however, the flutter/LCO onset is at a significantly higher dynamic pressure at Mach 0.90 than at lower Mach numbers. The Navier-Stokes simulations indicate a mild LCO onset at Mach 0.82, then a more rapidly growing instability at Mach 0.86 and 0.90. Finally, the modeling issues and their solution related to the use of a beam and pod finite element model to generate the Navier-Stokes structure mode shapes are discussed.

An overview of selected NASP aeroelastic studies at the NASA Langley Research Center

NASA Technical Reports Server (NTRS)

Spain, Charles V.; Soistmann, David L.; Parker, Ellen C.; Gibbons, Michael D.; Gilbert, Michael G.

1990-01-01

Following an initial discussion of the NASP flight environment, the results of recent aeroelastic testing of NASP-type highly swept delta-wing models in Langley's Transonic Dynamics Tunnel (TDT) are summarized. Subsonic and transonic flutter characteristics of a variety of these models are described, and several analytical codes used to predict flutter of these models are evaluated. These codes generally provide good, but conservative predictions of subsonic and transonic flutter. Also, test results are presented on a nonlinear transonic phenomena known as aileron buzz which occurred in the wind tunnel on highly swept delta wings with full-span ailerons. An analytical procedure which assesses the effects of hypersonic heating on aeroelastic instabilities (aerothermoelasticity) is also described. This procedure accurately predicted flutter of a heated aluminum wing on which experimental data exists. Results are presented on the application of this method to calculate the flutter characteristics of a fine-element model of a generic NASP configuration. Finally, it is demonstrated analytically that active controls can be employed to improve the aeroelastic stability and ride quality of a generic NASP vehicle flying at hypersonic speeds.

Aeroelasticity matters: Some reflections on two decades of testing in the NASA Langley transonic dynamics tunnel

NASA Technical Reports Server (NTRS)

Reed, W. H., III

1981-01-01

Testing of wind-tunnel aeroelastic models is a well established, widely used means of studying flutter trends, validating theory and investigating flutter margins of safety of new vehicle designs. The Langley Transonic Dynamics Tunnel was designed specifically for work on dynamics and aeroelastic problems of aircraft and space vehicles. A cross section of aeroelastic research and testing in the facility since it became operational more than two decades ago is presented. Examples selected from a large store of experience illustrate the nature and purpose of some major areas of work performed in the tunnel. These areas include: specialized experimental techniques; development testing of new aircraft and launch vehicle designs; evaluation of proposed "fixes" to solve aeroelastic problems uncovered during development testing; study of unexpected aeroelastic phenomena (i.e., "surprises"); control of aeroelastic effects by active and passive means; and, finally, fundamental research involving measurement of unsteady pressures on oscillating wings and control surface.

Further investigations of the aeroelastic behavior of the AFW wind-tunnel model using transonic small disturbance theory

NASA Technical Reports Server (NTRS)

Silva, Walter A.; Bennett, Robert M.

1992-01-01

The Computational Aeroelasticity Program-Transonic Small Disturbance (CAP-TSD) code, developed at LaRC, is applied to the active flexible wing wind-tunnel model for prediction of transonic aeroelastic behavior. A semi-span computational model is used for evaluation of symmetric motions, and a full-span model is used for evaluation of antisymmetric motions, and a full-span model is used for evaluation of antisymmetric motions. Static aeroelastic solutions using CAP-TSD are computed. Dynamic deformations are presented as flutter boundaries in terms of Mach number and dynamic pressure. Flutter boundaries that take into account modal refinements, vorticity and entropy corrections, antisymmetric motion, and sensitivity to the modeling of the wing tip ballast stores are also presented with experimental flutter results.

A Wind-Tunnel Parametric Investigation of Tiltrotor Whirl-Flutter Stability Boundaries

NASA Technical Reports Server (NTRS)

Piatak, David J.; Kvaternik, Raymond G.; Nixon, Mark W.; Langston, Chester W.; Singleton, Jeffrey D.; Bennett, Richard L.; Brown, Ross K.

2001-01-01

A wind-tunnel investigation of tiltrotor whirl-flutter stability boundaries has been conducted on a 1/5-size semispan tiltrotor model known as the Wing and Rotor Aeroelastic Test System (WRATS) in the NASA-Langley Transonic Dynamics Tunnel as part of a joint NASA/Army/Bell Helicopter Textron, Inc. (BHTI) research program. The model was first developed by BHTI as part of the JVX (V-22) research and development program in the 1980's and was recently modified to incorporate a hydraulically-actuated swashplate control system for use in active controls research. The modifications have changed the model's pylon mass properties sufficiently to warrant testing to re-establish its baseline stability boundaries. A parametric investigation of the effect of rotor design variables on stability was also conducted. The model was tested in both the on-downstop and off-downstop configurations, at cruise flight and hover rotor rotational speeds, and in both air and heavy gas (R-134a) test mediums. Heavy gas testing was conducted to quantify Mach number compressibility effects on tiltrotor stability. Experimental baseline stability boundaries in air are presented with comparisons to results from parametric variations of rotor pitch-flap coupling and control system stiffness. Increasing the rotor pitch-flap coupling (delta(sub 3) more negative) was found to have a destabilizing effect on stability, while a reduction in control system stiffness was found to have little effect on whirl-flutter stability. Results indicate that testing in R-134a, and thus matching full-scale tip Mach number, has a destabilizing effect, which demonstrates that whirl-flutter stability boundaries in air are unconservative.

Flutter Boundary Identification From Simulation Time Histories

NASA Technical Reports Server (NTRS)

Baker, Myles; Goggin, P. J.

1997-01-01

While there has been much recent progress in simulating nonlinear aeroelastic systems, and in predicting many of the aeroelastic phenomena of concern in transport aircraft design (i.e. transonic flutter buckets), the utility of a simulation in generating an understanding of the flutter behavior is limited. This is due in part to the high cost of generating these simulations; and the implied limitation on the number of conditions that can be analyzed, but there are also some difficulties introduced by the very nature of a simulation. Flutter engineers have traditionally worked in the frequency domain, and are accustomed to describing the flutter behavior of an airplane in terms of its V-G and V-F (or Q-G and Q-F) plots and flutter mode shapes. While the V-G and V-F plots give information about how the dynamic response of an airplane changes as the airspeed is increased, the simulation only gives information about one isolated condition (Mach, airspeed, altitude, etc.). Therefore, where a traditional flutter analysis can let the engineer determine an airspeed at which an airplane becomes unstable, while a simulation only serves as a binary check: either the airplane is fluttering at this condition, or it is not. In this document, a new technique is described in which system identification is used to easily extract modal frequencies and damping ratios from simulation time histories, and shows how the identified parameters can be used to determine the variation in frequency and dampin,o ratio as the airspeed is changed. This technique not only provides the flutter engineer with added insight into the aeroelastic behavior of the airplane, but it allows calculation of flutter mode shapes, and allows estimation of flutter boundaries while minimizing the number of simulations required.

Computations for the 16-foot transonic tunnel, NASA, Langley Research Center, revision 1

NASA Technical Reports Server (NTRS)

Mercer, Charles E.; Berrier, Bobby L.; Capone, Francis J.; Grayston, Alan M.; Sherman, C. D.

1987-01-01

The equations used by the 16 foot transonic tunnel in the data reduction programs are presented in eight modules. Each module consists of equations necessary to achieve a specific purpose. These modules are categorized in the following groups: tunnel parameters; jet exhaust measurements; skin friction drag; balance loads and model attitudes calculations; internal drag (or exit-flow distributions); pressure coefficients and integrated forces; thrust removal options; and turboprop options. This document is a companion document to NASA TM-83186, A User's Guide to the Langley 16 Foot Transonic Tunnel, August 1981.

Planform curvature effects on flutter characteristics of a wing with 56 deg leading-edge sweep and panel aspect ratio of 1.14

NASA Technical Reports Server (NTRS)

Keller, Donald F.; Sandford, Maynard C.; Pinkerton, Theresa L.

1991-01-01

An experimental and analytical investigation was initiated to determine the effects of planform curvature (curving the leading and trailing edges of a wing in the X-Y plane) on the transonic flutter characteristics of a series of three moderately swept wing models. Experimental flutter results were obtained in the Langley Transonic Dynamics Tunnel for Mach numbers from 0.60-1.00, with air as the test medium. The models were semispan cantilevered wings with a 3 percent biconvex airfoil and a panel aspect ratio of 1.14. The baseline model had straight leading and trailing edges (i.e., no planform curvature). The radii of curvature of the leading edges for these two models were 200 and 80 inches. The radii of curvature of the leading edges of the other two models were determined so that the root and tip chords were identical for all three models. Experimental results showed that flutter-speed index and flutter frequency ratio increased as planform curvature increase (radius of curvature of the leading edge was decreased) over the test range of Mach numbers. Analytical flutter results were calculated with a subsonic flutter-prediction program, and they agreed well with the experimental results.

Flutter parametric studies of cantilevered twin-engine-transport type wing with and without winglet. Volume 1: Low-speed investigations

NASA Technical Reports Server (NTRS)

Bhatia, K. G.; Nagaraja, K. S.

1984-01-01

Flutter characteristics of a cantilevered high aspect ratio wing with winglet were investigated. The configuration represented a current technology, twin-engine airplane. A low-speed and high-speed model were used to evaluate compressibility effects through transonic Mach numbers and a wide range of mass-density ratios. Four flutter mechanisms were obtained in test, as well as analysis from various combinations of configuration parameters. The coupling between wing tip vertical and chordwise motions was shown to have significant effect under some conditions. It is concluded that for the flutter model configurations studied, the winglet related flutter was amenable to the conventional flutter analysis techniques.

Real-time flutter analysis of an active flutter-suppression system on a remotely piloted research aircraft

NASA Technical Reports Server (NTRS)

Gilyard, G. B.; Edwards, J. W.

1983-01-01

Flight flutter-test results of the first aeroelastic research wing (ARW-1) of NASA's drones for aerodynamic and structural testing program are presented. The flight-test operation and the implementation of the active flutter-suppression system are described as well as the software techniques used to obtain real-time damping estimates and the actual flutter testing procedure. Real-time analysis of fast-frequency aileron excitation sweeps provided reliable damping estimates. The open-loop flutter boundary was well defined at two altitudes; a maximum Mach number of 0.91 was obtained. Both open-loop and closed-loop data were of exceptionally high quality. Although the flutter-suppression system provided augmented damping at speeds below the flutter boundary, an error in the implementation of the system resulted in the system being less stable than predicted. The vehicle encountered system-on flutter shortly after crossing the open-loop flutter boundary on the third flight and was lost. The aircraft was rebuilt. Changes made in real-time test techniques are included.

Evaluation of flow quality in two large NASA wind tunnels at transonic speeds

NASA Technical Reports Server (NTRS)

Harvey, W. D.; Stainback, P. C.; Owen, F. K.

1980-01-01

Wind tunnel testing of low drag airfoils and basic transition studies at transonic speeds are designed to provide high quality aerodynamic data at high Reynolds numbers. This requires that the flow quality in facilities used for such research be excellent. To obtain a better understanding of the characteristics of facility disturbances and identification of their sources for possible facility modification, detailed flow quality measurements were made in two prospective NASA wind tunnels. Experimental results are presented of an extensive and systematic flow quality study of the settling chamber, test section, and diffuser in the Langley 8 foot transonic pressure tunnel and the Ames 12 foot pressure wind tunnel. Results indicate that the free stream velocity and pressure fluctuation levels in both facilities are low at subsonic speeds and are so high as to make it difficult to conduct meaningful boundary layer control and transition studies at transonic speeds.

Experimental Investigations of the NASA Common Research Model in the NASA Langley National Transonic Facility and NASA Ames 11-Ft Transonic Wind Tunnel (Invited)

NASA Technical Reports Server (NTRS)

Rivers, S. M.; Dittberner, Ashley

2011-01-01

Experimental aerodynamic investigations of the NASA Common Research Model have been conducted in the NASA Langley National Transonic Facility and the NASA Ames 11-ft wind tunnel. Data have been obtained at chord Reynolds numbers of 5 million for five different configurations at both wind tunnels. Force and moment, surface pressure and surface flow visualization data were obtained in both facilities but only the force and moment data are presented herein. Nacelle/pylon, tail effects and tunnel to tunnel variations have been assessed. The data from both wind tunnels show that an addition of a nacelle/pylon gave an increase in drag, decrease in lift and a less nose down pitching moment around the design lift condition of 0.5 and that the tail effects also follow the expected trends. Also, all of the data shown fall within the 2-sigma limits for repeatability. The tunnel to tunnel differences are negligible for lift and pitching moment, while the drag shows a difference of less than ten counts for all of the configurations. These differences in drag may be due to the variation in the sting mounting systems at the two tunnels.

Status of the KTH-NASA Wind-Tunnel Test for Acquisition of Transonic Nonlinear Aeroelastic Data

NASA Technical Reports Server (NTRS)

Silva, Walter A.; Ringertz, Ulf; Stenfelt, Gloria; Eller, David; Keller, Donald F.; Chwalowski, Pawel

2016-01-01

This paper presents a status report on the collaboration between the Royal Institute of Technology (KTH) in Sweden and the NASA Langley Research Center regarding the design, fabrication, modeling, and testing of a full-span lighter configuration in the Transonic Dynamics Tunnel (TDT). The goal of the test is to acquire transonic limit-cycle- oscillation (LCO) data, including accelerations, strains, and unsteady pressures. Finite element models (FEMs) and aerodynamic models are presented and discussed along with results obtained to date.

Navier-Stokes, dynamics and aeroelastic computations for vortical flows, buffet and flutter applications

NASA Technical Reports Server (NTRS)

Kandil, Osama A.

1993-01-01

Research on Navier-Stokes, dynamics, and aeroelastic computations for vortical flows, buffet, and flutter applications was performed. Progress during the period from 1 Oct. 1992 to 30 Sep. 1993 is included. Papers on the following topics are included: vertical tail buffet in vortex breakdown flows; simulation of tail buffet using delta wing-vertical tail configuration; shock-vortex interaction over a 65-degree delta wing in transonic flow; supersonic vortex breakdown over a delta wing in transonic flow; and prediction and control of slender wing rock.

Recent Developments at the NASA Langley Research Center National Transonic Facility

NASA Technical Reports Server (NTRS)

Paryz, Roman W.

2011-01-01

Several upgrade projects have been completed or are just getting started at the NASA Langley Research Center National Transonic Facility. These projects include a new high capacity semi-span balance, model dynamics damping system, semi-span model check load stand, data acquisition system upgrade, facility automation system upgrade and a facility reliability assessment. This presentation will give a brief synopsis of each of these efforts.

Results of Two Free-fall Experiments on Flutter of Thin Unswept Wings in the Transonic Speed Range

NASA Technical Reports Server (NTRS)

Lauten, William T , Jr; Nelson, Herbert C

1957-01-01

Results of four thin, unswept, flutter airfoils attached to two freely falling bodies are reported. Two airfoils fluttered at a Mach number of 0.85, a third airfoil fluttered at a Mach number of 1.03, and a fourth fluttered at a Mach number of 1.07. Results of calculations of flutter speed using incompressible and compressible air-force coefficients, including a Mach number of 1.0, are presented.

Test Cases for the Benchmark Active Controls: Spoiler and Control Surface Oscillations and Flutter

NASA Technical Reports Server (NTRS)

Bennett, Robert M.; Scott, Robert C.; Wieseman, Carol D.

2000-01-01

As a portion of the Benchmark Models Program at NASA Langley, a simple generic model was developed for active controls research and was called BACT for Benchmark Active Controls Technology model. This model was based on the previously-tested Benchmark Models rectangular wing with the NACA 0012 airfoil section that was mounted on the Pitch and Plunge Apparatus (PAPA) for flutter testing. The BACT model had an upper surface spoiler, a lower surface spoiler, and a trailing edge control surface for use in flutter suppression and dynamic response excitation. Previous experience with flutter suppression indicated a need for measured control surface aerodynamics for accurate control law design. Three different types of flutter instability boundaries had also been determined for the NACA 0012/PAPA model, a classical flutter boundary, a transonic stall flutter boundary at angle of attack, and a plunge instability near M = 0.9. Therefore an extensive set of steady and control surface oscillation data was generated spanning the range of the three types of instabilities. This information was subsequently used to design control laws to suppress each flutter instability. There have been three tests of the BACT model. The objective of the first test, TDT Test 485, was to generate a data set of steady and unsteady control surface effectiveness data, and to determine the open loop dynamic characteristics of the control systems including the actuators. Unsteady pressures, loads, and transfer functions were measured. The other two tests, TDT Test 502 and TDT Test 5 18, were primarily oriented towards active controls research, but some data supplementary to the first test were obtained. Dynamic response of the flexible system to control surface excitation and open loop flutter characteristics were determined during Test 502. Loads were not measured during the last two tests. During these tests, a database of over 3000 data sets was obtained. A reasonably extensive subset of the data

Application of Approximate Unsteady Aerodynamics for Flutter Analysis

NASA Technical Reports Server (NTRS)

Pak, Chan-gi; Li, Wesley W.

2010-01-01

A technique for approximating the modal aerodynamic influence coefficient (AIC) matrices by using basis functions has been developed. A process for using the resulting approximated modal AIC matrix in aeroelastic analysis has also been developed. The method requires the unsteady aerodynamics in frequency domain, and this methodology can be applied to the unsteady subsonic, transonic, and supersonic aerodynamics. The flutter solution can be found by the classic methods, such as rational function approximation, k, p-k, p, root locus et cetera. The unsteady aeroelastic analysis using unsteady subsonic aerodynamic approximation is demonstrated herein. The technique presented is shown to offer consistent flutter speed prediction on an aerostructures test wing (ATW) 2 and a hybrid wing body (HWB) type of vehicle configuration with negligible loss in precision. This method computes AICs that are functions of the changing parameters being studied and are generated within minutes of CPU time instead of hours. These results may have practical application in parametric flutter analyses as well as more efficient multidisciplinary design and optimization studies.

Development and application of a program to calculate transonic flow around an oscillating three-dimensional wing using finite difference procedures

NASA Technical Reports Server (NTRS)

Weatherill, Warren H.; Ehlers, F. Edward

1989-01-01

A finite difference method for solving the unsteady transonic flow about harmonically oscillating wings is investigated. The procedure is based on separating the velocity potential into steady and unsteady parts and linearizing the resulting unsteady differential equation for small disturbances. The differential equation for the unsteady potential is linear with spatially varying coefficients and with the time variable eliminated by assuming harmonic motion. Difference equations are derived for harmonic transonic flow to include a coordinate transformation for swept and tapered planforms. A pilot program is developed for three-dimensional planar lifting surface configurations (including thickness) for the CRAY-XMP at Boeing Commercial Airplanes and for the CYBER VPS-32 at the NASA Langley Research Center. An investigation is made of the effect of the location of the outer boundaries on accuracy for very small reduced frequencies. Finally, the pilot program is applied to the flutter analysis of a rectangular wing.

Recent Transonic Flutter Investigations for Wings and External Stores

DTIC Science & Technology

1983-01-01

and difficult method? In the early days of high-speed air- craft design . the aeroelastician realized that non -compressible aerodynamic theory and... experimental aeroelastic model program that would provide insight into the effects of Reynolds number and angle of attack on various airfoil designs regarding...investigation is carried out both experimentally and analytically. The analytic modelling will be described in a later section. The flutter calculations

Unique Testing Capabilities of the NASA Langley Transonic Dynamics Tunnel, an Exercise in Aeroelastic Scaling

NASA Technical Reports Server (NTRS)

Ivanco, Thomas G.

2013-01-01

NASA Langley Research Center's Transonic Dynamics Tunnel (TDT) is the world's most capable aeroelastic test facility. Its large size, transonic speed range, variable pressure capability, and use of either air or R-134a heavy gas as a test medium enable unparalleled manipulation of flow-dependent scaling quantities. Matching these scaling quantities enables dynamic similitude of a full-scale vehicle with a sub-scale model, a requirement for proper characterization of any dynamic phenomenon, and many static elastic phenomena. Select scaling parameters are presented in order to quantify the scaling advantages of TDT and the consequence of testing in other facilities. In addition to dynamic testing, the TDT is uniquely well-suited for high risk testing or for those tests that require unusual model mount or support systems. Examples of recently conducted dynamic tests requiring unusual model support are presented. In addition to its unique dynamic test capabilities, the TDT is also evaluated in its capability to conduct aerodynamic performance tests as a result of its flow quality. Results of flow quality studies and a comparison to a many other transonic facilities are presented. Finally, the ability of the TDT to support future NASA research thrusts and likely vehicle designs is discussed.

Passive Wireless Vibration Sensing for Measuring Aerospace Structural Flutter

NASA Technical Reports Server (NTRS)

Wilson, William C.; Moore, Jason P.

2017-01-01

To reduce energy consumption, emissions, and noise, NASA is exploring the use of high aspect ratio wings on subsonic aircraft. Because high aspect ratio wings are susceptible to flutter events, NASA is also investigating methods of flutter detection and suppression. In support of that work a new remote, non-contact method for measuring flutter-induced vibrations has been developed. The new sensing scheme utilizes a microwave reflectometer to monitor the reflected response from an aeroelastic structure to ultimately characterize structural vibrations. To demonstrate the ability of microwaves to detect flutter vibrations, a carbon fiber-reinforced polymer (CFRP) composite panel was vibrated at various frequencies from 1Hz to 130Hz. The reflectometer response was found to closely resemble the sinusoidal response as measured with an accelerometer up to 100 Hz. The data presented demonstrate that microwaves can be used to measure flutter-induced aircraft vibrations.

Dynamic Deformation Measurements of an Aeroelastic Semispan Model. [conducted in the Transonic Dynamics Tunnel at the NASA Langley Research Center

NASA Technical Reports Server (NTRS)

Graves, Sharon S.; Burner, Alpheus W.; Edwards, John W.; Schuster, David M.

2001-01-01

The techniques used to acquire, reduce, and analyze dynamic deformation measurements of an aeroelastic semispan wind tunnel model are presented. Single-camera, single-view video photogrammetry (also referred to as videogrammetric model deformation, or VMD) was used to determine dynamic aeroelastic deformation of the semispan 'Models for Aeroelastic Validation Research Involving Computation' (MAVRIC) model in the Transonic Dynamics Tunnel at the NASA Langley Research Center. Dynamic deformation was determined from optical retroreflective tape targets at five semispan locations located on the wing from the root to the tip. Digitized video images from a charge coupled device (CCD) camera were recorded and processed to automatically determine target image plane locations that were then corrected for sensor, lens, and frame grabber spatial errors. Videogrammetric dynamic data were acquired at a 60-Hz rate for time records of up to 6 seconds during portions of this flutter/Limit Cycle Oscillation (LCO) test at Mach numbers from 0.3 to 0.96. Spectral analysis of the deformation data is used to identify dominant frequencies in the wing motion. The dynamic data will be used to separate aerodynamic and structural effects and to provide time history deflection data for Computational Aeroelasticity code evaluation and validation.

Development and demonstration of a flutter-suppression system using active controls. [wind tunnel tests

NASA Technical Reports Server (NTRS)

Sandford, M. C.; Abel, I.; Gray, D. L.

1975-01-01

The application of active control technology to suppress flutter was demonstrated successfully in the transonic dynamics tunnel with a delta-wing model. The model was a simplified version of a proposed supersonic transport wing design. An active flutter suppression method based on an aerodynamic energy criterion was verified by using three different control laws. The first two control laws utilized both leading-edge and trailing-edge active control surfaces, whereas the third control law required only a single trailing-edge active control surface. At a Mach number of 0.9 the experimental results demonstrated increases in the flutter dynamic pressure from 12.5 percent to 30 percent with active controls. Analytical methods were developed to predict both open-loop and closed-loop stability, and the results agreed reasonably well with the experimental results.

Investigation of Flow Separation in a Transonic-fan Linear Cascade Using Visualization Methods

NASA Technical Reports Server (NTRS)

Lepicovsky, Jan; Chima, Rodrick V.; Jett, Thomas A.; Bencic, Timothy J.; Weiland, Kenneth E.

2000-01-01

An extensive study into the nature of the separated flows on the suction side of modem transonic fan airfoils at high incidence is described in the paper. Suction surface.flow separation is an important flow characteristic that may significantly contribute to stall flutter in transonic fans. Flutter in axial turbomachines is a highly undesirable and dangerous self-excited mode of blade oscillations that can result in high cycle fatigue blade failure. The study basically focused on two visualization techniques: surface flow visualization using dye oils, and schlieren (and shadowgraph) flow visualization. The following key observations were made during the study. For subsonic inlet flow, the flow on the suction side of the blade is separated over a large portion of the blade, and the separated area increases with increasing inlet Mach number. For the supersonic inlet flow condition, the flow is attached from the leading edge up to the point where a bow shock from the upper neighboring blade hits the blade surface. Low cascade solidity, for the subsonic inlet flow, results in an increased area of separated flow. For supersonic flow conditions, a low solidity results in an improvement in flow over the suction surface. Finally, computational results modeling the transonic cascade flowfield illustrate our ability to simulate these flows numerically.

Heavy Gas Conversion of the NASA Langley Transonic Dynamics Tunnel

NASA Technical Reports Server (NTRS)

Corliss, James M.; Cole, Stanley, R.

1998-01-01

The heavy gas test medium has recently been changed in the Transonic Dynamics Tunnel (TDT) at the NASA Langley Research Center. A NASA Construction of Facilities project has converted the TDT heavy gas from dichlorodifluoromethane (R12) to 1,1,1,2 tetrafluoroethane (R134a). The facility s heavy gas processing system was extensively modified to implement the conversion to R134a. Additional system modifications have improved operator interfaces, hardware reliability, and quality of the research data. The facility modifications included improvements to the heavy gas compressor and piping, the cryogenic heavy gas reclamation system, and the heavy gas control room. A series of wind tunnel characterization and calibration tests are underway. Results of the flow characterization tests show the TDT operating envelope in R134a to be very similar to the previous operating envelope in R12.

Upgrades at the NASA Langley Research Center National Transonic Facility

NASA Technical Reports Server (NTRS)

Paryz, Roman W.

2012-01-01

Several projects have been completed or are nearing completion at the NASA Langley Research Center (LaRC) National Transonic Facility (NTF). The addition of a Model Flow-Control/Propulsion Simulation test capability to the NTF provides a unique, transonic, high-Reynolds number test capability that is well suited for research in propulsion airframe integration studies, circulation control high-lift concepts, powered lift, and cruise separation flow control. A 1992 vintage Facility Automation System (FAS) that performs the control functions for tunnel pressure, temperature, Mach number, model position, safety interlock and supervisory controls was replaced using current, commercially available components. This FAS upgrade also involved a design study for the replacement of the facility Mach measurement system and the development of a software-based simulation model of NTF processes and control systems. The FAS upgrades were validated by a post upgrade verification wind tunnel test. The data acquisition system (DAS) upgrade project involves the design, purchase, build, integration, installation and verification of a new DAS by replacing several early 1990's vintage computer systems with state of the art hardware/software. This paper provides an update on the progress made in these efforts. See reference 1.

Finite Element Analysis of a NASA National Transonic Facility Wind Tunnel Balance

NASA Technical Reports Server (NTRS)

Lindell, Michael C.

1996-01-01

This paper presents the results of finite element analyses and correlation studies performed on a NASA National Transonic Facility (NTF) Wind Tunnel balance. In the past NASA has relied primarily on classical hand analyses, coupled with relatively large safety factors, for predicting maximum stresses in wind tunnel balances. Now, with the significant advancements in computer technology and sophistication of general purpose analysis codes, it is more reasonable to pursue finite element analyses of these balances. The correlation studies of the present analyses show very good agreement between the analyses and data measured with strain gages and therefore the studies give higher confidence for using finite element analyses to analyze and optimize balance designs in the future.

Finite Element Analysis of a NASA National Transonic Facility Wide Tunnel Balance

NASA Technical Reports Server (NTRS)

Lindell, Michael C. (Editor)

1999-01-01

This paper presents the results of finite element analyses and correlation studies performed on a NASA National Transonic Facility (NTF) Wind Tunnel balance. In the past NASA has relied primarily on classical hand analyses, coupled with relatively large safety factors, for predicting maximum stresses in wind tunnel balances. Now, with the significant advancements in computer technology and sophistication of general purpose analysis codes, it is more reasonable to pursue finite element analyses of these balances. The correlation studies of the present analyses show very good agreement between the analyses and data measured with strain gages and therefore the studies give higher confidence for using finite element analyses to analyze and optimize balance designs in the future.

Issac, Jason Cherian ses in transonic flow

NASA Technical Reports Server (NTRS)

Issac, Jason Cherion; Kapania, Rakesh K.

1993-01-01

Flutter analysis of a two degree of freedom airfoil in compressible flow is performed using a state-space representation of the unsteady aerodynamic behavior. Indicial response functions are used to represent the normal force and moment response of the airfoil. The structural equations of motion of the airfoil with bending and torsional degrees of freedom are coupled to the unsteady air loads and the aeroelastic system so modelled is solved as an eigenvalue problem to determine the stability. The aeroelastic equations are also directly integrated with respect to time and the time-domain results compared with the results from the eigenanalysis. A good agreement is obtained. The derivatives of the flutter speed obtained from the eigenanalysis are calculated with respect to the mass and stiffness parameters by both analytical and finite-difference methods for various transonic Mach numbers. The experience gained from the two degree of freedom model is applied to study the sensitivity of the flutter response of a wing with respect to various shape parameters. The parameters being considered are as follows: (1) aspect ratio; (2) surface area of the wing; (3) taper ratio; and (4) sweep. The wing deflections are represented by Chebyshev polynomials. The compressible aerodynamic state-space model used for the airfoil section is extended to represent the unsteady aerodynamic forces on a generally laminated tapered skewed wing. The aeroelastic equations are solved as an eigenvalue problem to determine the flutter speed of the wing. The derivatives of the flutter speed with respect to the shape parameters are calculated by both analytical and finite difference methods.

Subsonic and Supersonic Flutter Analysis of a Highly Tapered Swept-Wing Planform, Including Effects of Density Variation and Finite Wing Thickness, and Comparison with Experiments

NASA Technical Reports Server (NTRS)

Yates, Carson, Jr.

1967-01-01

The flutter characteristics of several wings with an aspect-ratio of 4.0, a taper ratio of 0.2, and a quarter-chord sweepback of 45 deg. have been investigated analytically for Mach numbers up to 2.0. The calculations were based on the modified-strip-analysis method, the subsonic-kernel-function method, piston theory, and quasi-steady second-order theory. Results of t h e analysis and comparisons with experiment indicated that: (1) Flutter speeds were accurately predicted by the modified strip analysis, although accuracy at t h e highest Mach numbers required the use of nonlinear aerodynamic theory (which accounts for effects of wing thickness) for the calculation of the aerodynamic parameters. (2) An abrupt increase of flutter-speed coefficient with increasing Mach number, observed experimentally in the transonic range, was also indicated by the modified strip analysis. (3) In the low supersonic range for some densities, a discontinuous variation of flutter frequency with Mach number was indicated by the modified strip analysis. An abrupt change of frequency appeared experimentally in the transonic range. (4) Differences in flutter-speed-coefficient levels obtained from tests at low supersonic Mach numbers in two wind tunnels were also predicted by the modified strip analysis and were shown to be caused primarily by differences in mass ratio. (5) Flutter speeds calculated by the subsonic-kernel-function method were in good agreement with experiment and with the results of the modified strip analysis. (6) Flutter speed obtained from piston theory and from quasi-steady second-order theory were higher than experimental values by at least 38 percent.

Real-time flutter analysis

NASA Technical Reports Server (NTRS)

Walker, R.; Gupta, N.

1984-01-01

The important algorithm issues necessary to achieve a real time flutter monitoring system; namely, the guidelines for choosing appropriate model forms, reduction of the parameter convergence transient, handling multiple modes, the effect of over parameterization, and estimate accuracy predictions, both online and for experiment design are addressed. An approach for efficiently computing continuous-time flutter parameter Cramer-Rao estimate error bounds were developed. This enables a convincing comparison of theoretical and simulation results, as well as offline studies in preparation for a flight test. Theoretical predictions, simulation and flight test results from the NASA Drones for Aerodynamic and Structural Test (DAST) Program are compared.

Design, test, and evaluation of three active flutter suppression controllers

NASA Technical Reports Server (NTRS)

Adams, William M., Jr.; Christhilf, David M.; Waszak, Martin R.; Mukhopadhyay, Vivek; Srinathkumar, S.

1992-01-01

Three control law design techniques for flutter suppression are presented. Each technique uses multiple control surfaces and/or sensors. The first method uses traditional tools (such as pole/zero loci and Nyquist diagrams) for producing a controller that has minimal complexity and which is sufficiently robust to handle plant uncertainty. The second procedure uses linear combinations of several accelerometer signals and dynamic compensation to synthesize the model rate of the critical mode for feedback to the distributed control surfaces. The third technique starts with a minimum-energy linear quadratic Gaussian controller, iteratively modifies intensity matrices corresponding to input and output noise, and applies controller order reduction to achieve a low-order, robust controller. The resulting designs were implemented digitally and tested subsonically on the active flexible wing wind-tunnel model in the Langley Transonic Dynamics Tunnel. Only the traditional pole/zero loci design was sufficiently robust to errors in the nominal plant to successfully suppress flutter during the test. The traditional pole/zero loci design provided simultaneous suppression of symmetric and antisymmetric flutter with a 24-percent increase in attainable dynamic pressure. Posttest analyses are shown which illustrate the problems encountered with the other laws.

Current status of computational methods for transonic unsteady aerodynamics and aeroelastic applications

NASA Technical Reports Server (NTRS)

Edwards, John W.; Malone, John B.

1992-01-01

The current status of computational methods for unsteady aerodynamics and aeroelasticity is reviewed. The key features of challenging aeroelastic applications are discussed in terms of the flowfield state: low-angle high speed flows and high-angle vortex-dominated flows. The critical role played by viscous effects in determining aeroelastic stability for conditions of incipient flow separation is stressed. The need for a variety of flow modeling tools, from linear formulations to implementations of the Navier-Stokes equations, is emphasized. Estimates of computer run times for flutter calculations using several computational methods are given. Applications of these methods for unsteady aerodynamic and transonic flutter calculations for airfoils, wings, and configurations are summarized. Finally, recommendations are made concerning future research directions.

Feasibility study of the transonic biplane concept for transport aircraft application

NASA Technical Reports Server (NTRS)

Lange, R. H.; Cahill, J. F.; Bradley, E. S.; Eudaily, R. R.; Jenness, C. M.; Macwilkinson, D. G.

1974-01-01

Investigations were conducted to evaluate the feasibility of a transonic biplane consisting of a forward-mounted swept-back lower wing, a rear-mounted swept-forward upper wing, and a vertical fin connecting the wings at their tips. This wing arrangement results in significant reductions in induced drag relative to a monoplane designed with the same span, and it allows for a constant-section fuselage shape while closely matching an ideal area distribution curve for M = 0.95 cruise. However, no significant reductions in ramp weight were achieved for the biplane relative to a monoplane with the same mission capability. Flutter analyses of the biplane revealed both symmetric and antisymmetric instabilities that occur well below the required flutter speed. Further studies will be required to determine if acceptable flutter speeds can be achieved through the elimination of the instabilities by passive means or by active controls. Configurations designed for other missions, especially those with lower Mach numbers and lower dynamic pressures, should be examined since the geometries suitable for those design constraints might avoid the weight penalties and flutter instabilities which prevent exploitation of induced drag benefits for the configuration studied.

User's manual for XTRAN2L (version 1.2): A program for solving the general-frequency unsteady transonic small-disturbance equation

NASA Technical Reports Server (NTRS)

Seidel, D. A.; Batina, J. T.

1986-01-01

The development, use and operation of the XTRAN2L program that solves the two dimensional unsteady transonic small disturbance potential equation are described. The XTRAN2L program is used to calculate steady and unsteady transonic flow fields about airfoils and is capable of performing self contained transonic flutter calculations. Operation of the XTRAN2L code is described, and tables defining all input variables, including default values, are presented. Sample cases that use various program options are shown to illustrate operation of XTRAN2L. Computer listings containing input and selected output are included as an aid to the user.

Research and Applications in Aeroelasticity and Structural Dynamics at the NASA Langley Research Center

NASA Technical Reports Server (NTRS)

Abel, Irving

1997-01-01

An overview of recently completed programs in aeroelasticity and structural dynamics research at the NASA Langley Research Center is presented. Methods used to perform flutter clearance studies in the wind-tunnel on a high performance fighter are discussed. Recent advances in the use of smart structures and controls to solve aeroelastic problems, including flutter and gust response are presented. An aeroelastic models program designed to support an advanced high speed civil transport is described. An extension to transonic small disturbance theory that better predicts flows involving separation and reattachment is presented. The results of a research study to determine the effects of flexibility on the taxi and takeoff characteristics of a high speed civil transport are presented. The use of photogrammetric methods aboard Space Shuttle to measure spacecraft dynamic response is discussed. Issues associated with the jitter response of multi-payload spacecraft are discussed. Finally a Space Shuttle flight experiment that studied the control of flexible spacecraft is described.

Parametric Flutter Analysis of the TCA Configuration and Recommendation for FFM Design and Scaling

NASA Technical Reports Server (NTRS)

Baker, Myles; Lenkey, Peter

1997-01-01

The current HSR Aeroelasticity plan to design, build, and test a full span, free flying transonic flutter model in the TDT has many technical obstacles that must be overcome for a successful program. One technical obstacle is the determination of a suitable configuration and point in the sky to use in setting the scaling point for the ASE models program. Determining this configuration and point in the sky requires balancing several conflicting requirements, including model buildability, tunnel test safety, and the ability of the model to represent the flutter mechanisms of interest. As will be discussed in detail in subsequent sections, the current TCA design exhibits several flutter mechanisms of interest. It has been decided that the ASE models program will focus on the low frequency symmetric flutter mechanism, and will make no attempt to investigate high frequency flutter mechanisms. There are several reasons for this choice. First, it is believed that the high frequency flutter mechanisms are similar in nature to classical wing bending/torsion flutter, and therefore there is more confidence that this mechanism can be predicted using current techniques. The low frequency mode, on the other hand, is a highly coupled mechanism involving wing, body, tail, and engine motion which may be very difficult to predict. Second, the high frequency flutter modes result in very small weight penalties (several hundred pounds), while suppression of the low frequency mechanism inside the flight envelope causes thousands of pounds to be added to the structure. In order to successfully test the low frequency flutter mode of interest, a suitable starting configuration and point in the sky must be identified. The configuration and point in the sky must result in a wind tunnel model that (1) represents the low-frequency wing/body/engine/empennage flutter mechanisms that are unique to HSCT configurations, (2) flutters at an acceptably low frequency in the tunnel, (3) flutters at an

Characteristics of Control Laws Tested on the Semi-Span Super-Sonic Transport (S4T) Wind-Tunnel Model

NASA Technical Reports Server (NTRS)

Christhilf, David M.; Moulin, Boris; Ritz, Erich; Chen, P. C.; Roughen, Kevin M.; Perry, Boyd

2012-01-01

The Semi-Span Supersonic Transport (S4T) is an aeroelastically scaled wind-tunnel model built to test active controls concepts for large flexible supersonic aircraft in the transonic flight regime. It is one of several models constructed in the 1990's as part of the High Speed Research (HSR) Program. Control laws were developed for the S4T by M4 Engineering, Inc. and by Zona Technologies, Inc. under NASA Research Announcement (NRA) contracts. The model was tested in the NASA-Langley Transonic Dynamics Tunnel (TDT) four times from 2007 to 2010. The first two tests were primarily for plant identification. The third entry was used for testing control laws for Ride Quality Enhancement, Gust Load Alleviation, and Flutter Suppression. Whereas the third entry only tested FS subcritically, the fourth test demonstrated closed-loop operation above the open-loop flutter boundary. The results of the third entry are reported elsewhere. This paper reports on flutter suppression results from the fourth wind-tunnel test. Flutter suppression is seen as a way to provide stability margins while flying at transonic flight conditions without penalizing the primary supersonic cruise design condition. An account is given for how Controller Performance Evaluation (CPE) singular value plots were interpreted with regard to progressing open- or closed-loop to higher dynamic pressures during testing.

Emerging technology for transonic wind-tunnel-wall interference assessment and corrections

NASA Technical Reports Server (NTRS)

Newman, P. A.; Kemp, W. B., Jr.; Garriz, J. A.

1988-01-01

Several nonlinear transonic codes and a panel method code for wind tunnel/wall interference assessment and correction (WIAC) studies are reviewed. Contrasts between two- and three-dimensional transonic testing factors which affect WIAC procedures are illustrated with airfoil data from the NASA/Langley 0.3-meter transonic cyrogenic tunnel and Pathfinder I data. Also, three-dimensional transonic WIAC results for Mach number and angle-of-attack corrections to data from a relatively large 20 deg swept semispan wing in the solid wall NASA/Ames high Reynolds number Channel I are verified by three-dimensional thin-layer Navier-Stokes free-air solutions.

Vapor-screen flow-visualization experiments in the NASA Langley 0.3-m transonic cryogenic tunnel

NASA Technical Reports Server (NTRS)

Selby, G. V.

1986-01-01

The vortical flow on the leeward side of a delta-wing model has been visualized at several different tunnel conditions in the NASA Langley 0.3-Meter Transonic Cryogenic Tunnel using a vapor-screen flow-visualization technique. Vapor-screen photographs of the subject flow field are presented and interpreted relative to phenomenological implications. Results indicate that the use of nitrogen fog in conjunction with the vapor-screen technique is feasibile.

A New Forced Oscillation Capability for the Transonic Dynamics Tunnel

NASA Technical Reports Server (NTRS)

Piatak, David J.; Cleckner, Craig S.

2002-01-01

A new forced oscillation system has been installed and tested at NASA Langley Research Center's Transonic Dynamics Tunnel (TDT). The system is known as the Oscillating Turntable (OTT) and has been designed for the purpose of oscillating, large semispan models in pitch at frequencies up to 40 Hz to acquire high-quality unsteady pressure and loads data. Precisely controlled motions of a wind-tunnel model on the OTT can yield unsteady aerodynamic phenomena associated with flutter, limit cycle oscillations, shock dynamics, and non-linear aerodynamic effects on many vehicle configurations. This paper will discuss general design and components of the OTT and will present test data from performance testing and from research tests on two rigid semispan wind-tunnel models. The research tests were designed to challenge the OTT over a wide range of operating conditions while acquiring unsteady pressure data on a small rectangular supercritical wing and a large supersonic transport wing. These results will be presented to illustrate the performance capabilities, consistency of oscillations, and usefulness of the OTT as a research tool.

Investigation of the Flow Physics Driving Stall-Side Flutter in Advanced Forward Swept Fan Designs

NASA Technical Reports Server (NTRS)

Sanders, Albert J.; Liu, Jong S.; Panovsky, Josef; Bakhle, Milind A.; Stefko, George; Srivastava, Rakesh

2003-01-01

Flutter-free operation of advanced transonic fan designs continues to be a challenging task for the designers of aircraft engines. In order to meet the demands of increased performance and lighter weight, these modern fan designs usually feature low-aspect ratio shroudless rotor blade designs that make the task of achieving adequate flutter margin even more challenging for the aeroelastician. This is especially true for advanced forward swept designs that encompass an entirely new design space compared to previous experience. Fortunately, advances in unsteady computational fluid dynamic (CFD) techniques over the past decade now provide an analysis capability that can be used to quantitatively assess the aeroelastic characteristics of these next generation fans during the design cycle. For aeroelastic applications, Mississippi State University and NASA Glenn Research Center have developed the CFD code TURBO-AE. This code is a time-accurate three-dimensional Euler/Navier-Stokes unsteady flow solver developed for axial-flow turbomachinery that can model multiple blade rows undergoing harmonic oscillations with arbitrary interblade phase angles, i.e., nodal diameter patterns. Details of the code can be found in Chen et al. (1993, 1994), Bakhle et al. (1997, 1998), and Srivastava et al. (1999). To assess aeroelastic stability, the work-per-cycle from TURBO-AE is converted to the critical damping ratio since this value is more physically meaningful, with both the unsteady normal pressure and viscous shear forces included in the work-per-cycle calculation. If the total damping (aerodynamic plus mechanical) is negative, then the blade is unstable since it extracts energy from the flow field over the vibration cycle. TURBO-AE is an integral part of an aeroelastic design system being developed at Honeywell Engines, Systems & Services for flutter and forced response predictions, with test cases from development rig and engine tests being used to validate its predictive

Free-Flight Tests of 0.11-Scale North American F-100 Airplane Wings to Investigate the Possibility of Flutter in Transonic Speed Range at Varying Angles of Attack

NASA Technical Reports Server (NTRS)

O'Kelly, Burke R.

1954-01-01

Free-flight tests in the transonic speed range utilizing rocketpropelled models have been made on three pairs of 0.11-scale North American F-100 airplane wings having an aspect ratio of 3.47, a taper ratio of 0.308, 45 degree sweepback at the quarter-chord line, and thickness ratios of 31 and 5 percent to investigate the possibility of flutte r. Data from tests of two other rocket-propelled models which accidentally fluttered during a drag investigation of the North American F-100 airplane are also presented. The first set of wings (5 percent thick) was tested on a model which was disturbed in pitch by a moving tail and reached a maximum Mach number of 0.85. The wings encountered mild oscillations near the first - bending frequency at high lift coefficients. The second set of wings 9 percent thick was tested up to a maximum Mach number of 0.95 at (2) angles of attack provided by small rocket motors installed in the nose of the model. No oscillations resembling flutter were encountered during the coasting flight between separation from the booster and sustainer firing (Mach numbers from 0.86 to 0.82) or during the sustainer firing at accelerations of about 8g up to the maximum Mach number of the test (0.95). The third set of wings was similar to the first set and was tested up to a maximum Mach number of 1.24. A mild flutter at frequencies near the first-bending frequency of the wings was encountered between a Mach number of 1.15 and a Mach number of 1.06 during both accelerating and coasting flight. The two drag models, which were 0.ll-scale models of the North American F-100 airplane configuration, reached a maximum Mach number of 1.77. The wings of these models had bending and torsional frequencies which were 40 and 89 percent, respectively, of the calculated scaled frequencies of the full-scale 7-percent-thick wing. Both models experienced flutter of the same type as that experienced-by the third set of wings.

Economical Unsteady High-Fidelity Aerodynamics for Structural Optimization with a Flutter Constraint

NASA Technical Reports Server (NTRS)

Bartels, Robert E.; Stanford, Bret K.

2017-01-01

Structural optimization with a flutter constraint for a vehicle designed to fly in the transonic regime is a particularly difficult task. In this speed range, the flutter boundary is very sensitive to aerodynamic nonlinearities, typically requiring high-fidelity Navier-Stokes simulations. However, the repeated application of unsteady computational fluid dynamics to guide an aeroelastic optimization process is very computationally expensive. This expense has motivated the development of methods that incorporate aspects of the aerodynamic nonlinearity, classical tools of flutter analysis, and more recent methods of optimization. While it is possible to use doublet lattice method aerodynamics, this paper focuses on the use of an unsteady high-fidelity aerodynamic reduced order model combined with successive transformations that allows for an economical way of utilizing high-fidelity aerodynamics in the optimization process. This approach is applied to the common research model wing structural design. As might be expected, the high-fidelity aerodynamics produces a heavier wing than that optimized with doublet lattice aerodynamics. It is found that the optimized lower skin of the wing using high-fidelity aerodynamics differs significantly from that using doublet lattice aerodynamics.

Influence of Shock Wave on the Flutter Behavior of Fan Blades Investigated

NASA Technical Reports Server (NTRS)

Srivastava, Rakesh; Bakhle, Milind A.; Stefko, George L.

2003-01-01

Modern fan designs have blades with forward sweep; a lean, thin cross section; and a wide chord to improve performance and reduce noise. These geometric features coupled with the presence of a shock wave can lead to flutter instability. Flutter is a self-excited dynamic instability arising because of fluid-structure interaction, which causes the energy from the surrounding fluid to be extracted by the vibrating structure. An in-flight occurrence of flutter could be catastrophic and is a significant design issue for rotor blades in gas turbines. Understanding the flutter behavior and the influence of flow features on flutter will lead to a better and safer design. An aeroelastic analysis code, TURBO, has been developed and validated for flutter calculations at the NASA Glenn Research Center. The code has been used to understand the occurrence of flutter in a forward-swept fan design. The forward-swept fan, which consists of 22 inserted blades, encountered flutter during wind tunnel tests at part speed conditions.

Flutter Analysis of the Shuttle Tile Overlay Repair Concept

NASA Technical Reports Server (NTRS)

Bey, Kim S.; Scott, Robert C.; Bartels, Robert E.; Waters, William A.; Chen, Roger

2007-01-01

The Space Shuttle tile overlay repair concept, developed at the NASA Johnson Space Center, is designed for on-orbit installation over an area of damaged tile to permit safe re-entry. The thin flexible plate is placed over the damaged area and secured to tile at discreet points around its perimeter. A series of flutter analyses were performed to determine if the onset of flutter met the required safety margins. Normal vibration modes of the panel, obtained from a simplified structural analysis of the installed concept, were combined with a series of aerodynamic analyses of increasing levels of fidelity in terms of modeling the flow physics to determine the onset of flutter. Results from these analyses indicate that it is unlikely that the overlay installed at body point 1800 will flutter during re-entry.

Investigations for Supersonic Transports at Transonic and Supersonic Conditions

NASA Technical Reports Server (NTRS)

Rivers, S. Melissa B.; Owens, Lewis R.; Wahls, Richard A.

2007-01-01

Several computational studies were conducted as part of NASA s High Speed Research Program. Results of turbulence model comparisons from two studies on supersonic transport configurations performed during the NASA High-Speed Research program are given. The effects of grid topology and the representation of the actual wind tunnel model geometry are also investigated. Results are presented for both transonic conditions at Mach 0.90 and supersonic conditions at Mach 2.48. A feature of these two studies was the availability of higher Reynolds number wind tunnel data with which to compare the computational results. The transonic wind tunnel data was obtained in the National Transonic Facility at NASA Langley, and the supersonic data was obtained in the Boeing Polysonic Wind Tunnel. The computational data was acquired using a state of the art Navier-Stokes flow solver with a wide range of turbulence models implemented. The results show that the computed forces compare reasonably well with the experimental data, with the Baldwin-Lomax with Degani-Schiff modifications and the Baldwin-Barth models showing the best agreement for the transonic conditions and the Spalart-Allmaras model showing the best agreement for the supersonic conditions. The transonic results were more sensitive to the choice of turbulence model than were the supersonic results.

Whirl Flutter Studies for a SSTOL Transport Demonstrator

NASA Technical Reports Server (NTRS)

Acree, C. W., Jr.; Hoffman, Krishna

2004-01-01

A proposed new class of aircraft - the Advanced Theater Transport (ATT) will combine strategic range and high payload with 'Super-STOL' (short take-off and landing) capability. It is also proposed to modify a YC-15 into a technology demonstrator with a 20-deg tilt wing; four, eight-bladed propellers; cross-shafted gearboxes and V-22 engines. These constitute a unique combination of design features that potentially affect performance, loads and whirl-mode stability (whirl flutter). NASA Ames Research Center is working with Boeing and Hamilton Sundstrand on technology challenges presented by the concept; the purpose of NASA involvement is to establish requirements for the demonstrator and for early design guidance, with emphasis on whirl flutter. CAMRAD II is being used to study the effects of various design features on whirl flutter, with special attention to areas where such features differ from existing aircraft, notably tiltrotors. Although the stability margins appear to be more than adequate, the concept requires significantly different analytical methods, principally including far more blade modes, than typically used for tiltrotors.

An Experimental Evaluation of Advanced Rotorcraft Airfoils in the NASA Ames Eleven-foot Transonic Wind Tunnel

NASA Technical Reports Server (NTRS)

Flemming, Robert J.

1984-01-01

Five full scale rotorcraft airfoils were tested in the NASA Ames Eleven-Foot Transonic Wind Tunnel for full scale Reynolds numbers at Mach numbers from 0.3 to 1.07. The models, which spanned the tunnel from floor to ceiling, included two modern baseline airfoils, the SC1095 and SC1094 R8, which have been previously tested in other facilities. Three advanced transonic airfoils, designated the SSC-A09, SSC-A07, and SSC-B08, were tested to confirm predicted performance and provide confirmation of advanced airfoil design methods. The test showed that the eleven-foot tunnel is suited to two-dimensional airfoil testing. Maximum lift coefficients, drag coefficients, pitching moments, and pressure coefficient distributions are presented. The airfoil analysis codes agreed well with the data, with the Grumman GRUMFOIL code giving the best overall performance correlation.

Active Control of Wind-Tunnel Model Aeroelastic Response Using Neural Networks

NASA Technical Reports Server (NTRS)

Scott, Robert C.

2000-01-01

NASA Langley Research Center, Hampton, VA 23681 Under a joint research and development effort conducted by the National Aeronautics and Space Administration and The Boeing Company (formerly McDonnell Douglas) three neural-network based control systems were developed and tested. The control systems were experimentally evaluated using a transonic wind-tunnel model in the Langley Transonic Dynamics Tunnel. One system used a neural network to schedule flutter suppression control laws, another employed a neural network in a predictive control scheme, and the third employed a neural network in an inverse model control scheme. All three of these control schemes successfully suppressed flutter to or near the limits of the testing apparatus, and represent the first experimental applications of neural networks to flutter suppression. This paper will summarize the findings of this project.

Aerodynamic Measurements on a Large Splitter Plate for the NASA Langley Transonic Dynamics Tunnel

NASA Technical Reports Server (NTRS)

Schuster, David M.

2001-01-01

Tests conducted in the NASA Langley Research Center Transonic Dynamics Tunnel (TDT) assess the aerodynamic characteristics of a splitter plate used to test some semispan models in this facility. Aerodynamic data are analyzed to determine the effect of the splitter plate on the operating characteristics of the TDT, as well as to define the range of conditions over which the plate can be reasonably used to obtain aerodynamic data. Static pressures measurements on the splitter plate surface and the equipment fairing between the wind tunnel wall and the splitter plate are evaluated to determine the flow quality around the apparatus over a range of operating conditions. Boundary layer rake data acquired near the plate surface define the viscous characteristics of the flow over the plate. Data were acquired over a range of subsonic, transonic and supersonic conditions at dynamic pressures typical for models tested on this apparatus. Data from this investigation should be used as a guide for the design of TDT models and tests using the splitter plate, as well as to guide future splitter plate design for this facility.

Transonic empirical configuration design process

NASA Technical Reports Server (NTRS)

Whitcomb, R. T.

1983-01-01

This lecture describes some of the experimental research pertaining to transonic configuration development conducted by the Transonic Aerodynamics Branch of the NASA Langley Research Center. Discussions are presented of the following: use of florescent oil films for the study of surface boundary layer flows; the severe effect of wind tunnel wall interference on the measured configuration drag rise near the speed of sound as determined by a comparison between wind tunnel and free air results; the development of a near sonic transport configuration incorporating a supercritical wing and an indented fuselage, designed on the basis of the area rule with a modification to account for the presence of local supersonic flow above the wing; a device for improving the transonic pitch up of swept wings with very little added drag at the cruise condition; a means for reducing the large transonic aerodynamic interference between the wing, fuselage, nacelle and pylon for a for a fuselage mounted nacelle having the inlet above the wing; and methods for reducing the transonic interference between flows over a winglet and the wing.

Evaluation of Aeroservoelastic Effects on Flutter

NASA Technical Reports Server (NTRS)

Nagaraja, K. S.; Felt, Larry R.; Kraft, Raymond

1998-01-01

This report presents work performed by The Boeing Company to satisfy the deliverable "Evaluation of aeroservoelastic Effects on Symmetric Flutter" for Subtask 7 of Reference 1. The objective of this report is to incorporate the improved methods for studying the effects of a closed-loop control system on the aeroservoelastic behavior of the airplane planned under NASA HSR technical Integration Task 20 work. Also, a preliminary evaluation of the existing pitch control laws on symmetric flutter of the TCA configuration was addressed."The goal is to develop an improved modeling methodology and perform design studies that account for the aero-structures-systems interaction effects.

Active load control during rolling maneuvers. [performed in the Langley Transonic Dynamics Tunnel

NASA Technical Reports Server (NTRS)

Woods-Vedeler, Jessica A.; Pototzky, Anthony S.; Hoadley, Sherwood T.

1994-01-01

A rolling maneuver load alleviation (RMLA) system has been demonstrated on the active flexible wing (AFW) wind tunnel model in the Langley Transonic Dynamics Tunnel (TDT). The objective was to develop a systematic approach for designing active control laws to alleviate wing loads during rolling maneuvers. Two RMLA control laws were developed that utilized outboard control-surface pairs (leading and trailing edge) to counteract the loads and that used inboard trailing-edge control-surface pairs to maintain roll performance. Rolling maneuver load tests were performed in the TDT at several dynamic pressures that included two below and one 11 percent above open-loop flutter dynamic pressure. The RMLA system was operated simultaneously with an active flutter suppression system above open-loop flutter dynamic pressure. At all dynamic pressures for which baseline results were obtained, torsion-moment loads were reduced for both RMLA control laws. Results for bending-moment load reductions were mixed; however, design equations developed in this study provided conservative estimates of load reduction in all cases.

Surface Acoustic Wave Vibration Sensors for Measuring Aircraft Flutter

NASA Technical Reports Server (NTRS)

Wilson, William C.; Moore, Jason P.; Juarez, Peter D.

2016-01-01

Under NASA's Advanced Air Vehicles Program the Advanced Air Transport Technology (AATT) Project is investigating flutter effects on aeroelastic wings. To support that work a new method for measuring vibrations due to flutter has been developed. The method employs low power Surface Acoustic Wave (SAW) sensors. To demonstrate the ability of the SAW sensor to detect flutter vibrations the sensors were attached to a Carbon fiber-reinforced polymer (CFRP) composite panel which was vibrated at six frequencies from 1Hz to 50Hz. The SAW data was compared to accelerometer data and was found to resemble sine waves and match each other closely. The SAW module design and results from the tests are presented here.

Prediction of Flutter Boundary Using Flutter Margin for The Discrete-Time System

NASA Astrophysics Data System (ADS)

Dwi Saputra, Angga; Wibawa Purabaya, R.

2018-04-01

Flutter testing in a wind tunnel is generally conducted at subcritical speeds to avoid damages. Hence, The flutter speed has to be predicted from the behavior some of its stability criteria estimated against the dynamic pressure or flight speed. Therefore, it is quite important for a reliable flutter prediction method to estimates flutter boundary. This paper summarizes the flutter testing of a wing cantilever model in a wind tunnel. The model has two degree of freedom; they are bending and torsion modes. The flutter test was conducted in a subsonic wind tunnel. The dynamic data responses was measured by two accelerometers that were mounted on leading edge and center of wing tip. The measurement was repeated while the wind speed increased. The dynamic responses were used to determine the parameter flutter margin for the discrete-time system. The flutter boundary of the model was estimated using extrapolation of the parameter flutter margin against the dynamic pressure. The parameter flutter margin for the discrete-time system has a better performance for flutter prediction than the modal parameters. A model with two degree freedom and experiencing classical flutter, the parameter flutter margin for the discrete-time system gives a satisfying result in prediction of flutter boundary on subsonic wind tunnel test.

Fifteen Years of Operation at NASA's National Transonic Facility with the World's Largest Adjustable Speed Drive

NASA Technical Reports Server (NTRS)

Sydnor, George H.; Bhatia, Ram; Krattiger, Hansueli; Mylius, Justus; Schafer, D.

2012-01-01

In September 1995, a project was initiated to replace the existing drive line at NASA's most unique transonic wind tunnel, the National Transonic Facility (NTF), with a single 101 MW synchronous motor driven by a Load Commutated Inverter (LCI). This Adjustable Speed Drive (ASD) system also included a custom four-winding transformer, harmonic filter, exciter, switch gear, control system, and feeder cable. The complete system requirements and design details have previously been presented and published [1], as well as the commissioning and acceptance test results [2]. The NTF was returned to service in December 1997 with the new drive system powering the fan. Today, this installation still represents the world s largest horizontal single motor/drive combination. This paper describes some significant events that occurred with the drive system during the first 15 years of service. These noteworthy issues are analyzed and root causes presented. Improvements that have substantially increased the long term viability of the system are given.

Some applications of the NASTRAN level 16 subsonic flutter analysis capability. [to transport wing and arrow wing

NASA Technical Reports Server (NTRS)

Doggett, R. V., Jr.; Cunningham, H. J.

1976-01-01

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.

Flutter and Forced Response Analyses of Cascades using a Two-Dimensional Linearized Euler Solver

NASA Technical Reports Server (NTRS)

Reddy, T. S. R.; Srivastava, R.; Mehmed, O.

1999-01-01

Flutter and forced response analyses for a cascade of blades in subsonic and transonic flow is presented. The structural model for each blade is a typical section with bending and torsion degrees of freedom. The unsteady aerodynamic forces due to bending and torsion motions. and due to a vortical gust disturbance are obtained by solving unsteady linearized Euler equations. The unsteady linearized equations are obtained by linearizing the unsteady nonlinear equations about the steady flow. The predicted unsteady aerodynamic forces include the effect of steady aerodynamic loading due to airfoil shape, thickness and angle of attack. The aeroelastic equations are solved in the frequency domain by coupling the un- steady aerodynamic forces to the aeroelastic solver MISER. The present unsteady aerodynamic solver showed good correlation with published results for both flutter and forced response predictions. Further improvements are required to use the unsteady aerodynamic solver in a design cycle.

Rapid Aeroelastic Analysis of Blade Flutter in Turbomachines

NASA Technical Reports Server (NTRS)

Trudell, J. J.; Mehmed, O.; Stefko, G. L.; Bakhle, M. A.; Reddy, T. S. R.; Montgomery, M.; Verdon, J.

2006-01-01

The LINFLUX-AE computer code predicts flutter and forced responses of blades and vanes in turbomachines under subsonic, transonic, and supersonic flow conditions. The code solves the Euler equations of unsteady flow in a blade passage under the assumption that the blades vibrate harmonically at small amplitudes. The steady-state nonlinear Euler equations are solved by a separate program, then equations for unsteady flow components are obtained through linearization around the steady-state solution. A structural-dynamics analysis (see figure) is performed to determine the frequencies and mode shapes of blade vibrations, a preprocessor interpolates mode shapes from the structural-dynamics mesh onto the LINFLUX computational-fluid-dynamics mesh, and an interface code is used to convert the steady-state flow solution to a form required by LINFLUX. Then LINFLUX solves the linearized equations in the frequency domain to calculate the unsteady aerodynamic pressure distribution for a given vibration mode, frequency, and interblade phase angle. A post-processor uses the unsteady pressures to calculate generalized aerodynamic forces, response amplitudes, and eigenvalues (which determine the flutter frequency and damping). In comparison with the TURBO-AE aeroelastic-analysis code, which solves the equations in the time domain, LINFLUX-AE is 6 to 7 times faster.

Pressure-Sensitive Paint Measurements on the NASA Common Research Model in the NASA 11-ft Transonic Wind Tunnel

NASA Technical Reports Server (NTRS)

Bell, James H.

2011-01-01

The luminescence lifetime technique was used to make pressure-sensitive paint (PSP) measurements on a 2.7% Common Research Model in the NASA Ames 11ft Transonic Wind Tunnel. PSP data were obtained on the upper and lower surfaces of the wing and horizontal tail, as well as one side of the fuselage. Data were taken for several model attitudes of interest at Mach numbers between 0.70 and 0.87. Image data were mapped onto a three-dimensional surface grid suitable both for comparison with CFD and for integration of pressures to determine loads. Luminescence lifetime measurements were made using strobed LED (light-emitting diode) lamps to illuminate the PSP and fast-framing interline transfer cameras to acquire the PSP emission.

An experimental study of tip shape effects on the flutter of aft-swept, flat-plate wings

NASA Technical Reports Server (NTRS)

Dansberry, Bryan E.; Rivera, Jose A., Jr.; Farmer, Moses G.

1990-01-01

The effects of tip chord orientation on wing flutter are investigated experimentally using six cantilever-mounted, flat-plate wing models. Experimentally determined flutter characteristics of the six models are presented covering both the subsonic and transonic Mach number ranges. While all models have a 60 degree leading edge sweep, a 40.97 degree trailing edge sweep, and a root chord of 34.75 inches, they are subdivided into two series characterized by a higher aspect ratio and a lower aspect ratio. Each series is made up of three models with tip chord orientations which are parallel to the free-stream flow, perpendicular to the model mid-chord line, and perpendicular to the free-stream flow. Although planform characteristics within each series of models are held constant, structural characteristics such as mode shapes and natural frequencies are allowed to vary.

Turbulence Model Comparisons for Supersonic Transports at Transonic and Supersonic Conditions

NASA Technical Reports Server (NTRS)

Rivers, S. M. B.; Wahls, R. A.

2003-01-01

Results of turbulence model comparisons from two studies on supersonic transport configurations performed during the NASA High-speed Research program are given. Results are presented for both transonic conditions at Mach 0.90 and supersonic conditions at Mach 2.48. A feature of these two studies was the availability of higher Reynolds number wind tunnel data with which to compare the computational results. The transonic wind tunnel data was obtained in the National Transonic Facility at NASA Langley, and the supersonic data was obtained in the Boeing Polysonic Wind Tunnel. The computational data was acquired using a state of the art Navier-Stokes flow solver with a wide range of turbulence models implemented. The results show that the computed forces compare reasonably well with the experimental data, with the Baldwin- Lomax with Degani-Schiff modifications and the Baldwin-Barth models showing the best agreement for the transonic conditions and the Spalart-Allmaras model showing the best agreement for the supersonic conditions. The transonic results were more sensitive to the choice of turbulence model than were the supersonic results.

Turbomachinery aeroelasticity at NASA Lewis Research Center

NASA Technical Reports Server (NTRS)

Kaza, Krishna Rao V.

1989-01-01

The turbomachinery aeroelastic effort is focused on unstalled and stalled flutter, forced response, and whirl flutter of both single rotation and counter rotation propfans. It also includes forced response of the Space Shuttle Main Engine (SSME) turbopump blades. Because of certain unique features of propfans and the SSME turbopump blades, it is not possible to directly use the existing aeroelastic technology of conventional propellers, turbofans or helicopters. Therefore, reliable aeroelastic stability and response analysis methods for these propulsion systems must be developed. The development of these methods for propfans requires specific basic technology disciplines, such as 2-D and 3-D steady and unsteady aerodynamic theories in subsonic, transonic and supersonic flow regimes; modeling of composite blades; geometric nonlinear effects; and passive and active control of flutter and response. These methods are incorporated in a computer program, ASTROP. The program has flexibility such that new and future models in basic disciplines can be easily implemented.

Flow Quality Measurements in the NASA Ames Upgraded 11-by 11-Foot Transonic Wind Tunnel

NASA Technical Reports Server (NTRS)

Amaya, Max A.; Murthy, Sreedhara V.; George, M. W. (Technical Monitor)

2000-01-01

Among the many upgrades designed and implemented in the NASA Ames 11-by 11-Foot Transonic Wind Tunnel over the past few years, several directly affect flow quality in the test section: a turbulence reduction system with a honeycomb and two screens, a flow smoothing system in the back leg diffusers, an improved drive motor control system, and a full replacement set of composite blades for the compressor. Prior to the shut-down of the tunnel for construction activities, an 8-foot span rake populated with flow instrumentation was traversed in the test section to fully document the flow quality and establish a baseline against which the upgrades could be characterized. A similar set of measurements was performed during the recent integrated system test trials, but the scope was somewhat limited in accordance with the primary objective of such tests, namely to return the tunnel to a fully operational status. These measurements clearly revealed substantial improvements in flow angularity and significant reductions in turbulence level for both full-span and semi-span testing configurations, thus making the flow quality of the tunnel one of the best among existing transonic facilities.

Evaluation of Aeroservoelastic Effects on Flutter

NASA Technical Reports Server (NTRS)

Nagaraja, K. S.; Kraft, raymond; Felt, Larry

1998-01-01

The HSCT Flight Controls Group is developing a longitudinal control law, known as Gamma-dot / V, for the NASA HSR program. Currently, this control law is based on a quasi-steady aeroelastic (QSAE) model of the vehicle. This control law was implemented into the p-k flutter analysis process for closed loop aeroservoelastic analysis. The available flexible models, developed for the TCA aeroelastic analysis, were used to assess the effect of control laws on flutter at several different Mach numbers and mass conditions. Significant structures and flight control system interaction was observed during the initial assessment. Figures 1 and 2 present a summary of the effect of total closed loop gain and phase on flutter mechanisms, based on ideal sensors and real sensors, for Mach 0.95 and mass M02 condition. Control laws based on ideal sensors gave rise to increased coupling between the rigid body short period mode and the first symmetric elastic mode. This reduced the stability margins for the first elastic mode and does not meet the required 6 dB gain margin requirement. The effect of "real" sensors significantly increased the structures and control system interactions. This caused the elastic,modes to be highly unstable throughout most of the flight envelope. State-space models were developed for several conditions and then MATLAB program was used for the aeroservoelastic stability analysis. These results provided an independent verification of the p-k flutter analysis findings. Good overall agreement was observed between the p-k flutter analysis and state-space model results for both damping and frequency comparisons. These results are also included in this document.

Transonic Symposium: Theory, Application, and Experiment, volume 1, part 2

NASA Technical Reports Server (NTRS)

Foughner, Jerome T., Jr. (Compiler)

1989-01-01

In order to assess the state of the art in transonic flow disciplines and to glimpse at future directions, NASA-Langley held a Transonic Symposium. Emphasis was placed on steady, three dimensional external, transonic flow and its simulation, both numerically and experimentally. The symposium included technical sessions on wind tunnel and flight experiments; computational fluid dynamic applications; inviscid methods and grid generation; viscous methods and boundary layer stability; and wind tunnel techniques and wall interference. This, being volume 1, is unclassified.

Possible safety hazards associated with the operation of the 0.3-m transonic cryogenic tunnel at the NASA Langley Research Center

NASA Technical Reports Server (NTRS)

Webster, T. J.

1982-01-01

The 0.3 m Transonic Cryogenic Tunnel (TCT) at the NASA Langley Research Center was built in 1973 as a facility intended to be used for no more than 60 hours in order to verify the validity of the cryogenic wind tunnel concept at transonic speeds. The role of the 0.3 m TCT has gradually changed until now, after over 3000 hours of operation, it is classified as a major NASA research facility and, under the administration of the Experimental Techniques Branch, it is used extensively for the testing of airfoils at high Reynolds numbers and for the development of various technologies related to the efficient operation and use of cryogenic wind tunnels. The purpose of this report is to document the results of a recent safety analysis of the 0.3 m TCT facility. This analysis was made as part of an on going program with the Experimental Techniques Branch designed to ensure that the existing equipment and current operating procedures of the 0.3 m TCT facility are acceptable in terms of today's standards of safety for cryogenic systems.

Noise Prediction of NASA SR2 Propeller in Transonic Conditions

NASA Astrophysics Data System (ADS)

Gennaro, Michele De; Caridi, Domenico; Nicola, Carlo De

2010-09-01

In this paper we propose a numerical approach for noise prediction of high-speed propellers for Turboprop applications. It is based on a RANS approach for aerodynamic simulation coupled with Ffowcs Williams-Hawkings (FW-H) Acoustic Analogy for propeller noise prediction. The test-case geometry adopted for this study is the 8-bladed NASA SR2 transonic cruise propeller, and simulated Sound Pressure Levels (SPL) have been compared with experimental data available from Wind Tunnel and Flight Tests for different microphone locations in a range of Mach numbers between 0.78 and 0.85 and rotational velocities between 7000 and 9000 rpm. Results show the ability of this approach to predict noise to within a few dB of experimental data. Moreover corrections are provided to be applied to acoustic numerical results in order for them to be compared with Wind Tunnel and Flight Test experimental data, as well computational grid requirements and guidelines in order to perform complete aerodynamic and aeroacoustic calculations with highly competitive computational cost.

Aircraft Flutter Testing

NASA Technical Reports Server (NTRS)

1997-01-01

Wilmer Reed gained international recognition for his innovative research, contributions and patented ideas relating to flutter and aeroelasticity of aerospace vehicles at Langley Research Center. In the early 1980's, Reed retired from Langley and joined the engineering staff of Dynamic Engineering Inc. While at DEI, Reed conceived and patented the DEI Flutter Exciter, now used world-wide in flight flutter testing of new or modified aircraft designs. When activated, the DEI Flutter Exciter alternately deflects the airstream upward and downward in a rapid manner, creating a force similar to that produced by an oscillating trailing edge flap. The DEI Flutter Exciter is readily adaptable to a variety of aircraft.

Wall Boundary Layer Measurements for the NASA Langley Transonic Dynamics Tunnel

NASA Technical Reports Server (NTRS)

Wieseman, Carol D.; Bennett, Robert M.

2007-01-01

Measurements of the boundary layer parameters in the NASA Langley Transonic Dynamics tunnel were conducted during extensive calibration activities following the facility conversion from a Freon-12 heavy-gas test medium to R-134a. Boundary-layer rakes were mounted on the wind-tunnel walls, ceiling, and floor. Measurements were made over the range of tunnel operation envelope in both heavy gas and air and without a model in the test section at three tunnel stations. Configuration variables included open and closed east sidewall wall slots, for air and R134a test media, reentry flap settings, and stagnation pressures over the full range of tunnel operation. The boundary layer thickness varied considerably for the six rakes. The thickness for the east wall was considerably larger that the other rakes and was also larger than previously reported. There generally was some reduction in thickness at supersonic Mach numbers, but the effect of stagnation pressure, and test medium were not extensive.

Reduced-Order Modeling for Flutter/LCO Using Recurrent Artificial Neural Network

NASA Technical Reports Server (NTRS)

Yao, Weigang; Liou, Meng-Sing

2012-01-01

The present study demonstrates the efficacy of a recurrent artificial neural network to provide a high fidelity time-dependent nonlinear reduced-order model (ROM) for flutter/limit-cycle oscillation (LCO) modeling. An artificial neural network is a relatively straightforward nonlinear method for modeling an input-output relationship from a set of known data, for which we use the radial basis function (RBF) with its parameters determined through a training process. The resulting RBF neural network, however, is only static and is not yet adequate for an application to problems of dynamic nature. The recurrent neural network method [1] is applied to construct a reduced order model resulting from a series of high-fidelity time-dependent data of aero-elastic simulations. Once the RBF neural network ROM is constructed properly, an accurate approximate solution can be obtained at a fraction of the cost of a full-order computation. The method derived during the study has been validated for predicting nonlinear aerodynamic forces in transonic flow and is capable of accurate flutter/LCO simulations. The obtained results indicate that the present recurrent RBF neural network is accurate and efficient for nonlinear aero-elastic system analysis

Computational Results for the KTH-NASA Wind-Tunnel Model Used for Acquisition of Transonic Nonlinear Aeroelastic Data

NASA Technical Reports Server (NTRS)

Silva, Walter A.; Chwalowski, Pawel; Wieseman, Carol D.; Eller, David; Ringertz, Ulf

2017-01-01

A status report is provided on the collaboration between the Royal Institute of Technology (KTH) in Sweden and the NASA Langley Research Center regarding the aeroelastic analyses of a full-span fighter configuration wind-tunnel model. This wind-tunnel model was tested in the Transonic Dynamics Tunnel (TDT) in the summer of 2016. Large amounts of data were acquired including steady/unsteady pressures, accelerations, strains, and measured dynamic deformations. The aeroelastic analyses presented include linear aeroelastic analyses, CFD steady analyses, and analyses using CFD-based reduced-order models (ROMs).

Modeling Programs Increase Aircraft Design Safety

NASA Technical Reports Server (NTRS)

2012-01-01

Flutter may sound like a benign word when associated with a flag in a breeze, a butterfly, or seaweed in an ocean current. When used in the context of aerodynamics, however, it describes a highly dangerous, potentially deadly condition. Consider the case of the Lockheed L-188 Electra Turboprop, an airliner that first took to the skies in 1957. Two years later, an Electra plummeted to the ground en route from Houston to Dallas. Within another year, a second Electra crashed. In both cases, all crew and passengers died. Lockheed engineers were at a loss as to why the planes wings were tearing off in midair. For an answer, the company turned to NASA s Transonic Dynamics Tunnel (TDT) at Langley Research Center. At the time, the newly renovated wind tunnel offered engineers the capability of testing aeroelastic qualities in aircraft flying at transonic speeds near or just below the speed of sound. (Aeroelasticity is the interaction between aerodynamic forces and the structural dynamics of an aircraft or other structure.) Through round-the-clock testing in the TDT, NASA and industry researchers discovered the cause: flutter. Flutter occurs when aerodynamic forces acting on a wing cause it to vibrate. As the aircraft moves faster, certain conditions can cause that vibration to multiply and feed off itself, building to greater amplitudes until the flutter causes severe damage or even the destruction of the aircraft. Flutter can impact other structures as well. Famous film footage of the Tacoma Narrows Bridge in Washington in 1940 shows the main span of the bridge collapsing after strong winds generated powerful flutter forces. In the Electra s case, faulty engine mounts allowed a type of flutter known as whirl flutter, generated by the spinning propellers, to transfer to the wings, causing them to vibrate violently enough to tear off. Thanks to the NASA testing, Lockheed was able to correct the Electra s design flaws that led to the flutter conditions and return the

The NASA Langley 8-foot Transonic Pressure Tunnel calibration

NASA Technical Reports Server (NTRS)

Brooks, Cuyler W., Jr.; Harris, Charles D.; Reagon, Patricia G.

1994-01-01

The NASA Langley 8-Foot Transonic Pressure Tunnel is a continuous-flow, variable-pressure wind tunnel with control capability to independently vary Mach number, stagnation pressure, stagnation temperature, and humidity. The top and bottom walls of the test section are axially slotted to permit continuous variation of the test section Mach number from 0.2 to 1.2, the slot-width contour provides a gradient-free test section 50 in. long for Mach numbers equal to or greater than 1.0 and 100 in. long for Mach numbers less than 1.0. The stagnation pressure may be varied from 0.25 to 2.0 atm. The tunnel test section has been recalibrated to determine the relationship between the free-stream Mach number and the test chamber reference Mach number. The hardware was the same as that of an earlier calibration in 1972 but the pressure measurement instrumentation available for the recalibration was about an order of magnitude more precise. The principal result of the recalibration was a slightly different schedule of reentry flap settings for Mach numbers from 0.80 to 1.05 than that determined during the 1972 calibration. Detailed tunnel contraction geometry, test section geometry, and limited test section wall boundary layer data are presented.

Wind-tunnel evaluation of NASA developed control laws for flutter suppression on a DC-10 derivative wing

NASA Technical Reports Server (NTRS)

Abel, I.; Newsom, J. R.

1981-01-01

Two flutter suppression control laws were synthesized, implemented, and tested on a low speed aeroelastic wing model of a DC-10 derivative. The methodology used to design the control laws is described. Both control laws demonstrated increases in flutter speed in excess of 25 percent above the passive wing flutter speed. The effect of variations in gain and phase on the closed loop performance was measured and compared with analytical predictions. The analytical results are in good agreement with experimental data.

Status of wing flutter

NASA Technical Reports Server (NTRS)

Kussner, H G

1936-01-01

This report presents a survey of previous theoretical and experimental investigations on wing flutter covering thirteen cases of flutter observed on airplanes. The direct cause of flutter is, in the majority of cases, attributable to (mass-) unbalanced ailerons. Under the conservative assumption that the flutter with the phase angle most favorable for excitation occurs only in two degrees of freedom, the lowest critical speed can be estimated from the data obtained on the oscillation bench. Corrective measures for increasing the critical speed and for definite avoidance of wing flutter, are discussed.

A vapor generator for transonic flow visualization

NASA Technical Reports Server (NTRS)

Bruce, Robert A.; Hess, Robert W.; Rivera, Jose A., Jr.

1989-01-01

A vapor generator was developed for use in the NASA Langley Transonic Dynamics Tunnel (TDT). Propylene glycol was used as the vapor material. The vapor generator system was evaluated in a laboratory setting and then used in the TDT as part of a laser light sheet flow visualization system. The vapor generator provided satisfactory seeding of the air flow with visible condensate particles, smoke, for tests ranging from low subsonic through transonic speeds for tunnel total pressures from atmospheric pressure down to less than 0.1 atmospheric pressure.

Aerodynamic Modeling of Transonic Aircraft Using Vortex Lattice Coupled with Transonic Small Disturbance for Conceptual Design

NASA Technical Reports Server (NTRS)

Chaparro, Daniel; Fujiwara, Gustavo E. C.; Ting, Eric; Nguyen, Nhan

2016-01-01

The need to rapidly scan large design spaces during conceptual design calls for computationally inexpensive tools such as the vortex lattice method (VLM). Although some VLM tools, such as Vorview have been extended to model fully-supersonic flow, VLM solutions are typically limited to inviscid, subcritical flow regimes. Many transport aircraft operate at transonic speeds, which limits the applicability of VLM for such applications. This paper presents a novel approach to correct three-dimensional VLM through coupling of two-dimensional transonic small disturbance (TSD) solutions along the span of an aircraft wing in order to accurately predict transonic aerodynamic loading and wave drag for transport aircraft. The approach is extended to predict flow separation and capture the attenuation of aerodynamic forces due to boundary layer viscosity by coupling the TSD solver with an integral boundary layer (IBL) model. The modeling framework is applied to the NASA General Transport Model (GTM) integrated with a novel control surface known as the Variable Camber Continuous Trailing Edge Flap (VCCTEF).

Measurements of Flow Turbulence in the NASA Langley Transonic Dynamics Tunnel

NASA Technical Reports Server (NTRS)

Wiesman, Carol D.; Sleeper, Robert K.

2005-01-01

An assessment of the flow turbulence in the NASA Langley Transonic Dynamics Tunnel (TDT) was conducted during calibration activities following the facility conversion from a Freon-12 heavy-gas test medium to an R134a heavy-gas test medium. Total pressure, static pressure, and acoustic pressure levels were measured at several locations on a stingmounted rake. The test measured wall static pressures at several locations although this paper presents only those from one location. The test used two data acquisition systems, one sampling at 1000 Hz and the second sampling at 125 000 Hz, for acquiring time-domain data. This paper presents standard deviations and power spectral densities of the turbulence points throughout the wind tunnel envelope in air and R134a. The objective of this paper is to present the turbulence characteristics for the test section. No attempt is made to assess the causes of the turbulence. The present paper looks at turbulence in terms of pressure fluctuations. Reference 1 looked at tunnel turbulence in terms of velocity fluctuations.

Transonic aeroelastic analysis of launch vehicle configurations. Ph.D. Thesis

NASA Technical Reports Server (NTRS)

Filgueirasdeazevedo, Joao Luiz

1988-01-01

A numerical study of the aeroelastic stability of typical launch vehicle configurations in transonic flight is performed. Recent computational fluid dynamics techniques are used to simulate the transonic aerodynamic flow fields, as opposed to relying on experimental data for the unsteady aerodynamic pressures. The flow solver is coupled to an appropriate structural representation of the vehicle. The aerodynamic formulation is based on the thin layer approximation to the Reynolds-Averaged Navier-Stokes equations, where the account for turbulent mixing is done by the two-layer Baldwin and Lomax algebraic eddy viscosity model. The structural-dynamic equations are developed considering free-free flexural vibration of an elongated beam with variable properties and are cast in modal form. Aeroelastic analyses are performed by integrating simultaneously in the two sets of equations. By tracing the growth or decay of a perturbed oscillation, the aeroelastic stability of a given constant configuration can be ascertained. The method is described in detail, and results that indicate its application are presented. Applications include some validation cases for the algorithm developed, as well as the study of configurations known to have presented flutter programs in the past.

Flutter analysis using transversality theory

NASA Technical Reports Server (NTRS)

Afolabi, D.

1993-01-01

A new method of calculating flutter boundaries of undamped aeronautical structures is presented. The method is an application of the weak transversality theorem used in catastrophe theory. In the first instance, the flutter problem is cast in matrix form using a frequency domain method, leading to an eigenvalue matrix. The characteristic polynomial resulting from this matrix usually has a smooth dependence on the system's parameters. As these parameters change with operating conditions, certain critical values are reached at which flutter sets in. Our approach is to use the transversality theorem in locating such flutter boundaries using this criterion: at a flutter boundary, the characteristic polynomial does not intersect the axis of the abscissa transversally. Formulas for computing the flutter boundaries and flutter frequencies of structures with two degrees of freedom are presented, and extension to multi-degree of freedom systems is indicated. The formulas have obvious applications in, for instance, problems of panel flutter at supersonic Mach numbers.

Wind tunnel wall interference investigations in NAE/NRC High Reynolds Number 2D Facility and NASA Langley 0.3m Transonic Cryogenic Tunnel

NASA Technical Reports Server (NTRS)

Chan, Y. Y.; Nishimura, Y.; Mineck, R. E.

1989-01-01

Results are reported from a NAE/NRC and NASA cooperative program on two-dimensional wind-tunnel wall-interference research, aimed at developing the technology for correcting or eliminating wall interference effects in two-dimensional transonic wind-tunnel investigations. Both NASA Langley and NAE facilities are described, along with a NASA-designed and fabricated airfoil model. It is shown that data from the NAE facility, corrected for wall interference, agree with those obtained from the NASA tunnel, which has adaptive walls; the comparison of surface pressure data shows that the flowfield conditions in which the model is investigated appear to be nearly identical under most conditions. It is concluded that both approaches, posttest correction and an adaptive wall, adequately eliminate the tunnel-wall interference effects.

Aerodynamic performance and pressure distributions for a NASA SC(2)-0714 airfoil tested in the Langley 0.3-meter transonic cryogenic tunnel

NASA Technical Reports Server (NTRS)

Jenkins, Renaldo V.; Hill, Acquilla S.; Ray, Edward J.

1988-01-01

This report presents in graphic and tabular forms the aerodynamic coefficient and surface pressure distribution data for a NASA SC(2)-0714 airfoil tested in the Langley 0.3-Meter Transonic Cryogenic Tunnel. The test was another in a series of tests involved in the joint NASA/U.S. Industry Advanced Technology Airfoil Tests program. This 14% thick supercritical airfoil was tested at Mach numbers from 0.6 to 0.76 and angles of attack from -2.0 to 6.0 degrees. The test Reynolds numbers were 4 million, 6 million, 10 million, 15 million, 30 million, 40 million, and 45 million.

Multirate flutter suppression system design for the Benchmark Active Controls Technology Wing

NASA Technical Reports Server (NTRS)

Berg, Martin C.; Mason, Gregory S.

1994-01-01

To study the effectiveness of various control system design methodologies, the NASA Langley Research Center initiated the Benchmark Active Controls Project. In this project, the various methodologies will be applied to design a flutter suppression system for the Benchmark Active Controls Technology (BACT) Wing (also called the PAPA wing). Eventually, the designs will be implemented in hardware and tested on the BACT wing in a wind tunnel. This report describes a project at the University of Washington to design a multirate flutter suppression system for the BACT wing. The objective of the project was two fold. First, to develop a methodology for designing robust multirate compensators, and second, to demonstrate the methodology by applying it to the design of a multirate flutter suppression system for the BACT wing. The contributions of this project are (1) development of an algorithm for synthesizing robust low order multirate control laws (the algorithm is capable of synthesizing a single compensator which stabilizes both the nominal plant and multiple plant perturbations; (2) development of a multirate design methodology, and supporting software, for modeling, analyzing and synthesizing multirate compensators; and (3) design of a multirate flutter suppression system for NASA's BACT wing which satisfies the specified design criteria. This report describes each of these contributions in detail. Section 2.0 discusses our design methodology. Section 3.0 details the results of our multirate flutter suppression system design for the BACT wing. Finally, Section 4.0 presents our conclusions and suggestions for future research. The body of the report focuses primarily on the results. The associated theoretical background appears in the three technical papers that are included as Attachments 1-3. Attachment 4 is a user's manual for the software that is key to our design methodology.

Aeroelastic Calculations Using CFD for a Typical Business Jet Model

NASA Technical Reports Server (NTRS)

Gibbons, Michael D.

1996-01-01

Two time-accurate Computational Fluid Dynamics (CFD) codes were used to compute several flutter points for a typical business jet model. The model consisted of a rigid fuselage with a flexible semispan wing and was tested in the Transonic Dynamics Tunnel at NASA Langley Research Center where experimental flutter data were obtained from M(sub infinity) = 0.628 to M(sub infinity) = 0.888. The computational results were computed using CFD codes based on the inviscid TSD equation (CAP-TSD) and the Euler/Navier-Stokes equations (CFL3D-AE). Comparisons are made between analytical results and with experiment where appropriate. The results presented here show that the Navier-Stokes method is required near the transonic dip due to the strong viscous effects while the TSD and Euler methods used here provide good results at the lower Mach numbers.

Flutter Analysis for Turbomachinery Using Volterra Series

NASA Technical Reports Server (NTRS)

Liou, Meng-Sing; Yao, Weigang

2014-01-01

The objective of this paper is to describe an accurate and efficient reduced order modeling method for aeroelastic (AE) analysis and for determining the flutter boundary. Without losing accuracy, we develop a reduced order model based on the Volterra series to achieve significant savings in computational cost. The aerodynamic force is provided by a high-fidelity solution from the Reynolds-averaged Navier-Stokes (RANS) equations; the structural mode shapes are determined from the finite element analysis. The fluid-structure coupling is then modeled by the state-space formulation with the structural displacement as input and the aerodynamic force as output, which in turn acts as an external force to the aeroelastic displacement equation for providing the structural deformation. NASA's rotor 67 blade is used to study its aeroelastic characteristics under the designated operating condition. First, the CFD results are validated against measured data available for the steady state condition. Then, the accuracy of the developed reduced order model is compared with the full-order solutions. Finally the aeroelastic solutions of the blade are computed and a flutter boundary is identified, suggesting that the rotor, with the material property chosen for the study, is structurally stable at the operating condition, free of encountering flutter.

Turbine Blade and Endwall Heat Transfer Measured in NASA Glenn's Transonic Turbine Blade Cascade

NASA Technical Reports Server (NTRS)

Giel, Paul W.

2000-01-01

Higher operating temperatures increase the efficiency of aircraft gas turbine engines, but can also degrade internal components. High-pressure turbine blades just downstream of the combustor are particularly susceptible to overheating. Computational fluid dynamics (CFD) computer programs can predict the flow around the blades so that potential hot spots can be identified and appropriate cooling schemes can be designed. Various blade and cooling schemes can be examined computationally before any hardware is built, thus saving time and effort. Often though, the accuracy of these programs has been found to be inadequate for predicting heat transfer. Code and model developers need highly detailed aerodynamic and heat transfer data to validate and improve their analyses. The Transonic Turbine Blade Cascade was built at the NASA Glenn Research Center at Lewis Field to help satisfy the need for this type of data.

Modeling and control study of the NASA 0.3-meter transonic cryogenic tunnel for use with sulfur hexafluoride medium

NASA Technical Reports Server (NTRS)

Balakrishna, S.; Kilgore, W. Allen

1992-01-01

The NASA Langley 0.3-m Transonic Cryogenic Tunnel is to be modified to operate with sulfur hexafluoride gas while retaining its present capability to operate with nitrogen. The modified tunnel will provide high Reynolds number flow on aerodynamic models with two different test gases. The document details a study of the SF6 tunnel performance boundaries, thermodynamic modeling of the tunnel process, nonlinear dynamical simulation of math model to yield tunnel responses, the closed loop control requirements, control laws, and mechanization of the control laws on the microprocessor based controller.

Highly Maneuverable Aircraft Technology (HiMAT) flight-flutter test program

NASA Technical Reports Server (NTRS)

Kehoe, M. W.

1984-01-01

The highly maneuverable aircraft technology (HiMAT) vehicle was evaluated in a joint NASA and Air Force flight test program. The HiMAT vehicle is a remotely piloted research vehicle. Its design incorporates the use of advanced composite materials in the wings, and canards for aeroelastic tailoring. A flight-flutter test program was conducted to clear a sufficient flight envelope to allow for performance, stability and control, and loads testing. Testing was accomplished with and without flight control-surface dampers. Flutter clearance of the vehicle indicated satisfactory damping and damping trends for the structural modes of the HiMAT vehicle. The data presented include frequency and damping plotted as a function of Mach number.

Final design and fabrication of an active control system for flutter suppression on a supercritical aeroelastic research wing

NASA Technical Reports Server (NTRS)

Hodges, G. E.; Mcgehee, C. R.

1981-01-01

The final design and hardware fabrication was completed for an active control system capable of the required flutter suppression, compatible with and ready for installation in the NASA aeroelastic research wing number 1 (ARW-1) on Firebee II drone flight test vehicle. The flutter suppression system uses vertical acceleration at win buttock line 1.930 (76), with fuselage vertical and roll accelerations subtracted out, to drive wing outboard aileron control surfaces through appropriate symmetric and antisymmetric shaping filters. The goal of providing an increase of 20 percent above the unaugmented vehicle flutter velocity but below the maximum operating condition at Mach 0.98 is exceeded by the final flutter suppression system. Results indicate that the flutter suppression system mechanical and electronic components are ready for installation on the DAST ARW-1 wing and BQM-34E/F drone fuselage.

Reynolds Number Effects on a Supersonic Transport at Transonic Conditions

NASA Technical Reports Server (NTRS)

Wahls, R. N.; Owens, L. R.; Rivers, S. M. B.

2001-01-01

A High Speed Civil Transport configuration was tested in the National Transonic Facility at the NASA Langley Research Center as part of NASA's High Speed Research Program. The primary purposes of the tests were to assess Reynolds number scale effects and the high Reynolds number aerodynamic characteristics of a realistic, second generation supersonic transport while providing data for the assessment of computational methods. The tests included longitudinal and lateral/directional studies at low speed high-lift and transonic conditions across a range of Reynolds numbers from that available in conventional wind tunnels to near flight conditions. Results are presented which focus on both the Reynolds number and static aeroelastic sensitivities of longitudinal characteristics at Mach 0.90 for a configuration without an empennage.

Numerical design of streamlined tunnel walls for a two-dimensional transonic test

NASA Technical Reports Server (NTRS)

Newman, P. A.; Anderson, E. C.

1978-01-01

An analytical procedure is discussed for designing wall shapes for streamlined, nonporous, two-dimensional, transonic wind tunnels. It is based upon currently available 2-D inviscid transonic and boundary layer analysis computer programs. Predicted wall shapes are compared with experimental data obtained from the NASA Langley 6 by 19 inch Transonic Tunnel where the slotted walls were replaced by flexible nonporous walls. Comparisons are presented for the empty tunnel operating at a Mach number of 0.9 and for a supercritical test of an NACA 0012 airfoil at zero lift. Satisfactory agreement is obtained between the analytically and experimentally determined wall shapes.

Aeroservoelastic Modeling of Body Freedom Flutter for Control System Design

NASA Technical Reports Server (NTRS)

Ouellette, Jeffrey

2017-01-01

The communication of this method is being used by NASA in the ongoing collaborations with groups interested in the X-56A flight test program. Model generation for body freedom flutter Addressing issues in: State Consistency, Low frequency dynamics, Unsteady aerodynamics. Applied approach to X-56A MUTT: Comparing to flight test data.

Flutter Research on Skin Panels

NASA Technical Reports Server (NTRS)

Kordes, Eldon E.; Tuovila, Weimer J.; Guy, Lawrence D.

1960-01-01

Representative experimental results are presented to show the current status of the panel flutter problem. Results are presented for unstiffened rectangular panels and for rectangular panels stiffened by corrugated backing. Flutter boundaries are established for all types of panels when considered on the basis of equivalent isotropic plates. The effects of Mach number, differential pressure, and aerodynamic heating on panel flutter are discussed. A flutter analysis of orthotropic panels is presented in the appendix.

Flutter suppression control law synthesis for the Active Flexible Wing model

NASA Technical Reports Server (NTRS)

Mukhopadhyay, Vivek; Perry, Boyd, III; Noll, Thomas E.

1989-01-01

The Active Flexible Wing Project is a collaborative effort between the NASA Langley Research Center and Rockwell International. The objectives are the validation of methodologies associated with mathematical modeling, flutter suppression control law development and digital implementation of the control system for application to flexible aircraft. A flutter suppression control law synthesis for this project is described. The state-space mathematical model used for the synthesis included ten flexible modes, four control surface modes and rational function approximation of the doublet-lattice unsteady aerodynamics. The design steps involved developing the full-order optimal control laws, reducing the order of the control law, and optimizing the reduced-order control law in both the continuous and the discrete domains to minimize stochastic response. System robustness was improved using singular value constraints. An 8th order robust control law was designed to increase the symmetric flutter dynamic pressure by 100 percent. Preliminary results are provided and experiences gained are discussed.

Body-freedom flutter of a 1/2-scale forward-swept-wing model, an experimental and analytical study

NASA Technical Reports Server (NTRS)

Chipman, R.; Rauch, F.; Rimer, M.; Muniz, B.

1984-01-01

The aeroelastic phenomenon known as body-freedom flutter (BFF), a dynamic instability involving aircraft-pitch and wing-bending motions which, though rarely experienced on conventional vehicles, is characteristic of forward swept wing (FSW) aircraft was investigated. Testing was conducted in the Langley transonic dynamics tunnel on a flying, cable-mounted, 1/2-scale model of a FSW configuration with and without relaxed static stability (RSS). The BFF instability boundaries were found to occur at significantly lower airspeeds than those associated with aeroelastic wing divergence on the same model. For those cases with RSS, a canard-based stability augmentation system (SAS) was incorporated in the model. This SAS was designed using aerodynamic data measured during a preliminary tunnel test in which the model was attached to a force balance. Data from the subsequent flutter test indicated that BFF speed was not dependent on open-loop static margin but, rather, on the equivalent closed-loop dynamics provided by the SAS. Servo-aeroelastic stability analyses of the flying model were performed using a computer code known as SEAL and predicted the onset of BFF reasonably well.

Transonic Unsteady Aerodynamics and Aeroelasticity Held in San Diego, California on 7-11 October 1991 (L’Aerodynamique Instationnaire Transsonique el l’Aeroelasticite)

DTIC Science & Technology

1992-03-01

Specific goals were: these conditions was evident at the 1977 meeting in the shock oscillation calculations of Magnus and Yoshihara (Ref 7) and* Application... Magnus . for the minimum flutter speed at M = 0.96 and a favorable, - A case of self-excited shock oscillation about a 14% though unconservarive, agreement...Computational Test cases for Type 11 flows should be es- 7. Magnus , R. and Yoshihara, H.: The Transonic Oscillating tablished (unsteady separating and

Assessment of the National Transonic Facility for Laminar Flow Testing

NASA Technical Reports Server (NTRS)

Crouch, Jeffrey D.; Sutanto, Mary I.; Witkowski, David P.; Watkins, A. Neal; Rivers, Melissa B.; Campbell, Richard L.

2010-01-01

A transonic wing, designed to accentuate key transition physics, is tested at cryogenic conditions at the National Transonic Facility at NASA Langley. The collaborative test between Boeing and NASA is aimed at assessing the facility for high-Reynolds number testing of configurations with significant regions of laminar flow. The test shows a unit Reynolds number upper limit of 26 M/ft for achieving natural transition. At higher Reynolds numbers turbulent wedges emanating from the leading edge bypass the natural transition process and destroy the laminar flow. At lower Reynolds numbers, the transition location is well correlated with the Tollmien-Schlichting-wave N-factor. The low-Reynolds number results suggest that the flow quality is acceptable for laminar flow testing if the loss of laminar flow due to bypass transition can be avoided.

Aeroservoelastic Wind-Tunnel Test of the SUGAR Truss Braced Wing Wind-Tunnel Model

NASA Technical Reports Server (NTRS)

Scott, Robert C.; Allen, Timothy J.; Funk, Christie J.; Castelluccio, Mark A.; Sexton, Bradley W.; Claggett, Scott; Dykman, John; Coulson, David A.; Bartels, Robert E.

2015-01-01

The Subsonic Ultra Green Aircraft Research (SUGAR) Truss-Braced Wing (TBW) aeroservoelastic (ASE) wind-tunnel test was conducted in the NASA Langley Transonic Dynamics Tunnel (TDT) and was completed in April, 2014. The primary goals of the test were to identify the open-loop flutter boundary and then demonstrate flutter suppression. A secondary goal was to demonstrate gust load alleviation (GLA). Open-loop flutter and limit cycle oscillation onset boundaries were identified for a range of Mach numbers and various angles of attack. Two sets of control laws were designed for the model and both sets of control laws were successful in suppressing flutter. Control laws optimized for GLA were not designed; however, the flutter suppression control laws were assessed using the TDT Airstream Oscillation System. This paper describes the experimental apparatus, procedures, and results of the TBW wind-tunnel test. Acquired system ID data used to generate ASE models is also discussed.2 study.

Multirate Flutter Suppression System Design for the Benchmark Active Controls Technology Wing. Part 2; Methodology Application Software Toolbox

NASA Technical Reports Server (NTRS)

Mason, Gregory S.; Berg, Martin C.; Mukhopadhyay, Vivek

2002-01-01

To study the effectiveness of various control system design methodologies, the NASA Langley Research Center initiated the Benchmark Active Controls Project. In this project, the various methodologies were applied to design a flutter suppression system for the Benchmark Active Controls Technology (BACT) Wing. This report describes the user's manual and software toolbox developed at the University of Washington to design a multirate flutter suppression control law for the BACT wing.

Flight Reynolds Number Testing of the Orion Launch Abort Vehicle in the NASA Langley National Transonic Facility

NASA Technical Reports Server (NTRS)

Chan, David T.; Brauckmann, Gregory J.

2011-01-01

A 6%-scale unpowered model of the Orion Launch Abort Vehicle (LAV) ALAS-11-rev3c configuration was tested in the NASA Langley National Transonic Facility to obtain static aerodynamic data at flight Reynolds numbers. Subsonic and transonic data were obtained for Mach numbers between 0.3 and 0.95 for angles of attack from -4 to +22 degrees and angles of sideslip from -10 to +10 degrees. Data were also obtained at various intermediate Reynolds numbers between 2.5 million and 45 million depending on Mach number in order to examine the effects of Reynolds number on the vehicle. Force and moment data were obtained using a 6-component strain gauge balance that operated both at warm temperatures (+120 . F) and cryogenic temperatures (-250 . F). Surface pressure data were obtained with electronically scanned pressure units housed in heated enclosures designed to survive cryogenic temperatures. Data obtained during the 3-week test entry were used to support development of the LAV aerodynamic database and to support computational fluid dynamics code validation. Furthermore, one of the outcomes of the test was the reduction of database uncertainty on axial force coefficient for the static unpowered LAV. This was accomplished as a result of good data repeatability throughout the test and because of decreased uncertainty on scaling wind tunnel data to flight.

Parametric Evaluation of Thin, Transonic Circulation-Control Airfoils

NASA Technical Reports Server (NTRS)

Schlecht, Robin; Anders, Scott

2007-01-01

Wind-tunnel tests were conducted in the NASA Langley Transonic Dynamics Tunnel on a 6 percent-thick, elliptical circulation-control airfoil with upper-surface and lower-surface blowing capability. Results for elliptical Coanda trailing-edge geometries, biconvex Coanda trailing-edge geometries, and leading-edge geometries are reported. Results are presented at subsonic and transonic Mach numbers of 0.3 and 0.8, respectively. When considering one fixed trailing-edge geometry, for both the subsonic and transonic conditions it was found that the [3.0:1] ratio elliptical Coanda surface with the most rounded leading-edge [03] performed favorably and was determined to be the best compromise between comparable configurations that took advantage of the Coanda effect. This configuration generated a maximum. (Delta)C(sub 1) = 0.625 at a C(sub mu) = 0.06 at M = 0.3, alpha = 6deg. This same configuration generated a maximum (Delta)C(sub 1) = 0.275 at a C(sub mu) = 0.0085 at M = 0.8, alpha = 3deg.

Numerical Investigations of the Benchmark Supercritical Wing in Transonic Flow

NASA Technical Reports Server (NTRS)

Chwalowski, Pawel; Heeg, Jennifer; Biedron, Robert T.

2017-01-01

This paper builds on the computational aeroelastic results published previously and generated in support of the second Aeroelastic Prediction Workshop for the NASA Benchmark Supercritical Wing (BSCW) configuration. The computational results are obtained using FUN3D, an unstructured grid Reynolds-Averaged Navier-Stokes solver developed at the NASA Langley Research Center. The analysis results show the effects of the temporal and spatial resolution, the coupling scheme between the flow and the structural solvers, and the initial excitation conditions on the numerical flutter onset. Depending on the free stream condition and the angle of attack, the above parameters do affect the flutter onset. Two conditions are analyzed: Mach 0.74 with angle of attack 0 and Mach 0.85 with angle of attack 5. The results are presented in the form of the damping values computed from the wing pitch angle response as a function of the dynamic pressure or in the form of dynamic pressure as a function of the Mach number.

Technical activities of the configuration aeroelasticity branch

NASA Technical Reports Server (NTRS)

Cole, Stanley R. (Editor)

1991-01-01

A number of recent technical activities of the Configuration Aeroelasticity Branch of the NASA Langley Research Center are discussed in detail. The information on the research branch is compiled in twelve separate papers. The first of these topics is a summary of the purpose of the branch, including a full description of the branch and its associated projects and program efforts. The next ten papers cover specific projects and are as follows: Experimental transonic flutter characteristics of supersonic cruise configurations; Aeroelastic effects of spoiler surfaces mounted on a low aspect ratio rectangular wing; Planform curvature effects on flutter of 56 degree swept wing determined in Transonic Dynamics Tunnel (TDT); An introduction to rotorcraft testing in TDT; Rotorcraft vibration reduction research at the TDT; A preliminary study to determine the effects of tip geometry on the flutter of aft swept wings; Aeroelastic models program; NACA 0012 pressure model and test plan; Investigation of the use of extension twist coupling in composite rotor blades; and Improved finite element methods for rotorcraft structures. The final paper describes the primary facility operation by the branch, the Langley TDT.

Physical properties of the benchmark models program supercritical wing

NASA Technical Reports Server (NTRS)

Dansberry, Bryan E.; Durham, Michael H.; Bennett, Robert M.; Turnock, David L.; Silva, Walter A.; Rivera, Jose A., Jr.

1993-01-01

The goal of the Benchmark Models Program is to provide data useful in the development and evaluation of aeroelastic computational fluid dynamics (CFD) codes. To that end, a series of three similar wing models are being flutter tested in the Langley Transonic Dynamics Tunnel. These models are designed to simultaneously acquire model response data and unsteady surface pressure data during wing flutter conditions. The supercritical wing is the second model of this series. It is a rigid semispan model with a rectangular planform and a NASA SC(2)-0414 supercritical airfoil shape. The supercritical wing model was flutter tested on a flexible mount, called the Pitch and Plunge Apparatus, that provides a well-defined, two-degree-of-freedom dynamic system. The supercritical wing model and associated flutter test apparatus is described and experimentally determined wind-off structural dynamic characteristics of the combined rigid model and flexible mount system are included.

National Transonic Facility: A review of the operational plan

NASA Technical Reports Server (NTRS)

Liepmann, H. W.; Black, R. E.; Dietz, R. O.; Kirchner, M. E.; Sears, W. R.

1980-01-01

The proposed National Transonic Facility (NTF) operational plan is reviewed. The NTF will provide an aerodynamic test capability significantly exceeding that of other transonic regime wind tunnels now available. A limited number of academic research program that might use the NTF are suggested. It is concluded that the NTF operational plan is useful for management, technical, instrumentation, and model building techniques available in the specialized field of aerodynamic analysis and simulation. It is also suggested that NASA hold an annual conference to discuss wind tunnel research results and to report on developments that will further improve the utilization and cost effectiveness of the NTF and other wind tunnels.

Predicting Flutter and Forced Response in Turbomachinery

NASA Technical Reports Server (NTRS)

VanZante, Dale E.; Adamczyk, John J.; Srivastava, Rakesh; Bakhle, Milind A.; Shabbir, Aamir; Chen, Jen-Ping; Janus, J. Mark; To, Wai-Ming; Barter, John

2005-01-01

TURBO-AE is a computer code that enables detailed, high-fidelity modeling of aeroelastic and unsteady aerodynamic characteristics for prediction of flutter, forced response, and blade-row interaction effects in turbomachinery. Flow regimes that can be modeled include subsonic, transonic, and supersonic, with attached and/or separated flow fields. The three-dimensional Reynolds-averaged Navier-Stokes equations are solved numerically to obtain extremely accurate descriptions of unsteady flow fields in multistage turbomachinery configurations. Blade vibration is simulated by use of a dynamic-grid-deformation technique to calculate the energy exchange for determining the aerodynamic damping of vibrations of blades. The aerodynamic damping can be used to assess the stability of a blade row. TURBO-AE also calculates the unsteady blade loading attributable to such external sources of excitation as incoming gusts and blade-row interactions. These blade loadings, along with aerodynamic damping, are used to calculate the forced responses of blades to predict their fatigue lives. Phase-lagged boundary conditions based on the direct-store method are used to calculate nonzero interblade phase-angle oscillations; this practice eliminates the need to model multiple blade passages, and, hence, enables large savings in computational resources.

Transonic Semispan Aerodynamic Testing of the Hybrid Wing Body with Over Wing Nacelles in the National Transonic Facility

NASA Technical Reports Server (NTRS)

Chan, David T.; Hooker, John R.; Wick, Andrew; Plumley, Ryan W.; Zeune, Cale H.; Ol, Michael V.; DeMoss, Joshua A.

2017-01-01

A wind tunnel investigation of a 0.04-scale model of the Lockheed Martin Hybrid Wing Body (HWB) with Over Wing Nacelles (OWN) air mobility transport configuration was conducted in the National Transonic Facility at the NASA Langley Research Center under a collaborative partnership between NASA, the Air Force Research Laboratory, and Lockheed Martin Aeronautics Company. The wind tunnel test sought to validate the transonic aerodynamic performance of the HWB and to validate the efficiency benefits of the OWN installation as compared to the traditional under-wing installation. The semispan HWB model was tested in a clean wing configuration and also tested with two different nacelles representative of a modern turbofan engine and a future advanced high bypass ratio engine. The nacelles were installed in three different locations with two over-wing positions and one under-wing position. Five-component force and moment data, surface static pressure data, and aeroelastic deformation data were acquired. For the cruise configuration, the model was tested in an angle-of-attack range between -2 and 10 degrees at free-stream Mach numbers from 0.3 to 0.9 and at unit Reynolds numbers between 8 and 39 million per foot, achieving a maximum of 80% of flight Reynolds numbers across the Mach number range. The test results validated pretest computational fluid dynamic (CFD) simulations of the HWB performance including the OWN benefit and the results also exhibited excellent transonic drag data repeatability to within +/-1 drag count. This paper details the experimental setup and model overview, presents some sample data results, and describes the facility improvements that led to the success of the test.

Adjoint Method and Predictive Control for 1-D Flow in NASA Ames 11-Foot Transonic Wind Tunnel

NASA Technical Reports Server (NTRS)

Nguyen, Nhan; Ardema, Mark

2006-01-01

This paper describes a modeling method and a new optimal control approach to investigate a Mach number control problem for the NASA Ames 11-Foot Transonic Wind Tunnel. The flow in the wind tunnel is modeled by the 1-D unsteady Euler equations whose boundary conditions prescribe a controlling action by a compressor. The boundary control inputs to the compressor are in turn controlled by a drive motor system and an inlet guide vane system whose dynamics are modeled by ordinary differential equations. The resulting Euler equations are thus coupled to the ordinary differential equations via the boundary conditions. Optimality conditions are established by an adjoint method and are used to develop a model predictive linear-quadratic optimal control for regulating the Mach number due to a test model disturbance during a continuous pitch

Investigation of Seal-to-Floor Effects on Semi-Span Transonic Models

NASA Technical Reports Server (NTRS)

Sleppy, Mark A.; Engel, Eric A.; Watson, Kevin T.; Atler, Douglas M.

2009-01-01

In an effort to achieve the maximum possible Reynolds number (Re) when conducting production testing for flight loads aerodynamic databases, it has been the preferred practice of The Boeing Company / Commercial Airplanes (BCA) -- Loads and Dynamics Group since the early 1990's to test large scale semi-span models in the 11- By 11-Foot Transonic Wind Tunnel (TWT) leg of the Unitary Plan Wind Tunnel (UPWT) at the NASA Ames Research Center (ARC). There are many problems related to testing large scale semi-span models of high aspect ratio flexible transport wings, such as; floor boundary layer effects, wing spanwise wall effects, solid blockage buoyancy effects, floor mechanical interference effects, airflow under the model effects, or tunnel flow gradient effects. For most of these issues, BCA has developed and implemented either standard testing methods or numerical correction schemes and these will not be discussed in this document. Other researchers have reported on semi-span transonic testing correction issues, however most of the reported research has been for low Mach testing. Some of the reports for low Mach testing address the difficult problem of preventing undesirable airflow under a semi-span model while ensuring unrestricted main balance functionality, however, for transonic models this issue has gone unresolved. BCA has been cognizant for sometime that there are marked differences in wing pressure distributions from semi-span transonic model testing than from full model or flight testing. It has been suspected that these differences are at least in part due to airflow under the model. Previous efforts by BCA to address this issue have proven to be ineffective or inconclusive and in one situation resulted in broken hardware. This paper reports on a Boeing-NASA collaborative investigation based on a series of small tests conducted between June 2006 and November 2007 in the 11 by 11 foot Transonic Wind Tunnel at NASA Ames on three large commercial jet

Propfan test assessment testbed aircraft flutter model test report

NASA Technical Reports Server (NTRS)

Jenness, C. M. J.

1987-01-01

The PropFan Test Assessment (PTA) program includes flight tests of a propfan power plant mounted on the left wind of a modified Gulfstream II testbed aircraft. A static balance boom is mounted on the right wing tip for lateral balance. Flutter analyses indicate that these installations reduce the wing flutter stabilizing speed and that torsional stiffening and the installation of a flutter stabilizing tip boom are required on the left wing for adequate flutter safety margins. Wind tunnel tests of a 1/9th scale high speed flutter model of the testbed aircraft were conducted. The test program included the design, fabrication, and testing of the flutter model and the correlation of the flutter test data with analysis results. Excellent correlations with the test data were achieved in posttest flutter analysis using actual model properties. It was concluded that the flutter analysis method used was capable of accurate flutter predictions for both the (symmetric) twin propfan configuration and the (unsymmetric) single propfan configuration. The flutter analysis also revealed that the differences between the tested model configurations and the current aircraft design caused the (scaled) model flutter speed to be significantly higher than that of the aircraft, at least for the single propfan configuration without a flutter boom. Verification of the aircraft final design should, therefore, be based on flutter predictions made with the test validated analysis methods.

NASA Aeroelasticity Handbook Volume 2: Design Guides Part 2

NASA Technical Reports Server (NTRS)

Ramsey, John K. (Editor)

2006-01-01

The NASA Aeroelasticity Handbook comprises a database (in three formats) of NACA and NASA aeroelasticity flutter data through 1998 and a collection of aeroelasticity design guides. The Microsoft Access format provides the capability to search for specific data, retrieve it, and present it in a tabular or graphical form unique to the application. The full-text NACA and NASA documents from which the data originated are provided in portable document format (PDF), and these are hyperlinked to their respective data records. This provides full access to all available information from the data source. Two other electronic formats, one delimited by commas and the other by spaces, are provided for use with other software capable of reading text files. To the best of the author s knowledge, this database represents the most extensive collection of NACA and NASA flutter data in electronic form compiled to date by NASA. Volume 2 of the handbook contains a convenient collection of aeroelastic design guides covering fixed wings, turbomachinery, propellers and rotors, panels, and model scaling. This handbook provides an interactive database and design guides for use in the preliminary aeroelastic design of aerospace systems and can also be used in validating or calibrating flutter-prediction software.

Flight Flutter Testing of Rotary Wing Aircraft Using a Control System Oscillation Technique

NASA Technical Reports Server (NTRS)

Yen, J. G.; Viswanathan, S.; Matthys, C. G.

1976-01-01

A flight flutter testing technique is described in which the rotor controls are oscillated by series actuators to excite the rotor and airframe modes of interest, which are then allowed to decay. The moving block technique is then used to determine the damped frequency and damping variation with rotor speed. The method proved useful for tracking the stability of relatively well damped modes. The results of recently completed flight tests of an experimental soft-in-plane rotor are used to illustrate the technique. Included is a discussion of the application of this technique to investigation of the propeller whirl flutter stability characteristics of the NASA/Army XV-15 VTOL tilt rotor research aircraft.

Experimental flutter boundaries with unsteady pressure distributions for the NACA 0012 Benchmark Model

NASA Technical Reports Server (NTRS)

Rivera, Jose A., Jr.; Dansberry, Bryan E.; Farmer, Moses G.; Eckstrom, Clinton V.; Seidel, David A.; Bennett, Robert M.

1991-01-01

The Structural Dynamics Div. at NASA-Langley has started a wind tunnel activity referred to as the Benchmark Models Program. The objective is to acquire test data that will be useful for developing and evaluating aeroelastic type Computational Fluid Dynamics codes currently in use or under development. The progress is described which was achieved in testing the first model in the Benchmark Models Program. Experimental flutter boundaries are presented for a rigid semispan model (NACA 0012 airfoil section) mounted on a flexible mount system. Also, steady and unsteady pressure measurements taken at the flutter condition are presented. The pressure data were acquired over the entire model chord located at the 60 pct. span station.

Real-time flutter identification

NASA Technical Reports Server (NTRS)

Roy, R.; Walker, R.

1985-01-01

The techniques and a FORTRAN 77 MOdal Parameter IDentification (MOPID) computer program developed for identification of the frequencies and damping ratios of multiple flutter modes in real time are documented. Physically meaningful model parameterization was combined with state of the art recursive identification techniques and applied to the problem of real time flutter mode monitoring. The performance of the algorithm in terms of convergence speed and parameter estimation error is demonstrated for several simulated data cases, and the results of actual flight data analysis from two different vehicles are presented. It is indicated that the algorithm is capable of real time monitoring of aircraft flutter characteristics with a high degree of reliability.

[Typical atrial flutter: Diagnosis and therapy].

PubMed

Thomas, Dierk; Eckardt, Lars; Estner, Heidi L; Kuniss, Malte; Meyer, Christian; Neuberger, Hans-Ruprecht; Sommer, Philipp; Steven, Daniel; Voss, Frederik; Bonnemeier, Hendrik

2016-03-01

Typical, cavotricuspid-dependent atrial flutter is the most common atrial macroreentry tachycardia. The incidence of atrial flutter (typical and atypical forms) is age-dependent with 5/100,000 in patients less than 50 years and approximately 600/100,000 in subjects > 80 years of age. Concomitant heart failure or pulmonary disease further increases the risk of typical atrial flutter.Patients with atrial flutter may present with symptoms of palpitations, reduced exercise capacity, chest pain, or dyspnea. The risk of thromboembolism is probably similar to atrial fibrillation; therefore, the same antithrombotic prophylaxis is required in atrial flutter patients. Acutely symptomatic cases may be subjected to cardioversion or pharmacologic rate control to relieve symptoms. Catheter ablation of the cavotricuspid isthmus represents the primary choice in long-term therapy, associated with high procedural success (> 97 %) and low complication rates (0.5 %).This article represents the third part of a manuscript series designed to improve professional education in the field of cardiac electrophysiology. Mechanistic and clinical characteristics as well as management of isthmus-dependent atrial flutter are described in detail. Electrophysiological findings and catheter ablation of the arrhythmia are highlighted.

User guide for WIACX: A transonic wind-tunnel wall interference assessment and correction procedure for the NTF

NASA Technical Reports Server (NTRS)

Garriz, Javier A.; Haigler, Kara J.

1992-01-01

A three dimensional transonic Wind-tunnel Interference Assessment and Correction (WIAC) procedure developed specifically for use in the National Transonic Facility (NTF) at NASA Langley Research Center is discussed. This report is a user manual for the codes comprising the correction procedure. It also includes listings of sample procedures and input files for running a sample case and plotting the results.

Flutter suppression via piezoelectric actuation

NASA Technical Reports Server (NTRS)

Heeg, Jennifer

1991-01-01

Experimental flutter results obtained from wind tunnel tests of a two degree of freedom wind tunnel model are presented for the open and closed loop systems. The wind tunnel model is a two degree of freedom system which is actuated by piezoelectric plates configured as bimorphs. The model design was based on finite element structural analyses and flutter analyses. A control law was designed based on a discrete system model; gain feedback of strain measurements was utilized in the control task. The results show a 21 pct. increase in the flutter speed.

Report on a Cooperative Programme on Active Flutter Suppression,

DTIC Science & Technology

1980-08-01

soufflerie 16 pieds au frilon de is NASA A Langley Field. La maquette et les essais 6taient sous la responsabilitg de is firma Northrop (Fig. 1). Du c...ATLANTIC TREATY ORGANIZATION ADVISORY GROUP FOR AEROSPACE RESEARCH AND DEVELOPMENT (ORGANISATION DU TRAITE DE L’ATLANTIQUE NORD) A(ARI) Report No.689 REPORT...power unit was added to the console to provide for start/stop and bypass control of this unit from the wind tunnel control room. THE FLUTTER DETECTOR AND

A New High Channel-Count, High Scan-Rate, Data Acquisition System for the NASA Langley Transonic Dynamics Tunnel

NASA Technical Reports Server (NTRS)

Ivanco, Thomas G.; Sekula, Martin K.; Piatak, David J.; Simmons, Scott A.; Babel, Walter C.; Collins, Jesse G.; Ramey, James M.; Heald, Dean M.

2016-01-01

A data acquisition system upgrade project, known as AB-DAS, is underway at the NASA Langley Transonic Dynamics Tunnel. AB-DAS will soon serve as the primary data system and will substantially increase the scan-rate capabilities and analog channel count while maintaining other unique aeroelastic and dynamic test capabilities required of the facility. AB-DAS is configurable, adaptable, and enables buffet and aeroacoustic tests by synchronously scanning all analog channels and recording the high scan-rate time history values for each data quantity. AB-DAS is currently available for use as a stand-alone data system with limited capabilities while development continues. This paper describes AB-DAS, the design methodology, and the current features and capabilities. It also outlines the future work and projected capabilities following completion of the data system upgrade project.

Wind tunnel-sidewall-boundary-layer effects in transonic airfoil testing-some correctable, but some not

NASA Technical Reports Server (NTRS)

Lynch, F. T.; Johnson, C. B.

1988-01-01

The need to correct transonic airfoil wind tunnel test data for the influence of the tunnel sidewall boundary layers, in addition to the wall accepted corrections for the analytical investigation was carried out in order to evaluate sidewall boundary layer effects on transonic airfoil characteristics, and to validate proposed correction and the limit to their applications. This investigation involved testing of modern airfoil configurations in two different transonic airfoil test facilities, the 15 x 60 inch two-dimensional insert of the National Aeronautical Establishment (NAE) 5 foot tunnel in Ottawa, Canada, and the two-dimensional test section of the NASA Langley 0.3 m Transonic Cryogenic Tunnel (TCT). Results presented included effects of variations in sidewall-boundary layer bleed in both facilities, different sidewall boundary layer correction procedures, tunnel-to tunnel comparisons of correcte results, and flow conditions with and without separation.

Unsteady Aerodynamic Model Tuning for Precise Flutter Prediction

NASA Technical Reports Server (NTRS)

Pak, Chan-Gi

2011-01-01

A simple method for an unsteady aerodynamic model tuning is proposed in this study. This method is based on the direct modification of the aerodynamic influence coefficient matrices. The aerostructures test wing 2 flight-test data is used to demonstrate the proposed model tuning method. The flutter speed margin computed using only the test validated structural dynamic model can be improved using the additional unsteady aerodynamic model tuning, and then the flutter speed margin requirement of 15 % in military specifications can apply towards the test validated aeroelastic model. In this study, unsteady aerodynamic model tunings are performed at two time invariant flight conditions, at Mach numbers of 0.390 and 0.456. When the Mach number for the unsteady model tuning approaches to the measured fluttering Mach number, 0.502, at the flight altitude of 9,837 ft, the estimated flutter speed is approached to the measured flutter speed at this altitude. The minimum flutter speed difference between the estimated and measured flutter speed is -.14 %.

High Reynolds Number tests of the NASA SC(2)-0012 airfoil in the Langley 0.3-meter transonic cryogenic tunnel

NASA Technical Reports Server (NTRS)

Mineck, Raymond E.; Lawing, Pierce L.

1987-01-01

A wind-tunnel investigation of the NASA SC(2)-0012 airfoil has been conducted in the Langley 0.3-Meter Transonic Cryogenic Tunnel. This investigation supplements the two-dimensional airfoil studies of the Advanced Technology Airfoil Test Program. The Mach number was varied from 0.60 to 0.84. The stagnation temperature and pressure were varied to provide a Reynolds number range from 6 to 40 x 10 to the 6th power based on a 6.0-in. (15.24-cm) airfoil chord. No corrections for wind-tunnel wall interference have been made to the data. The aerodynamic results are presented as integrated force and moment coefficients and pressure distributions without any analysis.

Evaluation of diffuse-illumination holographic cinematography in a flutter cascade

NASA Technical Reports Server (NTRS)

Decker, A. J.

1986-01-01

Since 1979, the Lewis Research Center has examined holographic cinematography for three-dimensional flow visualization. The Nd:YAG lasers used were Q-switched, double-pulsed, and frequency-doubled, operating at 20 pulses per second. The primary subjects for flow visualization were the shock waves produced in two flutter cascades. Flow visualization was by diffuse-illumination, double-exposure, and holographic interferometry. The performances of the lasers, holography, and diffuse-illumination interferometry are evaluated in single-window wind tunnels. The fringe-contrast factor is used to evaluate the results. The effects of turbulence on shock-wave visualization in a transonic flow are discussed. The depth of field for visualization of a turbulent structure is demonstrated to be a measure of the relative density and scale of that structure. Other items discussed are the holographic emulsion, tests of coherence and polarization, effects of windows and diffusers, hologram bleaching, laser configurations, influence and handling of specular reflections, modes of fringe localization, noise sources, and coherence requirements as a function of the pulse energy. Holography and diffuse illumination interferometry are also reviewed.

Panel Flutter Emulation Using a Few Concentrated Forces

NASA Astrophysics Data System (ADS)

Dhital, Kailash; Han, Jae-Hung

2018-04-01

The objective of this paper is to study the feasibility of panel flutter emulation using a few concentrated forces. The concentrated forces are considered to be equivalent to aerodynamic forces. The equivalence is carried out using surface spline method and principle of virtual work. The structural modeling of the plate is based on the classical plate theory and the aerodynamic modeling is based on the piston theory. The present approach differs from the linear panel flutter analysis in scheming the modal aerodynamics forces with unchanged structural properties. The solutions for the flutter problem are obtained numerically using the standard eigenvalue procedure. A few concentrated forces were considered with an optimization effort to decide their optimal locations. The optimization process is based on minimizing the error between the flutter bounds from emulated and linear flutter analysis method. The emulated flutter results for the square plate of four different boundary conditions using six concentrated forces are obtained with minimal error to the reference value. The results demonstrated the workability and viability of using concentrated forces in emulating real panel flutter. In addition, the paper includes the parametric studies of linear panel flutter whose proper literatures are not available.

Active flutter suppression using dipole filters

NASA Technical Reports Server (NTRS)

Srinathkumar, S.; Waszak, Martin R.

1992-01-01

By using traditional control concepts of gain root locus, the active suppression of a flutter mode of a flexible wing is examined. It is shown that the attraction of the unstable mode towards a critical system zero determines the degree to which the flutter mode can be stabilized. For control situations where the critical zero is adversely placed in the complex plane, a novel compensation scheme called a 'Dipole' filter is proposed. This filter ensures that the flutter mode is stabilized with acceptable control energy. The control strategy is illustrated by designing flutter suppression laws for an active flexible wing (AFW) wind-tunnel model, where minimal control effort solutions are mandated by control rate saturation problems caused by wind-tunnel turbulence.

Blade Row Interaction Effects on the Performance of a Moderately Loaded NASA Transonic Compressor Stage

NASA Technical Reports Server (NTRS)

VanZante, Dale E.; To, Wai-Ming; Chen, Jen-Ping

2003-01-01

Blade row interaction effects on loss generation in compressors have received increased attention as compressor work-per-stage and blade loading have increased. Two dimensional Laser Doppler Velocimeter measurements of the velocity field in a NASA transonic compressor stage show the magnitude of interactions in the velocity field at the peak efficiency and near stall operating conditions. The experimental data are presented along with an assessment of the velocity field interactions. In the present study the experimental data are used to confirm the fidelity of a three-dimensional, time-accurate, Navier Stokes calculation of the stage using the MSU-TURBO code at the peak efficiency and near stall operating conditions. The simulations are used to quantify the loss generation associated with interaction phenomena. At the design point the stator pressure field has minimal effect on the rotor performance. The rotor wakes do have an impact on loss production in the stator passage at both operating conditions. A method for determining the potential importance of blade row interactions on performance is presented.

Unsteady Aerodynamic Model Tuning for Precise Flutter Prediction

NASA Technical Reports Server (NTRS)

Pak, Chan-gi

2011-01-01

A simple method for an unsteady aerodynamic model tuning is proposed in this study. This method is based on the direct modification of the aerodynamic influence coefficient matrices. The aerostructures test wing 2 flight-test data is used to demonstrate the proposed model tuning method. The flutter speed margin computed using only the test validated structural dynamic model can be improved using the additional unsteady aerodynamic model tuning, and then the flutter speed margin requirement of 15 percent in military specifications can apply towards the test validated aeroelastic model. In this study, unsteady aerodynamic model tunings are performed at two time invariant flight conditions, at Mach numbers of 0.390 and 0.456. When the Mach number for the unsteady aerodynamic model tuning approaches to the measured fluttering Mach number, 0.502, at the flight altitude of 9,837 ft, the estimated flutter speed is approached to the measured flutter speed at this altitude. The minimum flutter speed difference between the estimated and measured flutter speed is -0.14 percent.

Mechanism of Flutter A Theoretical and Experimental Investigation of the Flutter Problem

NASA Technical Reports Server (NTRS)

Theodorsen, Theodore; Garrick, I E

1940-01-01

The results of the basic flutter theory originally devised in 1934 and published as NACA Technical Report no. 496 are presented in a simpler and more complete form convenient for further studies. The paper attempts to facilitate the judgement of flutter problems by a systematic survey of the theoretical effects of the various parameters. A large number of experiments were conducted on cantilever wings, with and without ailerons, in the NACA high-speed wind tunnel for the purpose of verifying the theory and to study its adaptability to three-dimensional problems. The experiments included studies on wing taper ratios, nacelles, attached floats, and external bracings. The essential effects in the transition to the three-dimensional problem have been established. Of particular interest is the existence of specific flutter modes as distinguished from ordinary vibration modes. It is shown that there exists a remarkable agreement between theoretical and experimental results.

On the Application of Contour Bumps for Transonic Drag Reduction(Invited)

NASA Technical Reports Server (NTRS)

Milholen, William E., II; Owens, Lewis R.

2005-01-01

The effect of discrete contour bumps on reducing the transonic drag at off-design conditions on an airfoil have been examined. The research focused on fully-turbulent flow conditions, at a realistic flight chord Reynolds number of 30 million. State-of-the-art computational fluid dynamics methods were used to design a new baseline airfoil, and a family of fixed contour bumps. The new configurations were experimentally evaluated in the 0.3-m Transonic Cryogenic Tunnel at the NASA Langley Research center, which utilizes an adaptive wall test section to minimize wall interference. The computational study showed that transonic drag reduction, on the order of 12% - 15%, was possible using a surface contour bump to spread a normal shock wave. The computational study also indicated that the divergence drag Mach number was increased for the contour bump applications. Preliminary analysis of the experimental data showed a similar contour bump effect, but this data needed to be further analyzed for residual wall interference corrections.

An evaluation of three experimental processes for two-dimensional transonic tests

NASA Technical Reports Server (NTRS)

Zuppardi, Gennaro

1989-01-01

The aerodynamic measurements in conventional wind tunnels usually suffer from the interference effects of the sting supporting the model and the test section walls. These effects are particularly severe in the transonic regime. Sting interference effects can be overcome through the Magnetic Suspension technique. Wall effects can be alleviated by: testing airfoils in conventional, ventilated tunnels at relatively small model to tunnel size ratios; treatment of the tunnel wall boundary layers; or by utilization of the Adaptive Wall Test Section (AWTS) concept. The operating capabilities and results from two of the foremost two-dimensional, transonic, AWTS facilities in existence are assessed. These facilities are the NASA 0.3-Meter Transonic Cryogenic Tunnel and the ONERA T-2 facility located in Toulouse, France. In addition, the results derived from the well known conventional facility, the NAE 5 ft x 5 ft Canadian wind tunnel will be assessed. CAST10/D0A2 Airfoil results will be used in all of the evaluations.

Flight flutter testing of multi-jet aircraft

NASA Technical Reports Server (NTRS)

Bartley, J.

1975-01-01

Extensive flight flutter tests were conducted by BAC on B-52 and KC-135 prototype airplanes. The need for and importance of these flight flutter programs to Boeing airplane design are discussed. Basic concepts of flight flutter testing of multi-jet aircraft and analysis of the test data will be presented. Exciter equipment and instrumentation employed in these tests will be discussed.

Free-To-Roll Analysis of Abrupt Wing Stall on Military Aircraft at Transonic Speeds

NASA Technical Reports Server (NTRS)

Owens, D. Bruce; Capone, Francis J.; Brandon, Jay M.; Cunningham, Kevin; Chambers, Joseph R.

2003-01-01

Transonic free-to-roll and static wind tunnel tests for four military aircraft - the AV-8B, the F/A-18C, the preproduction F/A-18E, and the F-16C - have been analyzed. These tests were conducted in the NASA Langley 16-Foot Transonic Tunnel as a part of the NASA/Navy/Air Force Abrupt Wing Stall Program. The objectives were to evaluate the utility of the free-to-roll test technique as a tool for predicting areas of significant uncommanded lateral motions and for gaining insight into the wing-drop and wing-rock behavior of military aircraft at transonic conditions. The analysis indicated that the free-to-roll results had good agreement with flight data on all four models. A wide range of motions - limit cycle wing rock, occasional and frequent damped wing drop/rock and wing rock divergence - were observed. The analysis shows the effects that the static and dynamic lateral stability can have on the wing drop/rock behavior. In addition, a free-to-roll figure of merit was developed to assist in the interpretation of results and assessment of the severity of the motions.

CFD Predictions for Transonic Performance of the ERA Hybrid Wing-Body Configuration

NASA Technical Reports Server (NTRS)

Deere, Karen A.; Luckring, James M.; McMillin, S. Naomi; Flamm, Jeffrey D.; Roman, Dino

2016-01-01

A computational study was performed for a Hybrid Wing Body configuration that was focused at transonic cruise performance conditions. In the absence of experimental data, two fully independent computational fluid dynamics analyses were conducted to add confidence to the estimated transonic performance predictions. The primary analysis was performed by Boeing with the structured overset-mesh code OVERFLOW. The secondary analysis was performed by NASA Langley Research Center with the unstructured-mesh code USM3D. Both analyses were performed at full-scale flight conditions and included three configurations customary to drag buildup and interference analysis: a powered complete configuration, the configuration with the nacelle/pylon removed, and the powered nacelle in isolation. The results in this paper are focused primarily on transonic performance up to cruise and through drag rise. Comparisons between the CFD results were very good despite some minor geometric differences in the two analyses.

Suppression of flutter

NASA Technical Reports Server (NTRS)

Nissim, E. (Inventor)

1973-01-01

An active aerodynamic control system to control flutter over a large range of oscillatory frequencies is described. The system is not affected by mass, stiffness, elastic axis, or center of gravity location of the system, mode of vibration, or Mach number. The system consists of one or more pairs of leading edge and trailing edge hinged or deformable control surfaces, each pair operated in concert by a stability augmentation system. Torsion and bending motions are sensed and converted by the stability augmentation system into leading and trailing edge control surface deflections which produce lift forces and pitching moments to suppress flutter.

Model Deformation Measurements at NASA Langley Research Center

NASA Technical Reports Server (NTRS)

Burner, A. W.

1998-01-01

Only recently have large amounts of model deformation data been acquired in NASA wind tunnels. This acquisition of model deformation data was made possible by the development of an automated video photogrammetric system to measure the changes in wing twist and bending under aerodynamic load. The measurement technique is based upon a single view photogrammetric determination of two dimensional coordinates of wing targets with a fixed third dimensional coordinate, namely the spanwise location. A major consideration in the development of the measurement system was that use of the technique must not appreciably reduce wind tunnel productivity. The measurement technique has been used successfully for a number of tests at four large production wind tunnels at NASA and a dedicated system is nearing completion for a fifth facility. These facilities are the National Transonic Facility, the Transonic Dynamics Tunnel, and the Unitary Plan Wind Tunnel at NASA Langley, and the 12-FT Pressure Tunnel at NASA Ames. A dedicated system for the Langley 16-Foot Transonic Tunnel is scheduled to be used for the first time for a test in September. The advantages, limitations, and strategy of the technique as currently used in NASA wind tunnels are presented. Model deformation data are presented which illustrate the value of these measurements. Plans for further enhancements to the technique are presented.

Whirl Flutter Stability and Its Influence on the Design of the Distributed Electric Propeller Aircraft X- 57

NASA Technical Reports Server (NTRS)

Hoover, Christian B.; Shen, Jinwei; Kreshock, Andrew R.; Stanford, Bret K.; Piatak, David J.; Heeg, Jennifer

2017-01-01

This paper studies the whirl flutter stability of the NASA experimental electric propulsion aircraft designated the X-57 Maxwell. whirl flutter stability is studied at two flight conditions: sea level at 2700 RPM to represent take-off and landing and 8000 feet at 2250 RPM to represent cruise. Two multibody dynamics analyses are used: CAMRAD II and Dymore. The CAMRAD II model is a semi-span X-57 model with a modal representation for the wing/pylon system. The Dymore model is a semi-span wing with a propeller composed of beam elements for the wing/pylon system that airloads can be applied to. The two multibody dynamics analyses were verified by comparing structural properties between each other and the NASTRAN analysis. For whirl flutter, three design revisions of the wing and pylon mount system are studied. The predicted frequencies and damping ratio of the wing modes show good agreements between the two analyses. Dymore tended to predict a slightly lower damping ratio as velocity increased for all three dynamic modes presented. Whirl flutter for the semi-span model was not present up to 500 knots for the latest design, well above the operating range of the X-57.

Flutter analysis of composite box beams

NASA Technical Reports Server (NTRS)

Hodges, Dewey H.; Greenman, Matthew

1995-01-01

The dynamic aeroelastic instability of flutter is an important factor in the design of modern high-speed, flexible aircraft. The current trend is toward the creative use of composites to delay flutter. To obtain an optimum design, we need an accurate as well as efficient model. As a first step towards this goal, flutter analysis is carried out for an unswept composite box beam using a linear structural model and Theodorsen's unsteady aerodynamic theory. Structurally, the wing was modeled as a thin-walled box-beam of rectangular cross section. Theodorsen's theory was used to get 2-D unsteady aerodynamic forces, which were integrated over the span. A free-vibration analysis is carried out. These fundamental modes are used to get the flutter solution using the V-g method. Future work is intended to build on this foundation.

Multirate Flutter Suppression System Design for the Benchmark Active Controls Technology Wing. Part 1; Theory and Design Procedure

NASA Technical Reports Server (NTRS)

Mason, Gregory S.; Berg, Martin C.; Mukhopadhyay, Vivek

2002-01-01

To study the effectiveness of various control system design methodologies, the NASA Langley Research Center initiated the Benchmark Active Controls Project. In this project, the various methodologies were applied to design a flutter suppression system for the Benchmark Active Controls Technology (BACT) Wing. This report describes a project at the University of Washington to design a multirate suppression system for the BACT wing. The objective of the project was two fold. First, to develop a methodology for designing robust multirate compensators, and second, to demonstrate the methodology by applying it to the design of a multirate flutter suppression system for the BACT wing.

Boundary-layer measurements on a transonic low-aspect ratio wing

NASA Technical Reports Server (NTRS)

Keener, Earl R.

1985-01-01

Tabulations and plots are presented of boundary-layer velocity and flow-direction surveys from wind-tunnel tests of a large-scale (0.90 m semi-span) model of the NASA/Lockheed Wing C. This wing is a generic, transonic, supercritical, highly three-dimensional, low-aspect-ratio configuration designed with the use of a three-dimensional, transonic full-potential-flow wing code (FLO22). Tests were conducted at the design angle of attack of 5 deg over a Mach number range from 0.25 to 0.96 and a Reynolds number range of 3.4x10 to the 6th power. Wing pressures were measured at five span stations, and boundary-layer surveys were measured at the midspan station. The data are presented without analysis.

Reynolds Number Effects on Leading Edge Radius Variations of a Supersonic Transport at Transonic Conditions

NASA Technical Reports Server (NTRS)

Rivers, S. M. B.; Wahls, R. A.; Owens, L. R.

2001-01-01

A computational study focused on leading-edge radius effects and associated Reynolds number sensitivity for a High Speed Civil Transport configuration at transonic conditions was conducted as part of NASA's High Speed Research Program. The primary purposes were to assess the capabilities of computational fluid dynamics to predict Reynolds number effects for a range of leading-edge radius distributions on a second-generation supersonic transport configuration, and to evaluate the potential performance benefits of each at the transonic cruise condition. Five leading-edge radius distributions are described, and the potential performance benefit including the Reynolds number sensitivity for each is presented. Computational results for two leading-edge radius distributions are compared with experimental results acquired in the National Transonic Facility over a broad Reynolds number range.

Influence of mistuning on blade torsional flutter

NASA Technical Reports Server (NTRS)

Srinivasan, A. V.

1980-01-01

An analytical technique for the prediction of fan blade flutter was evaluated by utilizing first stage fan flutter data from tests on an advanced high performance engine. The formulation includes both aerodynamic and mechanical coupling among all the blades of the assembly. Mistuning is accounted for in the analysis so that individual blade inertias, frequencies, or damping can be considered. Airfoil stability was predicted by calculating a flutter determinant, the eigenvalues of which indicate the extent of susceptibility to flutter. When blade to blade differences in frequencies are considered, a stable system is predicted for the test points examined. For a tuned system, it was found that torsional flutter can be predicted at a limited number of interblade phase angles. Examination of these phase angles indicated that they were "close" to the condition of acoustic resonance. For the range of Mach numbers and reduced frequencies considered, the so called subcritical flutter cannot be predicted. The essential influence of mechanical coupling among the blades is to change the frequencies of the system with little or no change in damping; however, aerodynamic coupling together with mechanical coupling could change not only frequencies, but also damping in the system, with a trend toward instability.

LED's and the "Fluttering Heart" Phenomenon.

ERIC Educational Resources Information Center

Jewett, John W., Jr.

1993-01-01

Describes the nineteenth-century parlor trick entitled the Fluttering Heart phenomenon which uses a red heart on a bright blue background. Discusses theories concerning the apparent fluttering. Suggests doing the trick with a red light-emitting diode in a darkened room. (MVL)

Computational Modeling and Analysis of Aeroelastic Wing Flutter

NASA Astrophysics Data System (ADS)

Menon, Karthik; Katz, Joseph; Mittal, Rajat

2017-11-01

Aeroelastic flutter is ubiquitous in aeronautics; of particular relevance here is the flutter of aircraft wings, helicopter rotor blades, flexible wing MAVs and UAVs, and long-endurance aerial systems such as airships and solar powered air-vehicles. Here, we attempt to understand some fundamental aspects of this problem via immersed boundary method based numerical simulations of canonical bodies. We report findings on the effect of body geometry on the dynamics of flutter involving coupled pitch-heave oscillations. We also explore flow-induced flutter of airfoils in pre and post-stall configurations, including the effect of stiffness and pitch axis location. Finally, a novel force decomposition method is used to provide some insight into the flutter dynamics and associated unsteady flow physics. This work is supported by AFOSR Grant FA9550-16-1-0404.

Pressure distribution from high Reynolds number tests of a NASA SC(3)-0712(B) airfoil in the Langley 0.3-meter transonic cryogenic tunnel

NASA Technical Reports Server (NTRS)

Johnson, W. G., Jr.; Hill, A. S.; Eichmann, O.

1985-01-01

A wind tunnel investigation of a NASA 12-percent-thick, advanced-technology supercritical airfoil was conducted in the Langley 0.3-Meter Transonic Cryogenic Tunnel (TCT). This investigation represents another in the series of NASA/U.S. industry two-dimensional airfoil studies to be completed in the Advanced Technology Airfoil Tests program. Test temperature was varied from 220 K to 96 K at pressures ranging from 1.2 to 4.3 atm. Mach number was varied from 0.50 to 0.80. This investigation was designed to: (1) test a NASA advanced-technology airfoil from low to flight equivalent Reynolds numbers, (2) provide experience in cryogenic wind-tunnel model design and testing techniques, and (3) demonstrate the suitability of the 0.3-m TCT as an airfoil test facility. All the test objectives were met. The pressure data are presented without analysis in tabulated format and as plots of pressure coefficient versus position on the airfoil. This report was prepared for use in conjunction with the aerodynamic coefficient data published in NASA-TM-86371. Data are included which demonstrate the effects of fixed transition. Also included are remarks on the model design and fabrication.

An experimental and analytical investigation of proprotor whirl flutter

NASA Technical Reports Server (NTRS)

Kvaternik, R. G.; Kohn, J. S.

1977-01-01

The results of an experimental parametric investigation of whirl flutter are presented for a model consisting of a windmilling propeller-rotor, or proprotor, having blades with offset flapping hinges mounted on a rigid pylon with flexibility in pitch and yaw. The investigation was motivated by the need to establish a large data base from which to assess the predictability of whirl flutter for a proprotor since some question has been raised as to whether flutter in the forward whirl mode could be predicted with confidence. To provide the necessary data base, the parametric study included variation in the pylon pitch and yaw stiffnesses, flapping hinge offset, and blade kinematic pitch-flap coupling over a large range of advance ratios. Cases of forward whirl flutter and of backward whirl flutter are documented. Measured whirl flutter characteristics were shown to be in good agreement with predictions from two different linear stability analyses which employed simple, two dimensional, quasi-steady aerodynamics for the blade loading. On the basis of these results, it appears that proprotor whirl flutter, both forward and backward, can be predicted.

Transonic Drag Reduction Through Trailing-Edge Blowing on the FAST-MAC Circulation Control Model

NASA Technical Reports Server (NTRS)

Chan, David T.; Jones, Gregory S.; Milholen, William E., II; Goodliff, Scott L.

2017-01-01

A third wind tunnel test of the FAST-MAC circulation control semi-span model was completed in the National Transonic Facility at the NASA Langley Research Center where the model was configured for transonic testing of the cruise configuration with 0deg flap detection to determine the potential for transonic drag reduction with the circulation control blowing. The model allowed independent control of four circulation control plenums producing a high momentum jet from a blowing slot near the wing trailing edge that was directed over a 15% chord simple-hinged ap. Recent upgrades to transonic semi-span flow control testing at the NTF have demonstrated an improvement to overall data repeatability, particularly for the drag measurement, that allows for increased confidence in the data results. The static thrust generated by the blowing slot was removed from the wind-on data using force and moment balance data from wind-o thrust tares. This paper discusses the impact of the trailing-edge blowing to the transonic aerodynamics of the FAST-MAC model in the cruise configuration, where at flight Reynolds numbers, the thrust-removed corrected data showed that an overall drag reduction and increased aerodynamic efficiency was realized as a consequence of the blowing.

Historical Review of Uncommanded Lateral-Directional Motions at Transonic Conditions

NASA Technical Reports Server (NTRS)

Chambers, Joseph R.; Hall, Robert M.

2003-01-01

This paper presents the results of a survey of past experiences with uncommanded lateral-directional motions at transonic speeds during specific military aircraft programs. The effort was undertaken to provide qualitative and quantitative information on past airplane programs that might be of use to the participants in the joint NASA/Navy/Air Force Abrupt Wing Stall (AWS) Program. The AWS Program was initiated because of the experiences of the F/A-l8E/F development program, during which unexpected, severe wing-drop motions were encountered by preproduction aircraft at transonic conditions. These motions were judged to be significantly degrading to the primary mission requirements of the aircraft. Although the problem was subsequently solved for the production version of the F/A-l8E/F, a high-level review panel emphasized the poor understanding of such phenomena and issued a strong recommendation to: "Initiate a national research effort to thoroughly and systematically study the wing drop phenomena." A comprehensive, cooperative NASA/Navy/Air Force AWS Program was designed to respond to provide the required technology requirements. As part of the AWS Program, a work element was directed at a historical review of wing-drop experiences in past aircraft development programs at high subsonic and transonic speeds. In particular, information was requested regarding: specific aircraft configurations that exhibited uncommanded motions and the nature of the motions; geometric characteristics of the air- planes; flight conditions involved in occurrences; relevant data, including wind-tunnel, computational, and flight sources; figures of merit used for analyses; and approaches used to alleviate the problem. An attempt was also made to summarize some of the more important lessons learned from past experiences, and to recommend specific research efforts. In addition to providing technical information to assist the AWS research objectives, the study produced fundamental

Contributions of the NASA Langley Transonic Dynamics Tunnel to Launch Vehicle and Spacecraft Development

NASA Technical Reports Server (NTRS)

Cole, Stanley R.; Keller, Donald F.; Piatak, David J.

2000-01-01

The NASA Langley Transonic Dynamics Tunnel (TDT) has provided wind-tunnel experimental validation and research data for numerous launch vehicles and spacecraft throughout its forty year history. Most of these tests have dealt with some aspect of aeroelastic or unsteady-response testing, which is the primary purpose of the TDT facility. However, some space-related test programs that have not involved aeroelasticity have used the TDT to take advantage of specific characteristics of the wind-tunnel facility. In general. the heavy gas test medium, variable pressure, relatively high Reynolds number and large size of the TDT test section have made it the preferred facility for these tests. The space-related tests conducted in the TDT have been divided into five categories. These categories are ground wind loads, launch vehicle dynamics, atmospheric flight of space vehicles, atmospheric reentry. and planetary-probe testing. All known TDT tests of launch vehicles and spacecraft are discussed in this report. An attempt has been made to succinctly summarize each wind-tunnel test, or in the case of multiple. related tests, each wind-tunnel program. Most summaries include model program discussion, description of the physical wind-tunnel model, and some typical or significant test results. When available, references are presented to assist the reader in further pursuing information on the tests.

Flutter calculations in three degrees of freedom

NASA Technical Reports Server (NTRS)

Theodorsen, Theodore; Garrick, I E

1942-01-01

The present paper is a continuation of the general study of flutter published in NACA reports nos. 496 and 685. The paper is mainly devoted to flutter in three degrees of freedom (bending, torsion, and aileron) for which a number of selected cases have been calculated and presented in graphical form. The results are analyzed and discussed with regard to the effects of structural damping, of fractional-span ailerons, and of mass-balancing. The analysis shows that more emphasis should be put on the effect of structural damping and less on mass-balancing. The conclusion is drawn that a definite minimum amount of structural damping, which is usually found to be present, is essential in the calculations for an adequate description of the flutter case. Theoretical flutter predictions are thus brought into closer agreement with the facts of experience. A brief discussion is included of a particular biplane that had experienced flutter at about 200 miles per hour. Some simplifications have been achieved in the method of calculation. (author)

Effects of spoiler surfaces on the aeroelastic behavior of a low-aspect-ratio rectangular wing

NASA Technical Reports Server (NTRS)

Cole, Stanley R.

1990-01-01

An experimental research study to determine the effectiveness of spoiler surfaces in suppressing flutter onset for a low-aspect-ratio, rectangular wing was conducted in the Langley Transonic Dynamics Tunnel (TDT). The wing model used in this flutter test consisted of a rigid wing mounted to the wind-tunnel wall by a flexible, rectangular beam. The flexible beam was connected to the wing root and cantilever mounted to the wind-tunnel wall. The wing had a 1.5 aspect ratio based on wing semispan and a NACA 64A010 airfoil shape. The spoiler surfaces consisted of thin, rectangular aluminum plates that were vertically mounted to the wing surface. The spoiler surface geometry and location on the wing surface were varied to determine the effects of these parameters on the classical flutter of the wing model. Subsonically, the experiment showed that spoiler surfaces increased the flutter dynamic pressure with each successive increase in spoiler height or width. This subsonic increase in flutter dynamic pressure was approximately 15 percent for the maximum height spoiler configuration and for the maximum width spoiler configuration. At transonic Mach numbers, the flutter dynamic pressure conditions were increased even more substantially than at subsonic Mach numbers for some of the smaller spoiler surfaces. But greater than a certain spoiler size (in terms of either height or width) the spoilers forced a torsional instability in the transonic regime that was highly Mach number dependent. This detrimental torsional instability was found at dynamic pressures well below the expected flutter conditions. Variations in the spanwise location of the spoiler surfaces on the wing showed little effect on flutter. Flutter analysis was conducted for the basic configuration (clean wing with all spoiler surface mass properties included). The analysis correlated well with the clean wing experimental flutter results.

Magnetic-flutter-induced pedestal plasma transport

NASA Astrophysics Data System (ADS)

Callen, J. D.; Hegna, C. C.; Cole, A. J.

2013-11-01

Plasma toroidal rotation can limit reconnection of externally applied resonant magnetic perturbation (RMP) fields δB on rational magnetic flux surfaces. Hence it causes the induced radial perturbations δBρ to be small there, thereby inhibiting magnetic island formation and stochasticity at the top of pedestals in high (H-mode) confinement tokamak plasmas. However, the δBρs induced by RMPs increase away from rational surfaces and are shown to induce significant sinusoidal radial motion (flutter) of magnetic field lines with a radial extent that varies linearly with δBρ and inversely with distance from the rational surface because of the magnetic shear. This produces a radial electron thermal diffusivity that is (1/2)(δBρ/B0)2 times a kinetically derived, electron-collision-induced, magnetic-shear-reduced, effective parallel electron thermal diffusivity in the absence of magnetic stochasticity. These low collisionality flutter-induced transport processes and thin magnetic island effects are shown to be highly peaked in the vicinity of rational surfaces at the top of low collisionality pedestals. However, the smaller but finite level of magnetic-flutter-induced electron heat transport midway between rational surfaces is the primary factor that determines the electron temperature difference between rational surfaces at the pedestal top. The magnetic-flutter-induced non-ambipolar electron density transport can be large enough to push the plasma toward an electron density transport root. Requiring ambipolar density transport is shown to determine the radial electric field, the plasma toroidal rotation (via radial force balance), a reduced electron thermal diffusivity and increased ambipolar density transport in the pedestal. At high collisionality the various flutter effects are less strongly peaked at rational surfaces and generally less significant. They are thus less likely to exhibit flutter-induced resonant behaviour and transition toward an electron

High Reynolds number tests of a NASA SC(3)-0712(B) airfoil in the Langley 0.3-meter transonic cryogenic tunnel

NASA Technical Reports Server (NTRS)

Johnson, W. G., Jr.; Hill, A. S.; Eichmann, O.

1985-01-01

A wind tunnel investigation of a NASA 12-percent-thick, advanced-technology supercritical airfoil was conducted in the Langley 0.3-Meter Transonic Cryogenic Tunnel (TCT). This investigation represents another in the series of NASA/U.S. industry two-dimensional airfoil studies to be completed in the Advanced Technology Airfoil Tests program. Test temperature was varied from 220 K to 96 K at pressures ranging from 1.2 to 4.3 atm. Mach number was varied from 0.60 to 0.80. These variables provided a Reynolds number range from 4,400,000 to 40,000,000 based on a 15.24-cm (6.0-in.) airfoil chord. This investigation was designed to test a NASA advanced-technology airfoil from low to flight-equivalent Reynolds numbers, provide experience in cryogenic wind tunnel model design and testing techniques, and demonstrate the suitability of the 0.3-m TCT as an airfoil test facility. The aerodynamic results are presented as integrated force and moment coefficients and pressure distributions. Data are included which demonstrate the effects of fixed transition, Mach number, and Reynolds number on the aerodynamic characteristics. Also included are remarks on the model design, the model structural integrity, and the overall test experience.

Stall Flutter Control of a Smart Blade Section Undergoing Asymmetric Limit Oscillations

DOE PAGES

Li, Nailu; Balas, Mark J.; Nikoueeyan, Pourya; ...

2016-01-01

Stall flutter is an aeroelastic phenomenon resulting in unwanted oscillatory loads on the blade, such as wind turbine blade, helicopter rotor blade, and other flexible wing blades. While the stall flutter and related aeroelastic control have been studied theoretically and experimentally, microtab control of asymmetric limit cycle oscillations (LCOs) in stall flutter cases has not been generally investigated. This paper presents an aeroservoelastic model to study the microtab control of the blade section undergoing moderate stall flutter and deep stall flutter separately. The effects of different dynamic stall conditions and the consequent asymmetric LCOs for both stall cases are simulatedmore » and analyzed. Then, for the design of the stall flutter controller, the potential sensor signal for the stall flutter, the microtab control capability of the stall flutter, and the control algorithm for the stall flutter are studied. Lastly, the improvement and the superiority of the proposed adaptive stall flutter controller are shown by comparison with a simple stall flutter controller.« less

Exhaust Simulation Testing of a Hypersonic Airbreathing Model at Transonic Speeds

NASA Technical Reports Server (NTRS)

Huebner, Lawrence D.; Witte, David W.; Andrews, Earl H., Jr.

2004-01-01

An experimental study was performed to examine jet-effects for an airframe-integrated, scramjet-rocket combined-cycle vehicle configuration at transonic test conditions. This investigation was performed by testing an existing exhaust simulation wind tunnel model, known as Model 5B, in the NASA Langley 16-Ft. Transonic Tunnel. Tests were conducted at freestream Mach numbers from 0.7 to 1.2, at angles of attack from 2 to +14 degrees, and at up to seven nozzle static pressure ratio values for a set of horizontal-tail and body-flap deflections. The model aftbody, horizontal tails, and body flaps were extensively pressure instrumented to provide an understanding of jet-effects and control-surface/plume interactions, as well as for the development of analytical methodologies and calibration of computational fluid dynamic codes to predict this type of flow phenomenon. At all transonic test conditions examined, the exhaust flow at the exit of the internal nozzle was over-expanded, generating an exhaust plume that turned toward the aftbody. Pressure contour plots for the aftbody of Model 5B are presented for freestream transonic Mach numbers of 0.70, 0.95, and 1.20. These pressure data, along with shadowgraph images, indicated the impingement of an internal plume shock and at least one reflected shock onto the aftbody for all transonic conditions tested. These results also provided evidence of the highly three-dimensional nature of the aftbody exhaust flowfield. Parametric testing showed that angle-of-attack, static nozzle pressure ratio, and freestream Mach number all affected the exhaust-plume size, exhaust-flowfield shock structure, and the aftbody-pressure distribution, with Mach number having the largest effect. Integration of the aftbody pressure data showed large variations in the pitching moment throughout the transonic regime.

Experimental Investigation of a Preloaded Spring-tab Flutter Model

NASA Technical Reports Server (NTRS)

Smith, N H; Clevenson, S A; Barmby, J G

1947-01-01

An experimental investigation was made of a preloaded spring-tab flutter model to determine the effects on flutter speed of aspect ratio, tab frequency, and preloaded spring constant. The rudder was mass-balanced, and the flutter mode studied was essentially one of three degrees of freedom (fin bending coupled with rudder and tab oscillations). Inasmuch as the spring was preloaded, the tab-spring system was a nonlinear one. Frequency of the tab was the most significant parameter in this study, and an increase in flutter speed with increasing frequency is indicated. At a given frequency, the tab of high aspect ratio is shown to have a slightly lower flutter speed than the one of low aspect ratio. Because the frequency of the preloaded spring tab was found to vary radically with amplitude, the flutter speed decreased with increase in initial displacement of the tab.

Flight Flutter Testing of Supersonic Interceptors

NASA Technical Reports Server (NTRS)

Dublin, M.; Peller, R.

1975-01-01

A summary is presented of experiences in connection with flight flutter testing of supersonic interceptors. The planning and operational aspects involved are described along with the difficulties encountered, and the correlation between measurement and theory. Recommendations for future research and development to advance the science of flight flutter testing are included.

Characteristics of aeroelastic instabilities in turbomachinery - NASA full scale engine test results

NASA Technical Reports Server (NTRS)

Lubomski, J. F.

1979-01-01

Several aeromechanical programs were conducted in the NASA/USAF Joint Engine System Research Programs. The scope of these programs, the instrumentation, data acquisition and reduction, and the test results are discussed. Data pertinent to four different instabilities were acquired; two types of stall flutter, choke flutter and a system mode instability. The data indicates that each instability has its own unique characteristics. These characteristics are described.

NASA Environmentally Responsible Aviation High Overall Pressure Ratio Compressor Research Pre-Test CFD

NASA Technical Reports Server (NTRS)

Celestina, Mark L.; Fabian, John C.; Kulkarni, Sameer

2012-01-01

This paper describes a collaborative and cost-shared approach to reducing fuel burn under the NASA Environmentally Responsible Aviation project. NASA and General Electric (GE) Aviation are working together aa an integrated team to obtain compressor aerodynamic data that is mutually beneficial to both NASA and GE Aviation. The objective of the High OPR Compressor Task is to test a single stage then two stages of an advanced GE core compressor using state-of-the-art research instrumentation to investigate the loss mechanisms and interaction effects of embedded transonic highly-loaded compressor stages. This paper presents preliminary results from NASA's in-house multistage computational code, APNASA, in preparation for this advanced transonic compressor rig test.

Historical Review of Uncommanded Lateral-Directional Motions At Transonic Conditions (Invited)

NASA Technical Reports Server (NTRS)

Chambers, Joseph R.; Hall, Robert M.

2003-01-01

This paper presents the results of a survey of past experiences with uncommanded lateral-directional motions at transonic speeds during specific military aircraft programs. The effort was undertaken to provide qualitative and quantitative information on past airplane programs that might be of use to the participants in the joint NASA/Navy/Air Force Abrupt Wing Stall (AWS) Program. The AWS Program was initiated because of the experiences of the F/A-18E/F development program, during which unexpected, severe wing-drop motions were encountered by preproduction aircraft at transonic conditions. These motions were judged to be significantly degrading to the primary mission requirements of the aircraft. Although the problem was subsequently solved for the production version of the F/A-l8E/F, a high-level review panel emphasized the poor understanding of such phenomena and issued a strong recommendation to: Initiate a national research effort to thoroughly and systematically study the wing drop phenomena. A comprehensive, cooperative NASA/Navy/Air Force AWS Program was designed to respond to provide the required technology requirements. As part of the AWS Program, a work element was directed at a historical review of wing-drop experiences in past aircraft development programs at high subsonic and transonic speeds. In particular, information was requested regarding: specific aircraft configurations that exhibited uncommanded motions and the nature of the motions; geometric characteristics of the air- planes; flight conditions involved in occurrences; relevant data, including wind-tunnel, computational, and flight sources; figures of merit used for analyses; and approaches used to alleviate the problem. An attempt was also made to summarize some of the more important lessons learned from past experiences, and to recommend specific research efforts. In addition to providing technical information to assist the AWS research objectives, the study produced fundamental information

Ch-47C Fixed-System Stall-Flutter Damping

DTIC Science & Technology

1975-08-01

flutter. The steady and vibratory loads in the cyclic-trim linkage are so related that motions across the control system’s mechan- ical free play could...be a significant part of the stall-flutter motion, depending on the magnitude of the free play . For this reason it is recommended that future testing...include the deter- mination of the effects of control-system free play on the stall-flutter responses. , f ,**~ - ,***,- * **4 , - - *. i

Robust Flutter Margin Analysis that Incorporates Flight Data

NASA Technical Reports Server (NTRS)

Lind, Rick; Brenner, Martin J.

1998-01-01

An approach for computing worst-case flutter margins has been formulated in a robust stability framework. Uncertainty operators are included with a linear model to describe modeling errors and flight variations. The structured singular value, mu, computes a stability margin that directly accounts for these uncertainties. This approach introduces a new method of computing flutter margins and an associated new parameter for describing these margins. The mu margins are robust margins that indicate worst-case stability estimates with respect to the defined uncertainty. Worst-case flutter margins are computed for the F/A-18 Systems Research Aircraft using uncertainty sets generated by flight data analysis. The robust margins demonstrate flight conditions for flutter may lie closer to the flight envelope than previously estimated by p-k analysis.

Results for the hybrid laminar flow control experiment conducted in the NASA Langley 8-foot transonic pressure tunnel on a 7-foot chord model

NASA Technical Reports Server (NTRS)

Bobbitt, Percy J.; Ferris, James C.; Harvey, William D.; Goradia, Suresh H.

1992-01-01

A description is given of the development of, and results from, the hybrid laminar flow control (HLFC) experiment conducted in the NASA LaRC 8 ft Transonic Pressure Tunnel on a 7 ft chord, 23 deg swept model. The methods/codes used to obtain the contours of the HLFC model surface and to define the suction requirements are outlined followed by a discussion of the model construction, suction system, instrumentation, and some example results from the wind tunnel tests. Included in the latter are the effects of Mach number, suction level, and the extent of suction. An assessment is also given of the effect of the wind tunnel environment on the suction requirements. The data show that, at or near the design Mach number, large extents of laminar flow can be achieved with suction mass flows over the first 25 percent, or less, of the chord. Top surface drag coefficients with suction extending from the near leading edge to 20 percent of the chord were approximately 40 percent lower than those obtained with no suction. The results indicate that HLFC can be designed for transonic speeds with lift and drag coefficients approaching those of LFC designs but with much smaller extents and levels of suction.

A Bibliography of Transonic Dynamics Tunnel (TDT) Publications

NASA Technical Reports Server (NTRS)

Doggett, Robert V.

2016-01-01

The Transonic Dynamics Tunnel (TDT) at the National Aeronautics and Space Administration's (NASA) Langley Research Center began research operations in early 1960. Since that time, over 600 tests have been conducted, primarily in the discipline of aeroelasticity. This paper presents a bibliography of the publications that contain data from these tests along with other reports that describe the facility, its capabilities, testing techniques, and associated research equipment. The bibliography is divided by subject matter into a number of categories. An index by author's last name is provided.

Flutter-driven triboelectrification for harvesting wind energy

NASA Astrophysics Data System (ADS)

Bae, Jihyun; Lee, Jeongsu; Kim, Seongmin; Ha, Jaewook; Lee, Byoung-Sun; Park, Youngjun; Choong, Chweelin; Kim, Jin-Baek; Wang, Zhong Lin; Kim, Ho-Young; Park, Jong-Jin; Chung, U.-In

2014-09-01

Technologies to harvest electrical energy from wind have vast potentials because wind is one of the cleanest and most sustainable energy sources that nature provides. Here we propose a flutter-driven triboelectric generator that uses contact electrification caused by the self-sustained oscillation of flags. We study the coupled interaction between a fluttering flexible flag and a rigid plate. In doing so, we find three distinct contact modes: single, double and chaotic. The flutter-driven triboelectric generator having small dimensions of 7.5 × 5 cm at wind speed of 15 ms-1 exhibits high-electrical performances: an instantaneous output voltage of 200 V and a current of 60 μA with a high frequency of 158 Hz, giving an average power density of approximately 0.86 mW. The flutter-driven triboelectric generation is a promising technology to drive electric devices in the outdoor environments in a sustainable manner.

Investigations on precursor measures for aeroelastic flutter

NASA Astrophysics Data System (ADS)

Venkatramani, J.; Sarkar, Sunetra; Gupta, Sayan

2018-04-01

Wind tunnel experiments carried out on a pitch-plunge aeroelastic system in the presence of fluctuating flows reveal that flutter instability is presaged by a regime of intermittency. It is observed that as the flow speed gradually increases towards the flutter speed, there appears intermittent bursts of periodic oscillations which become more frequent as the wind speed increases and eventually the dynamics transition into fully developed limit cycle oscillations, marking the onset of flutter. The signature from these intermittent oscillations are exploited to develop measures that forewarn a transition to flutter and can serve as precursors. This study investigates a suite of measures that are obtained directly from the time history of measurements and are hence model independent. The dependence of these precursors on the size of the measured data set and the time required for their computation is investigated. These measures can be useful in structural health monitoring of aeroelastic structures.

Uncertainty Quantification of Turbulence Model Closure Coefficients for Transonic Wall-Bounded Flows

NASA Technical Reports Server (NTRS)

Schaefer, John; West, Thomas; Hosder, Serhat; Rumsey, Christopher; Carlson, Jan-Renee; Kleb, William

2015-01-01

The goal of this work was to quantify the uncertainty and sensitivity of commonly used turbulence models in Reynolds-Averaged Navier-Stokes codes due to uncertainty in the values of closure coefficients for transonic, wall-bounded flows and to rank the contribution of each coefficient to uncertainty in various output flow quantities of interest. Specifically, uncertainty quantification of turbulence model closure coefficients was performed for transonic flow over an axisymmetric bump at zero degrees angle of attack and the RAE 2822 transonic airfoil at a lift coefficient of 0.744. Three turbulence models were considered: the Spalart-Allmaras Model, Wilcox (2006) k-w Model, and the Menter Shear-Stress Trans- port Model. The FUN3D code developed by NASA Langley Research Center was used as the flow solver. The uncertainty quantification analysis employed stochastic expansions based on non-intrusive polynomial chaos as an efficient means of uncertainty propagation. Several integrated and point-quantities are considered as uncertain outputs for both CFD problems. All closure coefficients were treated as epistemic uncertain variables represented with intervals. Sobol indices were used to rank the relative contributions of each closure coefficient to the total uncertainty in the output quantities of interest. This study identified a number of closure coefficients for each turbulence model for which more information will reduce the amount of uncertainty in the output significantly for transonic, wall-bounded flows.

Status of the national transonic facility

NASA Technical Reports Server (NTRS)

Mckinney, L. W.; Gloss, B. B.

1982-01-01

The National Transonic Facility at NASA Langley Research Center, scheduled for completion in July, 1982, is described with emphasis on model and instrumentation activities, calibration plans and some initial research plans. Performance capabilities include a Mach number range of 0.2-1.2, a pressure range of 1-9 atmospheres, and a temperature range of 77-350 K, which will produce a maximum Reynolds number of 120 million at a Mach number of 1.0, based on a 0.25 m chord. A comprehensive tunnel calibration program is planned, which will cover basic tunnel calibration, data qualities, and data comparisons with other facilites and flights.

About the Effect of Control on Flutter and Post-Flutter of a Supersonic/Hypersonic Cross-Sectional Wing

NASA Technical Reports Server (NTRS)

Silva, Walter A.; Librescu, Liviu; Marzocca, Piergiovanni

2001-01-01

The control of the flutter instability and the conversion of the dangerous character of the flutter instability boundary into the undangerous one of a cross-sectional wing in a supersonic/hypersonic flow field is presented. The objective of this paper is twofold: i) to analyze the implications of nonlinear unsteady aerodynamics and physical nonlinearities on the character of the instability boundary in the presence of a control capability, and ii) to outline the effects played in the same respect by some important parameters of the aeroelastic system. As a by-product of this analysis, the implications of the active control on the linearized flutter behavior of the system are captured and emphasized. The bifurcation behavior of the open/closed loop aeroelastic system in the vicinity of the flutter boundary is studied via the use of a new methodology based on the Liapunov First Quantity. The expected outcome of this study is: a) to greatly enhance the scope and reliability of the aeroelastic analysis and design criteria of advanced supersonic/hypersonic flight vehicles and, b) provide a theoretical basis for the analysis of more complex nonlinear aeroelastic systems.

About the Effect of Control on Flutter and Post-Flutter of a Supersonic/Hypersonic Cross-Sectional Wing

NASA Technical Reports Server (NTRS)

Marzocca, Piergiovanni; Librescu, Liviu; Silva, Walter A.

2000-01-01

The control of the flutter instability and the conversion of the dangerous character of the flutter instability boundary into the undangerous one of a cross-sectional wing in a supersonic/hypersonic flow field is presented. The objective of this paper is twofold: i) to analyze the implications of nonlinear unsteady aerodynamics and physical nonlinearities on the character of the instability boundary in the presence of a control capability, and ii) to outline the effects played in the same respect by some important parameters of the aeroelastic system. As a by-product of this analysis, the implications of the active control on the linearized flutter behavior of the system are captured and emphasized. The bifurcation behavior of the open/closed loop aeroelastic system in the vicinity of the flutter boundary is studied via the use of a new methodology based on the Liapunov First Quantity. The expected outcome of this study is: a) to greatly enhance the scope and reliability of the aeroelastic analysis and design criteria of advanced supersonic/hypersonic flight vehicles and, b) provide a theoretical basis for the analysis of more complex nonlinear aeroelastic systems.

Preliminary study of effects of winglets on wing flutter

NASA Technical Reports Server (NTRS)

Doggett, R. V., Jr.; Farmer, M. G.

1976-01-01

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.

Nonlinear flutter analysis of composite panels

NASA Astrophysics Data System (ADS)

An, Xiaomin; Wang, Yan

2018-05-01

Nonlinear panel flutter is an interesting subject of fluid-structure interaction. In this paper, nonlinear flutter characteristics of curved composite panels are studied in very low supersonic flow. The composite panel with geometric nonlinearity is modeled by a nonlinear finite element method; and the responses are computed by the nonlinear Newmark algorithm. An unsteady aerodynamic solver, which contains a flux splitting scheme and dual time marching technology, is employed in calculating the unsteady pressure of the motion of the panel. Based on a half-step staggered coupled solution, the aeroelastic responses of two composite panels with different radius of R = 5 and R = 2.5 are computed and compared with each other at different dynamic pressure for Ma = 1.05. The nonlinear flutter characteristics comprising limited cycle oscillations and chaos are analyzed and discussed.

Unsteady Pressures in a Transonic Fan Cascade Due to a Single Oscillating Airfoil

NASA Technical Reports Server (NTRS)

Lepicovsky, J.; McFarland, E. R.; Capece, V. R.; Hayden, J.

2002-01-01

An extensive set of unsteady pressure data was acquired along the midspan of a modern transonic fan blade for simulated flutter conditions. The data set was acquired in a nine-blade linear cascade with an oscillating middle blade to provide a database for the influence coefficient method to calculate instantaneous blade loadings. The cascade was set for an incidence of 10 dg. The data were acquired on three stationary blades on each side of the middle blade that was oscillated at an amplitude of 0.6 dg. The matrix of test conditions covered inlet Mach numbers of 0.5, 0.8, and 1.1 and the oscillation frequencies of 200, 300, 400, and 500 Hz. A simple quasiunsteady two-dimensional computer simulation was developed to aid in the running of the experimental program. For high Mach number subsonic inlet flows the blade pressures exhibit very strong, low-frequency, self-induced oscillations even without forced blade oscillations, while for low subsonic and supersonic inlet Mach numbers the blade pressure unsteadiness is quite low. The amplitude of forced pressure fluctuations on neighboring stationary blades strongly depends on the inlet Mach number and forcing frequency. The flowfield behavior is believed to be governed by strong nonlinear effects due to a combination of viscosity, compressibility, and unsteadiness. Therefore, the validity of the quasi-unsteady simplified computer simulation is limited to conditions when the flowfield is behaving in a linear, steady manner. Finally, an extensive set of unsteady pressure data was acquired to help development and verification of computer codes for blade flutter effects.

Flutter suppression for the Active Flexible Wing - Control system design and experimental validation

NASA Technical Reports Server (NTRS)

Waszak, M. R.; Srinathkumar, S.

1992-01-01

The synthesis and experimental validation of a control law for an active flutter suppression system for the Active Flexible Wing wind-tunnel model is presented. The design was accomplished with traditional root locus and Nyquist methods using interactive computer graphics tools and with extensive use of simulation-based analysis. The design approach relied on a fundamental understanding of the flutter mechanism to formulate understanding of the flutter mechanism to formulate a simple control law structure. Experimentally, the flutter suppression controller succeeded in simultaneous suppression of two flutter modes, significantly increasing the flutter dynamic pressure despite errors in the design model. The flutter suppression controller was also successfully operated in combination with a rolling maneuver controller to perform flutter suppression during rapid rolling maneuvers.

Design and experimental validation of a flutter suppression controller for the active flexible wing

NASA Technical Reports Server (NTRS)

Waszak, Martin R.; Srinathkumar, S.

1992-01-01

The synthesis and experimental validation of an active flutter suppression controller for the Active Flexible Wing wind tunnel model is presented. The design is accomplished with traditional root locus and Nyquist methods using interactive computer graphics tools and extensive simulation based analysis. The design approach uses a fundamental understanding of the flutter mechanism to formulate a simple controller structure to meet stringent design specifications. Experimentally, the flutter suppression controller succeeded in simultaneous suppression of two flutter modes, significantly increasing the flutter dynamic pressure despite modeling errors in predicted flutter dynamic pressure and flutter frequency. The flutter suppression controller was also successfully operated in combination with another controller to perform flutter suppression during rapid rolling maneuvers.

An investigation of several factors involved in a finite difference procedure for analyzing the transonic flow about harmonically oscillating airfoils and wings

NASA Technical Reports Server (NTRS)

Ehlers, F. E.; Sebastian, J. D.; Weatherill, W. H.

1979-01-01

Analytical and empirical studies of a finite difference method for the solution of the transonic flow about harmonically oscillating wings and airfoils are presented. The procedure is based on separating the velocity potential into steady and unsteady parts and linearizing the resulting unsteady equations for small disturbances. Since sinusoidal motion is assumed, the unsteady equation is independent of time. Three finite difference investigations are discussed including a new operator for mesh points with supersonic flow, the effects on relaxation solution convergence of adding a viscosity term to the original differential equation, and an alternate and relatively simple downstream boundary condition. A method is developed which uses a finite difference procedure over a limited inner region and an approximate analytical procedure for the remaining outer region. Two investigations concerned with three-dimensional flow are presented. The first is the development of an oblique coordinate system for swept and tapered wings. The second derives the additional terms required to make row relaxation solutions converge when mixed flow is present. A finite span flutter analysis procedure is described using the two-dimensional unsteady transonic program with a full three-dimensional steady velocity potential.

Small Engine Technology (Set) Task 8 Aeroelastic Prediction Methods

NASA Technical Reports Server (NTRS)

Eick, Chris D.; Liu, Jong-Shang

1998-01-01

AlliedSignal Engines, in cooperation with NASA LeRC, completed an evaluation of recently developed aeroelastic computer codes using test cases from the AlliedSignal Engines fan blisk database. Test data for this task includes strain gage, light probe, performance, and steady-state pressure information obtained for conditions where synchronous or flutter vibratory conditions were found to occur. Aeroelastic codes evaluated include the quasi 3-D UNSFLO (developed at MIT and modified to include blade motion by AlliedSignal), the 2-D FREPS (developed by NASA LeRC), and the 3-D TURBO-AE (under development at NASA LeRC). Six test cases each where flutter and synchronous vibrations were found to occur were used for evaluation of UNSFLO and FREPS. In addition, one of the flutter cases was evaluated using TURBO-AE. The UNSFLO flutter evaluations were completed for 75 percent radial span and provided good agreement with the experimental test data. Synchronous evaluations were completed for UNSFLO but further enhancement needs to be added to the code before the unsteady pressures can be used to predict forced response vibratory stresses. The FREPS evaluations were hindered as the steady flow solver (SFLOW) was unable to converge to a solution for the transonic flow conditions in the fan blisk. This situation resulted in all FREPS test cases being attempted but no results were obtained during the present program. Currently, AlliedSignal is evaluating integrating FREPS with our existing steady flow solvers to bypass the SFLOW difficulties. ne TURBO-AE steady flow solution provided an excellent match with the AlliedSignal Engines calibrated DAWES 3-D viscous solver. Finally, the TURBO-AE unsteady analyses also matched experimental observations by predicting flutter for the single test case evaluated.

Advancing Test Capabilities at NASA Wind Tunnels

NASA Technical Reports Server (NTRS)

Bell, James

2015-01-01

NASA maintains twelve major wind tunnels at three field centers capable of providing flows at 0.1 M 10 and unit Reynolds numbers up to 45106m. The maintenance and enhancement of these facilities is handled through a unified management structure under NASAs Aeronautics and Evaluation and Test Capability (AETC) project. The AETC facilities are; the 11x11 transonic and 9x7 supersonic wind tunnels at NASA Ames; the 10x10 and 8x6 supersonic wind tunnels, 9x15 low speed tunnel, Icing Research Tunnel, and Propulsion Simulator Laboratory, all at NASA Glenn; and the National Transonic Facility, Transonic Dynamics Tunnel, LAL aerothermodynamics laboratory, 8 High Temperature Tunnel, and 14x22 low speed tunnel, all at NASA Langley. This presentation describes the primary AETC facilities and their current capabilities, as well as improvements which are planned over the next five years. These improvements fall into three categories. The first are operations and maintenance improvements designed to increase the efficiency and reliability of the wind tunnels. These include new (possibly composite) fan blades at several facilities, new temperature control systems, and new and much more capable facility data systems. The second category of improvements are facility capability advancements. These include significant improvements to optical access in wind tunnel test sections at Ames, improvements to test section acoustics at Glenn and Langley, the development of a Supercooled Large Droplet capability for icing research, and the development of an icing capability for large engine testing. The final category of improvements consists of test technology enhancements which provide value across multiple facilities. These include projects to increase balance accuracy, provide NIST-traceable calibration characterization for wind tunnels, and to advance optical instruments for Computational Fluid Dynamics (CFD) validation. Taken as a whole, these individual projects provide significant

Interactive flutter analysis and parametric study for conceptual wing design

NASA Technical Reports Server (NTRS)

Mukhopadhyay, Vivek

1995-01-01

An interactive computer program was developed for wing flutter analysis in the conceptual design stage. The objective was to estimate the flutter instability boundary of a flexible cantilever wing, when well defined structural and aerodynamic data are not available, and then study the effect of change in Mach number, dynamic pressure, torsional frequency, sweep, mass ratio, aspect ratio, taper ratio, center of gravity, and pitch inertia, to guide the development of the concept. The software was developed on MathCad (trademark) platform for Macintosh, with integrated documentation, graphics, database and symbolic mathematics. The analysis method was based on nondimensional parametric plots of two primary flutter parameters, namely Regier number and Flutter number, with normalization factors based on torsional stiffness, sweep, mass ratio, aspect ratio, center of gravity location and pitch inertia radius of gyration. The plots were compiled in a Vaught Corporation report from a vast database of past experiments and wind tunnel tests. The computer program was utilized for flutter analysis of the outer wing of a Blended Wing Body concept, proposed by McDonnell Douglas Corporation. Using a set of assumed data, preliminary flutter boundary and flutter dynamic pressure variation with altitude, Mach number and torsional stiffness were determined.

The New Heavy Gas Testing Capability in the NASA Langley Transonic Dynamics Tunnel

NASA Technical Reports Server (NTRS)

Cole, Stanley R.; Rivera, Jose A., Jr.

1997-01-01

The NASA Langley Transonic Dynamics Tunnel (TDT) has provided a unique capability for aeroelastic testing for over thirty-five years. The facility has a rich history of significant contributions to the design of many United States commercial transports and military aircraft. The facility has many features which contribute to its uniqueness for aeroelasticity testing; however, perhaps the most important facility capability is the use of a heavy gas test medium to achieve higher test densities. Higher test medium densities substantially improve model building requirements and therefore simplify the fabrication process for building aeroelastically scaled wind-tunnel models. The heavy gas also provides other testing benefits, including reduction in the power requirements to operate the facility during testing. Unfortunately, the use of the original heavy gas has been curtailed due to environmental concerns. A new gas, referred to as R-134a, has been identified as a suitable replacement for the former TDT heavy gas. The TDT is currently undergoing a facility upgrade to allow testing in R-134a heavy gas. This replacement gas will result in an operational test envelope, model scaling advantages, and general testing capabilities similar to those available with the former TDT heavy gas. As such, the TDT is expected to remain a viable facility for aeroelasticity research and aircraft dynamic clearance testing well into the 21st century. This paper describes the anticipated advantages and facility calibration plans for the new heavy gas and briefly reviews several past test programs that exemplify the possible benefits of heavy gas testing.

Optical detection of blade flutter. [in YF-100 turbofan engine

NASA Technical Reports Server (NTRS)

Nieberding, W. C.; Pollack, J. L.

1977-01-01

The paper examines the capabilities of photoelectric scanning (PES) and stroboscopic imagery (SI) as optical monitoring tools for detection of the onset of flutter in the fan blades of an aircraft gas turbine engine. Both optical techniques give visual data in real time as well as video-tape records. PES is shown to be an ideal flutter monitor, since a single cathode ray tube displays the behavior of all the blades in a stage simultaneously. Operation of the SI system continuously while searching for a flutter condition imposes severe demands on the flash tube and affects its reliability, thus limiting its use as a flutter monitor. A better method of operation is to search for flutter with the PES and limit the use of SI to those times when the PES indicates interesting blade activity.

Topics in structural dynamics: Nonlinear unsteady transonic flows and Monte Carlo methods in acoustics

NASA Technical Reports Server (NTRS)

Haviland, J. K.

1974-01-01

The results are reported of two unrelated studies. The first was an investigation of the formulation of the equations for non-uniform unsteady flows, by perturbation of an irrotational flow to obtain the linear Green's equation. The resulting integral equation was found to contain a kernel which could be expressed as the solution of the adjoint flow equation, a linear equation for small perturbations, but with non-constant coefficients determined by the steady flow conditions. It is believed that the non-uniform flow effects may prove important in transonic flutter, and that in such cases, the use of doublet type solutions of the wave equation would then prove to be erroneous. The second task covered an initial investigation into the use of the Monte Carlo method for solution of acoustical field problems. Computed results are given for a rectangular room problem, and for a problem involving a circular duct with a source located at the closed end.

Flutter suppression of plates using passive constrained viscoelastic layers

NASA Astrophysics Data System (ADS)

Cunha-Filho, A. G.; de Lima, A. M. G.; Donadon, M. V.; Leão, L. S.

2016-10-01

Flutter in aeronautical panels is a self-excited aeroelastic phenomenon which occurs during supersonic flights due to dynamic instability of inertia, elastic and aerodynamic forces of the system. In the flutter condition, when the critical aerodynamic pressure is reached, the vibration amplitudes of the panel become dynamically unstable and increase exponentially with time, significantly affecting the fatigue life of the existing aeronautical components. Thus, in this paper, the interest is to investigate the possibility reducing the effects of the supersonic aeroelastic instability of rectangular plates by applying passive constrained viscoelastic layers. The rationale for such study is the fact that as the addition of viscoelastic materials provides decreased vibration amplitudes it becomes important to quantify the suppression of plate flutter coalescence modes that can be obtained. Moreover, despite the fact that much research on the suppression of panel flutter has been carried out by using passive, semi-active and active control techniques, few works have been proposed to deal with the problem of predicting the flutter boundary of aeroviscoelastic systems, since they must conveniently account for the frequency- and temperature-dependent behavior of the viscoelastic material. After the presentation of the theoretical foundations of the methodology, the description of a numerical study on the flutter analysis of a three-layer sandwich plate is addressed.

Sidewall Mach Number Distributions for the NASA Langley Transonic Dynamics Tunnel

NASA Technical Reports Server (NTRS)

Florance, James R.; Rivera, Jose A., Jr.

2001-01-01

The Transonic Dynamics Tunnel(TDT) was recalibrated due to the conversion of the heavy gas test medium from R-12 to R-134a. The objectives of the tests were to determine the relationship between the free-stream Mach number and the measured test section Mach number, and to quantify any necessary corrections. Other tests included the measurement of pressure distributions along the test-section walls, test-section centerline, at certain tunnel stations via a rake apparatus, and in the tunnel settling chamber. Wall boundary layer, turbulence, and flow angularity measurements were also performed. This paper discusses the determination of sidewall Mach number distributions.

A Study of Wing Flutter

NASA Technical Reports Server (NTRS)

Zahm, A F; Bear, R M

1929-01-01

Part I describes vibration tests, in a wind tunnel, of simple airfoils and of the tail plane of an M0-1 airplane model; it also describes the air flow about this model. From these tests are drawn inferences as to the cause and cure of aerodynamic wing vibrations. Part II derives stability criteria for wing vibrations in pitch and roll, and gives design rules to obviate instability. Part III shows how to design spars to flex equally under a given wing loading and thereby economically minimize the twisting in pitch that permits cumulative flutter. Resonant flutter is not likely to ensue from turbulence of air flow along past wings and tail planes in usual flying conditions. To be flutterproof a wing must be void of reversible autorotation and not have its centroid far aft of its pitching axis, i. e., axis of pitching motion. Danger of flutter is minimized by so proportioning the wing's torsional resisting moment to the air pitching moment at high-speed angles that the torsional flexure is always small. (author)

Flutter suppression by active control and its benefits

NASA Technical Reports Server (NTRS)

Doggett, R. V., Jr.; Townsend, J. C.

1976-01-01

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.

Study of flutter related computational procedures for minimum weight structural sizing of advanced aircraft

NASA Technical Reports Server (NTRS)

Oconnell, R. F.; Hassig, H. J.; Radovcich, N. A.

1976-01-01

Results of a study of the development of flutter modules applicable to automated structural design of advanced aircraft configurations, such as a supersonic transport, are presented. Automated structural design is restricted to automated sizing of the elements of a given structural model. It includes a flutter optimization procedure; i.e., a procedure for arriving at a structure with minimum mass for satisfying flutter constraints. Methods of solving the flutter equation and computing the generalized aerodynamic force coefficients in the repetitive analysis environment of a flutter optimization procedure are studied, and recommended approaches are presented. Five approaches to flutter optimization are explained in detail and compared. An approach to flutter optimization incorporating some of the methods discussed is presented. Problems related to flutter optimization in a realistic design environment are discussed and an integrated approach to the entire flutter task is presented. Recommendations for further investigations are made. Results of numerical evaluations, applying the five methods of flutter optimization to the same design task, are presented.

Blended-Wing-Body Transonic Aerodynamics: Summary of Ground Tests and Sample Results

NASA Technical Reports Server (NTRS)

Carter, Melissa B.; Vicroy, Dan D.; Patel, Dharmendra

2009-01-01

The Blended-Wing-Body (BWB) concept has shown substantial performance benefits over conventional aircraft configuration with part of the benefit being derived from the absence of a conventional empennage arrangement. The configuration instead relies upon a bank of trailing edge devices to provide control authority and augment stability. To determine the aerodynamic characteristics of the aircraft, several wind tunnel tests were conducted with a 2% model of Boeing's BWB-450-1L configuration. The tests were conducted in the NASA Langley Research Center's National Transonic Facility and the Arnold Engineering Development Center s 16-Foot Transonic Tunnel. Characteristics of the configuration and the effectiveness of the elevons, drag rudders and winglet rudders were measured at various angles of attack, yaw angles, and Mach numbers (subsonic to transonic speeds). The data from these tests will be used to develop a high fidelity simulation model for flight dynamics analysis and also serve as a reference for CFD comparisons. This paper provides an overview of the wind tunnel tests and examines the effects of Reynolds number, Mach number, pitch-pause versus continuous sweep data acquisition and compares the data from the two wind tunnels.

Support System Effects on the NASA Common Research Model

NASA Technical Reports Server (NTRS)

Rivers, S. Melissa B.; Hunter, Craig A.

2012-01-01

An experimental investigation of the NASA Common Research Model was conducted in the NASA Langley National Transonic Facility and NASA Ames 11-Foot Transonic Wind Tunnel Facility for use in the Drag Prediction Workshop. As data from the experimental investigations was collected, a large difference in moment values was seen between the experimental and the computational data from the 4th Drag Prediction Workshop. This difference led to the present work. In this study, a computational assessment has been undertaken to investigate model support system interference effects on the Common Research Model. The configurations computed during this investigation were the wing/body/tail=0deg without the support system and the wing/body/tail=0deg with the support system. The results from this investigation confirm that the addition of the support system to the computational cases does shift the pitching moment in the direction of the experimental results.

Quiet High Speed Fan (QHSF) Flutter Calculations Using the TURBO Code

NASA Technical Reports Server (NTRS)

Bakhle, Milind A.; Srivastava, Rakesh; Keith, Theo G., Jr.; Min, James B.; Mehmed, Oral

2006-01-01

A scale model of the NASA/Honeywell Engines Quiet High Speed Fan (QHSF) encountered flutter wind tunnel testing. This report documents aeroelastic calculations done for the QHSF scale model using the blade vibration capability of the TURBO code. Calculations at design speed were used to quantify the effect of numerical parameters on the aerodynamic damping predictions. This numerical study allowed the selection of appropriate values of these parameters, and also allowed an assessment of the variability in the calculated aerodynamic damping. Calculations were also done at 90 percent of design speed. The predicted trends in aerodynamic damping corresponded to those observed during testing.

Wind Tunnel Measurements for Flutter of a Long-Afterbody Bridge Deck

PubMed Central

Chen, Zeng-Shun; Zhang, Cheng; Wang, Xu; Ma, Cun-Ming

2017-01-01

Bridges are an important component of transportation. Flutter is a self-excited, large amplitude vibration, which may lead to collapse of bridges. It must be understood and avoided. This paper takes the Jianghai Channel Bridge, which is a significant part of the Hong Kong-Zhuhai-Macao Bridge, as an example to investigate the flutter of the bridge deck. Firstly, aerodynamic force models for flutter of bridges were introduced. Then, wind tunnel tests of the bridge deck during the construction and the operation stages, under different wind attack angles and wind velocities, were carried out using a high frequency base balance (HFBB) system and laser displacement sensors. From the tests, the static aerodynamic forces and flutter derivatives of the bridge deck were observed. Correspondingly, the critical flutter wind speeds of the bridge deck were determined based on the derivatives, and they are compared with the directly measured flutter speeds. Results show that the observed derivatives are reasonable and applicable. Furthermore, the critical wind speeds in the operation stage is smaller than those in the construction stage. Besides, the flutter instabilities of the bridge in the construction and the operation stages are good. This study helps guarantee the design and the construction of the Jianghai Channel Bridge, and advances the understanding of flutter of long afterbody bridge decks. PMID:28208773

Wind Tunnel Measurements for Flutter of a Long-Afterbody Bridge Deck.

PubMed

Chen, Zeng-Shun; Zhang, Cheng; Wang, Xu; Ma, Cun-Ming

2017-02-09

Bridges are an important component of transportation. Flutter is a self-excited, large amplitude vibration, which may lead to collapse of bridges. It must be understood and avoided. This paper takes the Jianghai Channel Bridge, which is a significant part of the Hong Kong-Zhuhai-Macao Bridge, as an example to investigate the flutter of the bridge deck. Firstly, aerodynamic force models for flutter of bridges were introduced. Then, wind tunnel tests of the bridge deck during the construction and the operation stages, under different wind attack angles and wind velocities, were carried out using a high frequency base balance (HFBB) system and laser displacement sensors. From the tests, the static aerodynamic forces and flutter derivatives of the bridge deck were observed. Correspondingly, the critical flutter wind speeds of the bridge deck were determined based on the derivatives, and they are compared with the directly measured flutter speeds. Results show that the observed derivatives are reasonable and applicable. Furthermore, the critical wind speeds in the operation stage is smaller than those in the construction stage. Besides, the flutter instabilities of the bridge in the construction and the operation stages are good. This study helps guarantee the design and the construction of the Jianghai Channel Bridge, and advances the understanding of flutter of long afterbody bridge decks.

Analysis of NASA Common Research Model Dynamic Data

NASA Technical Reports Server (NTRS)

Balakrishna, S.; Acheson, Michael J.

2011-01-01

Recent NASA Common Research Model (CRM) tests at the Langley National Transonic Facility (NTF) and Ames 11-foot Transonic Wind Tunnel (11-foot TWT) have generated an experimental database for CFD code validation. The database consists of force and moment, surface pressures and wideband wing-root dynamic strain/wing Kulite data from continuous sweep pitch polars. The dynamic data sets, acquired at 12,800 Hz sampling rate, are analyzed in this study to evaluate CRM wing buffet onset and potential CRM wing flow separation.

Bayesian analysis of the flutter margin method in aeroelasticity

DOE PAGES

Khalil, Mohammad; Poirel, Dominique; Sarkar, Abhijit

2016-08-27

A Bayesian statistical framework is presented for Zimmerman and Weissenburger flutter margin method which considers the uncertainties in aeroelastic modal parameters. The proposed methodology overcomes the limitations of the previously developed least-square based estimation technique which relies on the Gaussian approximation of the flutter margin probability density function (pdf). Using the measured free-decay responses at subcritical (preflutter) airspeeds, the joint non-Gaussain posterior pdf of the modal parameters is sampled using the Metropolis–Hastings (MH) Markov chain Monte Carlo (MCMC) algorithm. The posterior MCMC samples of the modal parameters are then used to obtain the flutter margin pdfs and finally the fluttermore » speed pdf. The usefulness of the Bayesian flutter margin method is demonstrated using synthetic data generated from a two-degree-of-freedom pitch-plunge aeroelastic model. The robustness of the statistical framework is demonstrated using different sets of measurement data. In conclusion, it will be shown that the probabilistic (Bayesian) approach reduces the number of test points required in providing a flutter speed estimate for a given accuracy and precision.« less

NASTRAN flutter analysis of advanced turbopropellers

NASA Technical Reports Server (NTRS)

Elchuri, V.; Smith, G. C. C.

1982-01-01

An existing capability developed to conduct modal flutter analysis of tuned bladed-shrouded discs in NASTRAN was modified and applied to investigate the subsonic unstalled flutter characteristics of advanced turbopropellers. The modifications pertain to the inclusion of oscillatory modal aerodynamic loads of blades with large (backward and forward) variable sweep. The two dimensional subsonic cascade unsteady aerodynamic theory was applied in a strip theory manner with appropriate modifications for the sweep effects. Each strip is associated with a chord selected normal to any spanwise reference curve such as the blade leading edge. The stability of three operating conditions of a 10-bladed propeller is analyzed. Each of these operating conditions is iterated once to determine the flutter boundary. A 5-bladed propeller is also analyzed at one operating condition to investigate stability. Analytical results obtained are in very good agreement with those from wind tunnel tests.

Ocular flutter following Zika virus infection.

PubMed

Karam, Emely; Giraldo, Jose; Rodriguez, Flor; Hernandez-Pereira, Carlos E; Rodriguez-Morales, Alfonso J; Blohm, Gabriela M; Paniz-Mondolfi, Alberto E

2017-12-01

Zika virus (ZIKV) is an emerging flavivirus which has been linked to a number of neurologic manifestations such as Guillain-Barré syndrome (GBS), transverse myelitis, and meningo-encephalitis. Ophthalmologic manifestations are increasingly being reported; however, ocular dyskinesias have not been described in this context to date. Herein, we report a case of a 22-year-old female who presented with ocular flutter and associated Guillain-Barré syndrome following acute ZIKV infection. We speculate that although such symptoms may have originated from a direct viral insult, a post-infectious autoimmune mechanism may not be excluded. Physicians should include ZIKV as well as other flaviviruses in their diagnostic workup for all patients with ocular flutter/opsoclonus, after excluding other non-infectious causes of central nervous system pathology. To the best of our knowledge, this is the first report on the association of ocular flutter, GBS, and ZIKV infection.

Optical Detection of Blade Flutter

NASA Technical Reports Server (NTRS)

Nieberding, W. C.; Pollack, J. L.

1977-01-01

Dynamic strain gages mounted on rotor blades are used as the primary instrumentation for detecting the onset of flutter and defining the vibratory mode and frequency. Optical devices are evaluated for performing the same measurements as well as providing supplementary information on the vibratory characteristics. Two separate methods are studied: stroboscopic imagery of the blade tip and photoelectric scanning of blade tip motion. Both methods give visual data in real time as well as video tape records. The optical systems are described, and representative results are presented. The potential of this instrumentation in flutter research is discussed.

Tilt-rotor flutter control in cruise flight

NASA Technical Reports Server (NTRS)

Nasu, Ken-Ichi

1986-01-01

Tilt-rotor flutter control under cruising operation is analyzed. The rotor model consists of a straight fixed wing, a pylon attached to the wingtip, and a three-blade rotor. The wing is cantilevered to the fuselage and is allowed to bend forward and upward. It also has a torsional degree of freedom about the elastic axis. Each rotor blade has two bending degrees of freedom. Feedback of wingtip velocity and acceleration to cyclic pitch is investigated for flutter control, using strip theory and linearized equations of motion. To determine the feedback gain, an eigenvalue analysis is performed. A second, independent, timewise calculation is conducted to evaluate the control law while employing more sophisticated aerodynamics. The effectiveness of flutter control by cyclic pitch change was confirmed.

Real-time flutter boundary prediction based on time series models

NASA Astrophysics Data System (ADS)

Gu, Wenjing; Zhou, Li

2018-03-01

For the purpose of predicting the flutter boundary in real time during flutter flight tests, two time series models accompanied with corresponding stability criterion are adopted in this paper. The first method simplifies a long nonstationary response signal as many contiguous intervals and each is considered to be stationary. The traditional AR model is then established to represent each interval of signal sequence. While the second employs a time-varying AR model to characterize actual measured signals in flutter test with progression variable speed (FTPVS). To predict the flutter boundary, stability parameters are formulated by the identified AR coefficients combined with Jury's stability criterion. The behavior of the parameters is examined using both simulated and wind-tunnel experiment data. The results demonstrate that both methods show significant effectiveness in predicting the flutter boundary at lower speed level. A comparison between the two methods is also given in this paper.

Transonic Flow Past Cone Cylinders

NASA Technical Reports Server (NTRS)

Solomon, George E

1955-01-01

Experimental results are presented for transonic flow post cone-cylinder, axially symmetric bodies. The drag coefficient and surface Mach number are studied as the free-stream Mach number is varied and, wherever possible, the experimental results are compared with theoretical predictions. Interferometric results for several typical flow configurations are shown and an example of shock-free supersonic-to-subsonic compression is experimentally demonstrated. The theoretical problem of transonic flow past finite cones is discussed briefly and an approximate solution of the axially symmetric transonic equations, valid for a semi-infinite cone, is presented.

A computer program for automated flutter solution and matched point determination

NASA Technical Reports Server (NTRS)

Bhatia, K. G.

1973-01-01

The use of a digital computer program (MATCH) for automated determination of the flutter velocity and the matched-point flutter density is described. The program is based on the use of the modified Laguerre iteration formula to converge to a flutter crossing or a matched-point density. A general description of the computer program is included and the purpose of all subroutines used is stated. The input required by the program and various input options are detailed, and the output description is presented. The program can solve flutter equations formulated with up to 12 vibration modes and obtain flutter solutions for up to 10 air densities. The program usage is illustrated by a sample run, and the FORTRAN program listing is included.

Enhancements to the FAST-MAC Circulation Control Model and Recent High-Reynolds Number Testing in the National Transonic Facility

NASA Technical Reports Server (NTRS)

Milholen, William E., II; Jones, Gregory S.; Chan, David T.; Goodliff, Scott L.; Anders, Scott G.; Melton, Latunia P.; Carter, Melissa B.; Allan, Brian G.; Capone, Francis J.

2013-01-01

A second wind tunnel test of the FAST-MAC circulation control model was recently completed in the National Transonic Facility at the NASA Langley Research Center. The model was equipped with four onboard flow control valves allowing independent control of the circulation control plenums, which were directed over a 15% chord simple-hinged flap. The model was configured for low-speed high-lift testing with flap deflections of 30 and 60 degrees, along with the transonic cruise configuration with zero degree flap deflection. Testing was again conducted over a wide range of Mach numbers up to 0.88, and Reynolds numbers up to 30 million based on the mean chord. The first wind tunnel test had poor transonic force and moment data repeatability at mild cryogenic conditions due to inadequate thermal conditioning of the balance. The second test demonstrated that an improvement to the balance heating system significantly improved the transonic data repeatability, but also indicated further improvements are still needed. The low-speed highlift performance of the model was improved by testing various blowing slot heights, and the circulation control was again demonstrated to be effective in re-attaching the flow over the wing at off-design transonic conditions. A new tailored spanwise blowing technique was also demonstrated to be effective at transonic conditions with the benefit of reduced mass flow requirements.

NASTRAN documentation for flutter analysis of advanced turbopropellers

NASA Technical Reports Server (NTRS)

Elchuri, V.; Gallo, A. M.; Skalski, S. C.

1982-01-01

An existing capability developed to conduct modal flutter analysis of tuned bladed-shrouded discs was modified to facilitate investigation of the subsonic unstalled flutter characteristics of advanced turbopropellers. The modifications pertain to the inclusion of oscillatory modal aerodynamic loads of blades with large (backward and forward) varying sweep.

Further investigation of a finite difference procedure for analyzing the transonic flow about harmonically oscillating airfoils and wings

NASA Technical Reports Server (NTRS)

Weatherill, W. H.; Ehlers, F. E.; Yip, E.; Sebastian, J. D.

1980-01-01

Analytical and empirical studies of a finite difference method for the solution of the transonic flow about harmonically oscillating wings and airfoils are presented. The procedure is based on separating the velocity potential into steady and unsteady parts and linearizing the resulting unsteady equations for small disturbances. The steady velocity potential is obtained first from the well-known nonlinear equation for steady transonic flow. The unsteady velocity potential is then obtained from a linear differential equation in complex form with spatially varying coefficients. Since sinusoidal motion is assumed, the unsteady equation is independent of time. An out-of-core direct solution procedure was developed and applied to two-dimensional sections. Results are presented for a section of vanishing thickness in subsonic flow and an NACA 64A006 airfoil in supersonic flow. Good correlation is obtained in the first case at values of Mach number and reduced frequency of direct interest in flutter analyses. Reasonable results are obtained in the second case. Comparisons of two-dimensional finite difference solutions with exact analytic solutions indicate that the accuracy of the difference solution is dependent on the boundary conditions used on the outer boundaries. Homogeneous boundary conditions on the mesh edges that yield complex eigenvalues give the most accurate finite difference solutions. The plane outgoing wave boundary conditions meet these requirements.

Aeroservoelastic Testing of Free Flying Wind Tunnel Models Part 1: A Sidewall Supported Semispan Model Tested for Gust Load Alleviation and Flutter Suppression

NASA Technical Reports Server (NTRS)

Scott, Robert C.; Vetter, Travis K.; Penning, Kevin B.; Coulson, David A.; Heeg, Jennifer.

2013-01-01

of a two part document. Part 2 is titled: "Aeroservoelastic Testing of Free Flying Wind Tunnel Models, Part 2: A Centerline Supported Fullspan Model Tested for Gust Load Alleviation." A team comprised of the Air Force Research Laboratory (AFRL), Northrop Grumman, Lockheed Martin, and the NASA Langley Research Center conducted three aeroservoelastic wind tunnel tests in the Transonic Dynamics Tunnel to demonstrate active control technologies relevant to large, flexible vehicles. In the first of these three tests, a semispan, aeroelastically scaled, wind tunnel model of a flying wing SensorCraft vehicle was mounted to a force balance to demonstrate gust load alleviation. In the second and third tests, the same wing was mated to a new, multi-degree of freedom, sidewall mount. This mount allowed the half-span model to translate vertically and pitch at the wing root, allowing better simulation of the full span vehicle's rigid body modes. Gust load alleviation (GLA) and Body freedom flutter (BFF) suppression were successfully demonstrated. The rigid body degrees-of-freedom required that the model be flown in the wind tunnel using an active control system. This risky mode of testing necessitated that a model arrestment system be integrated into the new mount. The safe and successful completion of these free flying tests required the development and integration of custom hardware and software. This paper describes the many systems, software, and procedures that were developed as part of this effort.

An Interactive Software for Conceptual Wing Flutter Analysis and Parametric Study

NASA Technical Reports Server (NTRS)

Mukhopadhyay, Vivek

1996-01-01

An interactive computer program was developed for wing flutter analysis in the conceptual design stage. The objective was to estimate the flutter instability boundary of a flexible cantilever wing, when well-defined structural and aerodynamic data are not available, and then study the effect of change in Mach number, dynamic pressure, torsional frequency, sweep, mass ratio, aspect ratio, taper ratio, center of gravity, and pitch inertia, to guide the development of the concept. The software was developed for Macintosh or IBM compatible personal computers, on MathCad application software with integrated documentation, graphics, data base and symbolic mathematics. The analysis method was based on non-dimensional parametric plots of two primary flutter parameters, namely Regier number and Flutter number, with normalization factors based on torsional stiffness, sweep, mass ratio, taper ratio, aspect ratio, center of gravity location and pitch inertia radius of gyration. The parametric plots were compiled in a Vought Corporation report from a vast data base of past experiments and wind-tunnel tests. The computer program was utilized for flutter analysis of the outer wing of a Blended-Wing-Body concept, proposed by McDonnell Douglas Corp. Using a set of assumed data, preliminary flutter boundary and flutter dynamic pressure variation with altitude, Mach number and torsional stiffness were determined.

Transonic aerodynamic characteristics of a proposed wing-body reusable launch vehicle concept

NASA Technical Reports Server (NTRS)

Springer, A. M.

1995-01-01

A proposed wing-body reusable launch vehicle was tested in the NASA Marshall Space Flight Center's 14 x 14-inch trisonic wind tunnel during the winter of 1994. This test resulted in the vehicle's subsonic and transonic, Mach 0.3 to 1.96, longitudinal and lateral aerodynamic characteristics. The effects of control surface deflections on the basic vehicle's aerodynamics, including a body flap, elevons, ailerons, and tip fins, are presented.

Adaptive Modal Identification for Flutter Suppression Control

NASA Technical Reports Server (NTRS)

Nguyen, Nhan T.; Drew, Michael; Swei, Sean S.

2016-01-01

In this paper, we will develop an adaptive modal identification method for identifying the frequencies and damping of a flutter mode based on model-reference adaptive control (MRAC) and least-squares methods. The least-squares parameter estimation will achieve parameter convergence in the presence of persistent excitation whereas the MRAC parameter estimation does not guarantee parameter convergence. Two adaptive flutter suppression control approaches are developed: one based on MRAC and the other based on the least-squares method. The MRAC flutter suppression control is designed as an integral part of the parameter estimation where the feedback signal is used to estimate the modal information. On the other hand, the separation principle of control and estimation is applied to the least-squares method. The least-squares modal identification is used to perform parameter estimation.

Langley 16- Ft. Transonic Tunnel Pressure Sensitive Paint System

NASA Technical Reports Server (NTRS)

Sprinkle, Danny R.; Obara, Clifford J.; Amer, Tahani R.; Leighty, Bradley D.; Carmine, Michael T.; Sealey, Bradley S.; Burkett, Cecil G.

2001-01-01

This report describes the NASA Langley 16-Ft. Transonic Tunnel Pressure Sensitive Paint (PSP) System and presents results of a test conducted June 22-23, 2000 in the tunnel to validate the PSP system. The PSP system provides global surface pressure measurements on wind tunnel models. The system was developed and installed by PSP Team personnel of the Instrumentation Systems Development Branch and the Advanced Measurement and Diagnostics Branch. A discussion of the results of the validation test follows a description of the system and a description of the test.

Transonic stability and control characteristics of a 0.015 scale model 69-0 of the space shuttle orbiter with forebody RSI modification in the NASA/LaRC 8 foot TPT (LA72)

NASA Technical Reports Server (NTRS)

Ball, J. W.; Edwards, C. R.

1976-01-01

Tests were conducted in the NASA/LaRC 8 foot transonic wind tunnel from March 26 through 31, 1976. The model was a 0.015 scale SSV Orbiter with forebody modifications to simulate slight reductions in the reusable surface insulation (RSI) thickness. Six component aerodynamic force and moment data were obtained at Mach numbers from 0.35 to 1.20 over an angle of attack range from -2 deg to 20 deg at sideslip angles of 0 deg and 5 deg.

Multi-fractality in aeroelastic response as a precursor to flutter

NASA Astrophysics Data System (ADS)

Venkatramani, J.; Nair, Vineeth; Sujith, R. I.; Gupta, Sayan; Sarkar, Sunetra

2017-01-01

Wind tunnel tests on a NACA 0012 airfoil have been carried out to study the transition in aeroelastic response from an initial state characterised by low-amplitude aperiodic fluctuations to aeroelastic flutter when the system exhibits limit cycle oscillations. An analysis of the aeroelastic measurements reveals multi-fractal characteristics in the pre-flutter regime. This has not been studied in the literature. As the flow velocity approaches the flutter velocity from below, a gradual loss in multi-fractality is observed. Measures based on the generalised Hurst exponents are developed and are shown to have the potential to warn against impending aeroelastic flutter. The results of this study could be useful for health monitoring of aeroelastic structures.

Supersonic Stall Flutter of High Speed Fans. [in turbofan engines

NASA Technical Reports Server (NTRS)

Adamczyk, J. J.; Stevens, W.; Jutras, R.

1981-01-01

An analytical model is developed for predicting the onset of supersonic stall bending flutter in axial flow compressors. The analysis is based on a modified two dimensional, compressible, unsteady actuator disk theory. It is applied to a rotor blade row by considering a cascade of airfoils whose geometry and dynamic response coincide with those of a rotor blade element at 85 percent of the span height (measured from the hub). The rotor blades are assumed to be unshrouded (i.e., free standing) and to vibrate in their first flexural mode. The effects of shock waves and flow separation are included in the model through quasi-steady, empirical, rotor total-pressure-loss and deviation-angle correlations. The actuator disk model predicts the unsteady aerodynamic force acting on the cascade blading as a function of the steady flow field entering the cascade and the geometry and dynamic response of the cascade. Calculations show that the present model predicts the existence of a bending flutter mode at supersonic inlet Mach numbers. This flutter mode is suppressed by increasing the reduced frequency of the system or by reducing the steady state aerodynamic loading on the cascade. The validity of the model for predicting flutter is demonstrated by correlating the measured flutter boundary of a high speed fan stage with its predicted boundary. This correlation uses a level of damping for the blade row (i.e., the log decrement of the rotor system) that is estimated from the experimental flutter data. The predicted flutter boundary is shown to be in good agreement with the measured boundary.

Propeller Tip Flutter

NASA Technical Reports Server (NTRS)

Liebers, Fritz

1932-01-01

The present report is limited to a case of tip flutter recognized by experience as being important. It is the case where outside interferences force vibrations upon the propeller. Such interferences may be set up by the engine, or they may be the result of an unsymmetrical field of flow.

Evaluation of somatosensory cortical differences between flutter and vibration tactile stimuli.

PubMed

Han, Sang Woo; Chung, Yoon Gi; Kim, Hyung-Sik; Chung, Soon-Cheol; Park, Jang-Yeon; Kim, Sung-Phil

2013-01-01

In parallel with advances in haptic-based mobile computing systems, understanding of the neural processing of vibrotactile information becomes of great importance. In the human nervous system, two types of vibrotactile information, flutter and vibration, are delivered from mechanoreceptors to the somatosensory cortex through segregated neural afferents. To investigate how the somatosensory cortex differentiates flutter and vibration, we analyzed the cortical responses to vibrotactile stimuli with a wide range of frequencies. Specifically, we examined whether cortical activity changed most around 50 Hz, which is known as a boundary between flutter and vibration. We explored various measures to evaluate separability of cortical activity across frequency and found that the hypothesis margin method resulted in the greatest separability between flutter and vibration. This result suggests that flutter and vibration information may be processed by different neural processes in the somatosensory cortex.

Study of boundary-layer transition using transonic-cone preston tube data

NASA Technical Reports Server (NTRS)

Reed, T. D.; Moretti, P. M.

1980-01-01

The laminar boundary layer on a 10 degree cone in a transonic wind tunnel was studied. The inviscid flow and boundary layer development were simulated by computer programs. The effects of pitch and yaw angles on the boundary layer were examined. Preston-tube data, taken on the boundary-layer-transition cone in the NASA Ames 11 ft transonic wind tunnel, were used to develope a correlation which relates the measurements to theoretical values of laminar skin friction. The recommended correlation is based on a compressible form of the classical law-of-the-wall. The computer codes successfully simulates the laminar boundary layer for near-zero pitch and yaw angles. However, in cases of significant pitch and/or yaw angles, the flow is three dimensional and the boundary layer computer code used here cannot provide a satisfactory model. The skin-friction correlation is thought to be valid for body geometries other than cones.

Large Scale Flutter Data for Design of Rotating Blades Using Navier-Stokes Equations

NASA Technical Reports Server (NTRS)

Guruswamy, Guru P.

2012-01-01

A procedure to compute flutter boundaries of rotating blades is presented; a) Navier-Stokes equations. b) Frequency domain method compatible with industry practice. Procedure is initially validated: a) Unsteady loads with flapping wing experiment. b) Flutter boundary with fixed wing experiment. Large scale flutter computation is demonstrated for rotating blade: a) Single job submission script. b) Flutter boundary in 24 hour wall clock time with 100 cores. c) Linearly scalable with number of cores. Tested with 1000 cores that produced data in 25 hrs for 10 flutter boundaries. Further wall-clock speed-up is possible by performing parallel computations within each case.

Transonic Resonance Demonstrated To Be a Source of Internal Noise From Mixer-Ejector Nozzles

NASA Technical Reports Server (NTRS)

Zaman, Khairul B.

2002-01-01

During noise field studies with mixer-ejector nozzles in NASA's High-Speed Research program, tones were often encountered. The tones would persist in the simulated "cutback" condition (shortly after takeoff). Unfortunately, we did not understand their origin and, thus, could not develop a logical approach for suppressing them. We naturally questioned whether or not some of those tones were due to the transonic resonance. This was studied with a 1/13th scale model of the High-Speed Civil Transport nozzle. The first objective was to determine if indeed tones could be detected in the radiated noise. The next objective was to diagnose if those tones were due to the transonic resonance. Agreement of the frequencies with the correlation equation and the effect of boundary layer tripping were to be used in the diagnosis.

Supersonic unstalled flutter. [aerodynamic loading of thin airfoils induced by cascade motion

NASA Technical Reports Server (NTRS)

Adamczyk, J. J.; Goldstein, M. E.; Hartmann, M. J.

1978-01-01

Flutter analyses were developed to predict the onset of supersonic unstalled flutter of a cascade of two-dimensional airfoils. The first of these analyzes the onset of supersonic flutter at low levels of aerodynamic loading (i.e., backpressure), while the second examines the occurrence of supersonic flutter at moderate levels of aerodynamic loading. Both of these analyses are based on the linearized unsteady inviscid equations of gas dynamics to model the flow field surrounding the cascade. These analyses are utilized in a parametric study to show the effects of cascade geometry, inlet Mach number, and backpressure on the onset of single and multi degree of freedom unstalled supersonic flutter. Several of the results are correlated against experimental qualitative observation to validate the models.

High Speed Civil Transport (HSCT) Isolated Nacelle Transonic Boattail Drag Study and Results Using Computational Fluid Dynamics (CFD)

NASA Technical Reports Server (NTRS)

Midea, Anthony C.; Austin, Thomas; Pao, S. Paul; DeBonis, James R.; Mani, Mori

2005-01-01

Nozzle boattail drag is significant for the High Speed Civil Transport (HSCT) and can be as high as 25 percent of the overall propulsion system thrust at transonic conditions. Thus, nozzle boattail drag has the potential to create a thrust drag pinch and can reduce HSCT aircraft aerodynamic efficiencies at transonic operating conditions. In order to accurately predict HSCT performance, it is imperative that nozzle boattail drag be accurately predicted. Previous methods to predict HSCT nozzle boattail drag were suspect in the transonic regime. In addition, previous prediction methods were unable to account for complex nozzle geometry and were not flexible enough for engine cycle trade studies. A computational fluid dynamics (CFD) effort was conducted by NASA and McDonnell Douglas to evaluate the magnitude and characteristics of HSCT nozzle boattail drag at transonic conditions. A team of engineers used various CFD codes and provided consistent, accurate boattail drag coefficient predictions for a family of HSCT nozzle configurations. The CFD results were incorporated into a nozzle drag database that encompassed the entire HSCT flight regime and provided the basis for an accurate and flexible prediction methodology.

High Speed Civil Transport (HSCT) Isolated Nacelle Transonic Boattail Drag Study and Results Using Computational Fluid Dynamics (CFD)

NASA Technical Reports Server (NTRS)

Midea, Anthony C.; Austin, Thomas; Pao, S. Paul; DeBonis, James R.; Mani, Mori

1999-01-01

Nozzle boattail drag is significant for the High Speed Civil Transport (HSCT) and can be as high as 25% of the overall propulsion system thrust at transonic conditions. Thus, nozzle boattail drag has the potential to create a thrust-drag pinch and can reduce HSCT aircraft aerodynamic efficiencies at transonic operating conditions. In order to accurately predict HSCT performance, it is imperative that nozzle boattail drag be accurately predicted. Previous methods to predict HSCT nozzle boattail drag were suspect in the transonic regime. In addition, previous prediction methods were unable to account for complex nozzle geometry and were not flexible enough for engine cycle trade studies. A computational fluid dynamics (CFD) effort was conducted by NASA and McDonnell Douglas to evaluate the magnitude and characteristics of HSCT nozzle boattail drag at transonic conditions. A team of engineers used various CFD codes and provided consistent, accurate boattail drag coefficient predictions for a family of HSCT nozzle configurations. The CFD results were incorporated into a nozzle drag database that encompassed the entire HSCT flight regime and provided the basis for an accurate and flexible prediction methodology.

SWIFT Code Assessment for Two Similar Transonic Compressors

NASA Technical Reports Server (NTRS)

Chima, Rodrick V.

2009-01-01

One goal of the NASA Fundamental Aeronautics Program is the assessment of computational fluid dynamic (CFD) codes used for the design and analysis of many aerospace systems. This paper describes the assessment of the SWIFT turbomachinery analysis code for two similar transonic compressors, NASA rotor 37 and stage 35. The two rotors have identical blade profiles on the front, transonic half of the blade but rotor 37 has more camber aft of the shock. Thus the two rotors have the same shock structure and choking flow but rotor 37 produces a higher pressure ratio. The two compressors and experimental data are described here briefly. Rotor 37 was also used for test cases organized by ASME, IGTI, and AGARD in 1994-1998. Most of the participating codes over predicted pressure and temperature ratios, and failed to predict certain features of the downstream flowfield. Since then the AUSM+ upwind scheme and the k- turbulence model have been added to SWIFT. In this work the new capabilities were assessed for the two compressors. Comparisons were made with overall performance maps and spanwise profiles of several aerodynamic parameters. The results for rotor 37 were in much better agreement with the experimental data than the original blind test case results although there were still some discrepancies. The results for stage 35 were in very good agreement with the data. The results for rotor 37 were very sensitive to turbulence model parameters but the results for stage 35 were not. Comparison of the rotor solutions showed that the main difference between the two rotors was not blade camber as expected, but shock/boundary layer interaction on the casing.

HSCT Ref-H Transonic Flap Data Base: Wind-Tunnel Test and Comparison with Theory

NASA Technical Reports Server (NTRS)

Vijgen, Paul M.

1999-01-01

In cooperation with personnel from the Boeing ANP Laboratory and NASA Langley, a performance test was conducted using the Reference-H 1.675% model ("NASA Modular Model") without nacelles at the NASA Langley 16-Ft Transonic Tunnel. The main objective of the test was to determine the drag reduction achievable with leading-edge and trailing-edge flaps deflected along the outboard wing span at transonic Mach numbers (M = 0.9 to 1.2) for purpose of preliminary design and for comparison with computational predictions. The obtained drag data with flap deflections for Mach numbers of 1.07 to 1.20 are unique for the Reference H wing. Four leading-edge and two trailing-edge flap deflection angles were tested at a mean-wing chord-Reynolds number of about 5.7 million. An outboard-wing leading-edge flap deflection of 81 provides a 4.5 percent drag reduction at M = 1.2 A = 0.2), and much larger values at lower Mach numbers with larger flap deflections. The present results for the baseline (no flaps deflected) compare reasonably well with previous Boeing and NASA Ref-H tunnel tests, including high-Reynolds number NTF results. Viscous CFD simulations using the OVERFLOW thin-layer N.S. method properly predict the observed trend in drag reduction at M = 1.2 as function of leading-edge flap deflection. Modified linear theory properly predicts the flap effects on drag at subsonic conditions (Aero2S code), and properly predicts the absolute drag for the 40 and 80 leading-edge deflection at M = 1.2 (A389 code).

Coronary Sinus Activation and ECG Characteristics of Roof-Dependent Left Atrial Flutter After Pulmonary Vein Isolation.

PubMed

Casado Arroyo, Ruben; Laţcu, Decebal Gabriel; Maeda, Shingo; Kubala, Maciej; Santangeli, Pasquale; Garcia, Fermin Carlos; Enache, Bogdan; Eljamili, Mohammed; Hayashi, Tatsuya; Zado, Erica S; Saoudi, Nadir; Marchlinski, Francis E

2018-06-01

The electrocardiographic and intracardiac activation features of left atrial roof-dependent macroreentrant flutter have been incompletely characterized. Patients post-pulmonary vein (PV) isolation with roof-dependent atrial flutter based on activation and entrainment mapping were included. ECG and coronary sinus activation were compared with mitral annular (MA) flutter. The roof-dependent left atrial flutter circled the right PVs in 32 of 33 cases. Two forms of roof flutters were identified, posteroanterior, ascendant on posterior wall and descendant on anterior wall (n=24); and anteroposterior, ascendant on the anterior wall and descendent on the posterior wall (n=9). Both forms had positive large amplitude P waves in V 1 through V 2 with decreasing amplitude in V 3 through V 6 . Posteroanterior roof flutters had positive P wave in the inferior and negative P wave in leads I and aVL similar to counterclockwise MA flutter, but coronary sinus activation was simultaneous for roof and proximal to distal for counterclockwise. Anteroposterior roof flutters were similar to clockwise MA flutter with negative P in inferior leads and transition to flat or negative P in V 3 through V 6 . Coronary sinus activation time ≤39 ms identified roof versus MA flutter (sensitivity: 100% and specificity: 97%). Roof-dependent flutter around right PVs is more common than around left PVs. The ECG pattern for roof-dependent flutter around right PVs is similar to MA flutter with frontal plane axis dictated by septal activation. Roof-dependent flutter can be distinguished from MA flutter by more simultaneous rather than sequential coronary sinus activation. © 2018 American Heart Association, Inc.

The design, analysis, and testing of a low-budget wind-tunnel flutter model with active aerodynamic controls

NASA Technical Reports Server (NTRS)

Bolding, R. M.; Stearman, R. O.

1976-01-01

A low budget flutter model incorporating active aerodynamic controls for flutter suppression studies was designed as both an educational and research tool to study the interfering lifting surface flutter phenomenon in the form of a swept wing-tail configuration. A flutter suppression mechanism was demonstrated on a simple semirigid three-degree-of-freedom flutter model of this configuration employing an active stabilator control, and was then verified analytically using a doublet lattice lifting surface code and the model's measured mass, mode shapes, and frequencies in a flutter analysis. Preliminary studies were significantly encouraging to extend the analysis to the larger degree of freedom AFFDL wing-tail flutter model where additional analytical flutter suppression studies indicated significant gains in flutter margins could be achieved. The analytical and experimental design of a flutter suppression system for the AFFDL model is presented along with the results of a preliminary passive flutter test.

Further studies of stall flutter and nonlinear divergence of two-dimensional wings

NASA Technical Reports Server (NTRS)

Dugundji, J.; Chopra, I.

1975-01-01

An experimental investigation is made of the purely torsional stall flutter of a two-dimensional wing pivoted about the midchord, and also of the bending-torsion stall flutter of a two-dimensional wing pivoted about the quarterchord. For the purely torsional flutter case, large amplitude limit cycles ranging from + or - 11 to + or - 160 degrees were observed. Nondimensional harmonic coefficients were extracted from the free transient vibration tests for amplitudes up to 80 degrees. Reasonable nondimensional correlation was obtained for several wing configurations. For the bending-torsion flutter case, large amplitude coupled limit cycles were observed with torsional amplitudes as large as + or - 40 degrees. The torsion amplitudes first increased, then decreased with increasing velocity. Additionally, a small amplitude, predominantly torsional flutter was observed when the static equilibrium angle was near the stall angle.

Applications of potential theory computations to transonic aeroelasticity

NASA Technical Reports Server (NTRS)

Edwards, J. W.

1986-01-01

Unsteady aerodynamic and aeroelastic stability calculations based upon transonic small disturbance (TSD) potential theory are presented. Results from the two-dimensional XTRAN2L code and the three-dimensional XTRAN3S code are compared with experiment to demonstrate the ability of TSD codes to treat transonic effects. The necessity of nonisentropic corrections to transonic potential theory is demonstrated. Dynamic computational effects resulting from the choice of grid and boundary conditions are illustrated. Unsteady airloads for a number of parameter variations including airfoil shape and thickness, Mach number, frequency, and amplitude are given. Finally, samples of transonic aeroelastic calculations are given. A key observation is the extent to which unsteady transonic airloads calculated by inviscid potential theory may be treated in a locally linear manner.

Sensitivity Analysis of Flutter Response of a Wing Incorporating Finite-Span Corrections

NASA Technical Reports Server (NTRS)

Issac, Jason Cherian; Kapania, Rakesh K.; Barthelemy, Jean-Francois M.

1994-01-01

Flutter analysis of a wing is performed in compressible flow using state-space representation of the unsteady aerodynamic behavior. Three different expressions are used to incorporate corrections due to the finite-span effects of the wing in estimating the lift-curve slope. The structural formulation is based on a Rayleigh-Pitz technique with Chebyshev polynomials used for the wing deflections. The aeroelastic equations are solved as an eigen-value problem to determine the flutter speed of the wing. The flutter speeds are found to be higher in these cases, when compared to that obtained without accounting for the finite-span effects. The derivatives of the flutter speed with respect to the shape parameters, namely: aspect ratio, area, taper ratio and sweep angle, are calculated analytically. The shape sensitivity derivatives give a linear approximation to the flutter speed curves over a range of values of the shape parameter which is perturbed. Flutter and sensitivity calculations are performed on a wing using a lifting-surface unsteady aerodynamic theory using modules from a system of programs called FAST.

Development and Flight Test of an Active Flutter Suppression System for the F-4F with Stores. Part I. Design of the Active Flutter Suppression System.

DTIC Science & Technology

1982-09-01

of the wing-pylon-store changed considerably with excitation amplitude due to free play and preload. The active flutter suppression system worked well and provided an increase in flutter speed. (Author)

Trailing Edge Blowing on a Two-Dimensional Six-Percent Thick Elliptical Circulation Control Airfoil Up to Transonic Conditions

NASA Technical Reports Server (NTRS)

Alexander, Michael G.; Anders, Scott G.; Johnson, Stuart K.; Florance, Jennifer P.; Keller, Donald F.

2005-01-01

A wind tunnel test was conducted in the NASA Langley Transonic Dynamics Tunnel (TDT) on a six percent thick slightly cambered elliptical circulation control airfoil with both upper and lower surface blowing capability. Parametric evaluations of jet slot heights and Coanda surface shapes were conducted at momentum coefficients (Cm) from 0.0 to 0.12. Test data were acquired at Mach numbers of 0.3, 0.5, 0.7, 0.8, and 0.84 at Reynolds numbers per foot of 2.43 x 105 to 1.05 x 106. For a transonic condition, (Mach = 0.8 at alpha = 3 degrees), it was generally found the smaller slot and larger Coanda surface combination was overall more effective than other slot/Coanda surface combinations. Lower surface blowing was not as effective as the upper surface blowing over the same range of momentum coefficients. No appreciable Coanda surface, slot height, or slot blowing position preference was indicated transonically with the dual slot blowing.

Application of Linear and Non-Linear Harmonic Methods for Unsteady Transonic Flow

NASA Astrophysics Data System (ADS)

Gundevia, Rayomand

This thesis explores linear and non-linear computational methods for solving unsteady flow. The eventual goal is to apply these methods to two-dimensional and three-dimensional flutter predictions. In this study the quasi-one-dimensional nozzle is used as a framework for understanding these methods and their limitations. Subsonic and transonic cases are explored as the back-pressure is forced to oscillate with known amplitude and frequency. A steady harmonic approach is used to solve this unsteady problem for which perturbations are said to be small in comparison to the mean flow. The use of a linearized Euler equations (LEE) scheme is good at capturing the flow characteristics but is limited by accuracy to relatively small amplitude perturbations. The introduction of time-averaged second-order terms in the Non-Linear Harmonic (NLH) method means that a better approximation of the mean-valued solution, upon which the linearization is based, can be made. The nonlinear time-accurate Euler solutions are used for comparison and to establish the regimes of unsteadiness for which these schemes fails. The usefulness of the LEE and NLH methods lie in the gains in computational efficiency over the full equations.

Modernization and Activation of the NASA Ames 11- by 11-Foot Transonic Wind Tunnel

NASA Technical Reports Server (NTRS)

Kmak, Frank J.

2000-01-01

The Unitary Plan Wind Tunnel (UPWT) was modernized to improve performance, capability, productivity, and reliability. Automation systems were installed in all three UPWT tunnel legs and the Auxiliaries facility. Major improvements were made to the four control rooms, model support systems, main drive motors, and main drive speed control. Pressure vessel repairs and refurbishment to the electrical distribution system were also completed. Significant changes were made to improve test section flow quality in the 11-by 11-Foot Transonic leg. After the completion of the construction phase of the project, acceptance and checkout testing was performed to demonstrate the capabilities of the modernized facility. A pneumatic test of the tunnel circuit was performed to verify the structural integrity of the pressure vessel before wind-on operations. Test section turbulence, flow angularity, and acoustic parameters were measured throughout the tunnel envelope to determine the effects of the tunnel flow quality improvements. The new control system processes were thoroughly checked during wind-off and wind-on operations. Manual subsystem modes and automated supervisory modes of tunnel operation were validated. The aerodynamic and structural performance of both the new composite compressor rotor blades and the old aluminum rotor blades was measured. The entire subsonic and supersonic envelope of the 11-by 11-Foot Transonic leg was defined up to the maximum total pressure.

Computational Aeroelastic Analysis of the Semi-Span Super-Sonic Transport (S4T) Wind-Tunnel Model

NASA Technical Reports Server (NTRS)

Sanetrik, Mark D.; Silva, Walter A.; Hur, Jiyoung

2012-01-01

A summary of the computational aeroelastic analysis for the Semi-Span Super-Sonic Transport (S4T) wind-tunnel model is presented. A broad range of analysis techniques, including linear, nonlinear and Reduced Order Models (ROMs) were employed in support of a series of aeroelastic (AE) and aeroservoelastic (ASE) wind-tunnel tests conducted in the Transonic Dynamics Tunnel (TDT) at NASA Langley Research Center. This research was performed in support of the ASE element in the Supersonics Program, part of NASA's Fundamental Aeronautics Program. The analysis concentrated on open-loop flutter predictions, which were in good agreement with experimental results. This paper is one in a series that comprise a special S4T technical session, which summarizes the S4T project.

Overview of military technology at NASA Langley

NASA Technical Reports Server (NTRS)

Sawyer, Wallace C.; Jackson, Charlie M., Jr.

1989-01-01

The Langley Research Center began addressing major research topics pertinent to the design of military aircraft under the egis of The National Advisory Council on Aeronautics in 1917, until 1958, when it passed under the control of the newly-instituted NASA research facilities system. A historical account is presented of NASA-Langley's involvement in the experimental investigation of twin-engined jet aircraft nozzle interfairings, thrust reversers, high-efficiency supersonic cruise configurations, high-alpha aerodynamics, air-to-air combat handling qualities, wing/stores flutter suppression, and store carriage and separation characteristics.

Developing Uncertainty Models for Robust Flutter Analysis Using Ground Vibration Test Data

NASA Technical Reports Server (NTRS)

Potter, Starr; Lind, Rick; Kehoe, Michael W. (Technical Monitor)

2001-01-01

A ground vibration test can be used to obtain information about structural dynamics that is important for flutter analysis. Traditionally, this information#such as natural frequencies of modes#is used to update analytical models used to predict flutter speeds. The ground vibration test can also be used to obtain uncertainty models, such as natural frequencies and their associated variations, that can update analytical models for the purpose of predicting robust flutter speeds. Analyzing test data using the -norm, rather than the traditional 2-norm, is shown to lead to a minimum-size uncertainty description and, consequently, a least-conservative robust flutter speed. This approach is demonstrated using ground vibration test data for the Aerostructures Test Wing. Different norms are used to formulate uncertainty models and their associated robust flutter speeds to evaluate which norm is least conservative.

Flutter analysis of swept-wing subsonic aircraft with parameter studies of composite wings

NASA Technical Reports Server (NTRS)

Housner, J. M.; Stein, M.

1974-01-01

A computer program is presented for the flutter analysis, including the effects of rigid-body roll, pitch, and plunge of swept-wing subsonic aircraft with a flexible fuselage and engines mounted on flexible pylons. The program utilizes a direct flutter solution in which the flutter determinant is derived by using finite differences, and the root locus branches of the determinant are searched for the lowest flutter speed. In addition, a preprocessing subroutine is included which evaluates the variable bending and twisting stiffness properties of the wing by using a laminated, balanced ply, filamentary composite plate theory. The program has been substantiated by comparisons with existing flutter solutions. The program has been applied to parameter studies which examine the effect of filament orientation upon the flutter behavior of wings belonging to the following three classes: wings having different angles of sweep, wings having different mass ratios, and wings having variable skin thicknesses. These studies demonstrated that the program can perform a complete parameter study in one computer run. The program is designed to detect abrupt changes in the lowest flutter speed and mode shape as the parameters are varied.

Labyrinth Seal Flutter Analysis and Test Validation in Support of Robust Rocket Engine Design

NASA Technical Reports Server (NTRS)

El-Aini, Yehia; Park, John; Frady, Greg; Nesman, Tom

2010-01-01

High energy-density turbomachines, like the SSME turbopumps, utilize labyrinth seals, also referred to as knife-edge seals, to control leakage flow. The pressure drop for such seals is order of magnitude higher than comparable jet engine seals. This is aggravated by the requirement of tight clearances resulting in possible unfavorable fluid-structure interaction of the seal system (seal flutter). To demonstrate these characteristics, a benchmark case of a High Pressure Oxygen Turbopump (HPOTP) outlet Labyrinth seal was studied in detail. First, an analytical assessment of the seal stability was conducted using a Pratt & Whitney legacy seal flutter code. Sensitivity parameters including pressure drop, rotor-to-stator running clearances and cavity volumes were examined and modeling strategies established. Second, a concurrent experimental investigation was undertaken to validate the stability of the seal at the equivalent operating conditions of the pump. Actual pump hardware was used to construct the test rig, also referred to as the (Flutter Rig). The flutter rig did not include rotational effects or temperature. However, the use of Hydrogen gas at high inlet pressure provided good representation of the critical parameters affecting flutter especially the speed of sound. The flutter code predictions showed consistent trends in good agreement with the experimental data. The rig test program produced a stability threshold empirical parameter that separated operation with and without flutter. This empirical parameter was used to establish the seal build clearances to avoid flutter while providing the required cooling flow metering. The calibrated flutter code along with the empirical flutter parameter was used to redesign the baseline seal resulting in a flutter-free robust configuration. Provisions for incorporation of mechanical damping devices were introduced in the redesigned seal to ensure added robustness

Results of the AFRSI rewaterproofing systems screening test in the NASA/Ames Research Center (ARC) 2 x 2-foot transonic wind tunnel

NASA Technical Reports Server (NTRS)

Marroquin, J.; Kingsland, R. B.

1985-01-01

An experimental investigation was conducted in the NASA/Ames Research Center 2x2-foot Transonic Wind Tunnel to evaluate two AFRSI rewaterproofing systems and to investigate films as a means of reducing blanket joint distortion. The wind tunnel wall slot configuration influenced on the flow field over the test panel was investigated; primarily using oil flow data, and resulted in a closed slot configuration to provide a satisfactory screening environment flow field for the test. Sixteen AFRSI test panels, configured to represent the test system or film, were subjected to this screening environment (a flow field of separated and reattached flow at a freestream Mach numnber of 0.65 and q = 650 or 900 psf). Each condition was held until damage to the test article was observed or 55 minutes if no damage was incurred. All objectives related to AFRSI rewaterproofing and to the use of films to stiffen the blanket fibers were achieved.

Worst-Case Flutter Margins from F/A-18 Aircraft Aeroelastic Data

NASA Technical Reports Server (NTRS)

Lind, Rick; Brenner, Marty

1997-01-01

An approach for computing worst-case flutter margins has been formulated in a robust stability framework. Uncertainty operators are included with a linear model to describe modeling errors and flight variations. The structured singular value, micron, computes a stability margin which directly accounts for these uncertainties. This approach introduces a new method of computing flutter margins and an associated new parameter for describing these margins. The micron margins are robust margins which indicate worst-case stability estimates with respect to the defined uncertainty. Worst-case flutter margins are computed for the F/A-18 SRA using uncertainty sets generated by flight data analysis. The robust margins demonstrate flight conditions for flutter may lie closer to the flight envelope than previously estimated by p-k analysis.

Analysis of stall flutter of a helicopter radar blade

NASA Technical Reports Server (NTRS)

Crimi, P.

1973-01-01

A study of rotor blade aeroelastic stability was carried out, using an analytic model of a two-dimensional airfoil undergoing dynamic stall and an elastomechanical representation including flapping, flapwise bending and torsional degrees of freedom. Results for a hovering rotor demonstrated that the models used are capable of reproducing both classical and stall flutter. The minimum rotor speed for the occurrence of stall flutter in hover, was found to be determined from coupling between torsion and flapping. Instabilities analogous to both classical and stall flutter were found to occur in forward flight. However, the large stall-related torsional oscillations which commonly limit aircraft forward speed appear to be the response to rapid changes in aerodynamic moment which accompany stall and unstall, rather than the result of an aeroelastic instability. The severity of stall-related instabilities and response was found to depend to some extent on linear stability. Increasing linear stability lessens the susceptibility to stall flutter and reduced the magnitude of the torsional response to stall and unstall.

A Nonlinear Modal Aeroelastic Solver for FUN3D

NASA Technical Reports Server (NTRS)

Goldman, Benjamin D.; Bartels, Robert E.; Biedron, Robert T.; Scott, Robert C.

2016-01-01

A nonlinear structural solver has been implemented internally within the NASA FUN3D computational fluid dynamics code, allowing for some new aeroelastic capabilities. Using a modal representation of the structure, a set of differential or differential-algebraic equations are derived for general thin structures with geometric nonlinearities. ODEPACK and LAPACK routines are linked with FUN3D, and the nonlinear equations are solved at each CFD time step. The existing predictor-corrector method is retained, whereby the structural solution is updated after mesh deformation. The nonlinear solver is validated using a test case for a flexible aeroshell at transonic, supersonic, and hypersonic flow conditions. Agreement with linear theory is seen for the static aeroelastic solutions at relatively low dynamic pressures, but structural nonlinearities limit deformation amplitudes at high dynamic pressures. No flutter was found at any of the tested trajectory points, though LCO may be possible in the transonic regime.

Supersonic Panel Flutter Test Results for Flat Fiber-Glass Sandwich Panels with Foamed Cores

NASA Technical Reports Server (NTRS)

Tuovila, W. J.; Presnell, John G., Jr.

1961-01-01

Flutter tests have been made on flat panels having a 1/4 inch-thick plastic-foam core covered with thin fiber-glass laminates. The testing was done in the Langley Unitary Plan wind tunnel at Mach numbers from 1.76 t o 2.87. The flutter boundary for these panels was found to be near the flutter boundary of thin metal panels when compared on the basis of an equivalent panel stiffness. The results also demonstrated that the depth of the cavity behind the panel has a pronounced influence on flutter. Changing the cavity depth from 1 1/2 inches to 1/2 inch reduced the dynamic pressure at start of flutter by 40 percent. No flutter was obtained when the spacers on the back of the panel were against the bottom of the cavity.

Panel-flutter analysis of a thermal protection-shield concept for the space shuttle.

NASA Technical Reports Server (NTRS)

Cunningham, H. J.

1972-01-01

Analysis of the panel flutter characteristics of a candidate thermal protection system (TPS) for the space shuttle, using piston theory aerodynamics and Lagrange equations. The results show the TPS candidate panel array to be deep in the 'no-flutter' region during launch and, therefore, safe from panel flutter.

TAIR: A transonic airfoil analysis computer code

NASA Technical Reports Server (NTRS)

Dougherty, F. C.; Holst, T. L.; Grundy, K. L.; Thomas, S. D.

1981-01-01

The operation of the TAIR (Transonic AIRfoil) computer code, which uses a fast, fully implicit algorithm to solve the conservative full-potential equation for transonic flow fields about arbitrary airfoils, is described on two levels of sophistication: simplified operation and detailed operation. The program organization and theory are elaborated to simplify modification of TAIR for new applications. Examples with input and output are given for a wide range of cases, including incompressible, subcritical compressible, and transonic calculations.

Flutter performance of bend-twist coupled large-scale wind turbine blades

NASA Astrophysics Data System (ADS)

Hayat, Khazar; de Lecea, Alvaro Gorostidi Martinez; Moriones, Carlos Donazar; Ha, Sung Kyu

2016-05-01

The bend-twist coupling (BTC) is proven to be effective in mitigating the fatigue loads for large-scale wind turbine blades, but at the same time it may cause the risk of flutter instability. The BTC is defined as a feature of twisting of the blade induced by the primary bending deformation. In the classical flutter, the BTC arises from the aerodynamic loads changing with the angle of attack. In this study, the effects of the structural BTC on the flutter are investigated by considering the layup unbalances (ply angle, material and thickness of the composite laminates) in the NREL 5-MW wind turbine rotor blade of glass fiber/epoxy [02/+45/-45]S laminates. It is numerically shown that the flutter speed may decrease by about 5 percent with unbalanced ply-angle only (one side angle, from 45° to 25°). It was then demonstrated that the flutter performance of the wind turbine blade can be increased by using lighter and stiffer carbon fibers which ensures the higher structural BTC at the same time.

A curve fitting method for solving the flutter equation. M.S. Thesis

NASA Technical Reports Server (NTRS)

Cooper, J. L.

1972-01-01

A curve fitting approach was developed to solve the flutter equation for the critical flutter velocity. The psi versus nu curves are approximated by cubic and quadratic equations. The curve fitting technique utilized the first and second derivatives of psi with respect to nu. The method was tested for two structures, one structure being six times the total mass of the other structure. The algorithm never showed any tendency to diverge from the solution. The average time for the computation of a flutter velocity was 3.91 seconds on an IBM Model 50 computer for an accuracy of five per cent. For values of nu close to the critical root of the flutter equation the algorithm converged on the first attempt. The maximum number of iterations for convergence to the critical flutter velocity was five with an assumed value of nu relatively distant from the actual crossover.

Beam Flutter and Energy Harvesting in Internal Flow

NASA Astrophysics Data System (ADS)

Tosi, Luis Phillipe; Colonius, Tim; Sherrit, Stewart; Lee, Hyeong Jae

2017-11-01

Aeroelastic flutter, largely studied for causing engineering failures, has more recently been used as a means of extracting energy from the flow. Particularly, flutter of a cantilever or an elastically mounted plate in a converging-diverging flow passage has shown promise as an energy harvesting concept for internal flow applications. The instability onset is observed as a function of throat velocity, internal wall geometry, fluid and structure material properties. To enable these devices, our work explores features of the fluid-structure coupled dynamics as a function of relevant nondimensional parameters. The flutter boundary is examined through stability analysis of a reduced order model, and corroborated with numerical simulations at low Reynolds number. Experiments for an energy harvester design are qualitatively compared to results from analytical and numerical work, suggesting a robust limit cycle ensues due to a subcritical Hopf bifurcation. Bosch Corporation.

Improvements to the fastex flutter analysis computer code

NASA Technical Reports Server (NTRS)

Taylor, Ronald F.

1987-01-01

Modifications to the FASTEX flutter analysis computer code (UDFASTEX) are described. The objectives were to increase the problem size capacity of FASTEX, reduce run times by modification of the modal interpolation procedure, and to add new user features. All modifications to the program are operable on the VAX 11/700 series computers under the VAX operating system. Interfaces were provided to aid in the inclusion of alternate aerodynamic and flutter eigenvalue calculations. Plots can be made of the flutter velocity, display and frequency data. A preliminary capability was also developed to plot contours of unsteady pressure amplitude and phase. The relevant equations of motion, modal interpolation procedures, and control system considerations are described and software developments are summarized. Additional information documenting input instructions, procedures, and details of the plate spline algorithm is found in the appendices.

Long-term endurance sport is a risk factor for development of lone atrial flutter.

PubMed

Claessen, Guido; Colyn, Erwin; La Gerche, André; Koopman, Pieter; Alzand, Becker; Garweg, Christophe; Willems, Rik; Nuyens, Dieter; Heidbuchel, Hein

2011-06-01

To evaluate whether in a population of patients with 'lone atrial flutter', the proportion of those engaged in long-term endurance sports is higher than that observed in the general population. An age and sex-matched retrospective case-control study. A database with 638 consecutive patients who underwent ablation for atrial flutter at the University of Leuven. Sixty-one patients (55 men, 90%) fitted the inclusion criteria of 'lone atrial flutter', ie, aged 65 years or less, without documented atrial fibrillation and without identifiable underlying disease (including hypertension). Sex, age and inclusion criteria-matched controls, two for each flutter patient, were selected in a general practice in the same geographical region. Sports activity was evaluated by detailed questionnaires, which were available in 58 flutter patients (95%). A transthoracic echocardiogram was performed in all lone flutter patients. Types of sports, number of years of participation and average number of hours per week. The proportion of regular sportsmen (≥3 h of sports practice per week) among patients with lone atrial flutter was significantly higher than that observed in the general population (50% vs 17%; p<0.0001). The proportion of sportsmen engaged in long-term endurance sports (participation in cycling, running or swimming for ≥3 h/week) was also significantly higher in lone flutter patients than in controls (31% vs 8%; p=0.0003). Those flutter patients performing endurance sports had a larger left atrium than non-sportsmen (p=0.04, by one-way analysis of variance). A history of endurance sports and subsequent left atrial remodelling may be a risk factor for the development of atrial flutter.

Correlation of transonic-cone preston-tube data and skin friction

NASA Technical Reports Server (NTRS)

Abu-Mostafa, A. S.; Reed, T. D.

1984-01-01

Preston-tube measurements obtained on the Arnold Engineering Development Center (AEDC) Transition Cone have been correlated with theoretical skin friction coefficients in transitional and turbulent flow. This has been done for the NASA Ames 11-Ft Transonic Wind Tunnel (11 TWT) and flight tests. The developed semi-empirical correlations of Preston-tube data have been used to derive a calibration procedure for the 11 TWT flow quality. This procedure has been applied to the corrected laminar data, and an effective freestream unit Reynolds number is defined by requiring a matching of the average Preston-tube pressure in flight and in the tunnel. This study finds that the operating Reynolds number is below the effective value required for a match in laminar Preston-tube data. The distribution of this effective Reynolds number with Mach number correlates well with the freestream noise level in this tunnel. Analyses of transitional and turbulent data, however, did not result in effective Reynolds numbers that can be correlated with background noise. This is a result of the fact that vorticity fluctuations present in transitional and turbulent boundary layers dominate Preston-tube pressure fluctuations and, therefore, mask the tunnel noise eff ects. So, in order to calibrate the effects of noise on transonic wind tunnel tests only laminar data should be used, preferably at flow conditions similar to those in flight tests. To calibrate the effects of transonic wind-tunnel noise on drag measurements, however, the Preston-tube data must be supplemented with direct measurements of skin friction.

High-Reynolds Number Circulation Control Testing in the National Transonic Facility

NASA Technical Reports Server (NTRS)

Milholen, William E., II; Jones, Gregory S.; Chan, David T.; Goodliff, Scott L.

2012-01-01

A new capability to test active flow control concepts and propulsion simulations at high Reynolds numbers in the National Transonic Facility at the NASA Langley Research Center is being developed. The first active flow control experiment was completed using the new FAST-MAC semi-span model to study Reynolds number scaling effects for several circulation control concepts. Testing was conducted over a wide range of Mach numbers, up to chord Reynolds numbers of 30 million. The model was equipped with four onboard flow control valves allowing independent control of the circulation control plenums, which were directed over a 15% chord simple-hinged flap. Preliminary analysis of the uncorrected lift data showed that the circulation control increased the low-speed maximum lift coefficient by 33%. At transonic speeds, the circulation control was capable of positively altering the shockwave pattern on the upper wing surface and reducing flow separation. Furthermore, application of the technique to only the outboard portion of the wing demonstrated the feasibility of a pneumatic based roll control capability.

Study of flutter related computational procedures for minimum weight structural sizing of advanced aircraft, supplemental data

NASA Technical Reports Server (NTRS)

Oconnell, R. F.; Hassig, H. J.; Radovcich, N. A.

1975-01-01

Computational aspects of (1) flutter optimization (minimization of structural mass subject to specified flutter requirements), (2) methods for solving the flutter equation, and (3) efficient methods for computing generalized aerodynamic force coefficients in the repetitive analysis environment of computer-aided structural design are discussed. Specific areas included: a two-dimensional Regula Falsi approach to solving the generalized flutter equation; method of incremented flutter analysis and its applications; the use of velocity potential influence coefficients in a five-matrix product formulation of the generalized aerodynamic force coefficients; options for computational operations required to generate generalized aerodynamic force coefficients; theoretical considerations related to optimization with one or more flutter constraints; and expressions for derivatives of flutter-related quantities with respect to design variables.

The status of analytical preparation for 2-dimensional testing at high transonic speeds in the University of Southampton transonic self-streamlining wind tunnel

NASA Technical Reports Server (NTRS)

Lewis, M. C.

1984-01-01

Validation data from the Transonic Self-Streamlining Wind Tunnel has proved the feasibility of streamlining two dimensional flexible walls at low speeds and up to transonic speeds, the upper limit being the speed where the flexible walls are just supercritical. At this condition, breakdown of the wall setting strategy is evident in that convergence is neither as rapid nor as stable as for lower speeds, and wall streamlining criteria are not always completely satisfied. The only major step necessary to permit the extension of two dimensional testing into higher transonic speeds is the provision of a rapid algorithm to solve for mixed flow in the imagery flow fields. The status of two dimensional high transonic testing in the Transonic Self-Streamlining Wind Tunnel is outlined and, in particular, the progress of adapting an algorithm, which solves the Transonic Small Perturbation Equation, for predicting the imagery flow fields is detailed.

The NASA Langley Research Center 0.3-meter transonic cryogenic tunnel microcomputer controller source code

NASA Technical Reports Server (NTRS)

Kilgore, W. Allen; Balakrishna, S.

1991-01-01

The 0.3 m Transonic Cryogenic Tunnel (TCT) microcomputer based controller has been operating for several thousand hours in a safe and efficient manner. A complete listing is provided of the source codes for the tunnel controller and tunnel simulator. Included also is a listing of all the variables used in these programs. Several changes made to the controller are described. These changes are to improve the controller ease of use and safety.

Development and Flight Test of an Active Flutter Suppression System for the F-4F with Stores. Part 3. Flight Demonstration of the Active Flutter Suppression System.

DTIC Science & Technology

1983-06-01

that the dynamic behavior of the wing-pylon-store changed considerably with excitation amplitude due to free play and preload. The active flutter suppression system worked well and provided an increase in flutter speed.

The application of measurement techniques to track flutter testing

NASA Technical Reports Server (NTRS)

Roglin, H. R.

1975-01-01

The application is discussed of measurement techniques to captive flight flutter tests at the Supersonic Naval Ordnance Research Track (SNORT), U. S. Naval Ordnance Test Station, China Lake, California. The high-speed track, by its ability to prove the validity of design and to accurately determine the actual margin of safety, offers a unique method of flutter testing for the aircraft design engineer.

Transonic Symposium: Theory, Application and Experiment, volume 2

NASA Technical Reports Server (NTRS)

Foughner, Jerome T., Jr. (Compiler)

1989-01-01

Papers presented at the Transonic Symposium are compiled. The following subject areas are covered: National Transonic Facility status; transonic aerodynamics of slender wing-body configuration; laminar flow flight experiments; laminar flow wind tunnel experiments; computational support of X-29A flight experiment; transition location on a clean-up glove installed on a F-14 aircraft; and design studies for a laminar glove for the X-29 aircraft.

Recommendations for Achieving Accurate Numerical Simulation of Tip Clearance Flows in Transonic Compressor Rotors

NASA Technical Reports Server (NTRS)

VanZante, Dale E.; Strazisar, Anthony J.; Wood, Jerry R,; Hathaway, Michael D.; Okiishi, Theodore H.

2000-01-01

The tip clearance flows of transonic compressor rotors are important because they have a significant impact on rotor and stage performance. While numerical simulations of these flows are quite sophisticated. they are seldom verified through rigorous comparisons of numerical and measured data because these kinds of measurements are rare in the detail necessary to be useful in high-speed machines. In this paper we compare measured tip clearance flow details (e.g. trajectory and radial extent) with corresponding data obtained from a numerical simulation. Recommendations for achieving accurate numerical simulation of tip clearance flows are presented based on this comparison. Laser Doppler Velocimeter (LDV) measurements acquired in a transonic compressor rotor, NASA Rotor 35, are used. The tip clearance flow field of this transonic rotor was simulated using a Navier-Stokes turbomachinery solver that incorporates an advanced k-epsilon turbulence model derived for flows that are not in local equilibrium. Comparison between measured and simulated results indicates that simulation accuracy is primarily dependent upon the ability of the numerical code to resolve important details of a wall-bounded shear layer formed by the relative motion between the over-tip leakage flow and the shroud wall. A simple method is presented for determining the strength of this shear layer.

Computational design of low aspect ratio wing-winglet configurations for transonic wind-tunnel tests

NASA Technical Reports Server (NTRS)

Kuhlman, John M.; Brown, Christopher K.

1989-01-01

Computational designs were performed for three different low aspect ratio wing planforms fitted with nonplanar winglets; one of the three configurations was selected to be constructed as a wind tunnel model for testing in the NASA LaRC 8-foot transonic pressure tunnel. A design point of M = 0.8, C(sub L) is approximate or = to 0.3 was selected, for wings of aspect ratio equal to 2.2, and leading edge sweep angles of 45 deg and 50 deg. Winglet length is 15 percent of the wing semispan, with a cant angle of 15 deg, and a leading edge sweep of 50 deg. Winglet total area equals 2.25 percent of the wing reference area. The design process and the predicted transonic performance are summarized for each configuration. In addition, a companion low-speed design study was conducted, using one of the transonic design wing-winglet planforms but with different camber and thickness distributions. A low-speed wind tunnel model was constructed to match this low-speed design geometry, and force coefficient data were obtained for the model at speeds of 100 to 150 ft/sec. Measured drag coefficient reductions were of the same order of magnitude as those predicted by numerical subsonic performance predictions.

Gravity effects on wind-induced flutter of leaves

NASA Astrophysics Data System (ADS)

Clemmer, Nickalaus; Kopperstad, Karsten; Solano, Tomas; Shoele, Kourosh; Ordonez, Juan

2017-11-01

Wind-Induced flutter of leaves depends on both wind velocity and the gravity. To study the gravitational effects on the oscillatory behavior of leaves in the wind, a wind tunnel that can be tilted about the center of the test section is created. This unique rotation capability allows systematic investigation of gravitational effects on the fluttering response of leaves. The flow-induced vibration will be studied for three different leaves at several different tilting angles including the wind travels horizontally, vertically downward and vertically upward. In each situation, the long axis of a leaf is placed parallel to the wind direction and its response is studied at different flow speed. Oscillation of the leaf is recorded via high-speed camera at each of setup, and the effect of the gravity on stabilizing or destabilizing the fluttering response is investigated. Summer REU student at Florida State University.

Self streamlining wind tunnel: Further low speed testing and final design studies for the transonic facility

NASA Technical Reports Server (NTRS)

Wolf, S. W. D.

1977-01-01

Work has continued with the low speed self streamlining wind tunnel (SSWT) using the NACA 0012-64 airfoil in an effort to explain the discrepancies between the NASA Langley low turbulence pressure tunnel (LTPT) and SSWT results obtained with the airfoil stalled. Conventional wind tunnel corrections were applied to straight wall SSWT airfoil data, to illustrate the inadequacy of standard correction techniques in circumstances of high blockage. Also one SSWT test was re-run at different air speeds to investigate the effects of such changes on airfoil data and wall contours. Mechanical design analyses for the transonic self streamlining wind tunnel (TSWT) were completed by the application of theoretical airfoil flow field data to the elastic beam and streamline analysis. The control system for the transonic facility is outlined.

NASA High-Reynolds Number Circulation Control Research - Overview of CFD and Planned Experiments

NASA Technical Reports Server (NTRS)

Milholen, W. E., II; Jones, Greg S.; Cagle, Christopher M.

2010-01-01

A new capability to test active flow control concepts and propulsion simulations at high Reynolds numbers in the National Transonic Facility at the NASA Langley Research Center is being developed. This technique is focused on the use of semi-span models due to their increased model size and relative ease of routing high-pressure air to the model. A new dual flow-path high-pressure air delivery station has been designed, along with a new high performance transonic sem -si pan wing model. The modular wind tunnel model is designed for testing circulation control concepts at both transonic cruise and low-speed high-lift conditions. The ability of the model to test other active flow control techniques will be highlighted. In addition, a new higher capacity semi-span force and moment wind tunnel balance has been completed and calibrated to enable testing at transonic conditions.

The flutter-by effect: a comprehensive study of the fluttering cusps of the Perceval heart valve prosthesis.

PubMed

Aljalloud, Ali; Shoaib, Mohamed; Egron, Sandrine; Arias, Jessica; Tewarie, Lachmandath; Schnoering, Heike; Lotfi, Shahram; Goetzenich, Andreas; Hatam, Nima; Pott, Desiree; Zhong, Zhaoyang; Steinseifer, Ulrich; Zayat, Rachad; Autschbach, Ruediger

2018-05-17

Sutureless aortic valve prostheses are gaining popularity due to the substantial reduction in cross-clamp time. In this study, we report our observations on the cusp-fluttering phenomenon of the Perceval bioprosthesis (LivaNova, London, UK) using a combination of technical and medical perspectives. Between August 2014 and December 2016, a total of 108 patients (69% women) with a mean age of 78 years had aortic valve replacement using the Perceval bioprosthesis (34 combined procedures). All patients underwent transoesophageal echocardiography (TOE) intraoperatively. TOE was performed postoperatively to detect paravalvular leakage and to measure gradients, acceleration time, Doppler velocity indices (Vmax and LVOT/Vmax AV) and effective orifice area indices. In addition, a TOE examination was performed in 21 patients postoperatively. Data were collected retrospectively from our hospital database. The retrospective evaluation of the intraoperative TOE examinations revealed consistent fluttering in all patients with the Perceval bioprosthesis. The echocardiographic postoperative measurements showed a mean effective orifice area index of 0.91 ± 0.12 cm2/m2. The overall mean pressure and peak pressure gradients were in a higher range (13.5 ± 5.1 mmHg and 25.5 ± 8.6 mmHg, respectively), whereas acceleration time (62.8 ± 16.4 ms) and Doppler velocity indices (0.43 ± 0.11) were within the normal range according to the American Society of Echocardiography or european association of echocardiography (EAE) guidelines. The 2-dimensional TOE in Motion Mode (M-Mode) that was performed in patients with elevated lactate dehydrogenase (LDH) levels revealed remarkable fluttering of the cusps of the Perceval bioprosthesis. In our study cohort, we observed the fluttering phenomenon in all patients who received the Perceval bioprosthesis, which was correlated with elevated LDH levels and higher pressure gradients.

Optimization of cascade blade mistuning under flutter and forced response constraints

NASA Technical Reports Server (NTRS)

Murthy, D. V.; Haftka, R. T.

1984-01-01

In the development of modern turbomachinery, problems of flutter instabilities and excessive forced response of a cascade of blades that were encountered have often turned out to be extremely difficult to eliminate. The study of these instabilities and the forced response is complicated by the presence of mistuning; that is, small differences among the individual blades. The theory of mistuned cascade behavior shows that mistuning can have a beneficial effect on the stability of the rotor. This beneficial effect is produced by the coupling between the more stable and less stable flutter modes introduced by mistuning. The effect of mistuning on the forced response can be either beneficial or adverse. Kaza and Kielb have studied the effects of two types of mistuning on the flutter and forced response: alternate mistuning where alternte blades are identical and random mistuning. The objective is to investigate other patterns of mistuning which maximize the beneficial effects on the flutter and forced response of the cascade. Numerical optimization techniques are employed to obtain optimal mistuning patterns. The optimization program seeks to minimize the amount of mistuning required to satisfy constraints on flutter speed and forced response.

Investigation of Reynolds Number Effects on a Generic Fighter Configuration in the National Transonic Facility

NASA Technical Reports Server (NTRS)

Tomek, W. G.; Hall, R. M.; Wahls, R. A.; Luckring, J. M.; Owens, L. R.

2002-01-01

A wind tunnel test of a generic fighter configuration was tested in the National Transonic Facility through a cooperative agreement between NASA Langley Research Center and McDonnell Douglas. The primary purpose of the test was to assess Reynolds number scale effects on a thin-wing, fighter-type configuration up to full-scale flight conditions (that is, Reynolds numbers of the order of 60 million). The test included longitudinal and lateral/directional studies at subsonic and transonic conditions across a range of Reynolds numbers from that available in conventional wind tunnels to flight conditions. Results are presented for three Mach numbers (0.6, 0.8, and 0.9) and three configurations: (1) Fuselage/Wing; (2) Fuselage/Wing/Centerline Vertical Tail/Horizontal Tail; and (3) Fuselage/Wing/Trailing-Edge Extension/Twin Vertical Tails. Reynolds number effects on the longitudinal aerodynamic characteristics are presented herein.

Transonic wing DFVLR-F4 as European test model

NASA Technical Reports Server (NTRS)

Redeker, G.; Schmidt, N.

1980-01-01

A transonic wing, the DFVLR-F4 was designed and tested as a model in European transonic wind tunnels and was found to give performance improvements over conventional wings. One reason for the improvement was the reduction of compression shocks in the transonic region as the result of improved wing design.

Analytical and experimental investigation of flutter suppression by piezoelectric actuation

NASA Technical Reports Server (NTRS)

Heeg, Jennifer

1993-01-01

The objective of this research was to analytically and experimentally study the capabilities of piezoelectric plate actuators for suppressing flutter. Piezoelectric materials are characterized by their ability to produce voltage when subjected to a mechanical strain. The converse piezoelectric effect can be utilized to actuate a structure by applying a voltage. For this investigation, a two-degree-of-freedom wind tunnel model was designed, analyzed, and tested. The model consisted of a rigid wing and a flexible mount system that permitted a translational and a rotational degree of freedom. The model was designed such that flutter was encountered within the testing envelope of the wind tunnel. Actuators made of piezoelectric material were affixed to leaf springs of the mount system. Command signals, applied to the piezoelectric actuators, exerted control over the damping and stiffness properties. A mathematical aeroservoelastic model was constructed by using finite element methods, laminated plate theory, and aeroelastic analysis tools. Plant characteristics were determined from this model and verified by open loop experimental tests. A flutter suppression control law was designed and implemented on a digital control computer. Closed loop flutter testing was conducted. The experimental results represent the first time that adaptive materials have been used to actively suppress flutter. They demonstrate that small, carefully placed actuating plates can be used effectively to control aeroelastic response.

Aerothermoelastic Topology Optimization with Flutter and Buckling Metrics (Postprint)

DTIC Science & Technology

2013-07-01

topologies of an unheated panel, thermal buckling-optimal topologies, and flutter- optimality of a heated panel (where the latter case presents a...topological compromise between the former two). The effect of various constraint boundaries, temperature gradients, and (for the flutter of the heated panel...optimality of a heated panel (where the latter case presents a topological compromise between the former two). The effect of various constraint boundaries

Fan Flutter Computations Using the Harmonic Balance Method

NASA Technical Reports Server (NTRS)

Bakhle, Milind A.; Thomas, Jeffrey P.; Reddy, T.S.R.

2009-01-01

An experimental forward-swept fan encountered flutter at part-speed conditions during wind tunnel testing. A new propulsion aeroelasticity code, based on a computational fluid dynamics (CFD) approach, was used to model the aeroelastic behavior of this fan. This threedimensional code models the unsteady flowfield due to blade vibrations using a harmonic balance method to solve the Navier-Stokes equations. This paper describes the flutter calculations and compares the results to experimental measurements and previous results from a time-accurate propulsion aeroelasticity code.

Recent transonic unsteady pressure measurements at the NASA Langley Research Center

NASA Technical Reports Server (NTRS)

Sandford, M. C.; Ricketts, R. H.; Hess, R. W.

1985-01-01

Four semispan wing model configurations were studied in the Transonic Dynamics Tunnel (TDT). The first model had a clipped delta planform with a circular arc airfoil, the second model had a high aspect ratio planform with a supercritical airfoil, the third model has a rectangular planform with a supercritical airfoil and the fourth model had a high aspect ratio planform with a supercritical airfoil. To generate unsteady flow, the first and third models were equipped with pitch oscillation mechanisms and the first, second and fourth models were equipped with control surface oscillation mechanisms. The fourth model was similar in planform and airfoil shape to the second model, but it is the only one of the four models that has an elastic wing structure. The unsteady pressure studies of the four models are described and some typical results for each model are presented. Comparison of selected experimental data with analytical results also are included.

Sedimentation and fluttering of a cylinder in a confined liquid

NASA Astrophysics Data System (ADS)

D'Angelo, Maria Veronica; Cachile, Mario; Hulin, Jean-Pierre; Auradou, Harold

2017-10-01

The sedimentation and fluttering (angular oscillation of the axis) of straight cylinders are studied in a viscous fluid at rest filling a vertical Hele-Shaw cell for different density contrasts ρs-ρf and fluid viscosities μf and for two cylinder densities ρs and diameters D . The influence of confinement in the cell is studied by comparing the present results to those of the literature for nonconfined fluids. While the confinement and the cylinder length L both influence strongly the mean sedimentation velocity Vs, the characteristics of the fluttering instability are much more similar in the confined and nonconfined cases. While the drag coefficient is nearly constant in a nonconfined fluid, it is larger here and depends both on L (due to flow blockage) and on the Reynolds number ReD=VsD ρf/μf ; the inertial and viscous drag components have equal magnitudes for ReD≃40 . For fluttering, instead, the key parameter is the Froude number Fr=Vs/Vg [Vg=√{(ρs-ρf) g L /ρf }] , and the fluttering oscillations vanish below Fr˜0.07 for all cylinders and fluids investigated. Above this threshold, the angular amplitude increases with Fr up to a plateau value, while that of the horizontal oscillations is, at first, very large and then decreases; both amplitudes are reduced when the viscous drag is dominant, but, if inertial drag is dominant, all data points follow a common trend. For all fluids and cylinders, too, the fluttering frequency varies as f =0.102 Vg/L . These features of fluttering are generally qualitatively similar to those reported in nonconfined fluids, but this instability is observable down to lower ReD values (≃24 instead of ˜200 ).

Data reduction formulas for the 16-foot transonic tunnel: NASA Langley Research Center, revision 2

NASA Technical Reports Server (NTRS)

Mercer, Charles E.; Berrier, Bobby L.; Capone, Francis J.; Grayston, Alan M.

1992-01-01

The equations used by the 16-Foot Transonic Wind Tunnel in the data reduction programs are presented in nine modules. Each module consists of equations necessary to achieve a specific purpose. These modules are categorized in the following groups: (1) tunnel parameters; (2) jet exhaust measurements; (3) skin friction drag; (4) balance loads and model attitudes calculations; (5) internal drag (or exit-flow distribution); (6) pressure coefficients and integrated forces; (7) thrust removal options; (8) turboprop options; and (9) inlet distortion.

Electrophysiological determinant for induction of isthmus dependent counterclockwise and clockwise atrial flutter in humans.

PubMed

Lin, J L; Lai, L P; Lin, L J; Tseng, Y Z; Lien, W P; Huang, S K

1999-01-01

To investigate the electrophysiological determinant underlying the electrical induction of counterclockwise and clockwise isthmus dependent atrial flutter. The isthmus bordered by the inferior vena caval orifice-tricuspid annulus-coronary sinus ostium (IVCO-TA-CSO) has been assumed to be the site of both slow conduction and unidirectional block critical to the initiation of atrial flutter. Trans-isthmus and the global atrial conduction were studied in 25 patients with isthmus dependent atrial flutter (group A) and in 21 patients without atrial flutter (group B), by pacing at the coronary sinus ostium and the low lateral right atrium (LLRA) and mapping with a 20 pole Halo catheter in the right atrium. Mean (SD) fluoroscopic isthmus length between the coronary sinus ostium and LLRA sites was 28.1 (4.0) mm in group A and 28.0 (3.9) mm in group B (p = 0.95), but the trans-isthmus conduction velocity of both directions at various pacing cycle lengths was nearly halved in group A compared with group B (mean 0.39-0.46 m/s v 0.83-0.89 m/s, p < 0.0001). Pacing at coronary sinus ostium directly induced counterclockwise atrial flutter in 14 patients and pacing at LLRA induced clockwise atrial flutter in 11 patients, following abrupt unidirectional trans-isthmus block. Transient atrial tachyarrhythmias preceded the onset of atrial flutter in 10 counterclockwise and six clockwise cases of atrial flutter. None of the group B patients had inducible atrial flutter even in the presence of trans-isthmus block. The intra- and interatrial conduction times, as well as the conduction velocities at the right atrial free wall and the septum, were similar and largely within the normal range in both groups. Critical slowing of the trans-IVCO-TA-CSO isthmus conduction, but not the unidirectional block or the global atrial performance, is the electrophysiological determinant of the induction of counterclockwise and clockwise isthmus dependent atrial flutter in man.

NASA Common Research Model Test Envelope Extension With Active Sting Damping at NTF

NASA Technical Reports Server (NTRS)

Rivers, Melissa B.; Balakrishna, S.

2014-01-01

The NASA Common Research Model (CRM) high Reynolds number transonic wind tunnel testing program was established to generate an experimental database for applied Computational Fluid Dynamics (CFD) validation studies. During transonic wind tunnel tests, the CRM encounters large sting vibrations when the angle of attack approaches the second pitching moment break, which can sometimes become divergent. CRM transonic test data analysis suggests that sting divergent oscillations are related to negative net sting damping episodes associated with flow separation instability. The National Transonic Facility (NTF) has been addressing remedies to extend polar testing up to and beyond the second pitching moment break point of the test articles using an active piezoceramic damper system for both ambient and cryogenic temperatures. This paper reviews CRM test results to gain understanding of sting dynamics with a simple model describing the mechanics of a sting-model system and presents the performance of the damper under cryogenic conditions.

A Numerical Simulator for Three-Dimensional Flows Through Vibrating Blade Rows

NASA Technical Reports Server (NTRS)

Chuang, H. Andrew; Verdon, Joseph M.

1998-01-01

The three-dimensional, multi-stage, unsteady, turbomachinery analysis, TURBO, has been extended to predict the aeroelastic and aeroacoustic response behaviors of a single blade row operating within a cylindrical annular duct. In particular, a blade vibration capability has been incorporated so that the TURBO analysis can be applied over a solution domain that deforms with a vibratory blade motion. Also, unsteady far-field conditions have been implemented to render the computational boundaries at inlet and exit transparent to outgoing unsteady disturbances. The modified TURBO analysis is applied herein to predict unsteady subsonic and transonic flows. The intent is to partially validate this nonlinear analysis for blade flutter applications, via numerical results for benchmark unsteady flows, and to demonstrate the analysis for a realistic fan rotor. For these purposes, we have considered unsteady subsonic flows through a 3D version of the 10th Standard Cascade, and unsteady transonic flows through the first stage rotor of the NASA Lewis, Rotor 67, two-stage fan.

Turbulence and modeling in transonic flow

NASA Technical Reports Server (NTRS)

Rubesin, Morris W.; Viegas, John R.

1989-01-01

A review is made of the performance of a variety of turbulence models in the evaluation of a particular well documented transonic flow. This is done to supplement a previous attempt to calibrate and verify transonic airfoil codes by including many more turbulence models than used in the earlier work and applying the calculations to an experiment that did not suffer from uncertainties in angle of attack and was free of wind tunnel interference. It is found from this work, as well as in the earlier study, that the Johnson-King turbulence model is superior for transonic flows over simple aerodynamic surfaces, including moderate separation. It is also shown that some field equation models with wall function boundary conditions can be competitive with it.

Generalized Reduced Order Modeling of Aeroservoelastic Systems

NASA Astrophysics Data System (ADS)

Gariffo, James Michael

Transonic aeroelastic and aeroservoelastic (ASE) modeling presents a significant technical and computational challenge. Flow fields with a mixture of subsonic and supersonic flow, as well as moving shock waves, can only be captured through high-fidelity CFD analysis. With modern computing power, it is realtively straightforward to determine the flutter boundary for a single structural configuration at a single flight condition, but problems of larger scope remain quite costly. Some such problems include characterizing a vehicle's flutter boundary over its full flight envelope, optimizing its structural weight subject to aeroelastic constraints, and designing control laws for flutter suppression. For all of these applications, reduced-order models (ROMs) offer substantial computational savings. ROM techniques in general have existed for decades, and the methodology presented in this dissertation builds on successful previous techniques to create a powerful new scheme for modeling aeroelastic systems, and predicting and interpolating their transonic flutter boundaries. In this method, linear ASE state-space models are constructed from modal structural and actuator models coupled to state-space models of the linearized aerodynamic forces through feedback loops. Flutter predictions can be made from these models through simple eigenvalue analysis of their state-transition matrices for an appropriate set of dynamic pressures. Moreover, this analysis returns the frequency and damping trend of every aeroelastic branch. In contrast, determining the critical dynamic pressure by direct time-marching CFD requires a separate run for every dynamic pressure being analyzed simply to obtain the trend for the critical branch. The present ROM methodology also includes a new model interpolation technique that greatly enhances the benefits of these ROMs. This enables predictions of the dynamic behavior of the system for flight conditions where CFD analysis has not been explicitly

Development of Background-Oriented Schlieren for NASA Langley Research Center Ground Test Facilities

NASA Technical Reports Server (NTRS)

Bathel, Brett F.; Borg, Stephen; Jones, Stephen; Overmeyer, Austin; Walker, Eric; Goad, William; Clem, Michelle; Schairer, Edward T.; Mizukaki, Toshiharu

2015-01-01

This paper provides an overview of recent wind tunnel tests performed at the NASA Langley Research Center where the Background-Oriented Schlieren (BOS) technique was used to provide information pertaining to flow-field density disturbances. The facilities in which the BOS technique was applied included the National Transonic Facility (NTF), Transonic Dynamics Tunnel (TDT), 31-Inch Mach 10 Air Tunnel, 15-Inch Mach 6 High-Temperature Air Tunnel, Rotor Test Cell at the 14 by 22 Subsonic Tunnel, and a 13-Inch Low-Speed Tunnel.

Recent Productivity Improvements to the National Transonic Facility

NASA Technical Reports Server (NTRS)

Popernack, Thomas G., Jr.; Sydnor, George H.

1998-01-01

Productivity gains have recently been made at the National Transonic Facility wind tunnel at NASA Langley Research Center. A team was assigned to assess and set productivity goals to achieve the desired operating cost and output of the facility. Simulations have been developed to show the sensitivity of selected process productivity improvements in critical areas to reduce overall test cycle times. The improvements consist of an expanded liquid nitrogen storage system, a new fan drive, a new tunnel vent stack heater, replacement of programmable logic controllers, an increased data communications speed, automated test sequencing, and a faster model changeout system. Where possible, quantifiable results of these improvements are presented. Results show that in most cases, improvements meet the productivity gains predicted by the simulations.

Review of design and operational characteristics of the 0.3-meter transonic cryogenic tunnel

NASA Technical Reports Server (NTRS)

Ray, E. J.; Ladson, C. L.; Adcock, J. B.; Lawing, P. L.; Hall, R. M.

1979-01-01

The past 6 years of operation with the NASA Langley 0.3 m transonic cryogenic tunnel (TCT) show that there are no insurmountable problems associated with cryogenic testing with gaseous nitrogen at transonic Mach numbers. The fundamentals of the concept were validated both analytically and experimentally and the 0.3 m TCT, with its unique Reynolds number capability, was used for a wide variety of aerodynamic tests. Techniques regarding real-gas effects were developed and cryogenic tunnel conditions can be set and maintained accurately. Cryogenic cooling by injecting liquid nitrogen directly into the tunnel circuit imposes no problems with temperature distribution or dynamic response characteristics. Experience with the 0.3 m TCT, indicates that there is a significant learning process associated with cryogenic, high Reynolds number testing. Many of the questions have already been answered; however, factors such as tunnel control, run logic, economics, instrumentation, and model technology present many new and challenging problems.

User's Guide for a Modular Flutter Analysis Software System (Fast Version 1.0)

NASA Technical Reports Server (NTRS)

Desmarais, R. N.; Bennett, R. M.

1978-01-01

The use and operation of a group of computer programs to perform a flutter analysis of a single planar wing are described. This system of programs is called FAST for Flutter Analysis System, and consists of five programs. Each program performs certain portions of a flutter analysis and can be run sequentially as a job step or individually. FAST uses natural vibration modes as input data and performs a conventional V-g type of solution. The unsteady aerodynamics programs in FAST are based on the subsonic kernel function lifting-surface theory although other aerodynamic programs can be used. Application of the programs is illustrated by a sample case of a complete flutter calculation that exercises each program.

Comparisons of Flutter Analyses for an Experimental Fan

NASA Technical Reports Server (NTRS)

Bakhle, Milind A.; Reddy, T. S. R.; Stefko, George L.

2010-01-01

Two propulsion aeroelasticity codes were used to model the aeroelastic characteristics of an experimental forward-swept fan that encountered flutter during wind tunnel testing. Both of these three-dimensional codes model the unsteady flowfield due to blade vibrations using the Navier-Stokes equations. In the first approach, the unsteady flow equations are solved using an implicit time-marching approach. In the second approach, the unsteady flow equations are converted to a harmonic balance form and solved using a pseudo-time marching method. This paper describes the flutter calculations and compares the results to experimental measurements.

Jump conditions in transonic equilibria

DOE Office of Scientific and Technical Information (OSTI.GOV)

Guazzotto, L.; Betti, R.; Jardin, S. C.

2013-04-15

In the present paper, the numerical calculation of transonic equilibria, first introduced with the FLOW code in Guazzotto et al.[Phys. Plasmas 11, 604 (2004)], is critically reviewed. In particular, the necessity and effect of imposing explicit jump conditions at the transonic discontinuity are investigated. It is found that 'standard' (low-{beta}, large aspect ratio) transonic equilibria satisfy the correct jump condition with very good approximation even if the jump condition is not explicitly imposed. On the other hand, it is also found that high-{beta}, low aspect ratio equilibria require the correct jump condition to be explicitly imposed. Various numerical approaches aremore » described to modify FLOW to include the jump condition. It is proved that the new methods converge to the correct solution even in extreme cases of very large {beta}, while they agree with the results obtained with the old implementation of FLOW in lower-{beta} equilibria.« less

Flutter suppression digital control law design and testing for the AFW wind tunnel model

NASA Technical Reports Server (NTRS)

Mukhopadhyay, Vivek

1994-01-01

The design of a control law for simultaneously suppressing the symmetric and antisymmetric flutter modes of a sting mounted fixed-in-roll aeroelastic wind-tunnel model is described. The flutter suppression control law was designed using linear quadratic Gaussian theory, and it also involved control law order reduction, a gain root-locus study, and use of previous experimental results. A 23 percent increase in the open-loop flutter dynamic pressure was demonstrated during the wind-tunnel test. Rapid roll maneuvers at 11 percent above the symmetric flutter boundary were also performed when the model was in a free-to-roll configuration.

Flutter suppression digital control law design and testing for the AFW wind-tunnel model

NASA Technical Reports Server (NTRS)

Mukhopadhyay, Vivek

1992-01-01

Design of a control law for simultaneously suppressing the symmetric and antisymmetric flutter modes of a string mounted fixed-in-roll aeroelastic wind tunnel model is described. The flutter suppression control law was designed using linear quadratic Gaussian theory and involved control law order reduction, a gain root-locus study, and the use of previous experimental results. A 23 percent increase in open-loop flutter dynamic pressure was demonstrated during the wind tunnel test. Rapid roll maneuvers at 11 percent above the symmetric flutter boundary were also performed when the model was in a free-to-roll configuration.

Flutter suppression digital control law design and testing for the AFW wind tunnel model

NASA Technical Reports Server (NTRS)

Mukhopadhyay, Vivek

1992-01-01

Design of a control law for simultaneously suppressing the symmetric and antisymmetric flutter modes of a sting mounted fixed-in-roll aeroelastic wind tunnel model is described. The flutter suppression control law was designed using linear quadratic Gaussian theory, and involved control law order reduction, a gain root-locus study and use of previous experimental results. A 23 percent increase in the open-loop flutter dynamic pressure was demonstrated during the wind tunnel test. Rapid roll maneuvers at 11 percent above the symmetric flutter boundary were also performed when the model was in a free-to-roll configuration.

Transonic Shock-Wave/Boundary-Layer Interactions on an Oscillating Airfoil

NASA Technical Reports Server (NTRS)

Davis, Sanford S.; Malcolm, Gerald N.

1980-01-01

Unsteady aerodynamic loads were measured on an oscillating NACA 64A010 airfoil In the NASA Ames 11 by 11 ft Transonic Wind Tunnel. Data are presented to show the effect of the unsteady shock-wave/boundary-layer interaction on the fundamental frequency lift, moment, and pressure distributions. The data show that weak shock waves induce an unsteady pressure distribution that can be predicted quite well, while stronger shock waves cause complex frequency-dependent distributions due to flow separation. An experimental test of the principles of linearity and superposition showed that they hold for weak shock waves while flows with stronger shock waves cannot be superimposed.

NASA and CFD - Making investments for the future

NASA Technical Reports Server (NTRS)

Hessenius, Kristin A.; Richardson, P. F.

1992-01-01

From a NASA perspective, CFD is a new tool for fluid flow simulation and prediction with virtually none of the inherent limitations of other ground-based simulation techniques. A primary goal of NASA's CFD research program is to develop efficient and accurate computational techniques for utilization in the design and analysis of aerospace vehicles. The program in algorithm development has systematically progressed through the hierarchy of engineering simplifications of the Navier-Stokes equations, starting with the inviscid formulations such as transonic small disturbance, full potential, and Euler.

Flutter and forced response of mistuned rotors using standing wave analysis

NASA Technical Reports Server (NTRS)

Dugundji, J.; Bundas, D. J.

1983-01-01

A standing wave approach is applied to the analysis of the flutter and forced response of tuned and mistuned rotors. The traditional traveling wave cascade airforces are recast into standing wave arbitrary motion form using Pade approximants, and the resulting equations of motion are written in the matrix form. Applications for vibration modes, flutter, and forced response are discussed. It is noted that the standing wave methods may prove to be more versatile for dealing with certain applications, such as coupling flutter with forced response and dynamic shaft problems, transient impulses on the rotor, low-order engine excitation, bearing motions, and mistuning effects in rotors.

Flutter and forced response of mistuned rotors using standing wave analysis

NASA Technical Reports Server (NTRS)

Bundas, D. J.; Dungundji, J.

1983-01-01

A standing wave approach is applied to the analysis of the flutter and forced response of tuned and mistuned rotors. The traditional traveling wave cascade airforces are recast into standing wave arbitrary motion form using Pade approximants, and the resulting equations of motion are written in the matrix form. Applications for vibration modes, flutter, and forced response are discussed. It is noted that the standing wave methods may prove to be more versatile for dealing with certain applications, such as coupling flutter with forced response and dynamic shaft problems, transient impulses on the rotor, low-order engine excitation, bearing motion, and mistuning effects in rotors.

Laser Displacement Measurements of Fan Blades in Resonance and Flutter During the Boundary Layer Ingesting Inlet and Distortion-Tolerant Fan Test

NASA Technical Reports Server (NTRS)

Duffy, Kirsten P.; Provenza, Andrew J.; Bakhle, Milind A.; Min, James B.; Abdul-Aziz, Ali

2018-01-01

NASA's Advanced Air Transport Technology Project is investigating boundary layer ingesting propulsors for future subsonic commercial aircraft to improve aircraft efficiency, thereby reducing fuel burn. To that end, a boundary layer ingesting inlet and distortion-tolerant fan stage was designed, fabricated, and tested within the 8' x 6' Supersonic Wind Tunnel at NASA Glenn Research Center. Because of the distortion in the air flow over the fan, the blades were designed to withstand a much higher aerodynamic forcing than for a typical clean flow. The blade response for several resonance modes were measured during start-up and shutdown, as well as at near 85% design speed. Flutter in the first bending mode was also observed in the fan at the design speed, at an off-design condition, although instabilities were difficult to instigate with this fan in general. Blade vibrations were monitored through twelve laser displacement probes that were placed around the inner circumference of the casing, at the blade leading and trailing edges. These probes captured the movement of all the blades during the entire test. Results are presented for various resonance mode amplitudes, frequencies and damping, as well as flutter amplitudes and frequency. Benefits and disadvantages of laser displacement probe measurements versus strain gage measurements are discussed.

Aeroservoelastic Modeling of Body Freedom Flutter for Control System Design

NASA Technical Reports Server (NTRS)

Ouellette, Jeffrey

2017-01-01

One of the most severe forms of coupling between aeroelasticity and flight dynamics is an instability called freedom flutter. The existing tools often assume relatively weak coupling, and are therefore unable to accurately model body freedom flutter. Because the existing tools were developed from traditional flutter analysis models, inconsistencies in the final models are not compatible with control system design tools. To resolve these issues, a number of small, but significant changes have been made to the existing approaches. A frequency domain transformation is used with the unsteady aerodynamics to ensure a more physically consistent stability axis rational function approximation of the unsteady aerodynamic model. The aerodynamic model is augmented with additional terms to account for limitations of the baseline unsteady aerodynamic model and to account for the gravity forces. An assumed modes method is used for the structural model to ensure a consistent definition of the aircraft states across the flight envelope. The X-56A stiff wing flight-test data were used to validate the current modeling approach. The flight-test data does not show body-freedom flutter, but does show coupling between the flight dynamics and the aeroelastic dynamics and the effects of the fuel weight.

Flutter Phenomenon in Flow Driven Energy Harvester-A Unified Theoretical Model for "Stiff" and "Flexible" Materials.

PubMed

Chen, Yu; Mu, Xiaojing; Wang, Tao; Ren, Weiwei; Yang, Ya; Wang, Zhong Lin; Sun, Chengliang; Gu, Alex Yuandong

2016-10-14

Here, we report a stable and predictable aero-elastic motion in the flow-driven energy harvester, which is different from flapping and vortex-induced-vibration (VIV). A unified theoretical frame work that describes the flutter phenomenon observed in both "stiff" and "flexible" materials for flow driven energy harvester was presented in this work. We prove flutter in both types of materials is the results of the coupled effects of torsional and bending modes. Compared to "stiff" materials, which has a flow velocity-independent flutter frequency, flexible material presents a flutter frequency that almost linearly scales with the flow velocity. Specific to "flexible" materials, pre-stress modulates the frequency range in which flutter occurs. It is experimentally observed that a double-clamped "flexible" piezoelectric P(VDF-TrFE) thin belt, when driven into the flutter state, yields a 1,000 times increase in the output voltage compared to that of the non-fluttered state. At a fixed flow velocity, increase in pre-stress level of the P(VDF-TrFE) thin belt up-shifts the flutter frequency. In addition, this work allows the rational design of flexible piezoelectric devices, including flow-driven energy harvester, triboelectric energy harvester, and self-powered wireless flow speed sensor.

The NASA Langley 16-Foot Transonic Tunnel: Historical Overview, Facility Description, Calibration, Flow Characteristics, and Test Capabilities

NASA Technical Reports Server (NTRS)

Capone, Francis J.; Bangert, Linda S.; Asbury, Scott C.; Mills, Charles T. L.; Bare, E. Ann

1995-01-01

The Langley 16-Foot Transonic Tunnel is a closed-circuit single-return atmospheric wind tunnel that has a slotted octagonal test section with continuous air exchange. The wind tunnel speed can be varied continuously over a Mach number range from 0.1 to 1.3. Test-section plenum suction is used for speeds above a Mach number of 1.05. Over a period of some 40 years, the wind tunnel has undergone many modifications. During the modifications completed in 1990, a new model support system that increased blockage, new fan blades, a catcher screen for the first set of turning vanes, and process controllers for tunnel speed, model attitude, and jet flow for powered models were installed. This report presents a complete description of the Langley 16-Foot Transonic Tunnel and auxiliary equipment, the calibration procedures, and the results of the 1977 and the 1990 wind tunnel calibration with test section air removal. Comparisons with previous calibrations showed that the modifications made to the wind tunnel had little or no effect on the aerodynamic characteristics of the tunnel. Information required for planning experimental investigations and the use of test hardware and model support systems is also provided.

Half Wing N219 Aircraft Model Clean Configuration for Flutter Test On Low Speed Wind Tunnel

NASA Astrophysics Data System (ADS)

Syamsuar, Sayuti; Sampurno, Budi; Mayang Mahasti, Katia; Bayu Sakti Pratama, Muchamad; Widi Sasongko, Triyono; Kartika, Nina; Suksmono, Adityo; Aji Saputro, Mohamad Ivan; Bahtera Eskayudha, Dimas

2018-04-01

Flutter is a rapid self-feeding motion which is caused by the interaction of aerodynamic, structural and inertial forces. Flutter can cause major damage on aircraft structure which can lead to fatal accident in aviation. Several methods have been evolved to avoid the flutter phenomena occur during the flight envelope of aircraft design. On this study, method was developed by Indonesian Aerospace which consist of Finite Element Method (FEM) analysis, Ground Vibration Test (GVT), and Wind Tunnel Flutter Test (WTT). Based on the study, FEM have similar results toward to Wind Tunnel Flutter Test conjunction the clean configuration of N219 aircraft half wing model.

Two degree-of-freedom flutter solution for a personal computer

NASA Technical Reports Server (NTRS)

Turnock, D. L.

1985-01-01

A computer programmed flutter solution has been written in the BASIC language for a personal computer. The program is for two degree-of-freedom bending torsion flutter applications and utilizes two dimensional Theodorsen aerodynamics. The aerodynamics were modified to include approximations for Mach number (compressibility) effects and aspect ratio (finite span) effects. Input options, user instructions, program listing, and a test case application are included.

Application of a flight test and data analysis technique to flutter of a drone aircraft

NASA Technical Reports Server (NTRS)

Bennett, R. M.; Abel, I.

1981-01-01

Modal identification results are presented that were obtained from recent flight flutter tests of a drone vehicle with a research wing equipped with an active flutter suppression system (FSS). Frequency and damping of several modes are determined by a time domain modal analysis of the impulse response function obtained by Fourier transformations of data from fast swept sine wave excitation by the FSS control surfaces on the wing. Flutter points are determined for two different altitudes with the FSS off. Data are given for near the flutter boundary with the FSS on.

Aerodynamic characteristics of three slender sharp-edge 74 degrees swept wings at subsonic, transonic, and supersonic Mach numbers

NASA Technical Reports Server (NTRS)

Davenport, E. E.

1974-01-01

Slender sharp-edge wings having leading-edge sweep angles of 74 deg have been studied at Mach numbers from 0.60 to 2.80, at angles of attack from about minus 4 deg to 22 deg, and at angles of sideslip from 0 deg to 5 deg. The wings had delta, arrow, and diamond planforms. The experimental tests were made in the Langley 8-foot transonic pressure tunnel and the Langley Unitary Plan wind tunnel test section number 1. The theoretical predictions were made using the theories of NASA TN D-3767 and NASA TN D-6243. The results of the study indicated that the lift and drag characteristics as affected by planform and Mach number could be reasonably well predicted for the delta wing in the subsonic and transonic Mach number range. In the supersonic range, the delta and diamond wings were about equally good in the degree of agreement between experiment and theory. In making drag-due-to-lift predictions the vortex lift effects must be taken into account if reasonable results are to be obtained at moderate or high lift coefficients.

Flight-Test Evaluation of Flutter-Prediction Methods

NASA Technical Reports Server (NTRS)

Lind, RIck; Brenner, Marty

2003-01-01

The flight-test community routinely spends considerable time and money to determine a range of flight conditions, called a flight envelope, within which an aircraft is safe to fly. The cost of determining a flight envelope could be greatly reduced if there were a method of safely and accurately predicting the speed associated with the onset of an instability called flutter. Several methods have been developed with the goal of predicting flutter speeds to improve the efficiency of flight testing. These methods include (1) data-based methods, in which one relies entirely on information obtained from the flight tests and (2) model-based approaches, in which one relies on a combination of flight data and theoretical models. The data-driven methods include one based on extrapolation of damping trends, one that involves an envelope function, one that involves the Zimmerman-Weissenburger flutter margin, and one that involves a discrete-time auto-regressive model. An example of a model-based approach is that of the flutterometer. These methods have all been shown to be theoretically valid and have been demonstrated on simple test cases; however, until now, they have not been thoroughly evaluated in flight tests. An experimental apparatus called the Aerostructures Test Wing (ATW) was developed to test these prediction methods.

Asymptotic research of transonic gas flows

NASA Astrophysics Data System (ADS)

Velmisov, Petr A.; Tamarova, Yuliya A.

2017-12-01

The article is dedicated to the development asymptotic theory of gas flowing at speed next to sound velocity, particularly of gas transonic flows, i.e. the flows, containing both, subsonic and supersonic areas. The main issue, when styding such flows, are nonlinearity and combined type of equations, describing the transonic flow. Based on asymptotic nonlinear equation obtained in the article, the gas transonic flows is studied, considering transverse disturbance with respect to the main flow. The asymptotic conditions at shock-wave front and conditions on the streamlined surface are found. Moreover, the equation of sound surface and asymptotic formula defining the pressure are recorded. Several exact particular solutions of such equation are given, and their application to solve several tasks of transonic aerodynamics is indicated. Specifically, the polynomial form solution describing gas axisymmetric flows in Laval nozzles with constant acceleration in direction of the nozzle's axis and flow swirling is obtained. The solutions describing the unsteady flow along the channels between spinning surfaces are presented. The asymptotic equation is obtained, describing the flow, appearing during non-separated and separated flow past, closely approximated to cylindrical one. Specific solutions are given, based on which the examples of steady flow are formed.

Robust Flutter Analysis for Aeroservoelastic Systems

NASA Astrophysics Data System (ADS)

Kotikalpudi, Aditya

The dynamics of a flexible air vehicle are typically described using an aeroservoelastic model which accounts for interaction between aerodynamics, structural dynamics, rigid body dynamics and control laws. These subsystems can be individually modeled using a theoretical approach and experimental data from various ground tests can be combined into them. For instance, a combination of linear finite element modeling and data from ground vibration tests may be used to obtain a validated structural model. Similarly, an aerodynamic model can be obtained using computational fluid dynamics or simple panel methods and partially updated using limited data from wind tunnel tests. In all cases, the models obtained for these subsystems have a degree of uncertainty owing to inherent assumptions in the theory and errors in experimental data. Suitable uncertain models that account for these uncertainties can be built to study the impact of these modeling errors on the ability to predict dynamic instabilities known as flutter. This thesis addresses the methods used for modeling rigid body dynamics, structural dynamics and unsteady aerodynamics of a blended wing design called the Body Freedom Flutter vehicle. It discusses the procedure used to incorporate data from a wide range of ground based experiments in the form of model uncertainties within these subsystems. Finally, it provides the mathematical tools for carrying out flutter analysis and sensitivity analysis which account for these model uncertainties. These analyses are carried out for both open loop and controller in the loop (closed loop) cases.

Reduced Uncertainties in the Flutter Analysis of the Aerostructures Test Wing

NASA Technical Reports Server (NTRS)

Pak, Chan-Gi; Lung, Shun Fat

2011-01-01

Tuning the finite element model using measured data to minimize the model uncertainties is a challenging task in the area of structural dynamics. A test validated finite element model can provide a reliable flutter analysis to define the flutter placard speed to which the aircraft can be flown prior to flight flutter testing. Minimizing the difference between numerical and experimental results is a type of optimization problem. Through the use of the National Aeronautics and Space Administration Dryden Flight Research Center's (Edwards, California) multidisciplinary design, analysis, and optimization tool to optimize the objective function and constraints; the mass properties, the natural frequencies, and the mode shapes are matched to the target data, and the mass matrix orthogonality is retained. The approach in this study has been applied to minimize the model uncertainties for the structural dynamic model of the aerostructures test wing, which was designed, built, and tested at the National Aeronautics and Space Administration Dryden Flight Research Center. A 25 percent change in flutter speed has been shown after reducing the uncertainties.

Reduced Uncertainties in the Flutter Analysis of the Aerostructures Test Wing

NASA Technical Reports Server (NTRS)

Pak, Chan-gi; Lung, Shun-fat

2010-01-01

Tuning the finite element model using measured data to minimize the model uncertainties is a challenging task in the area of structural dynamics. A test validated finite element model can provide a reliable flutter analysis to define the flutter placard speed to which the aircraft can be flown prior to flight flutter testing. Minimizing the difference between numerical and experimental results is a type of optimization problem. Through the use of the National Aeronautics and Space Administration Dryden Flight Research Center s (Edwards, California, USA) multidisciplinary design, analysis, and optimization tool to optimize the objective function and constraints; the mass properties, the natural frequencies, and the mode shapes are matched to the target data and the mass matrix orthogonality is retained. The approach in this study has been applied to minimize the model uncertainties for the structural dynamic model of the aerostructures test wing, which was designed, built, and tested at the National Aeronautics and Space Administration Dryden Flight Research Center. A 25-percent change in flutter speed has been shown after reducing the uncertainties

Flow Disturbance Characterization Measurements in the National Transonic Facility

NASA Technical Reports Server (NTRS)

King, Rudolph A.; Andino, Marlyn Y.; Melton, Latunia; Eppink, Jenna; Kegerise, Michael A.; Tsoi, Andrew

2012-01-01

Recent flow measurements have been acquired in the National Transonic Facility (NTF) to assess the unsteady flow environment in the test section. The primary purpose of the test is to determine the feasibility of the NTF to conduct laminar-flow-control testing and boundary-layer transition sensitive testing. The NTF can operate in two modes, warm (air) and cold/cryogenic (nitrogen) test conditions for testing full and semispan scaled models. The warm-air mode enables low to moderately high Reynolds numbers through the use of high tunnel pressure, and the nitrogen mode enables high Reynolds numbers up to flight conditions, depending on aircraft type and size, utilizing high tunnel pressure and cryogenic temperatures. NASA's Environmentally Responsible Aviation (ERA) project is interested in demonstrating different laminar-flow technologies at flight-relevant operating conditions throughout the transonic Mach number range and the NTF is well suited for the initial ground-based demonstrations. Roll polar data at selected test conditions were obtained to look at the uniformity of the flow disturbance field in the test section. Data acquired from the rake probes included mean total temperatures, mean and fluctuating static/total pressures, and mean and fluctuating hot-wire measurements. . Based on the current measurements and previous data, an assessment was made that the NTF is a suitable facility for ground-based demonstrations of laminar-flow technologies at flight-relevant conditions in the cryogenic mode.

Self streamlining wind tunnel: Further low speed testing and final design studies for the transonic facility

NASA Technical Reports Server (NTRS)

Wolf, S. W. D.

1978-01-01

Work was continued with the low speed self streamlining wind tunnel (SSWT) using the NACA 0012-64 airfoil in an effort to explain the discrepancies between the NASA Langley low turbulence pressure tunnel (LTPT) and SSWT results obtained with the airfoil stalled. Conventional wind tunnel corrections were applied to straight wall SSWT airfoil data, to illustrate the inadequacy of standard correction techniques in circumstances of high blockage. Also one SSWT test was re-run at different air speeds to investigate the effects of such changes (perhaps through changes in Reynold's number and freestream turbulence levels) on airfoil data and wall contours. Mechanical design analyses for the transonic self-streamlining wind tunnel (TSWT) were completed by the application of theoretical airfoil flow field data to the elastic beam and streamline analysis. The control system for the transonic facility, which will eventually allow on-line computer operation of the wind tunnel, was outlined.

Computational Analysis of the Transonic Dynamics Tunnel Using FUN3D

NASA Technical Reports Server (NTRS)

Chwalowski, Pawel; Quon, Eliot; Brynildsen, Scott E.

2016-01-01

This paper presents results from an exploratory two-year effort of applying Computational Fluid Dynamics (CFD) to analyze the empty-tunnel flow in the NASA Langley Research Center Transonic Dynamics Tunnel (TDT). The TDT is a continuous-flow, closed circuit, 16- x 16-foot slotted-test-section wind tunnel, with capabilities to use air or heavy gas as a working fluid. In this study, experimental data acquired in the empty tunnel using the R-134a test medium was used to calibrate the computational data. The experimental calibration data includes wall pressures, boundary-layer profiles, and the tunnel centerline Mach number profiles. Subsonic and supersonic flow regimes were considered, focusing on Mach 0.5, 0.7 and Mach 1.1 in the TDT test section. This study discusses the computational domain, boundary conditions, and initial conditions selected and the resulting steady-state analyses using NASA's FUN3D CFD software.

Development of Reduced-Order Models for Aeroelastic and Flutter Prediction Using the CFL3Dv6.0 Code

NASA Technical Reports Server (NTRS)

Silva, Walter A.; Bartels, Robert E.

2002-01-01

A reduced-order model (ROM) is developed for aeroelastic analysis using the CFL3D version 6.0 computational fluid dynamics (CFD) code, recently developed at the NASA Langley Research Center. This latest version of the flow solver includes a deforming mesh capability, a modal structural definition for nonlinear aeroelastic analyses, and a parallelization capability that provides a significant increase in computational efficiency. Flutter results for the AGARD 445.6 Wing computed using CFL3D v6.0 are presented, including discussion of associated computational costs. Modal impulse responses of the unsteady aerodynamic system are then computed using the CFL3Dv6 code and transformed into state-space form. Important numerical issues associated with the computation of the impulse responses are presented. The unsteady aerodynamic state-space ROM is then combined with a state-space model of the structure to create an aeroelastic simulation using the MATLAB/SIMULINK environment. The MATLAB/SIMULINK ROM is used to rapidly compute aeroelastic transients including flutter. The ROM shows excellent agreement with the aeroelastic analyses computed using the CFL3Dv6.0 code directly.

Transonic Free-To-Roll Analysis of the F/A-18E and F-35 Configurations

NASA Technical Reports Server (NTRS)

Owens, D. Bruce; McConnell, Jeffrey K.; Brandon, Jay M.; Hall, Robert M.

2004-01-01

The free-to-roll technique is used as a tool for predicting areas of uncommanded lateral motions. Recently, the NASA/Navy/Air Force Abrupt Wing Stall Program extended the use of this technique to the transonic speed regime. Using this technique, this paper evaluates various wing configurations on the pre-production F/A-18E aircraft and the Joint Strike Fighter (F-35) aircraft. The configurations investigated include leading and trailing edge flap deflections, fences, leading edge flap gap seals, and vortex generators. These tests were conducted in the NASA Langley 16-Foot Transonic Tunnel. The analysis used a modification of a figure-of-merit developed during the Abrupt Wing Stall Program to discern configuration effects. The results showed how the figure-of-merit can be used to schedule wing flap deflections to avoid areas of uncommanded lateral motion. The analysis also used both static and dynamic wind tunnel data to provide insight into the uncommanded lateral behavior. The dynamic data was extracted from the time history data using parameter identification techniques. In general, modifications to the pre-production F/A-18E resulted in shifts in angle-of-attack where uncommanded lateral activity occurred. Sealing the gap between the inboard and outboard leading-edge flaps on the Navy version of the F-35 eliminated uncommanded lateral activity or delayed the activity to a higher angle-of-attack.

Unsteady transonic flow calculations for realistic aircraft configurations

NASA Technical Reports Server (NTRS)

Batina, John T.; Seidel, David A.; Bland, Samuel R.; Bennett, Robert M.

1987-01-01

A transonic unsteady aerodynamic and aeroelasticity code has been developed for application to realistic aircraft configurations. The new code is called CAP-TSD which is an acronym for Computational Aeroelasticity Program - Transonic Small Disturbance. The CAP-TSD code uses a time-accurate approximate factorization (AF) algorithm for solution of the unsteady transonic small-disturbance equation. The AF algorithm is very efficient for solution of steady and unsteady transonic flow problems. It can provide accurate solutions in only several hundred time steps yielding a significant computational cost savings when compared to alternative methods. The new code can treat complete aircraft geometries with multiple lifting surfaces and bodies including canard, wing, tail, control surfaces, launchers, pylons, fuselage, stores, and nacelles. Applications are presented for a series of five configurations of increasing complexity to demonstrate the wide range of geometrical applicability of CAP-TSD. These results are in good agreement with available experimental steady and unsteady pressure data. Calculations for the General Dynamics one-ninth scale F-16C aircraft model are presented to demonstrate application to a realistic configuration. Unsteady results for the entire F-16C aircraft undergoing a rigid pitching motion illustrated the capability required to perform transonic unsteady aerodynamic and aeroelastic analyses for such configurations.

Experimental uncertainty and drag measurements in the national transonic facility

NASA Technical Reports Server (NTRS)

Batill, Stephen M.

1994-01-01

This report documents the results of a study which was conducted in order to establish a framework for the quantitative description of the uncertainty in measurements conducted in the National Transonic Facility (NTF). The importance of uncertainty analysis in both experiment planning and reporting results has grown significantly in the past few years. Various methodologies have been proposed and the engineering community appears to be 'converging' on certain accepted practices. The practical application of these methods to the complex wind tunnel testing environment at the NASA Langley Research Center was based upon terminology and methods established in the American National Standards Institute (ANSI) and the American Society of Mechanical Engineers (ASME) standards. The report overviews this methodology.

Development of a Free-to-Roll Transonic Test Capability (Invited)

NASA Technical Reports Server (NTRS)

Capone, F. J.; Owens, D. B.; Hall, R. M.

2003-01-01

As part of the NASA/Navy Abrupt Wing Stall Program, a relatively low-cost, rapid-access wind-tunnel free-to-roll rig was developed. This rig combines the use of conventional models and test apparatuses to evaluate both transonic performance and wing-drop/rock tendencies in a single tunnel entry. A description of the test hardware as well as a description of the experimental procedures is given. The free-to-roll test rig has been used successfully to assess the static and dynamic characteristics of three different configurations--two configurations that exhibit uncommanded lateral motions, (pre-production F/A-18E and AV-8B), and one that did not (F/A-18C).

Flutter suppression and gust alleviation using active controls

NASA Technical Reports Server (NTRS)

Nissim, E.

1974-01-01

The effects of active controls on the suppression of flutter and gust alleviation of two different types of subsonic aircraft (the Arava, twin turboprop STOL transport, and the Westwind twin-jet business transport) are investigated. The active controls are introduced in pairs which include, in any chosen wing strip, a leading-edge (LE) control and a trailing-edge (TE) control. Each control surface is allowed to be driven by a combined linear-rotational sensor system, located on the activated strip. The control law, which translates the sensor signals into control surface rotations, is based on the concept of aerodynamic energy. The results indicate the extreme effectiveness of the active systems in controlling flutter. A single system spanning 10% of the wing semispan made the Arava flutter-free, and a similar active system, for the Westwind aircraft, yielded a reduction of 75% in the maximum bending moment of the wing and a reduction of 90% in the acceleration of the cg of the aircraft. Results for simultaneous activation of several LE - TE systems are presented. Further work needed to bring the investigation to completion is also discussed.

(NTF) National Transonic Facility Test 213-SFW Flow Control II,

NASA Image and Video Library

2012-11-19

(NTF) National Transonic Facility Test 213-SFW Flow Control II, Fast-MAC Model: The fundamental Aerodynamics Subsonic Transonic-Modular Active Control (Fast-MAC) Model was tested for the 2nd time in the NTF. The objectives were to document the effects of Reynolds numbers on circulation control aerodynamics and to develop and open data set for CFD code validation. Image taken in building 1236, National Transonic Facility

Eigenspace techniques for active flutter suppression

NASA Technical Reports Server (NTRS)

Garrard, William L.; Liebst, Bradley S.; Farm, Jerome A.

1987-01-01

The use of eigenspace techniques for the design of an active flutter suppression system for a hypothetical research drone is discussed. One leading edge and two trailing edge aerodynamic control surfaces and four sensors (accelerometers) are available for each wing. Full state control laws are designed by selecting feedback gains which place closed loop eigenvalues and shape closed loop eigenvectors so as to stabilize wing flutter and reduce gust loads at the wing root while yielding accepatable robustness and satisfying constrains on rms control surface activity. These controllers are realized by state estimators designed using an eigenvalue placement/eigenvector shaping technique which results in recovery of the full state loop transfer characteristics. The resulting feedback compensators are shown to perform almost as well as the full state designs. They also exhibit acceptable performance in situations in which the failure of an actuator is simulated.

Calibration of transonic and supersonic wind tunnels

NASA Technical Reports Server (NTRS)

Reed, T. D.; Pope, T. C.; Cooksey, J. M.

1977-01-01

State-of-the art instrumentation and procedures for calibrating transonic (0.6 less than M less than 1.4) and supersonic (M less than or equal to 3.5) wind tunnels were reviewed and evaluated. Major emphasis was given to transonic tunnels. Continuous, blowdown and intermittent tunnels were considered. The required measurements of pressure, temperature, flow angularity, noise and humidity were discussed, and the effects of measurement uncertainties were summarized. A comprehensive review of instrumentation currently used to calibrate empty tunnel flow conditions was included. The recent results of relevant research are noted and recommendations for achieving improved data accuracy are made where appropriate. It is concluded, for general testing purposes, that satisfactory calibration measurements can be achieved in both transonic and supersonic tunnels. The goal of calibrating transonic tunnels to within 0.001 in centerline Mach number appears to be feasible with existing instrumentation, provided correct calibration procedures are carefully followed. A comparable accuracy can be achieved off-centerline with carefully designed, conventional probes, except near Mach 1. In the range 0.95 less than M less than 1.05, the laser Doppler velocimeter appears to offer the most promise for improved calibration accuracy off-centerline.

Application of a flight test and data analysis technique to flutter of a drone aircraft

NASA Technical Reports Server (NTRS)

Bennett, R. M.

1981-01-01

Modal identification results presented were obtained from recent flight flutter tests of a drone vehicle with a research wing (DAST ARW-1 for Drones for Aerodynamic and Structural Testing, Aeroelastic Research Wing-1). This vehicle is equipped with an active flutter suppression system (FSS). Frequency and damping of several modes are determined by a time domain modal analysis of the impulse response function obtained by Fourier transformations of data from fast swept sine wave excitation by the FSS control surface on the wing. Flutter points are determined for two different altitudes with the FSS off. Data are given for near the flutter boundary with the FSS on.

Active controls for flutter suppression and gust alleviation in supersonic aircraft. [YF-17 flutter model

NASA Technical Reports Server (NTRS)

Nissim, E.

1980-01-01

Results of work done on active controls on the modified YF-17 flutter model are summarized. The basic derivation of a suitable control law is discussed. It is shown that discrepencies found between analysis and wind tunnel tests originate from the lack of proper implementation of the desired control law. Program capabilities are described.

Beam-modulation methods in quantitative and flow-visualization holographic interferometry

NASA Technical Reports Server (NTRS)

Decker, Arthur J.

1986-01-01

Heterodyne holographic interferometry and time-average holography with a frequency shifted reference beam are discussed. Both methods will be used for the measurement and visualization of internal transonic flows where the target facility is a flutter cascade. The background and experimental requirements for both methods are reviewed. Measurements using heterodyne holographic interferometry are presented. The performance of the laser required for time-average holography of time-varying transonic flows is discussed.

Transonic Axial Splittered Rotor Tandem Stator Stage

DTIC Science & Technology

2016-12-01

CODE 13. ABSTRACT (maximum 200 words) Development of a procedure to model the hot shape of a rotor blade and a comparison analysis of the transonic...fluid-structure interaction. Rotational forces as well as gas loading forces were observed as an influence on blade deformation. Utilizing the...Turbomachinery, splittered rotor, tandem stator, transonic compressor, blade deformation, fluid-structure interaction 15. NUMBER OF PAGES 87 16. PRICE

Analysis of viscous transonic flow over airfoil sections

NASA Technical Reports Server (NTRS)

Huff, Dennis L.; Wu, Jiunn-Chi; Sankar, L. N.

1987-01-01

A full Navier-Stokes solver has been used to model transonic flow over three airfoil sections. The method uses a two-dimensional, implicit, conservative finite difference scheme for solving the compressible Navier-Stokes equations. Results are presented as prescribed for the Viscous Transonic Airfoil Workshop to be held at the AIAA 25th Aerospace Sciences Meeting. The NACA 0012, RAE 2822 and Jones airfoils have been investigated for both attached and separated transonic flows. Predictions for pressure distributions, loads, skin friction coefficients, boundary layer displacement thickness and velocity profiles are included and compared with experimental data when possible. Overall, the results are in good agreement with experimental data.

Mach Stability Improvements Using an Existing Second Throat Capability at the National Transonic Facility

NASA Technical Reports Server (NTRS)

Chan, David T.; Balakrishna, Sundareswara; Walker, Eric L.; Goodliff, Scott L.

2015-01-01

Recent data quality improvements at the National Transonic Facility have an intended goal of reducing the Mach number variation in a data point to within plus or minus 0.0005, with the ultimate goal of reducing the data repeatability of the drag coefficient for full-span subsonic transport models at transonic speeds to within half a drag count. This paper will discuss the Mach stability improvements achieved through the use of an existing second throat capability at the NTF to create a minimum area at the end of the test section. These improvements were demonstrated using both the NASA Common Research Model and the NTF Pathfinder-I model in recent experiments. Sonic conditions at the throat were verified using sidewall static pressure data. The Mach variation levels from both experiments in the baseline tunnel configuration and the choked tunnel configuration will be presented and the correlation between Mach number and drag will also be examined. Finally, a brief discussion is given on the consequences of using the second throat in its location at the end of the test section.

Mach Stability Improvements Using an Existing Second Throat Capability at the National Transonic Facility

NASA Technical Reports Server (NTRS)

Chan, David T.

2015-01-01

Recent data quality improvements at the National Transonic Facility (NTF) have an intended goal of reducing the Mach number variation in a data point to within unit vector A plus or minus 0.0005, with the ultimate goal of reducing the data repeatability of the drag coefficient for full-span subsonic transport models at transonic speeds to within half of a drag count. This paper will discuss the Mach stability improvements achieved through the use of an existing second throat capability at the NTF to create a minimum area at the end of the test section. These improvements were demonstrated using both the NASA Common Research Model and the NTF Pathfinder-I model in recent experiments. Sonic conditions at the throat were verified using sidewall static pressure data. The Mach variation levels from both experiments in the baseline tunnel configuration and the choked tunnel configuration will be presented. Finally, a brief discussion is given on the consequences of using the second throat in its location at the end of the test section.

Computational Analysis of the Transonic Dynamics Tunnel Using FUN3D

DOE Office of Scientific and Technical Information (OSTI.GOV)

Chwalowski, Pawel; Quon, Eliot; Brynildsen, Scott E.

This paper presents results from an explanatory two-year effort of applying Computational Fluid Dynamics (CFD) to analyze the empty-tunnel flow in the NASA Langley Research Center Transonic Dynamics Tunnel (TDT). The TDT is a continuous-flow, closed circuit, 16- x 16-foot slotted-test-section wind tunnel, with capabilities to use air or heavy gas as a working fluid. In this study, experimental data acquired in the empty tunnel using the R-134a test medium was used to calibrate the computational data. The experimental calibration data includes wall pressures, boundary-layer profiles, and the tunnel centerline Mach number profiles. Subsonic and supersonic flow regimes were considered,more » focusing on Mach 0.5, 0.7 and Mach 1.1 in the TDT test section. This study discusses the computational domain, boundary conditions, and initial conditions selected in the resulting steady-state analyses using NASA's FUN3D CFD software.« less

Empirical Flutter Prediction Method.

DTIC Science & Technology

1988-03-05

been used in this way to discover species or subspecies of animals, and to discover different types of voter or comsumer requiring different persuasions...respect to behavior or performance or response variables. Once this were done, corresponding clusters might be sought among descriptive or predictive or...jump in a response. The first sort of usage does not apply to the flutter prediction problem. Here the types of behavior are the different kinds of

Oscillating Cascade Aerodynamics at Large Mean Incidence Angles

NASA Technical Reports Server (NTRS)

Buffum, Daniel H.

1997-01-01

In a cooperative program with Pratt & Whitney, researchers obtained fundamental separated flow unsteady aerodynamic data in the NASA Lewis Research Center's Oscillating Cascade. These data fill a void that has hindered the understanding and prediction of subsonic and transonic stall flutter. For small-amplitude torsional oscillations, unsteady pressure distributions were measured on airfoils with cross sections representative of an advanced, low-aspect-ratio fan blade. Data were obtained for two mean incidence angles with a subsonic inflow. At high mean incidence angles (alpha = 10 deg), the mean flow separated at the leading edge and reattached at about 40 percent of the chord. For comparison purposes, data were also obtained for a low incidence angle (a = 0 deg) attached flow.

Design and test of three active flutter suppression controllers

NASA Technical Reports Server (NTRS)

Christhilf, David M.; Waszak, Martin R.; Adams, William M.; Srinathkumar, S.; Mukhopadhyay, Vivek

1991-01-01

Three flutter suppression control law design techniques are presented. Each uses multiple control surfaces and/or sensors. The first uses linear combinations of several accelerometer signals together with dynamic compensation to synthesize the modal rate of the critical mode for feedback to distributed control surfaces. The second uses traditional tools (pole/zero loci and Nyquist diagrams) to develop a good understanding of the flutter mechanism and produce a controller with minimal complexity and good robustness to plant uncertainty. The third starts with a minimum energy Linear Quadratic Gaussian controller, applies controller order reduction, and then modifies weight and noise covariance matrices to improve multi-variable robustness. The resulting designs were implemented digitally and tested subsonically on the Active Flexible Wing (AFW) wind tunnel model. Test results presented here include plant characteristics, maximum attained closed-loop dynamic pressure, and Root Mean Square control surface activity. A key result is that simultaneous symmetric and antisymmetric flutter suppression was achieved by the second control law, with a 24 percent increase in attainable dynamic pressure.

Auditory fovea and Doppler shift compensation: adaptations for flutter detection in echolocating bats using CF-FM signals.

PubMed

Schnitzler, Hans-Ulrich; Denzinger, Annette

2011-05-01

Rhythmical modulations in insect echoes caused by the moving wings of fluttering insects are behaviourally relevant information for bats emitting CF-FM signals with a high duty cycle. Transmitter and receiver of the echolocation system in flutter detecting foragers are especially adapted for the processing of flutter information. The adaptations of the transmitter are indicated by a flutter induced increase in duty cycle, and by Doppler shift compensation (DSC) that keeps the carrier frequency of the insect echoes near a reference frequency. An adaptation of the receiver is the auditory fovea on the basilar membrane, a highly expanded frequency representation centred to the reference frequency. The afferent projections from the fovea lead to foveal areas with an overrepresentation of sharply tuned neurons with best frequencies near the reference frequency throughout the entire auditory pathway. These foveal neurons are very sensitive to stimuli with natural and simulated flutter information. The frequency range of the foveal areas with their flutter processing neurons overlaps exactly with the frequency range where DS compensating bats most likely receive echoes from fluttering insects. This tight match indicates that auditory fovea and DSC are adaptations for the detection and evaluation of insects flying in clutter.

Development and application of an optimization procedure for flutter suppression using the aerodynamic energy concept

NASA Technical Reports Server (NTRS)

Nissim, E.; Abel, I.

1978-01-01

An optimization procedure is developed based on the responses of a system to continuous gust inputs. The procedure uses control law transfer functions which have been partially determined by using the relaxed aerodynamic energy approach. The optimization procedure yields a flutter suppression system which minimizes control surface activity in a gust environment. The procedure is applied to wing flutter of a drone aircraft to demonstrate a 44 percent increase in the basic wing flutter dynamic pressure. It is shown that a trailing edge control system suppresses the flutter instability over a wide range of subsonic mach numbers and flight altitudes. Results of this study confirm the effectiveness of the relaxed energy approach.

Pressure-distribution measurements on a transonic low-aspect ratio wing

NASA Technical Reports Server (NTRS)

Keener, E. R.

1985-01-01

Experimental surface pressure distributions and oil flow photographs are presented for a 0.90 m semispan model of NASA/Lockheed Wing C, a generic transonic, supercritical, low aspect ratio, highly 3-dimensional configuration. This wing was tested at the design angle of attack of 5 deg over a Mach number range from 0.25 to 0.96, and a Reynolds number range from 3.4 x 1,000,000 to 10 x 1,000,000. Pressures were measured with both the tunnel floor and ceiling suction slots open for most of the tests but taped closed for some tests to simulate solid walls. A comparison is made with the measured pressures from a small model in high Reynolds number facility and with predicted pressures using two three dimesional, transonic full potential flow wing codes: design code FLO22 (nonconservative) and TWING code (conservative). At the given design condition, a small region of flow separation occurred. At a Mach number of 0.82 the flow was unseparated and the surface flow angles were less than 10 deg, indicating that the boundary layer flow was not 3-D. Evidence indicate that wings that are optimized for mild shock waves and mild pressure recovery gradients generally have small 3-D boundary layer flow at design conditions for unseparated flow.

Thrust Removal Scheme for the FAST-MAC Circulation Control Model Tested in the National Transonic Facility

NASA Technical Reports Server (NTRS)

Chan, David T.; Milholen, William E., II; Jones, Gregory S.; Goodliff, Scott L.

2014-01-01

A second wind tunnel test of the FAST-MAC circulation control semi-span model was recently completed in the National Transonic Facility at the NASA Langley Research Center. The model allowed independent control of four circulation control plenums producing a high momentum jet from a blowing slot near the wing trailing edge that was directed over a 15% chord simple-hinged flap. The model was configured for transonic testing of the cruise configuration with 0deg flap deflection to determine the potential for drag reduction with the circulation control blowing. Encouraging results from analysis of wing surface pressures suggested that the circulation control blowing was effective in reducing the transonic drag on the configuration, however this could not be quantified until the thrust generated by the blowing slot was correctly removed from the force and moment balance data. This paper will present the thrust removal methodology used for the FAST-MAC circulation control model and describe the experimental measurements and techniques used to develop the methodology. A discussion on the impact to the force and moment data as a result of removing the thrust from the blowing slot will also be presented for the cruise configuration, where at some Mach and Reynolds number conditions, the thrust-removed corrected data showed that a drag reduction was realized as a consequence of the blowing.

Flutter Phenomenon in Flow Driven Energy Harvester–A Unified Theoretical Model for “Stiff” and “Flexible” Materials

PubMed Central

Chen, Yu; Mu, Xiaojing; Wang, Tao; Ren, Weiwei; Yang, Ya; Wang, Zhong Lin; Sun, Chengliang; Gu, Alex Yuandong

2016-01-01

Here, we report a stable and predictable aero-elastic motion in the flow-driven energy harvester, which is different from flapping and vortex-induced-vibration (VIV). A unified theoretical frame work that describes the flutter phenomenon observed in both “stiff” and “flexible” materials for flow driven energy harvester was presented in this work. We prove flutter in both types of materials is the results of the coupled effects of torsional and bending modes. Compared to “stiff” materials, which has a flow velocity-independent flutter frequency, flexible material presents a flutter frequency that almost linearly scales with the flow velocity. Specific to “flexible” materials, pre-stress modulates the frequency range in which flutter occurs. It is experimentally observed that a double-clamped “flexible” piezoelectric P(VDF-TrFE) thin belt, when driven into the flutter state, yields a 1,000 times increase in the output voltage compared to that of the non-fluttered state. At a fixed flow velocity, increase in pre-stress level of the P(VDF-TrFE) thin belt up-shifts the flutter frequency. In addition, this work allows the rational design of flexible piezoelectric devices, including flow-driven energy harvester, triboelectric energy harvester, and self-powered wireless flow speed sensor. PMID:27739484

Implementation of the WICS Wall Interference Correction System at the National Transonic Facility

NASA Technical Reports Server (NTRS)

Iyer, Venkit; Everhart, Joel L.; Bir, Pamela J.; Ulbrich, Norbert

2000-01-01

The Wall Interference Correction System (WICS) is operational at the National Transonic Facility (NTF) of NASA Langley Research Center (NASA LaRC) for semispan and full span tests in the solid wall (slots covered) configuration. The method is based on the wall pressure signature method for computing corrections to the measured parameters. It is an adaptation of the WICS code operational at the 12 ft pressure wind tunnel (12ft PWT) of NASA Ames Research Center (NASA ARC). This paper discusses the details of implementation of WICS at the NTF including tunnel calibration, code modifications for tunnel and support geometry, changes made for the NTF wall orifices layout, details of interfacing with the tunnel data processing system, and post-processing of results. Example results of applying WICS to a semispan test and a full span test are presented. Comparison with classical correction results and an analysis of uncertainty in the corrections are also given. As a special application of the code, the Mach number calibration data from a centerline pipe test was computed by WICS. Finally, future work for expanding the applicability of the code including online implementation is discussed.

Implementation of the WICS Wall Interference Correction System at the National Transonic Facility

NASA Technical Reports Server (NTRS)

Iyer, Venkit; Martin, Lockheed; Everhart, Joel L.; Bir, Pamela J.; Ulbrich, Norbert

2000-01-01

The Wall Interference Correction System (WICS) is operational at the National Transonic Facility (NTF) of NASA Langley Research Center (NASA LaRC) for semispan and full span tests in the solid wall (slots covered) configuration, The method is based on the wall pressure signature method for computing corrections to the measured parameters. It is an adaptation of the WICS code operational at the 12 ft pressure wind tunnel (12ft PWT) of NASA Ames Research Center (NASA ARC). This paper discusses the details of implementation of WICS at the NTF including, tunnel calibration, code modifications for tunnel and support geometry, changes made for the NTF wall orifices layout, details of interfacing with the tunnel data processing system, and post-processing of results. Example results of applying WICS to a semispan test and a full span test are presented. Comparison with classical correction results and an analysis of uncertainty in the corrections are also given. As a special application of the code, the Mach number calibration data from a centerline pipe test was computed by WICS. Finally, future work for expanding the applicability of the code including online implementation is discussed.

14 CFR 23.629 - Flutter.

Code of Federal Regulations, 2013 CFR

2013-01-01

... range up to VD/MD, or VDF/MDF for jets; (2) The vibratory response of the structure during the test indicates freedom from flutter; (3) A proper margin of damping exists at VD/MD, or VDF/MDF for jets; and (4) As VD/MD (or VDF/MDF for jets) is approached, there is no large or rapid reduction in damping. (c...

14 CFR 23.629 - Flutter.

Code of Federal Regulations, 2014 CFR

2014-01-01

... range up to VD/MD, or VDF/MDF for jets; (2) The vibratory response of the structure during the test indicates freedom from flutter; (3) A proper margin of damping exists at VD/MD, or VDF/MDF for jets; and (4) As VD/MD (or VDF/MDF for jets) is approached, there is no large or rapid reduction in damping. (c...

Flutter suppression and stability analysis for a variable-span wing via morphing technology

NASA Astrophysics Data System (ADS)

Li, Wencheng; Jin, Dongping

2018-01-01

A morphing wing can enhance aerodynamic characteristics and control authority as an alternative to using ailerons. To use morphing technology for flutter suppression, the dynamical behavior and stability of a variable-span wing subjected to the supersonic aerodynamic loads are investigated numerically in this paper. An axially moving cantilever plate is employed to model the variable-span wing, in which the governing equations of motion are established via the Kane method and piston theory. A morphing strategy based on axially moving rates is proposed to suppress the flutter that occurs beyond the critical span length, and the flutter stability is verified by Floquet theory. Furthermore, the transient stability during the morphing motion is analyzed and the upper bound of the morphing rate is obtained. The simulation results indicate that the proposed morphing law, which is varying periodically with a proper amplitude, could accomplish the flutter suppression. Further, the upper bound of the morphing speed decreases rapidly once the span length is close to its critical span length.

A Conceptual Wing Flutter Analysis Tool for Systems Analysis and Parametric Design Study

NASA Technical Reports Server (NTRS)

Mukhopadhyay, Vivek

2003-01-01

An interactive computer program was developed for wing flutter analysis in the conceptual design stage. The objective was to estimate flutt er instability boundaries of a typical wing, when detailed structural and aerodynamic data are not available. Effects of change in key flu tter parameters can also be estimated in order to guide the conceptual design. This userfriendly software was developed using MathCad and M atlab codes. The analysis method was based on non-dimensional paramet ric plots of two primary flutter parameters, namely Regier number and Flutter number, with normalization factors based on wing torsion stiffness, sweep, mass ratio, taper ratio, aspect ratio, center of gravit y location and pitch-inertia radius of gyration. These parametric plo ts were compiled in a Chance-Vought Corporation report from database of past experiments and wind tunnel test results. An example was prese nted for conceptual flutter analysis of outer-wing of a Blended-Wing- Body aircraft.

Design optimization of axisymmetric bodies in nonuniform transonic flow

NASA Technical Reports Server (NTRS)

Lan, C. Edward

1989-01-01

An inviscid transonic code capable of designing an axisymmetric body in a uniform or nonuniform flow was developed. The design was achieved by direct optimiation by coupling an analysis code with an optimizer. Design examples were provided for axisymmetric bodies with fineness ratios of 8.33 and 5 at different Mach numbers. It was shown that by reducing the nose radius and increasing the afterbody thickness of initial shapes obtained from symmetric NACA four-digit airfoil contours, wave drag could be reduced by 29 percent for a body of fineness ratio 8.33 in a nonuniform transonic flow of M = 0.98 to 0.995. The reduction was 41 percent for a body of fineness ratio 5 in a uniform transonic flow of M = 0.925 and 65 percent for the same body but in a nonuniform transonic flow of M = 0.90 to 0.95.

The NASA aircraft noise prediction program improved propeller analysis system

NASA Technical Reports Server (NTRS)

Nguyen, L. Cathy

1991-01-01

The improvements and the modifications of the NASA Aircraft Noise Prediction Program (ANOPP) and the Propeller Analysis System (PAS) are described. Comparisons of the predictions and the test data are included in the case studies for the flat plate model in the Boundary Layer Module, for the effects of applying compressibility corrections to the lift and pressure coefficients, for the use of different weight factors in the Propeller Performance Module, for the use of the improved retarded time equation solution, and for the effect of the number grids in the Transonic Propeller Noise Module. The DNW tunnel test data of a propeller at different angles of attack and the Dowty Rotol data are compared with ANOPP predictions. The effect of the number of grids on the Transonic Propeller Noise Module predictions and the comparison of ANOPP TPN and DFP-ATP codes are studied. In addition to the above impact studies, the transonic propeller noise predictions for the SR-7, the UDF front rotor, and the support of the enroute noise test program are included.

Comparison of Computational and Experimental Results for a Transonic Variable-Speed Power-Turbine Blade Operating with Low Inlet Turbulence Levels

NASA Technical Reports Server (NTRS)

Booth, David; Flegel, Ashlie

2015-01-01

A computational assessment of the aerodynamic performance of the midspan section of a variable-speed power-turbine blade is described. The computation comprises a periodic single blade that represents the 2-D Midspan section VSPT blade that was tested in the NASA Glenn Research Center Transonic Turbine Blade Cascade Facility. A commercial, off-the-shelf (COTS) software package, Pointwise and CFD++, was used for the grid generation and RANS and URANS computations. The CFD code, which offers flexibility in terms of turbulence and transition modeling options, was assessed in terms of blade loading, loss, and turning against test data from the transonic tunnel. Simulations were assessed at positive and negative incidence angles that represent the turbine cruise and take-off design conditions. The results indicate that the secondary flow induced at the positive incidence cruise condition results in a highly loaded case and transitional flow on the blade is observed. The negative incidence take-off condition is unloaded and the flow is very two-dimensional. The computational results demonstrate the predictive capability of the gridding technique and COTS software for a linear transonic turbine blade cascade with large incidence angle variation.

Comparison of Computational and Experimental Results for a Transonic Variable-speed Power-Turbine Blade Operating with Low Inlet Turbulence Levels

NASA Technical Reports Server (NTRS)

Booth, David T.; Flegel, Ashlie B.

2015-01-01

A computational assessment of the aerodynamic performance of the midspan section of a variable-speed power-turbine blade is described. The computation comprises a periodic single blade that represents the 2-D Midspan section VSPT blade that was tested in the NASA Glenn Research Center Transonic Turbine Blade Cascade Facility. A commercial, off-the-shelf (COTS) software package, Pointwise and CFD++, was used for the grid generation and RANS and URANS computations. The CFD code, which offers flexibility in terms of turbulence and transition modeling options, was assessed in terms of blade loading, loss, and turning against test data from the transonic tunnel. Simulations were assessed at positive and negative incidence angles that represent the turbine cruise and take-off design conditions. The results indicate that the secondary flow induced at the positive incidence cruise condition results in a highly loaded case and transitional flow on the blade is observed. The negative incidence take-off condition is unloaded and the flow is very two-dimensional. The computational results demonstrate the predictive capability of the gridding technique and COTS software for a linear transonic turbine blade cascade with large incidence angle variation.

An analytical and experimental investigation of flutter suppression via piezoelectric actuation

NASA Technical Reports Server (NTRS)

Heeg, Jennifer

1992-01-01

The objective of this research was to analytically and experimentally study the capabilities of adaptive material plate actuators for suppressing flutter. Piezoelectrics are materials which are characterized by their ability to produce voltage when subjected to a mechanical strain. The converse piezoelectric effect can be utilized to actuate a structure by applying a voltage. For this investigation, a two degree of freedom wind-tunnel model was designed, analyzed, and tested. The model consisted of a rigid wing and a flexible mount system which permitted translational and rotational degrees of freedom. Actuators, made of piezoelectric material were affixed to leaf springs on the mount system. Command signals, applied to the piezoelectric actuators, exerted control over the closed-loop damping and stiffness properties. A mathematical aeroservoelastic model was constructed using finite element and stiffness properties. A mathematical aeroservoelastic model was constructed using finite element methods, laminated plate theory, and aeroelastic analysis tools. A flutter suppression control law was designed, implemented on a digital control computer, and tested to conditions 20 percent above the passive flutter speed of the model. The experimental results represent the first time that adaptive materials have been used to actively suppress flutter. It demonstrates that small, carefully-placed actuating plates can be used effectively to control aeroelastic response.

Testing and checkout experiences in the National Transonic Facility since becoming operational

NASA Technical Reports Server (NTRS)

Bruce, W. E., Jr.; Gloss, B. B.; Mckinney, L. W.

1988-01-01

The U.S. National Transonic Facility, constructed by NASA to meet the national needs for High Reynolds Number Testing, has been operational in a checkout and test mode since the operational readiness review (ORR) in late 1984. During this time, there have been problems centered around the effect of large temperature excursions on the mechanical movement of large components, the reliable performance of instrumentation systems, and an unexpected moisture problem with dry insulation. The more significant efforts since the ORR are reviewed and NTF status concerning hardware, instrumentation and process controls systems, operating constraints imposed by the cryogenic environment, and data quality and process controls is summarized.

Experiences in using the CYBER 203 for three-dimensional transonic flow calculations

NASA Technical Reports Server (NTRS)

Melson, N. D.; Keller, J. D.

1982-01-01

In this paper, the authors report on some of their experiences modifying two three-dimensional transonic flow programs (FLO22 and FLO27) for use on the NASA Langley Research Center CYBER 203. Both of the programs discussed were originally written for use on serial machines. Several methods were attempted to optimize the execution of the two programs on the vector machine, including: (1) leaving the program in a scalar form (i.e., serial computation) with compiler software used to optimize and vectorize the program, (2) vectorizing parts of the existing algorithm in the program, and (3) incorporating a new vectorizable algorithm (ZEBRA I or ZEBRA II) in the program.

Laser anemometer measurements in a transonic axial-flow fan rotor

NASA Technical Reports Server (NTRS)

Strazisar, Anthony J.; Wood, Jerry R.; Hathaway, Michael D.; Suder, Kenneth L.

1989-01-01

Laser anemometer surveys were made of the 3-D flow field in NASA rotor 67, a low aspect ratio transonic axial-flow fan rotor. The test rotor has a tip relative Mach number of 1.38. The flowfield was surveyed at design speed at near peak efficiency and near stall operating conditions. Data is presented in the form of relative Mach number and relative flow angle distributions on surfaces of revolution at nine spanwise locations evenly spaced from hub to tip. At each spanwise location, data was acquired upstream, within, and downstream of the rotor. Aerodynamic performance measurements and detailed rotor blade and annulus geometry are also presented so that the experimental results can be used as a test case for 3-D turbomachinery flow analysis codes.

Rotor design of high tip speed low loading transonic fan.

NASA Technical Reports Server (NTRS)

Erwin, J. R.; Vitale, N. G.

1972-01-01

This paper describes the design concepts, principles and details of a high tip speed transonic rotor having low aerodynamic loading. The purpose of the NASA sponsored investigation was to determine whether good efficiency and large stall margin could be obtained by designing a rotor to avoid flow separation associated with strong normal shocks. Fully supersonic flow through the outboard region of the rotor with compression accomplished by weak oblique shocks were major design concepts employed. Computer programs were written and used to derive blade sections consistent from the all-supersonic tip region to the all-subsonic hub region. Preliminary test results indicate attainment of design pressure ratio and design flow at design speed with about a 1.6 point decrement in efficiency and large stall margin.

Inlet flow field investigation. Part 1: Transonic flow field survey

NASA Technical Reports Server (NTRS)

Yetter, J. A.; Salemann, V.; Sussman, M. B.

1984-01-01

A wind tunnel investigation was conducted to determine the local inlet flow field characteristics of an advanced tactical supersonic cruise airplane. A data base for the development and validation of analytical codes directed at the analysis of inlet flow fields for advanced supersonic airplanes was established. Testing was conducted at the NASA-Langley 16-foot Transonic Tunnel at freestream Mach numbers of 0.6 to 1.20 and angles of attack from 0.0 to 10.0 degrees. Inlet flow field surveys were made at locations representative of wing (upper and lower surface) and forebody mounted inlet concepts. Results are presented in the form of local inlet flow field angle of attack, sideflow angle, and Mach number contours. Wing surface pressure distributions supplement the flow field data.

Cruise Speed Sensitivity Study for Transonic Truss Braced Wing

NASA Technical Reports Server (NTRS)

Wells, Douglas P.

2017-01-01

NASA's investment and research in aviation has led to new technologies and concepts that make aircraft more efficient and environmentally friendly. One aircraft design operational concept is the reduction of cruise speed to reduce fuel burned during a mission. Although this is not a new idea, it was used by all of the contractors involved in a 2008 NASA sponsored study that solicited concept and technology ideas to reduce environmental impacts for future subsonic passenger transports. NASA is currently improving and building new analysis capabilities to analyze advanced concepts. To test some of these new capabilities, a transonic truss braced wing configuration was used as a test case. This paper examines the effects due to changes in the design cruise speed and other tradeoffs in the design space. The analysis was baselined to the Boeing SUGAR High truss braced wing concept. An optimization was run at five different design cruise Mach numbers. These designs are compared to provide an initial assessment space and the parameters that should be considered when selecting a design cruise speed. A discussion of the design drivers is also included. The results show that the wing weight in the current analysis has more influence on the takeoff gross weight than expected. This effect caused lower than expected wing sweep angle values for higher cruise speed designs.

Beam-modulation methods in quantitative and flow visualization holographic interferometry

NASA Technical Reports Server (NTRS)

Decker, A.

1986-01-01

This report discusses heterodyne holographic interferometry and time-average holography with a frequency shifted reference beam. Both methods will be used for the measurement and visualization of internal transonic flows, where the target facility is a flutter cascade. The background and experimental requirements for both methods are reviewed. Measurements using heterodyne holographic interferometry are presented. The performance of the laser required for time-average holography of time-varying transonic flows is discussed.

Suppression of background noise in a transonic wind-tunnel test section

NASA Technical Reports Server (NTRS)

Schutzenhofer, L. A.; Howard, P. W.

1975-01-01

Some exploratory tests were recently performed in the transonic test section of the NASA Marshall Space Flight Center 14-in. wind tunnel to suppress the background noise. In these tests, the perforated walls of the test section were covered with fine wire screens. The screens eliminated the edge tones generated by the holes in the perforated walls and significantly reduced the tunnel background noise. The tunnel noise levels were reduced to such a degree by this simple modification at Mach numbers 0.75, 0.9, 1.1, 1.2, and 1.46 that the fluctuating pressure levels of a turbulent boundary layer could be measured on a 5-deg half-angle cone.

Overview of Transonic to Hypersonic Stage Separation Tool Development for Multi-Stage-to-Orbit Concepts

NASA Technical Reports Server (NTRS)

Murphy, Kelly J.; Bunning, Pieter G.; Pamadi, Bandu N.; Scallion, William I.; Jones, Kenneth M.

2004-01-01

An overview of research efforts at NASA in support of the stage separation and ascent aerothermodynamics research program is presented. The objective of this work is to develop a synergistic suite of experimental, computational, and engineering tools and methods to apply to vehicle separation across the transonic to hypersonic speed regimes. Proximity testing of a generic bimese wing-body configuration is on-going in the transonic (Mach numbers 0.6, 1.05, and 1.1), supersonic (Mach numbers 2.3, 3.0, and 4.5) and hypersonic (Mach numbers 6 and 10) speed regimes in four wind tunnel facilities at the NASA Langley Research Center. An overset grid, Navier-Stokes flow solver has been enhanced and demonstrated on a matrix of proximity cases and on a dynamic separation simulation of the bimese configuration. Steady-state predictions with this solver were in excellent agreement with wind tunnel data at Mach 3 as were predictions via a Cartesian-grid Euler solver. Experimental and computational data have been used to evaluate multi-body enhancements to the widely-used Aerodynamic Preliminary Analysis System, an engineering methodology, and to develop a new software package, SepSim, for the simulation and visualization of vehicle motions in a stage separation scenario. Web-based software will be used for archiving information generated from this research program into a database accessible to the user community. Thus, a framework has been established to study stage separation problems using coordinated experimental, computational, and engineering tools.

FLEET Velocimetry Measurements on a Transonic Airfoil

NASA Technical Reports Server (NTRS)

Burns, Ross A.; Danehy, Paul M.

2017-01-01

Femtosecond laser electronic excitation tagging (FLEET) velocimetry was used to study the flowfield around a symmetric, transonic airfoil in the NASA Langley 0.3-m TCT facility. A nominal Mach number of 0.85 was investigated with a total pressure of 125 kPa and total temperature of 280 K. Two-components of velocity were measured along vertical profiles at different locations above, below, and aft of the airfoil at angles of attack of 0 deg, 3.5 deg, and 7deg. Measurements were assessed for their accuracy, precision, dynamic range, spatial resolution, and overall measurement uncertainty in the context of the applied flowfield. Measurement precisions as low as 1 m/s were observed, while overall uncertainties ranged from 4 to 5 percent. Velocity profiles within the wake showed sufficient accuracy, precision, and sensitivity to resolve both the mean and fluctuating velocities and general flow physics such as shear layer growth. Evidence of flow separation is found at high angles of attack.

Higher-Order Spectral Analysis of F-18 Flight Flutter Data

NASA Technical Reports Server (NTRS)

Silva, Walter A.; Dunn, Shane

2005-01-01

Royal Australian Air Force (RAAF) F/A-18 flight flutter test data is presented and analyzed using various techniques. The data includes high-quality measurements of forced responses and limit cycle oscillation (LCO) phenomena. Standard correlation and power spectral density (PSD) techniques are applied to the data and presented. Novel applications of experimentally-identified impulse responses and higher-order spectral techniques are also applied to the data and presented. The goal of this research is to develop methods that can identify the onset of nonlinear aeroelastic phenomena, such as LCO, during flutter testing.

Influences of Models on the Unsteady Pressure Characteristics of the NASA National Transonic Facility

NASA Technical Reports Server (NTRS)

Jones, Gregory; Balakrishna, Sundareswara; DeMoss, Joshua; Goodliff, Scott; Bailey, Matthew

2015-01-01

Pressure fluctuations have been measured over the course of several tests in the National Transonic Facility to study unsteady phenomenon both with and without the influence of a model. Broadband spectral analysis will be used to characterize the length scales of the tunnel. Special attention will be given to the large-scale, low frequency data that influences the Mach number and force and moment variability. This paper will also discuss the significance of the vorticity and sound fields that can be related to the Common Research Model and will also highlight the comparisons to an empty tunnel configuration. The effectiveness of vortex generators placed at the interface of the test section and wind tunnel diffuser showed promise in reducing the empty tunnel unsteadiness, however, the vortex generators were ineffective in the presence of a model.

Activities in Aeroelasticity at NASA Langley Research Center

NASA Technical Reports Server (NTRS)

Perry, Boyd, III; Noll, Thomas E.

1997-01-01

This paper presents the results of recently-completed research and presents status reports of current research being performed within the Aeroelasticity Branch of the NASA Langley Research Center. Within the paper this research is classified as experimental, analytical, and theoretical aeroelastic research. The paper also describes the Langley Transonic Dynamics Tunnel, its features, capabilities, a new open-architecture data acquisition system, ongoing facility modifications, and the subsequent calibration of the facility.

On the application of transonic similarity rules to wings of finite span

NASA Technical Reports Server (NTRS)

Spreiter, John R

1953-01-01

The transonic aerodynamic characteristics of wings of finite span are discussed from the point of view of a unified small perturbation theory for subsonic, transonic, and supersonic flows about thin wings. This approach avoids certain ambiguities which appear if one studies transonic flows by means of equations derived under the more restrictive assumption that the local velocities are everywhere close to sonic velocity. The relation between the two methods of analysis of transonic flow is examined, the similarity rules and known solutions of transonic flow theory are reviewed, and the asymptotic behavior of the lift, drag, and pitching-moment characteristics of wings of large and small aspect ratio is discussed. It is shown that certain methods of data presentation are advantageous for the effective display of these characteristics.

Numerical calculations of two dimensional, unsteady transonic flows with circulation

NASA Technical Reports Server (NTRS)

Beam, R. M.; Warming, R. F.

1974-01-01

The feasibility of obtaining two-dimensional, unsteady transonic aerodynamic data by numerically integrating the Euler equations is investigated. An explicit, third-order-accurate, noncentered, finite-difference scheme is used to compute unsteady flows about airfoils. Solutions for lifting and nonlifting airfoils are presented and compared with subsonic linear theory. The applicability and efficiency of the numerical indicial function method are outlined. Numerically computed subsonic and transonic oscillatory aerodynamic coefficients are presented and compared with those obtained from subsonic linear theory and transonic wind-tunnel data.

Stability analysis of nonlinear autonomous systems - General theory and application to flutter

NASA Technical Reports Server (NTRS)

Smith, L. L.; Morino, L.

1975-01-01

The analysis makes use of a singular perturbation method, the multiple time scaling. Concepts of stable and unstable limit cycles are introduced. The solution is obtained in the form of an asymptotic expansion. Numerical results are presented for the nonlinear flutter of panels and airfoils in supersonic flow. The approach used is an extension of a method for analyzing nonlinear panel flutter reported by Morino (1969).

Changing axis deviation and paroxysmal atrial flutter associated with subclinical hyperthyroidism.

PubMed

Patanè, Salvatore; Marte, Filippo

2010-10-08

Subclinical hyperthyroidism is an increasingly recognized entity that is defined as a normal serum free thyroxine and free triiodothyronine levels with a thyroid-stimulating hormone level suppressed below the normal range and usually undetectable. It has been reported that subclinical hyperthyroidism is not associated with coronary heart disease or mortality from cardiovascular causes but it is sufficient to induce arrhythmias including atrial fibrillation and atrial flutter. It has also been reported that increased factor X activity in patients with subclinical hyperthyroidism represents a potential hypercoagulable state. Rarely, it has also been reported intermittent changing axis deviation during atrial fibrillation and during atrial flutter. We present a case of paroxysmal atrial flutter and changing axis deviation associated with subclinical hyperthyroidism, in a 76-year-old Italian man. Also this case focuses attention on the importance of a correct evaluation of subclinical hyperthyroidism. Copyright © 2008 Elsevier Ireland Ltd. All rights reserved.

Hypersonic panel flutter in a rarefied atmosphere

NASA Technical Reports Server (NTRS)

Resende, Hugo B.

1993-01-01

Panel flutter is a form of dynamic aeroelastic instability resulting from the interaction between motion of an aircraft structural panel and the aerodynamic loads exerted on that panel by air flowing past one of the faces. It differs from lifting surface flutter in the sense that it is not usually catastrophic, the panel's motion being limited by nonlinear membrane stresses produced by the transverse displacement. Above some critical airflow condition, the linear instability grows to a limit cycle . The present investigation studies panel flutter in an aerodynamic regime known as 'free molecule flow', wherein intermolecular collisions can be neglected and loads are caused by interactions between individual molecules and the bounding surface. After collision with the panel, molecules may be reflected specularly or reemitted in diffuse fashion. Two parameters characterize this process: the 'momentum accommodation coefficient', which is the fraction of the specularly reflected molecules; and the ratio between the panel temperature and that of the free airstream. This model is relevant to the case of hypersonic flight vehicles traveling at very high altitudes and especially for panels oriented parallel to the airstream or in the vehicle's lee. Under these conditions the aerodynamic shear stress turns out to be considerably larger than the surface pressures, and shear effects must be included in the model. This is accomplished by means of distributed longitudinal and bending loads. The former can cause the panel to buckle. In the example of a simply-supported panel, it turns out that the second mode of free vibration tends to dominate the flutter solution, which is carried out by a Galerkin analysis. Several parametric studies are presented. They include the effects of (1) temperature ratio; (2) momentum accommodation coefficient; (3) spring parameters, which are associated with how the panel is connected to adjacent structures; (4) a parameter which relates compressive

Critical and post-critical behaviour of two-degree-of-freedom flutter-based generators

NASA Astrophysics Data System (ADS)

Pigolotti, Luca; Mannini, Claudio; Bartoli, Gianni; Thiele, Klaus

2017-09-01

Energy harvesting from flow-induced vibrations is a recent research field, which considers a diverse range of systems, among which two-degree-of-freedom flutter-based solutions were individuated as good candidates to obtain high energy performance. In the present work, numerical linear analyses and wind-tunnel tests were conducted on a flat-plate sectional model. The aim is to identify some design guidelines for generators exploiting the classical-flutter instability, through the investigation of the critical condition and the response during the post-critical regime. Many sets of governing parameters of interest from the energy-harvesting point of view were considered, including high levels of heaving damping to simulate the operation of a conversion apparatus. In particular, eccentricity of the elastic centre and small downstream mass unbalance can be introduced as solutions aiming at optimal operative ranges. The collected results suggest the high potentiality of flutter-based generators, and a significant enhancement of performance can be envisaged. Moreover, they contribute to improve the knowledge of the flutter excitation mechanism and to widen the dataset of measurements in the post-critical regime.

Theoretical Studies of Three Dimensional Transonic Flow through a Compressor Blade Row.

DTIC Science & Technology

1980-11-30

Row", Calspan Report No. AB-5487-A-l, AFOSR-TR-76- 1082 , AD-A031234, (August 1976). 2 Rae, W.J., "Relaxation Solutions for Three-Dimensional Transonic...S487-A-1, AFOSR-TR-76- 1082 , AD-A031234, (August 1976). 2. Rae, W.J., "Relaxation Solutions for Three-Dimensional Transonic Flow Through a Compressor

Modernization of the Transonic Axial Compressor Test Rig

DTIC Science & Technology

2017-12-01

13. ABSTRACT (maximum 200 words) This work presents the design and simulation process of modernizing the Naval Postgraduate School’s transonic...fabricate the materials. Stiffness tests and modal analysis were conducted via Finite Element Analysis (FEA) software. This analysis was used to design ...work presents the design and simulation process of modernizing the Naval Postgraduate School’s transonic compressor test rig (TCR). The TCR, which

Transonic Wing Shape Optimization Using a Genetic Algorithm

NASA Technical Reports Server (NTRS)

Holst, Terry L.; Pulliam, Thomas H.; Kwak, Dochan (Technical Monitor)

2002-01-01

A method for aerodynamic shape optimization based on a genetic algorithm approach is demonstrated. The algorithm is coupled with a transonic full potential flow solver and is used to optimize the flow about transonic wings including multi-objective solutions that lead to the generation of pareto fronts. The results indicate that the genetic algorithm is easy to implement, flexible in application and extremely reliable.

Further Investigation of the Support System Effects and Wing Twist on the NASA Common Research Model

NASA Technical Reports Server (NTRS)

Rivers, Melissa B.; Hunter, Craig A.; Campbell, Richard L.

2012-01-01

An experimental investigation of the NASA Common Research Model was conducted in the NASA Langley National Transonic Facility and NASA Ames 11-foot Transonic Wind Tunnel Facility for use in the Drag Prediction Workshop. As data from the experimental investigations was collected, a large difference in moment values was seen between the experiment and computational data from the 4th Drag Prediction Workshop. This difference led to a computational assessment to investigate model support system interference effects on the Common Research Model. The results from this investigation showed that the addition of the support system to the computational cases did increase the pitching moment so that it more closely matched the experimental results, but there was still a large discrepancy in pitching moment. This large discrepancy led to an investigation into the shape of the as-built model, which in turn led to a change in the computational grids and re-running of all the previous support system cases. The results of these cases are the focus of this paper.

Computation of transonic flow about helicopter rotor blades

NASA Technical Reports Server (NTRS)

Arieli, R.; Tauber, M. E.; Saunders, D. A.; Caughey, D. A.

1986-01-01

An inviscid, nonconservative, three-dimensional full-potential flow code, ROT22, has been developed for computing the quasi-steady flow about a lifting rotor blade. The code is valid throughout the subsonic and transonic regime. Calculations from the code are compared with detailed laser velocimeter measurements made in the tip region of a nonlifting rotor at a tip Mach number of 0.95 and zero advance ratio. In addition, comparisons are made with chordwise surface pressure measurements obtained in a wind tunnel for a nonlifting rotor blade at transonic tip speeds at advance ratios from 0.40 to 0.50. The overall agreement between theoretical calculations and experiment is very good. A typical run on a CRAY X-MP computer requires about 30 CPU seconds for one rotor position at transonic tip speed.

Transonic Dynamics Tunnel Force and Pressure Data Acquired on the HSR Rigid Semispan Model

NASA Technical Reports Server (NTRS)

Schuster, David M.; Rausch, Russ D.

1999-01-01

This report describes the aerodynamic data acquired on the High Speed Research Rigid Semispan Model (HSR-RSM) during NASA Langley Transonic Dynamics Tunnel (TDT) Test 520 conducted from 18 March to 4 April, 1996. The purpose of this test was to assess the aerodynamic character of a rigid high speed civil transport wing. The wing was fitted with a single trailing edge control surface which was both steadily deflected and oscillated during the test to investigate the response of the aerodynamic data to steady and unsteady control motion. Angle-of-attack and control surface deflection polars at subsonic, transonic and low-supersonic Mach numbers were obtained in the tunnel?s heavy gas configuration. Unsteady pressure and steady loads data were acquired on the wing, while steady pressures were measured on the fuselage. These data were reduced using a variety of methods, programs and computer systems. The reduced data was ultimately compiled onto a CD-ROM volume which was distributed to HSR industry team members in July, 1996. This report documents the methods used to acquire and reduce the data, and provides an assessment of the quality, repeatability, and overall character of the aerodynamic data measured during this test.

Aerodynamic Measurements of an Incidence Tolerant Blade in a Transonic Turbine Cascade

NASA Technical Reports Server (NTRS)

McVetta, Ashlie B.; Giel, Paul W.

2012-01-01

An overview of the recent facility modifications to NASA s Transonic Turbine Blade Cascade Facility and aerodynamic measurements on the VSPT incidence-tolerant blade are presented. This work supports the development of variable-speed power turbine (VSPT) speed-change technology for the NASA Large Civil Tilt Rotor (LCTR) vehicle. In order to maintain acceptable main rotor propulsive efficiency, the VSPT operates over a nearly 50% speed range from takeoff to altitude cruise. This results in 50 or more variations in VSPT blade incidence angles. The Transonic Turbine Blade Cascade Facility has the ability to operate over a wide range of Reynolds numbers and Mach numbers, but had to be modified in order to accommodate the negative incidence angle variation required by the LCTR VSPT operation. Details of the modifications are described. An incidence-tolerant blade was developed under an RTPAS study contract and tested in the cascade to look at the effects of large incidence angle and Reynolds number variations. Recent test results are presented which include midspan exit total pressure and flow angle measurements obtained at three inlet angles representing the cruise, take-off, and maximum incidence flight mission points. For each inlet angle, data were obtained at five flow conditions with exit Reynolds numbers varying from 2.12 106 to 2.12 105 and two isentropic exit Mach numbers of 0.72 and 0.35. Three-dimensional flowfield measurements were also acquired at the cruise and take-off points. The flowfield measurements were acquired using a five-hole and three-hole pneumatic probe located in a survey plane 8.6% axial chord downstream of the blade trailing edge plane and covering three blade passages. Blade and endwall static pressure distributions were also acquired for each flow condition.

Three-Dimensional Flow Field Measurements in a Transonic Turbine Cascade

NASA Technical Reports Server (NTRS)

Giel, P. W.; Thurman, D. R.; Lopez, I.; Boyle, R. J.; VanFossen, G. J.; Jett, T. A.; Camperchioli, W. P.; La, H.

1996-01-01

Three-dimensional flow field measurements are presented for a large scale transonic turbine blade cascade. Flow field total pressures and pitch and yaw flow angles were measured at an inlet Reynolds number of 1.0 x 10(exp 6) and at an isentropic exit Mach number of 1.3 in a low turbulence environment. Flow field data was obtained on five pitchwise/spanwise measurement planes, two upstream and three downstream of the cascade, each covering three blade pitches. Three-hole boundary layer probes and five-hole pitch/yaw probes were used to obtain data at over 1200 locations in each of the measurement planes. Blade and endwall static pressures were also measured at an inlet Reynolds number of 0.5 x 10(exp 6) and at an isentropic exit Mach number of 1.0. Tests were conducted in a linear cascade at the NASA Lewis Transonic Turbine Blade Cascade Facility. The test article was a turbine rotor with 136 deg of turning and an axial chord of 12.7 cm. The flow field in the cascade is highly three-dimensional as a result of thick boundary layers at the test section inlet and because of the high degree of flow turning. The large scale allowed for very detailed measurements of both flow field and surface phenomena. The intent of the work is to provide benchmark quality data for CFD code and model verification.

NASA Rotor 37 CFD Code Validation: Glenn-HT Code

NASA Technical Reports Server (NTRS)

Ameri, Ali A.

2010-01-01

In order to advance the goals of NASA aeronautics programs, it is necessary to continuously evaluate and improve the computational tools used for research and design at NASA. One such code is the Glenn-HT code which is used at NASA Glenn Research Center (GRC) for turbomachinery computations. Although the code has been thoroughly validated for turbine heat transfer computations, it has not been utilized for compressors. In this work, Glenn-HT was used to compute the flow in a transonic compressor and comparisons were made to experimental data. The results presented here are in good agreement with this data. Most of the measures of performance are well within the measurement uncertainties and the exit profiles of interest agree with the experimental measurements.

Remodeling of sinus node function after catheter ablation of right atrial flutter.

PubMed

Daoud, Emile G; Weiss, Raul; Augostini, Ralph S; Kalbfleisch, Steven J; Schroeder, Jason; Polsinelli, Georgia; Hummel, John D

2002-01-01

The purpose of this study was to investigate the effect of ablation of right atrial flutter upon sinus node function in humans. This study enrolled 35 patients. Twenty-four patients (16 men and 8 women; age 68 +/- 11 years) were referred for ablation of persistent atrial flutter (duration 8 +/- 11 months). After ablation, there was abnormal sinus node function defined as a corrected sinus node recovery time (CSNRT) > or = 550 msec. The control group consisted of 11 patients who were undergoing pacemaker implantation for sinus node disease but did not have a history of atrial dysrhythmias or ablation. Within 24 hours of ablation or pacemaker implantation, baseline maximal CSNRT was measured through a permanent pacemaker by AAI pacing at six cycle lengths: 600, 550, 500, 450, 400, and 350 msec. CSNRT then was measured in the same manner at 48 hours, 14 days, and 3 months after ablation/pacemaker implantation. P wave amplitude and duration, and percent atrial sensing also were assessed at the same intervals. For patients undergoing atrial flutter ablation, there was progressive temporal recovery of CSNRT (1,204 +/- 671 msec at baseline vs 834 +/- 380 msec at 3 months; P < 0.001) and a significant increase in the percent atrial sensing and P wave amplitude at 3 months compared with baseline (P < 0.001). In control subjects, there was no change in the CSNRT, percent atrial pacing, or P wave amplitude. After ablation of persistent atrial flutter, there is temporal recovery of CSNRT and increase in spontaneous atrial activity. These findings suggest that atrial flutter induces reversible changes in sinus node function.

Finite elements and finite differences for transonic flow calculations

NASA Technical Reports Server (NTRS)

Hafez, M. M.; Murman, E. M.; Wellford, L. C.

1978-01-01

The paper reviews the chief finite difference and finite element techniques used for numerical solution of nonlinear mixed elliptic-hyperbolic equations governing transonic flow. The forms of the governing equations for unsteady two-dimensional transonic flow considered are the Euler equation, the full potential equation in both conservative and nonconservative form, the transonic small-disturbance equation in both conservative and nonconservative form, and the hodograph equations for the small-disturbance case and the full-potential case. Finite difference methods considered include time-dependent methods, relaxation methods, semidirect methods, and hybrid methods. Finite element methods include finite element Lax-Wendroff schemes, implicit Galerkin method, mixed variational principles, dual iterative procedures, optimal control methods and least squares.

Active flutter suppression using optical output feedback digital controllers

NASA Technical Reports Server (NTRS)

1982-01-01

A method for synthesizing digital active flutter suppression controllers using the concept of optimal output feedback is presented. A convergent algorithm is employed to determine constrained control law parameters that minimize an infinite time discrete quadratic performance index. Low order compensator dynamics are included in the control law and the compensator parameters are computed along with the output feedback gain as part of the optimization process. An input noise adjustment procedure is used to improve the stability margins of the digital active flutter controller. Sample rate variation, prefilter pole variation, control structure variation and gain scheduling are discussed. A digital control law which accommodates computation delay can stabilize the wing with reasonable rms performance and adequate stability margins.

A Worst-Case Approach for On-Line Flutter Prediction

NASA Technical Reports Server (NTRS)

Lind, Rick C.; Brenner, Martin J.

1998-01-01

Worst-case flutter margins may be computed for a linear model with respect to a set of uncertainty operators using the structured singular value. This paper considers an on-line implementation to compute these robust margins in a flight test program. Uncertainty descriptions are updated at test points to account for unmodeled time-varying dynamics of the airplane by ensuring the robust model is not invalidated by measured flight data. Robust margins computed with respect to this uncertainty remain conservative to the changing dynamics throughout the flight. A simulation clearly demonstrates this method can improve the efficiency of flight testing by accurately predicting the flutter margin to improve safety while reducing the necessary flight time.

Flutter instability of freely hanging articulated pipes conveying fluid

NASA Astrophysics Data System (ADS)

Schouveiler, Lionel; Chermette, Félix

2018-03-01

We experimentally investigate the stability of freely hanging articulated pipes made of rigid segments connected by flexible joints and with their displacements constrained in a vertical plane. When the velocity of the fluid conveyed by the pipe is increased, flutter-type instability occurs above a critical value. The critical velocity and the characteristics of the flutter modes (frequency, amplitude, and shape) are determined as a function of the number n of segments into the pipe which is varied from 2 to 5. Experimental results are compared to predictions from linear stability analysis extending previous studies by taking into account damping due to the dissipation in the joints. Qualitative agreement is found and the limits of the analysis are discussed.

NACA0012 benchmark model experimental flutter results with unsteady pressure distributions

NASA Technical Reports Server (NTRS)

Rivera, Jose A., Jr.; Dansberry, Bryan E.; Bennett, Robert M.; Durham, Michael H.; Silva, Walter A.

1992-01-01

The Structural Dynamics Division at NASA Langley Research Center has started a wind tunnel activity referred to as the Benchmark Models Program. The primary objective of this program is to acquire measured dynamic instability and corresponding pressure data that will be useful for developing and evaluating aeroelastic type computational fluid dynamics codes currently in use or under development. The program is a multi-year activity that will involve testing of several different models to investigate various aeroelastic phenomena. This paper describes results obtained from a second wind tunnel test of the first model in the Benchmark Models Program. This first model consisted of a rigid semispan wing having a rectangular planform and a NACA 0012 airfoil shape which was mounted on a flexible two degree of freedom mount system. Experimental flutter boundaries and corresponding unsteady pressure distribution data acquired over two model chords located at the 60 and 95 percent span stations are presented.

Pressure- and Temperature-Sensitive Paint at 0.3-m Transonic Cryogenic Tunnel

NASA Technical Reports Server (NTRS)

Watkins, A. Neal; Leighty, Bradley D.; Lipford, William E.; Goodman, Kyle Z.

2015-01-01

Recently both Pressure- and Temperature-Sensitive Paint experiments were conducted at cryogenic conditions in the 0.3-m Transonic Cryogenic Tunnel at NASA Langley Research Center. This represented a re-introduction of the techniques to the facility after more than a decade, and provided a means to upgrade the measurements using newer technology as well as demonstrate that the techniques were still viable in the facility. Temperature-Sensitive Paint was employed on a laminar airfoil for transition detection and Pressure-Sensitive Paint was employed on a supercritical airfoil. This report will detail the techniques and their unique challenges that need to be overcome in cryogenic environments. In addition, several optimization strategies will also be discussed.

Flutter and oscillating air-force calculations for an airfoil in two-dimensional supersonic flow

NASA Technical Reports Server (NTRS)

Garrick, I E; Rubinow, S I

1946-01-01

A connected account is given of the Possio theory of non-stationary flow for small disturbances in a two-dimensional supersonic flow and of its application to the determination of the aerodynamic forces on an oscillating airfoil. Further application is made to the problem of wing flutter in the degrees of freedom - torsion, bending, and aileron rotations. Numerical tables for flutter calculations are provided for various values of the Mach number greater than unity. Results for bending-torsion wing flutter are shown in figures and are discussed. The static instabilities of divergence and aileron reversal are examined as is a one-degree-of-freedom case of torsional oscillatory instability.

Subsonic flutter analysis addition to NASTRAN. [for use with CDC 6000 series digital computers

NASA Technical Reports Server (NTRS)

Doggett, R. V., Jr.; Harder, R. L.

1973-01-01

A subsonic flutter analysis capability has been developed for NASTRAN, and a developmental version of the program has been installed on the CDC 6000 series digital computers at the Langley Research Center. The flutter analysis is of the modal type, uses doublet lattice unsteady aerodynamic forces, and solves the flutter equations by using the k-method. Surface and one-dimensional spline functions are used to transform from the aerodynamic degrees of freedom to the structural degrees of freedom. Some preliminary applications of the method to a beamlike wing, a platelike wing, and a platelike wing with a folded tip are compared with existing experimental and analytical results.

Prospective Observational Cohort Study of Fetal Atrial Flutter & Supraventricular Tachycardia

ClinicalTrials.gov

2017-12-15

Atrial Flutter; Tachycardia, Supraventricular; Tachycardia, Atrial Ectopic; Tachycardia, Reciprocating; Tachycardia Atrial; Tachycardia, Atrioventricular Nodal Reentry; Tachycardia, Paroxysmal; Fetal Hydrops

Unsteady Aerodynamic Response of a Linear Cascade of Airfoils in Separated Flow

NASA Technical Reports Server (NTRS)

Capece, Vincent R.; Ford, Christopher; Bone, Christopher; Li, Rui

2004-01-01

The overall objective of this research program was to investigate methods to modify the leading edge separation region, which could lead to an improvement in aeroelastic stability of advanced airfoil designs. The airfoil section used is representative of current low aspect ratio fan blade tip sections. The experimental potion of this study investigated separated zone boundary layer from removal through suction slots. Suction applied to a cavity in the vicinity of the separation onset point was found to be the most effective location. The computational study looked into the influence of front camber on flutter stability. To assess the influence of the change in airfoil shape on stability the work-per-cycle was evaluated for torsion mode oscillations. It was shown that the front camberline shape can be an important factor for stabilizing the predicted work-per-cycle and reducing the predicted extent of the separation zone. In addition, data analysis procedures are discussed for reducing data acquired in experiments that involve periodic unsteady data. This work was conducted in support of experiments being conducted in the NASA Glenn Research Center Transonic Flutter Cascade. The spectral block averaging method is presented. This method is shown to be able to account for variations in airfoil oscillation frequency that can occur in experiments that force oscillate the airfoils to simulate flutter.

Analysis of Limit Cycle Oscillation Data from the Aeroelastic Test of the SUGAR Truss-Braced Wing Model

NASA Technical Reports Server (NTRS)

Bartels, Robert E.; Funk, Christie; Scott, Robert C.

2015-01-01

Research focus in recent years has been given to the design of aircraft that provide significant reductions in emissions, noise and fuel usage. Increases in fuel efficiency have also generally been attended by overall increased wing flexibility. The truss-braced wing (TBW) configuration has been forwarded as one that increases fuel efficiency. The Boeing company recently tested the Subsonic Ultra Green Aircraft Research (SUGAR) Truss-Braced Wing (TBW) wind-tunnel model in the NASA Langley Research Center Transonic Dynamics Tunnel (TDT). This test resulted in a wealth of accelerometer data. Other publications have presented details of the construction of that model, the test itself, and a few of the results of the test. This paper aims to provide a much more detailed look at what the accelerometer data says about the onset of aeroelastic instability, usually known as flutter onset. Every flight vehicle has a location in the flight envelope of flutter onset, and the TBW vehicle is not different. For the TBW model test, the flutter onset generally occurred at the conditions that the Boeing company analysis said it should. What was not known until the test is that, over a large area of the Mach number dynamic pressure map, the model displayed wing/engine nacelle aeroelastic limit cycle oscillation (LCO). This paper dissects that LCO data in order to provide additional insights into the aeroelastic behavior of the model.

Solution of steady and unsteady transonic-vortex flows using Euler and full-potential equations

NASA Technical Reports Server (NTRS)

Kandil, Osama A.; Chuang, Andrew H.; Hu, Hong

1989-01-01

Two methods are presented for inviscid transonic flows: unsteady Euler equations in a rotating frame of reference for transonic-vortex flows and integral solution of full-potential equation with and without embedded Euler domains for transonic airfoil flows. The computational results covered: steady and unsteady conical vortex flows; 3-D steady transonic vortex flow; and transonic airfoil flows. The results are in good agreement with other computational results and experimental data. The rotating frame of reference solution is potentially efficient as compared with the space fixed reference formulation with dynamic gridding. The integral equation solution with embedded Euler domain is computationally efficient and as accurate as the Euler equations.

Flutter Sensitivity to Boundary Layer Thickness, Structural Damping, and Static Pressure Differential for a Shuttle Tile Overlay Repair Concept

NASA Technical Reports Server (NTRS)

Scott, Robert C.; Bartels, Robert E.

2009-01-01

This paper examines the aeroelastic stability of an on-orbit installable Space Shuttle patch panel. CFD flutter solutions were obtained for thick and thin boundary layers at a free stream Mach number of 2.0 and several Mach numbers near sonic speed. The effect of structural damping on these flutter solutions was also examined, and the effect of structural nonlinearities associated with in-plane forces in the panel was considered on the worst case linear flutter solution. The results of the study indicated that adequate flutter margins exist for the panel at the Mach numbers examined. The addition of structural damping improved flutter margins as did the inclusion of nonlinear effects associated with a static pressure difference across the panel.

Transonic Unsteady Aerodynamics and Aeroelasticity 1987, part 1

NASA Technical Reports Server (NTRS)

Bland, Samuel R. (Compiler)

1989-01-01

Computational fluid dynamics methods have been widely accepted for transonic aeroelastic analysis. Previously, calculations with the TSD methods were used for 2-D airfoils, but now the TSD methods are applied to the aeroelastic analysis of the complete aircraft. The Symposium papers are grouped into five subject areas, two of which are covered in this part: (1) Transonic Small Disturbance (TSD) theory for complete aircraft configurations; and (2) Full potential and Euler equation methods.

Transonic CFD applications at Boeing

NASA Technical Reports Server (NTRS)

Tinoco, E. N.

1989-01-01

The use of computational methods for three dimensional transonic flow design and analysis at the Boeing Company is presented. A range of computational tools consisting of production tools for every day use by project engineers, expert user tools for special applications by computational researchers, and an emerging tool which may see considerable use in the near future are described. These methods include full potential and Euler solvers, some coupled to three dimensional boundary layer analysis methods, for transonic flow analysis about nacelle, wing-body, wing-body-strut-nacelle, and complete aircraft configurations. As the examples presented show, such a toolbox of codes is necessary for the variety of applications typical of an industrial environment. Such a toolbox of codes makes possible aerodynamic advances not previously achievable in a timely manner, if at all.

Vibration and flutter of mistuned bladed-disk assemblies

NASA Technical Reports Server (NTRS)

Kaza, K. R. V.; Kielb, R. E.

1984-01-01

An analytical model for investigating vibration and flutter of mistuned bladed disk assemblies is presented. This model accounts for elastic, inertial and aerodynamic coupling between bending and torsional motions of each individual blade, elastic and inertial couplings between the blades and the disk, and aerodynamic coupling among the blades. The disk was modeled as a circular plate with constant thickness and each blade was represented by a twisted, slender, straight, nonuniform, elastic beam with a symmetric cross section. The elastic axis, inertia axis, and the tension axis were taken to be noncoincident and the structural warping of the section was explicitly considered. The blade aerodynamic loading in the subsonic and supersonic flow regimes was obtained from two-dimensional unsteady, cascade theories. All the possible standing wave modes of the disk and traveling wave modes of the blades were included. The equations of motion were derived by using the energy method in conjunction with the assumed mode shapes for the disk and the blades. Continuities of displacement and slope at the blade-disk junction were maintained. The equations were solved to investigate the effects of blade-disk coupling and blade frequency mistuning on vibration and flutter. Results showed that the flexibility of practical disks such as those used for current generation turbofans did not have a significant influence on either the tuned or mistuned flutter characteristics. However, the disk flexibility may have a strong influence on some of the system frequencies and on forced response.

Vibration and flutter of mistuned bladed-disk assemblies

NASA Technical Reports Server (NTRS)

Rao, K.; Kaza, V.; Kielb, R. E.

1984-01-01

An analytical model for investigating vibration and flutter of mistuned bladed disk assemblies is presented. This model accounts for elastic, inertial and aerodynamic coupling between bending and torsional motions of each individual blade, elastic and inertial couplings between the blades and the disk, and aerodynamic coupling among the blades. The disk was modeled as a circular plate with constant thickness and each blade was represented by a twisted, slender, straight, nonuniform, elastic beam with a symmetric cross section. The elastic axis, inertia axis, and the tension axis were taken to be noncoincident and the structural warping of the section was explicitly considered. The blade aerodynamic loading in the subsonic and supersonic flow regimes was obtained from two-dimensional unsteady, cascade theories. All the possible standing wave modes of the disk and traveling wave modes of the blades were included. The equations of motion were derived by using the energy method in conjunction with the assumed mode shapes for the disk and the blades. Continuities of displacement and slope at the blade-disk junction were maintained. The equations were solved to investigate the effects of blade-disk coupling and blade frequency mistuning on vibration and flutter. Results showed that the flexibility of practical disks such as those used for current generation turbufans did not have a significant influence on either the tuned or mistuned flutter characteristics. However, the disk flexibility may have a strong influence on some of the system frequencies and on forced response.

Calculative techniques for transonic flows about certain classes of wing body combinations

NASA Technical Reports Server (NTRS)

Stahara, S. S.; Spreiter, J. R.

1972-01-01

Procedures based on the method of local linearization and transonic equivalence rule were developed for predicting properties of transonic flows about certain classes of wing-body combinations. The procedures are applicable to transonic flows with free stream Mach number in the ranges near one, below the lower critical and above the upper critical. Theoretical results are presented for surface and flow field pressure distributions for both lifting and nonlifting situations.

Semi-empirical model for prediction of unsteady forces on an airfoil with application to flutter

NASA Technical Reports Server (NTRS)

Mahajan, Aparajit J.; Kaza, Krishna Rao V.

1992-01-01

A semi-empirical model is described for predicting unsteady aerodynamic forces on arbitrary airfoils under mildly stalled and unstalled conditions. Aerodynamic forces are modeled using second order ordinary differential equations for lift and moment with airfoil motion as the input. This model is simultaneously integrated with structural dynamics equations to determine flutter characteristics for a two degrees-of-freedom system. Results for a number of cases are presented to demonstrate the suitability of this model to predict flutter. Comparison is made to the flutter characteristics determined by a Navier-Stokes solver and also the classical incompressible potential flow theory.

Semi-empirical model for prediction of unsteady forces on an airfoil with application to flutter

NASA Technical Reports Server (NTRS)

Mahajan, A. J.; Kaza, K. R. V.; Dowell, E. H.

1993-01-01

A semi-empirical model is described for predicting unsteady aerodynamic forces on arbitrary airfoils under mildly stalled and unstalled conditions. Aerodynamic forces are modeled using second order ordinary differential equations for lift and moment with airfoil motion as the input. This model is simultaneously integrated with structural dynamics equations to determine flutter characteristics for a two degrees-of-freedom system. Results for a number of cases are presented to demonstrate the suitability of this model to predict flutter. Comparison is made to the flutter characteristics determined by a Navier-Stokes solver and also the classical incompressible potential flow theory.

Comparison of analytical and wind-tunnel results for flutter and gust response of a transport wing with active controls

NASA Technical Reports Server (NTRS)

Abel, I.; Perry, B., III; Newsom, J. R.

1982-01-01

Two flutter suppression control laws wre designed and tested on a low speed aeroelastic model of a DC-10 derivative wing. Both control laws demontrated increases in flutter speed in excess of 25 percent above the passive wing flutter speed. In addition, one of the control laws was effective in reducing loads due to turbulence generated in the wind tunnel. The effect of variations in gain and phase on the closed-loop performance was measured and is compared with predictions. In general, both flutter and gust response predictions agree reasonably well with experimental data.

Modifications to the Langley 8-foot transonic pressure tunnel for the laminar flow control experiment

NASA Technical Reports Server (NTRS)

Harris, Charles D.; Brooks, Cuyler W., Jr.

1988-01-01

Modifications to the NASA Langley 8 Foot Transonic Pressure Tunnel in support of the Lamina Flow Control (LFC) Experiment included the installation of a honeymoon and five screens in the settling chamber upstream of the test section 41-long test section liner that extended from the upstream end of the test section contraction region, through the best section, and into the diffuser. The honeycomb and screens were installed as permanent additions to the facility, and the liner was a temporary addition to be removed at the conclusion of the LFC Experiment. These modifications are briefly described.

Axisymmetric & non-axisymmetric exhaust jet induced-effects on a V/STOL vehicle design. Part 1: Data presentation. [conducted in Ames 11-foot transonic tunnel

NASA Technical Reports Server (NTRS)

Schnell, W. C.; Ordonez, G. W.

1981-01-01

A 1/8 scale jet-effects model was tested in the NASA Ames 11 ft transonic tunnel at static conditions and over a range of Mach numbers from 0.4 to 1.4. The data presented show that significant differences in aeropropulsion performance can be expected by varying the exhaust nozzle type and its geometric parameters on a V/STOL underwing nacelle installation.

Shape sensitivity analysis of flutter response of a laminated wing

NASA Technical Reports Server (NTRS)

Bergen, Fred D.; Kapania, Rakesh K.

1988-01-01

A method is presented for calculating the shape sensitivity of a wing aeroelastic response with respect to changes in geometric shape. Yates' modified strip method is used in conjunction with Giles' equivalent plate analysis to predict the flutter speed, frequency, and reduced frequency of the wing. Three methods are used to calculate the sensitivity of the eigenvalue. The first method is purely a finite difference calculation of the eigenvalue derivative directly from the solution of the flutter problem corresponding to the two different values of the shape parameters. The second method uses an analytic expression for the eigenvalue sensitivities of a general complex matrix, where the derivatives of the aerodynamic, mass, and stiffness matrices are computed using a finite difference approximation. The third method also uses an analytic expression for the eigenvalue sensitivities, but the aerodynamic matrix is computed analytically. All three methods are found to be in good agreement with each other. The sensitivities of the eigenvalues were used to predict the flutter speed, frequency, and reduced frequency. These approximations were found to be in good agreement with those obtained using a complete reanalysis.

Flutter and divergence instability of supported piezoelectric nanotubes conveying fluid

NASA Astrophysics Data System (ADS)

Bahaadini, Reza; Hosseini, Mohammad; Jamali, Behnam

2018-01-01

In this paper, divergence and flutter instabilities of supported piezoelectric nanotubes containing flowing fluid are investigated. To take the size effects into account, the nonlocal elasticity theory is implemented in conjunction with the Euler-Bernoulli beam theory incorporating surface stress effects. The Knudsen number is applied to investigate the slip boundary conditions between the flow and wall of nanotube. The nonlocal governing equations of nanotube are obtained using Newtonian method, including the influence of piezoelectric voltage, surface effects, Knudsen number and nonlocal parameter. Applying Galerkin approach to transform resulting equations into a set of eigenvalue equations under the simple-simple (S-S) and clamped-clamped (C-C) boundary conditions. The effects of the piezoelectric voltage, surface effects, Knudsen number, nonlocal parameter and boundary conditions on the divergence and flutter boundaries of nanotubes are discussed. It is observed that the fluid-conveying nanotubes with both ends supported lose their stability by divergence first and then by flutter with increase in fluid velocity. Results indicate the importance of using piezoelectric voltage, nonlocal parameter and Knudsen number in decrease of critical flow velocities of system. Moreover, the surface effects have a significant role on the eigenfrequencies and critical fluid velocity.

Flutter of High-Speed Civil Transport Flexible Semispan Model: Time-Frequency Analysis

NASA Technical Reports Server (NTRS)

Chabalko, Christopher C.; Hajj, Muhammad R.; Silva, Walter A.

2006-01-01

Time/frequency analysis of fluctuations measured by pressure taps and strain gauges in the experimental studies of the flexible semispan model of a high-speed civil transport wing configuration is performed. The interest is in determining the coupling between the aerodynamic loads and structural motions that led to the hard flutter conditions and loss of the model. The results show that, away from the hard flutter point, the aerodynamic loads at all pressure taps near the wing tip and the structural motions contained the same frequency components. On the other hand, in the flow conditions leading to the hard flutter, the frequency content of the pressure fluctuations near the leading and trailing edges varied significantly. This led to contribution to the structural motions over two frequency ranges. The ratio of these ranges was near 2:1, which suggests the possibility of nonlinear structural coupling.

Application of FLEET Velocimetry in the NASA Langley 0.3-meter Transonic Cryogenic Tunnel

NASA Technical Reports Server (NTRS)

Burns, Ross A.; Danehy, Paul M.; Halls, Benjamin R.; Jiang, Naibo

2015-01-01

Femtosecond laser electronic excitation and tagging (FLEET) velocimetry is demonstrated in a large-scale transonic cryogenic wind tunnel. Test conditions include total pressures, total temperatures, and Mach numbers ranging from 15 to 58 psia, 200 to 295 K, and 0.2 to 0.75, respectively. Freestream velocity measurements exhibit accuracies within 1 percent and precisions better than 1 m/s. The measured velocities adhere closely to isentropic flow theory over the domain of temperatures and pressures that were tested. Additional velocity measurements are made within the tunnel boundary layer; virtual trajectories traced out by the FLEET signal are indicative of the characteristic turbulent behavior in this region of the flow, where the unsteadiness increases demonstrably as the wall is approached. Mean velocities taken within the boundary layer are in agreement with theoretical velocity profiles, though the fluctuating velocities exhibit a greater deviation from theoretical predictions.

Role of computational fluid dynamics in unsteady aerodynamics for aeroelasticity

NASA Technical Reports Server (NTRS)

Guruswamy, Guru P.; Goorjian, Peter M.

1989-01-01

In the last two decades there have been extensive developments in computational unsteady transonic aerodynamics. Such developments are essential since the transonic regime plays an important role in the design of modern aircraft. Therefore, there has been a large effort to develop computational tools with which to accurately perform flutter analysis at transonic speeds. In the area of Computational Fluid Dynamics (CFD), unsteady transonic aerodynamics are characterized by the feature of modeling the motion of shock waves over aerodynamic bodies, such as wings. This modeling requires the solution of nonlinear partial differential equations. Most advanced codes such as XTRAN3S use the transonic small perturbation equation. Currently, XTRAN3S is being used for generic research in unsteady aerodynamics and aeroelasticity of almost full aircraft configurations. Use of Euler/Navier Stokes equations for simple typical sections has just begun. A brief history of the development of CFD for aeroelastic applications is summarized. The development of unsteady transonic aerodynamics and aeroelasticity are also summarized.

Aerodynamic design for improved manueverability by use of three-dimensional transonic theory

NASA Technical Reports Server (NTRS)

Mann, M. J.; Campbell, R. L.; Ferris, J. C.

1984-01-01

Improvements in transonic maneuver performance by the use of three-dimensional transonic theory and a transonic design procedure were examined. The FLO-27 code of Jameson and Caughey was used to design a new wing for a fighter configuration with lower drag at transonic maneuver conditions. The wing airfoil sections were altered to reduce the upper-surface shock strength by means of a design procedure which is based on the iterative application of the FLO-27 code. The plan form of the fighter configuration was fixed and had a leading edge sweep of 45 deg and an aspect ratio of 3.28. Wind-tunnel tests were conducted on this configuration at Mach numbers from 0.60 to 0.95 and angles of attack from -2 deg to 17 deg. The transonic maneuver performance of this configuration was evaluated by comparison with a wing designed by empirical methods and a wing designed primarily by two-dimensional transonic theory. The configuration designed by the use of FLO-27 had the same or lower drag than the empirical wing and, for some conditions, lower drag than the two-dimensional design. From some maneuver conditions, the drag of the two-dimensional design was somewhat lower.

An initial investigation into methods of computing transonic aerodynamic sensitivity coefficients

NASA Technical Reports Server (NTRS)

Carlson, Leland A.

1994-01-01

The primary accomplishments of the project are as follows: (1) Using the transonic small perturbation equation as a flowfield model, the project demonstrated that the quasi-analytical method could be used to obtain aerodynamic sensitivity coefficients for airfoils at subsonic, transonic, and supersonic conditions for design variables such as Mach number, airfoil thickness, maximum camber, angle of attack, and location of maximum camber. It was established that the quasi-analytical approach was an accurate method for obtaining aerodynamic sensitivity derivatives for airfoils at transonic conditions and usually more efficient than the finite difference approach. (2) The usage of symbolic manipulation software to determine the appropriate expressions and computer coding associated with the quasi-analytical method for sensitivity derivatives was investigated. Using the three dimensional fully conservative full potential flowfield model, it was determined that symbolic manipulation along with a chain rule approach was extremely useful in developing a combined flowfield and quasi-analytical sensitivity derivative code capable of considering a large number of realistic design variables. (3) Using the three dimensional fully conservative full potential flowfield model, the quasi-analytical method was applied to swept wings (i.e. three dimensional) at transonic flow conditions. (4) The incremental iterative technique has been applied to the three dimensional transonic nonlinear small perturbation flowfield formulation, an equivalent plate deflection model, and the associated aerodynamic and structural discipline sensitivity equations; and coupled aeroelastic results for an aspect ratio three wing in transonic flow have been obtained.

The NASA Langley Research Center 0.3-meter transonic cryogenic tunnel T-P/Re-M controller manual

NASA Technical Reports Server (NTRS)

Balakrishna, S.; Kilgore, W. Allen

1989-01-01

A new microcomputer based controller for the 0.3-m Transonic Cryogenic Tunnel (TCT) has been commissioned in 1988 and has reliably operated for more than a year. The tunnel stagnation pressure, gas stagnation temperature, tunnel wall structural temperature and flow Mach number are precisely controlled by the new controller in a stable manner. The tunnel control hardware, software, and the flow chart to assist in calibration of the sensors, actuators, and the controller real time functions are described. The software installation details are also presented. The report serves as the maintenance and trouble shooting manual for the 0.3-m TCT controller.

An iterative transformation procedure for numerical solution of flutter and similar characteristics-value problems

NASA Technical Reports Server (NTRS)

Gossard, Myron L

1952-01-01

An iterative transformation procedure suggested by H. Wielandt for numerical solution of flutter and similar characteristic-value problems is presented. Application of this procedure to ordinary natural-vibration problems and to flutter problems is shown by numerical examples. Comparisons of computed results with experimental values and with results obtained by other methods of analysis are made.

Comparison of driven and simulated "free" stall flutter in a wind tunnel

NASA Astrophysics Data System (ADS)

Culler, Ethan; Farnsworth, John; Fagley, Casey; Seidel, Jurgen

2016-11-01

Stall flutter and dynamic stall have received a significant amount of attention over the years. To experimentally study this problem, the body undergoing stall flutter is typically driven at a characteristic, single frequency sinusoid with a prescribed pitching amplitude and mean angle of attack offset. This approach allows for testing with repeatable kinematics, however it effectively decouples the structural motion from the aerodynamic forcing. Recent results suggest that this driven approach could misrepresent the forcing observed in a "free" stall flutter scenario. Specifically, a dynamically pitched rigid NACA 0018 wing section was tested in the wind tunnel under two modes of operation: (1) Cyber-Physical where "free" stall flutter was physically simulated through a custom motor-control system modeling a torsional spring and (2) Direct Motor-Driven Dynamic Pitch at a single frequency sinusoid representative of the cyber-physical motion. The time-resolved pitch angle and moment were directly measured and compared for each case. It was found that small deviations in the pitch angle trajectory between these two operational cases generate significantly different aerodynamic pitching moments on the wing section, with the pitching moments nearly 180o out of phase in some cases. This work is supported by the Air Force Office of Scientific Research through the Flow Interactions and Control Program and by the National Defense Science and Engineering Graduate Fellowship Program.

Application of a multi-level grid method to transonic flow calculations

NASA Technical Reports Server (NTRS)

South, J. C., Jr.; Brandt, A.

1976-01-01

A multi-level grid method was studied as a possible means of accelerating convergence in relaxation calculations for transonic flows. The method employs a hierarchy of grids, ranging from very coarse to fine. The coarser grids are used to diminish the magnitude of the smooth part of the residuals. The method was applied to the solution of the transonic small disturbance equation for the velocity potential in conservation form. Nonlifting transonic flow past a parabolic arc airfoil is studied with meshes of both constant and variable step size.

Simplified and refined structural modeling for economical flutter analysis and design

NASA Technical Reports Server (NTRS)

Ricketts, R. H.; Sobieszczanski, J.

1977-01-01

A coordinated use of two finite-element models of different levels of refinement is presented to reduce the computer cost of the repetitive flutter analysis commonly encountered in structural resizing to meet flutter requirements. One model, termed a refined model (RM), represents a high degree of detail needed for strength-sizing and flutter analysis of an airframe. The other model, called a simplified model (SM), has a relatively much smaller number of elements and degrees-of-freedom. A systematic method of deriving an SM from a given RM is described. The method consists of judgmental and numerical operations to make the stiffness and mass of the SM elements equivalent to the corresponding substructures of RM. The structural data are automatically transferred between the two models. The bulk of analysis is performed on the SM with periodical verifications carried out by analysis of the RM. In a numerical example of a supersonic cruise aircraft with an arrow wing, this approach permitted substantial savings in computer costs and acceleration of the job turn-around.

Panel flutter optimization by gradient projection

NASA Technical Reports Server (NTRS)

Pierson, B. L.

1975-01-01

A gradient projection optimal control algorithm incorporating conjugate gradient directions of search is described and applied to several minimum weight panel design problems subject to a flutter speed constraint. New numerical solutions are obtained for both simply-supported and clamped homogeneous panels of infinite span for various levels of inplane loading and minimum thickness. The minimum thickness inequality constraint is enforced by a simple transformation of variables.

Generalization of the subsonic kernel function in the s-plane, with applications to flutter analysis

NASA Technical Reports Server (NTRS)

Cunningham, H. J.; Desmarais, R. N.

1984-01-01

A generalized subsonic unsteady aerodynamic kernel function, valid for both growing and decaying oscillatory motions, is developed and applied in a modified flutter analysis computer program to solve the boundaries of constant damping ratio as well as the flutter boundary. Rates of change of damping ratios with respect to dynamic pressure near flutter are substantially lower from the generalized-kernel-function calculations than from the conventional velocity-damping (V-g) calculation. A rational function approximation for aerodynamic forces used in control theory for s-plane analysis gave rather good agreement with kernel-function results, except for strongly damped motion at combinations of high (subsonic) Mach number and reduced frequency.

Forward-swept wing configuration designed for high maneuverability by use of a transonic computational method

NASA Technical Reports Server (NTRS)

Mann, M. J.; Mercer, C. E.

1986-01-01

A transonic computational analysis method and a transonic design procedure have been used to design the wing and the canard of a forward-swept-wing fighter configuration for good transonic maneuver performance. A model of this configuration was tested in the Langley 16-Foot Transonic Tunnel. Oil-flow photographs were obtained to examine the wind flow patterns at Mach numbers from 0.60 to 0.90. The transonic theory gave a reasonably good estimate of the wing pressure distributions at transonic maneuver conditions. Comparison of the forward-swept-wing configuration with an equivalent aft-swept-wing-configuration showed that, at a Mach number of 0.90 and a lift coefficient of 0.9, the two configurations have the same trimmed drag. The forward-swept wing configuration was also found to have trimmed drag levels at transonic maneuver conditions which are comparable to those of the HiMAT (highly maneuverable aircraft technology) configuration and the X-29 forward-swept-wing research configuration. The configuration of this study was also tested with a forebody strake.

Prediction and control of coupled-mode flutter in future wind turbine blades

NASA Astrophysics Data System (ADS)

Modarres-Sadeghi, Yahya; Currier, Todd; Caracoglia, Luca; Lackner, Matthew; Hollot, Christopher

2017-11-01

Coupled-mode flutter can be observed in future offshore wind turbine blades. We have shown this fact by considering various candidate blade designs, in all of which the blade's first torsional mode couples with one of its flapwise modes, resulting in coupled-mode flutter. We have shown how the ratio of these two natural frequencies can result in blades with a critical flutter speed even lower than their rated speed, especially for blades with low torsional natural frequencies. We have also shown how the stochastic nature of the system parameters (as an example, due to uncertainties in the manufacturing process) can significantly influence the onset of instability. We have proposed techniques to predict the onset of these instabilities and the resulting limit-cycle response, and strategies to control them, by either postponing the onset of instability, or lowering the magnitude of the limit-cycle response. The work is supported by the National Science Foundation, Award CBET-1437988 and Collaborative Awards CMMI-1462646 and CMMI-1462774.

On curve veering and flutter of rotating blades

NASA Technical Reports Server (NTRS)

Afolabi, Dare; Mehmed, Oral

1993-01-01

The eigenvalues of rotating blades usually change with rotation speed according to the Stodola-Southwell criterion. Under certain circumstances, the loci of eigenvalues belonging to two distinct modes of vibration approach each other very closely, and it may appear as if the loci cross each other. However, our study indicates that the observable frequency loci of an undamped rotating blade do not cross, but must either repel each other (leading to 'curve veering'), or attract each other (leading to 'frequency coalescence'). Our results are reached by using standard arguments from algebraic geometry--the theory of algebraic curves and catastrophe theory. We conclude that it is important to resolve an apparent crossing of eigenvalue loci into either a frequency coalescence or a curve veering, because frequency coalescence is dangerous since it leads to flutter, whereas curve veering does not precipitate flutter and is, therefore, harmless with respect to elastic stability.

An initial investigation into methods of computing transonic aerodynamic sensitivity coefficients

NASA Technical Reports Server (NTRS)

Carlson, Leland A.

1988-01-01

The initial effort was concentrated on developing the quasi-analytical approach for two-dimensional transonic flow. To keep the problem computationally efficient and straightforward, only the two-dimensional flow was considered and the problem was modeled using the transonic small perturbation equation.

Probabilistic Design of a Plate-Like Wing to Meet Flutter and Strength Requirements

NASA Technical Reports Server (NTRS)

Stroud, W. Jefferson; Krishnamurthy, T.; Mason, Brian H.; Smith, Steven A.; Naser, Ahmad S.

2002-01-01

An approach is presented for carrying out reliability-based design of a metallic, plate-like wing to meet strength and flutter requirements that are given in terms of risk/reliability. The design problem is to determine the thickness distribution such that wing weight is a minimum and the probability of failure is less than a specified value. Failure is assumed to occur if either the flutter speed is less than a specified allowable or the stress caused by a pressure loading is greater than a specified allowable. Four uncertain quantities are considered: wing thickness, calculated flutter speed, allowable stress, and magnitude of a uniform pressure load. The reliability-based design optimization approach described herein starts with a design obtained using conventional deterministic design optimization with margins on the allowables. Reliability is calculated using Monte Carlo simulation with response surfaces that provide values of stresses and flutter speed. During the reliability-based design optimization, the response surfaces and move limits are coordinated to ensure accuracy of the response surfaces. Studies carried out in the paper show the relationship between reliability and weight and indicate that, for the design problem considered, increases in reliability can be obtained with modest increases in weight.

Vibration and flutter characteristics of the SR7L large-scale propfan

NASA Technical Reports Server (NTRS)

August, Richard; Kaza, Krishna Rao V.