Aeroelastic stability of periodic systems with application to rotor blade flutter
NASA Technical Reports Server (NTRS)
Friedmann, P.; Silverthorn, L. J.
1974-01-01
The dynamics of a helicopter blade in forward flight are described by a system of linear differential equations with periodic coefficients. The stability of this periodic aeroelastic system is determined, using multivariable Floquet-Liapunov theory. The transition matrix at the end of the period is evaluated by: (1) direct numerical integration, and (2) a new, approximate method, which consists in approximating a periodic function by a series of step functions. The numerical accuracy and efficiency of the methods is compared, and the second method is shown to be superior by far. Results illustrating the effect of the periodic coefficients and various blade parameters are presented.
Aeroelastic flutter produces hummingbird feather songs.
Clark, Christopher J; Elias, Damian O; Prum, Richard O
2011-09-09
During courtship flights, males of some hummingbird species produce diverse sounds with tail feathers of varying shapes. We show that these sounds are produced by air flowing past a feather, causing it to aeroelastically flutter and generate flutter-induced sound. Scanning laser doppler vibrometery and high-speed video of individual feathers of different sizes and shapes in a wind tunnel revealed multiple vibratory modes that produce a range of acoustic frequencies and harmonic structures. Neighboring feathers can be aerodynamically coupled and flutter either at the same frequency, resulting in sympathetic vibrations that increase loudness, or at different frequencies, resulting in audible interaction frequencies. Aeroelastic flutter is intrinsic to stiff airfoils such as feathers and thus explains tonal sounds that are common in bird flight.
Flutter and Divergence Analysis using the Generalized Aeroelastic Analysis Method
NASA Technical Reports Server (NTRS)
Edwards, John W.; Wieseman, Carol D.
2003-01-01
The Generalized Aeroelastic Analysis Method (GAAM) is applied to the analysis of three well-studied checkcases: restrained and unrestrained airfoil models, and a wing model. An eigenvalue iteration procedure is used for converging upon roots of the complex stability matrix. For the airfoil models, exact root loci are given which clearly illustrate the nature of the flutter and divergence instabilities. The singularities involved are enumerated, including an additional pole at the origin for the unrestrained airfoil case and the emergence of an additional pole on the positive real axis at the divergence speed for the restrained airfoil case. Inconsistencies and differences among published aeroelastic root loci and the new, exact results are discussed and resolved. The generalization of a Doublet Lattice Method computer code is described and the code is applied to the calculation of root loci for the wing model for incompressible and for subsonic flow conditions. The error introduced in the reduction of the singular integral equation underlying the unsteady lifting surface theory to a linear algebraic equation is discussed. Acknowledging this inherent error, the solutions of the algebraic equation by GAAM are termed 'exact.' The singularities of the problem are discussed and exponential series approximations used in the evaluation of the kernel function shown to introduce a dense collection of poles and zeroes on the negative real axis. Again, inconsistencies and differences among published aeroelastic root loci and the new 'exact' results are discussed and resolved. In all cases, aeroelastic flutter and divergence speeds and frequencies are in good agreement with published results. The GAAM solution procedure allows complete control over Mach number, velocity, density, and complex frequency. Thus all points on the computed root loci can be matched-point, consistent solutions without recourse to complex mode tracking logic or dataset interpolation, as in the k and p
Aeroelastic flutter in axial flow-The continuum theory
NASA Astrophysics Data System (ADS)
Balakrishnan, A. V.; Tuffaha, A. M.
2012-11-01
We present a mathematical continuum model for aeroelastic flutter of a Goland type structure subject to axial airflow. The model consists of a linearized Euler full potential equation for the airflow and a second order linear structure equation in two degrees of freedom plunge and pitch (bending and torsion). These are coupled through velocity matching type conditions and Kutta type condition describing the pressure jump. The approach mimics the approach used to study aeroelastic flutter in the normal flow case [?], which deals with aircraft applications. We layout the theoretical framework for determining the aeroelastic modes and the flutter point of the structure at any given mode. We will focus on the torsion aeroelastic modes and consider bending modes in future work. The importance of studying aeroelastic flutter in the axial flow case has come to attention in the recent years in light of non aircraft applications of which we mention two: the problem of snoring or apnea, which can be characterized as palattal flutter and secondly power generation from structures placed in axial flow.
Harmonic Balance Computations of Fan Aeroelastic Stability
NASA Technical Reports Server (NTRS)
Bakhle, Milind A.; Reddy, T. S. R.
2010-01-01
A harmonic balance (HB) aeroelastic analysis, which has been recently developed, was used to determine the aeroelastic stability (flutter) characteristics of an experimental fan. To assess the numerical accuracy of this HB aeroelastic analysis, a time-domain aeroelastic analysis was also used to determine the aeroelastic stability characteristics of the same fan. Both of these three-dimensional analysis codes model the unsteady flowfield due to blade vibrations using the Reynolds-averaged Navier-Stokes (RANS) equations. In the HB analysis, the unsteady flow equations are converted to a HB form and solved using a pseudo-time marching method. In the time-domain analysis, the unsteady flow equations are solved using an implicit time-marching approach. Steady and unsteady computations for two vibration modes were carried out at two rotational speeds: 100 percent (design) and 70 percent (part-speed). The steady and unsteady results obtained from the two analysis methods compare well, thus verifying the recently developed HB aeroelastic analysis. Based on the results, the experimental fan was found to have no aeroelastic instability (flutter) at the conditions examined in this study.
Bayesian analysis of the flutter margin method in aeroelasticity
NASA Astrophysics Data System (ADS)
Khalil, Mohammad; Poirel, Dominique; Sarkar, Abhijit
2016-12-01
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 flutter 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. 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.
Bayesian analysis of the flutter margin method in aeroelasticity
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 flutter 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.
Aeroelastic stability analysis of a Darrieus wind turbine
Popelka, D.
1982-02-01
An aeroelastic stability analysis has been developed for predicting flutter instabilities on vertical axis wind turbines. The analytical model and mathematical formulation of the problem are described as well as the physical mechanism that creates flutter in Darrieus turbines. Theoretical results are compared with measured experimental data from flutter tests of the Sandia 2 Meter turbine. Based on this comparison, the analysis appears to be an adequate design evaluation tool.
Bayesian analysis of the flutter margin method in aeroelasticity
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
New Flutter Analysis Technique for Time-Domain Computational Aeroelasticity
NASA Technical Reports Server (NTRS)
Pak, Chan-Gi; Lung, Shun-Fat
2017-01-01
A new time-domain approach for computing flutter speed is presented. Based on the time-history result of aeroelastic simulation, the unknown unsteady aerodynamics model is estimated using a system identification technique. The full aeroelastic model is generated via coupling the estimated unsteady aerodynamic model with the known linear structure model. The critical dynamic pressure is computed and used in the subsequent simulation until the convergence of the critical dynamic pressure is achieved. The proposed method is applied to a benchmark cantilevered rectangular wing.
Rotorcraft aeroelastic stability
NASA Technical Reports Server (NTRS)
Ormiston, Robert A.; Warmbrodt, William G.; Hodges, Dewey H.; Peters, David A.
1988-01-01
Theoretical and experimental developments in the aeroelastic and aeromechanical stability of helicopters and tilt-rotor aircraft are addressed. Included are the underlying nonlinear structural mechanics of slender rotating beams, necessary for accurate modeling of elastic cantilever rotor blades, and the development of dynamic inflow, an unsteady aerodynamic theory for low-frequency aeroelastic stability applications. Analytical treatment of isolated rotor stability in hover and forward flight, coupled rotor-fuselage stability in hover and forward flight, and analysis of tilt-rotor dynamic stability are considered. Results of parametric investigations of system behavior are presented, and correlation between theoretical results and experimental data from small and large scale wind tunnel and flight testing are discussed.
NASA Astrophysics Data System (ADS)
Tosi, Luis Phillipe; Colonius, Tim; Lee, Hyeong Jae; Sherrit, Stewart; Jet Propulsion Laboratory Collaboration; California Institute of Technology Collaboration
2016-11-01
Aeroelastic flutter arises when the motion of a structure and its surrounding flowing fluid are coupled in a constructive manner, causing large amplitudes of vibration in the immersed solid. A cantilevered beam in axial flow within a nozzle-diffuser geometry exhibits interesting resonance behavior that presents good prospects for internal flow energy harvesting. Different modes can be excited as a function of throat velocity, nozzle geometry, fluid and cantilever material parameters. Similar behavior has been also observed in elastically mounted rigid plates, enabling new designs for such devices. This work explores the relationship between the aeroelastic flutter instability boundaries and relevant non-dimensional parameters via experiments, numerical, and stability analyses. Parameters explored consist of a non-dimensional stiffness, a non-dimensional mass, non-dimensional throat size, and Reynolds number. A map of the system response in this parameter space may serve as a guide to future work concerning possible electrical output and failure prediction in harvesting devices.
Triboelectret-based aeroelastic flutter energy harvesters
NASA Astrophysics Data System (ADS)
Perez, Matthias; Boisseau, Sebastien; Geisler, Matthias; Despesse, Ghislain; Reboud, Jean Luc
2016-11-01
This paper highlights some experimental results on several electrostatic membranes tested in a wind tunnel between 0 and 20m.s-1 for airflow energy harvesting. The main idea is to use the aeroelastic behavior of thin flexible films to induce simultaneously the capacitance variations and the polarization required by the triboelectric/electrostatic conversion. This technology provides thin and flexible devices and avoids the issue of electrets discharge. Our prototypes (<16cm2) allowed a quick startup (from 3ms-1), an electrical power-flux density from 0.1μW.cm-2 to 60μW.cm-2. In order to complete the energy harvesting chain, we have used a wireless sensor with temperature and acceleration measures coupled to a low power transmission (Bluetooth Low Energy) with reception on a smartphone.
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.
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.
Aeroelastic stability and response of rotating structures
NASA Technical Reports Server (NTRS)
Keith, Theo G., Jr.
1993-01-01
A summary of the work performed during the progress period is presented. Analysis methods for predicting loads and instabilities of wind turbines were developed. Three new areas of research to aid the Advanced Turboprop Project (ATP) were initiated and developed. These three areas of research are aeroelastic analysis methods for cascades including blade and disk flexibility; stall flutter analysis; and computational aeroelasticity.
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.
New Flutter Analysis Technique for CFD-based Unsteady Aeroelasticity
NASA Technical Reports Server (NTRS)
Pak, Chan-gi; Jutte, Christine V.
2009-01-01
This paper presents a flutter analysis technique for the transonic flight regime. The technique uses an iterative approach to determine the critical dynamic pressure for a given mach number. Unlike other CFD-based flutter analysis methods, each iteration solves for the critical dynamic pressure and uses this value in subsequent iterations until the value converges. This process reduces the iterations required to determine the critical dynamic pressure. To improve the accuracy of the analysis, the technique employs a known structural model, leaving only the aerodynamic model as the unknown. The aerodynamic model is estimated using unsteady aeroelastic CFD analysis combined with a parameter estimation routine. The technique executes as follows. The known structural model is represented as a finite element model. Modal analysis determines the frequencies and mode shapes for the structural model. At a given mach number and dynamic pressure, the unsteady CFD analysis is performed. The output time history of the surface pressure is converted to a nodal aerodynamic force vector. The forces are then normalized by the given dynamic pressure. A multi-input multi-output parameter estimation software, ERA, estimates the aerodynamic model through the use of time histories of nodal aerodynamic forces and structural deformations. The critical dynamic pressure is then calculated using the known structural model and the estimated aerodynamic model. This output is used as the dynamic pressure in subsequent iterations until the critical dynamic pressure is determined. This technique is demonstrated on the Aerostructures Test Wing-2 model at NASA's Dryden Flight Research Center.
Hummingbird feather sounds are produced by aeroelastic flutter, not vortex-induced vibration.
Clark, Christopher J; Elias, Damian O; Prum, Richard O
2013-09-15
Males in the 'bee' hummingbird clade produce distinctive, species-specific sounds with fluttering tail feathers during courtship displays. Flutter may be the result of vortex shedding or aeroelastic interactions. We investigated the underlying mechanics of flutter and sound production of a series of different feathers in a wind tunnel. All feathers tested were capable of fluttering at frequencies varying from 0.3 to 10 kHz. At low airspeeds (Uair) feather flutter was highly damped, but at a threshold airspeed (U*) the feathers abruptly entered a limit-cycle vibration and produced sound. Loudness increased with airspeed in most but not all feathers. Reduced frequency of flutter varied by an order of magnitude, and declined with increasing Uair in all feathers. This, along with the presence of strong harmonics, multiple modes of flutter and several other non-linear effects indicates that flutter is not simply a vortex-induced vibration, and that the accompanying sounds are not vortex whistles. Flutter is instead aeroelastic, in which structural (inertial/elastic) properties of the feather interact variably with aerodynamic forces, producing diverse acoustic results.
Aeroelastic Stability Computations for Turbomachinery
NASA Technical Reports Server (NTRS)
Srivastava, R.; Bakhle, M. A.; Keith, T. G., Jr.; Stefko, G. L.
2001-01-01
This paper describes an aeroelastic analysis program for turbomachines. Unsteady Navier-Stokes equations are solved on dynamically deforming, body fitted, grid to obtain the aeroelastic characteristics. Blade structural response is modeled using a modal representation of the blade and the work-per-cycle method is used to evaluate the stability characteristics. Nonzero interblade phase angle is modeled using phase-lagged boundary conditions. Results obtained showed good correlation with existing experimental, analytical, and numerical results. Numerical analysis also showed that given the computational resources available today, engineering solutions with good accuracy are possible using higher fidelity analyses.
Aeroelastic Stability and Response of Rotating Structures
NASA Technical Reports Server (NTRS)
Keith, Theo G., Jr.; Reddy, T. S. R.
1998-01-01
A summary of the work performed from 1996 to 1997 is presented. More details can be found in the cited references. This grant led to the development of aeroelastic analyses methods for predicting flutter and forced response in fans, compressors, and turbines using computational
Aeroelastic Stability & Response of Rotating Structures
NASA Technical Reports Server (NTRS)
Keith, Theo G., Jr.; Reddy, T. S. R.
2001-01-01
A summary of the work performed under NASA grant NCC3-605 is presented. More details can be found in the cited references. This grant led to the development of relatively faster aeroelastic analyses methods for predicting flutter and forced response in fans, compressors, and turbines using computational fluid dynamic (CFD) methods.
Aeroelastic flutter of feathers, flight and the evolution of non-vocal communication in birds.
Clark, Christopher J; Prum, Richard O
2015-11-01
Tonal, non-vocal sounds are widespread in both ordinary bird flight and communication displays. We hypothesized these sounds are attributable to an aerodynamic mechanism intrinsic to flight feathers: aeroelastic flutter. Individual wing and tail feathers from 35 taxa (from 13 families) that produce tonal flight sounds were tested in a wind tunnel. In the wind tunnel, all of these feathers could flutter and generate tonal sound, suggesting that the capacity to flutter is intrinsic to flight feathers. This result implies that the aerodynamic mechanism of aeroelastic flutter is potentially widespread in flight of birds. However, the sounds these feathers produced in the wind tunnel replicated the actual flight sounds of only 15 of the 35 taxa. Of the 20 negative results, we hypothesize that 10 are false negatives, as the acoustic form of the flight sound suggests flutter is a likely acoustic mechanism. For the 10 other taxa, we propose our negative wind tunnel results are correct, and these species do not make sounds via flutter. These sounds appear to constitute one or more mechanism(s) we call 'wing whirring', the physical acoustics of which remain unknown. Our results document that the production of non-vocal communication sounds by aeroelastic flutter of flight feathers is widespread in birds. Across all birds, most evolutionary origins of wing- and tail-generated communication sounds are attributable to three mechanisms: flutter, percussion and wing whirring. Other mechanisms of sound production, such as turbulence-induced whooshes, have evolved into communication sounds only rarely, despite their intrinsic ubiquity in ordinary flight.
Active and passive techniques for tiltrotor aeroelastic stability augmentation
NASA Astrophysics Data System (ADS)
Hathaway, Eric L.
Tiltrotors are susceptible to whirl flutter, an aeroelastic instability characterized by a coupling of rotor-generated aerodynamic forces and elastic wing modes in high speed airplane-mode flight. The conventional approach to ensuring adequate whirl flutter stability will not scale easily to larger tiltrotor designs. This study constitutes an investigation of several alternatives for improving tiltrotor aerolastic stability. A whirl flutter stability analysis is developed that does not rely on more complex models to determine the variations in crucial input parameters with flight condition. Variation of blade flap and lag frequency, and pitch-flap, pitch-lag, and flap-lag couplings, are calculated from physical parameters, such as blade structural flap and lag stiffness distribution (inboard or outboard of pitch bearing), collective pitch, and precone. The analysis is used to perform a study of the influence of various design parameters on whirl flutter stability. While previous studies have investigated the individual influence of various design parameters, the present investigation uses formal optimization techniques to determine a unique combination of parameters that maximizes whirl flutter stability. The optimal designs require only modest changes in the key rotor and wing design parameters to significantly increase flutter speed. When constraints on design parameters are relaxed, optimized configurations are obtained that allow large values of kinematic pitch-flap (delta3) coupling without degrading aeroelastic stability. Larger values of delta3 may be desirable for advanced tiltrotor configurations. An investigation of active control of wing flaperons for stability augmentation is also conducted. Both stiff- and soft-inplane tiltrotor configurations are examined. Control systems that increase flutter speed and wing mode sub-critical damping are designed while observing realistic limits on flaperon deflection. The flaperon is shown to be particularly
Aeroelastic Stability and Response of Rotating Structures
NASA Technical Reports Server (NTRS)
Keith, Theo G., Jr.; Reddy, Tondapu
2004-01-01
A summary of the work performed under NASA grant is presented. More details can be found in the cited references. This grant led to the development of relatively faster aeroelastic analysis methods for predicting flutter and forced response in fans, compressors, and turbines using computational fluid dynamic (CFD) methods. These methods are based on linearized two- and three-dimensional, unsteady, nonlinear aerodynamic equations. During the period of the grant, aeroelastic analysis that includes the effects of uncertainties in the design variables has also been developed.
Aeroelastic stability of wind turbine blades
NASA Technical Reports Server (NTRS)
Kaza, K. R. V.
1928-01-01
The second degree nonlinear aeroelastic equations for a flexible, twisted, nonuniform wind turbine blade were developed using Hamilton's principle. The derivation of these equations has its basis in the geometric nonlinear theory of elasticity. These equations with periodic coefficients are suitable for determining the aeroelastic stability and response of large wind turbine blades. Methods for solving these equations are discussed.
Aeroelastic stability analysis of the AD-1 manned oblique-wing aircraft
NASA Technical Reports Server (NTRS)
Rutkowski, M. J.
1977-01-01
The AD-1 manned flight test program was conducted to evaluate the stability, control and handling characteristics of oblique wing aircraft. The results of the aeroelastic stability analysis are presented for both the wing alone and the wing with ailerons. A comparison was made between the results obtained using the traditional k-method of flutter analysis and the results using the PK or British method of flutter analysis. Studies were performed using the latest version of the NASTRAN computer code as well as the PASS/FLUT program.
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.
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.
Aeroelastic stability of wind turbine blade/aileron systems
NASA Technical Reports Server (NTRS)
Strain, J. C.; Mirandy, L.
1995-01-01
Aeroelastic stability analyses have been performed for the MOD-5A blade/aileron system. Various configurations having different aileron torsional stiffness, mass unbalance, and control system damping have been investigated. The analysis was conducted using a code recently developed by the General Electric Company - AILSTAB. The code extracts eigenvalues for a three degree of freedom system, consisting of: (1) a blade flapwise mode; (2) a blade torsional mode; and (3) an aileron torsional mode. Mode shapes are supplied as input and the aileron can be specified over an arbitrary length of the blade span. Quasi-steady aerodynamic strip theory is used to compute aerodynamic derivatives of the wing-aileron combination as a function of spanwise position. Equations of motion are summarized herein. The program provides rotating blade stability boundaries for torsional divergence, classical flutter (bending/torsion) and wing/aileron flutter. It has been checked out against fixed-wing results published by Theodorsen and Garrick. The MOD-5A system is stable with respect to divergence and classical flutter for all practical rotor speeds. Aileron torsional stiffness must exceed a minimum critical value to prevent aileron flutter. The nominal control system stiffness greatly exceeds this minimum during normal operation. The basic system, however, is unstable for the case of a free (or floating) aileron. The instability can be removed either by the addition of torsional damping or mass-balancing the ailerons. The MOD-5A design was performed by the General Electric Company, Advanced Energy Program Department under Contract DEN3-153 with NASA Lewis Research Center and sponsored by the Department of Energy.
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.
ASTROP2 Users Manual: A Program for Aeroelastic Stability Analysis of Propfans
NASA Technical Reports Server (NTRS)
Reddy, T. S. R.; Lucero, John M.
1996-01-01
This manual describes the input data required for using the second version of the ASTROP2 (Aeroelastic STability and Response Of Propulsion systems - 2 dimensional analysis) computer code. In ASTROP2, version 2.0, the program is divided into two modules: 2DSTRIP, which calculates the structural dynamic information; and 2DASTROP, which calculates the unsteady aerodynamic force coefficients from which the aeroelastic stability can be determined. In the original version of ASTROP2, these two aspects were performed in a single program. The improvements to version 2.0 include an option to account for counter rotation, improved numerical integration, accommodation for non-uniform inflow distribution, and an iterative scheme to flutter frequency convergence. ASTROP2 can be used for flutter analysis of multi-bladed structures such as those found in compressors, turbines, counter rotating propellers or propfans. The analysis combines a two-dimensional, unsteady cascade aerodynamics model and a three dimensional, normal mode structural model using strip theory. The flutter analysis is formulated in the frequency domain resulting in an eigenvalue determinant. The flutter frequency and damping can be inferred from the eigenvalues.
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.
Aeroelastic stability analyses of two counter rotating propfan designs for a cruise missile model
NASA Technical Reports Server (NTRS)
Mahajan, Aparajit J.; Lucero, John M.; Mehmed, Oral; Stefko, George L.
1992-01-01
Aeroelastic stability analyses were performed to insure structural integrity of two counterrotating propfan blade designs for a NAVY/Air Force/NASA cruise missile model wind tunnel test. This analysis predicted if the propfan designs would be flutter free at the operating conditions of the wind tunnel test. Calculated stability results are presented for the two blade designs with rotational speed and Mach number as the parameters. A aeroelastic analysis code ASTROP2 (Aeroelastic Stability and Response of Propulsion Systems - 2 Dimensional Analysis), developed at LeRC, was used in this project. The aeroelastic analysis is a modal method and uses the combination of a finite element structural model and two dimensional steady and unsteady cascade aerodynamic models. This code was developed to analyze single rotation propfans but was modified and applied to counterrotating propfans for the present work. Modifications were made to transform the geometry and rotation of the aft rotor to the same reference frame as the forward rotor, to input a non-uniform inflow into the rotor being analyzed, and to automatically converge to the least stable aeroelastic mode.
Smithornis broadbills produce loud wing song by aeroelastic flutter of medial primary wing feathers.
Clark, Christopher J; Kirschel, Alexander N G; Hadjioannou, Louis; Prum, Richard O
2016-04-01
Broadbills in the genus Smithornis produce a loud brreeeeet during a distinctive flight display. It has been posited that this klaxon-like sound is generated non-vocally with the outer wing feathers (P9, P10), but no scientific studies have previously addressed this hypothesis. Although most birds that make non-vocal communication sounds have feathers with a shape distinctively modified for sound production, Smithornis broadbills do not. We investigated whether this song is produced vocally or with the wings in rufous-sided broadbill (S. rufolateralis) and African broad bill (S. capensis). In support of the wing song hypothesis, synchronized high-speed video and sound recordings of displays demonstrated that sound pulses were produced during the downstroke, subtle gaps sometimes appeared between the outer primary feathers P6-P10, and wing tip speed reached 16 m s(-1) Tests of a spread wing in a wind tunnel demonstrated that at a specific orientation, P6 and P7 flutter and produce sound. Wind tunnel tests on individual feathers P5-P10 from a male of each species revealed that while all of these feathers can produce sound via aeroelastic flutter, P6 and P7 produce the loudest sounds, which are similar in frequency to the wing song, at airspeeds achievable by the wing tip during display flight. Consistent with the wind tunnel experiments, field manipulations of P6, P7 and P8 changed the timbre of the wing song, and reduced its tonality, demonstrating that P6 and P7 are together the sound source, and not P9 or P10. The resultant wing song appears to have functionally replaced vocal song.
NASA Astrophysics Data System (ADS)
Taylor, I. J.; Vezza, M.
2009-01-01
The results of a numerical investigation into the aerodynamic characteristics and aeroelastic stability of a proposed footbridge across a highway in the north of England are presented. The longer than usual span, along with the unusual nature of the pedestrian barriers, indicated that the deck configuration was likely to be beyond the reliable limits of the British design code BD 49/01. The calculations were performed using the discrete vortex method, DIVEX, developed at the Universities of Glasgow and Strathclyde. DIVEX has been successfully validated on a wide range of problems, including the aeroelastic response of bridge deck sections. In particular, the investigation focussed on the effects of non-standard pedestrian barriers on the structural integrity of the bridge. The proposed deck configuration incorporated a barrier comprised of angled flat plates, and the bridge was found to be unstable at low wind speeds, with the plates having a strong turning effect on the flow at the leading edge of the deck. These effects are highlighted in both a static and dynamic analysis of the bridge deck, along with modifications to the design that aim to improve the aeroelastic stability of the deck. Proper orthogonal decomposition (POD) was also used to investigate the unsteady pressure field on the upper surface of the static bridge deck. The results of the flutter investigation and the POD analysis highlight the strong influence of the pedestrian barriers on the overall aerodynamic characteristics and aeroelastic stability of the bridge.
NASA Astrophysics Data System (ADS)
Jha, Sourabh; Crittenden, Thomas; Glezer, Ari
2016-11-01
Heat transport within high aspect ratio, rectangular mm-scale channels that model segments of a high-performance, air-cooled heat sink is enhanced by the formation of unsteady small-scale vortical motions induced by autonomous, aeroelastic fluttering of cantilevered planar thin-film reeds. The flow mechanisms and scaling of the interactions between the reed and the channel flow are explored to overcome the limits of forced convection heat transport from air-side heat exchangers. High-resolution PIV measurements in a testbed model show that undulations of the reed's surface lead to formation and advection of vorticity concentrations, and to alternate shedding of spanwise CW and CCW vortices. These vortices scale with the reed motion amplitude, and ultimately result in motions of decreasing scales and enhanced dissipation that are reminiscent of a turbulent flow. The vorticity shedding lead to strong enhancement in heat transfer that increases with the Reynolds number of the base flow (e.g., the channel's thermal coefficient of performance is enhanced by 2.4-fold and 9-fold for base flow Re = 4,000 and 17,400, respectively, with corresponding decreases of 50 and 77% in the required channel flow rates). This is demonstrated in heat sinks for improving the thermal performance of low-Re thermoelectric power plant air-cooled condensers, where the global air-side pressure losses can be significantly reduced by lowering the required air volume flow rate at a given heat flux and surface temperature. AFOSR and NSF-EPRI.
NASA Technical Reports Server (NTRS)
Goldman, Benjamin D.; Dowell, Earl H.; Scott, Robert C.
2014-01-01
Conical shell theory and piston theory aerodynamics are used to study the aeroelastic stability of the thermal protection system (TPS) on the NASA Hypersonic Inflatable Aerodynamic Decelerator (HIAD). Structural models of the TPS consist of single or multiple orthotropic conical shell systems resting on several circumferential linear elastic supports. The shells in each model may have pinned (simply-supported) or elastically-supported edges. The Lagrangian is formulated in terms of the generalized coordinates for all displacements and the Rayleigh-Ritz method is used to derive the equations of motion. The natural modes of vibration and aeroelastic stability boundaries are found by calculating the eigenvalues and eigenvectors of a large coefficient matrix. When the in-flight configuration of the TPS is approximated as a single shell without elastic supports, asymmetric flutter in many circumferential waves is observed. When the elastic supports are included, the shell flutters symmetrically in zero circumferential waves. Structural damping is found to be important in this case. Aeroelastic models that consider the individual TPS layers as separate shells tend to flutter asymmetrically at high dynamic pressures relative to the single shell models. Several parameter studies also examine the effects of tension, orthotropicity, and elastic support stiffness.
NASA Technical Reports Server (NTRS)
Adams, W. M., Jr.; Tiffany, S. H.
1983-01-01
A control law is developed to suppress symmetric flutter for a mathematical model of an aeroelastic research vehicle. An implementable control law is attained by including modified LQG (linear quadratic Gaussian) design techniques, controller order reduction, and gain scheduling. An alternate (complementary) design approach is illustrated for one flight condition wherein nongradient-based constrained optimization techniques are applied to maximize controller robustness.
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.
2015-01-05
The aeroelastic behavior of a finite aspect ratio (AR=6) NACA0018 wing is computationally analyzed. HPCMP CREATE(trademark)-AV Kestrel, a fully...aeroelastically deforming wing . Externally controlled blowing slots distributed along the span of the wing are used to inject mass into the flow field to...coefficients. For the rigid wing , the lift is increased, as are the pitching and rolling moments. When aeroelastic deformation is considered, the
Aeroelastic Stability of a Four-Bladed Semi-Articulated Soft-Inplane Tiltrotor Model
NASA Technical Reports Server (NTRS)
Nixon, Mark W.; Langston, Chester W.; Singleton, Jeffrey D.; Piatak, David J.; Kvaternik, Raymond G.; Corso, Lawrence M.; Brown, Ross
2003-01-01
A new four-bladed, semi-articulated, soft-inplane rotor system, designed as a candidate for future heavy-lift rotorcraft, was tested at model scale on the Wing and Rotor Aeroelastic Testing System (WRATS), a 1/5-size aeroelastic wind-tunnel model based on the V-22. The experimental investigation included a hover test with the model in helicopter mode subject to ground resonance conditions, and a forward flight test with the model in airplane mode subject to whirl-flutter conditions. An active control system designed to augment system damping was also tested as part of this investigation. Results of this study indicate that the new four-bladed, soft-inplane rotor system in hover has adequate damping characteristics and is stable throughout its rotor-speed envelope. However, in airplane mode it produces very low damping in the key wing beam-bending mode, and has a low whirl-flutter stability boundary with respect to airspeed. The active control system was successful in augmenting the damping of the fundamental system modes, and was found to be robust with respect to changes in rotor-speed and airspeed. Finally, conversion-mode dynamic loads were measured on the rotor and these were found to be significantly lower for the new soft-inplane hub than for the previous baseline stiff-inplane hub.
Aeroelastic Stability of a Four-Bladed Semi-Articulated Soft-Inplane Tiltrotor Model
NASA Technical Reports Server (NTRS)
Nixon, Mark W.; Langston, Chester W.; Singleton, Jeffrey D.; Piatak, David J.; Kvaternik, Raymond G.; Corso, Lawrence M.; Brown, Ross K.
2003-01-01
A new four-bladed, semi-articulated, soft-inplane rotor system, designed as a candidate for future heavy-lift rotorcraft, was tested at model scale on the Wing and Rotor Aeroelastic Testing System (WRATS), a 1/5-size aeroelastic wind-tunnel model based on the V-22. The experimental investigation included a hover test with the model in helicopter mode subject to ground resonance conditions, and a forward flight test with the model in airplane mode subject to whirl-flutter conditions. An active control system designed to augment system damping was also tested as part of this investigation. Results of this study indicate that the new four-bladed, soft-inplane rotor system in hover has adequate damping characteristics and is stable throughout its rotor-speed envelope. However, in airplane mode it produces very low damping in the key wing beam-bending mode, and has a low whirl-flutter stability boundary with respect to airspeed. The active control system was successful in augmenting the damping of the fundamental system modes, and was found to be robust with respect to changes in rotor speed and airspeed. Finally, conversion-mode dynamic loads were measured on the rotor and these were found to be signi.cantly lower for the new soft-inplane hub than for the previous baseline stiff - inplane hub.
Helicopter aeroelastic stability and response - Current topics and future trends
NASA Technical Reports Server (NTRS)
Friedmann, Peretz P.
1990-01-01
This paper presents several current topics in rotary wing aeroelasticity and concludes by attempting to anticipate future trends and developments. These topics are: (1) the role of geometric nonlinearities; (2) structural modeling, and aeroelastic analysis of composite rotor blades; (3) aeroelastic stability and response in forward flight; (4) modeling of coupled rotor/fuselage aeromechanical problems and their active control; and (5) the coupled rotor-fuselage vibration problem and its alleviation by higher harmonic control. Selected results illustrating the fundamental aspects of these topics are presented. Future developments are briefly discussed.
Aeroelastic Stability of Idling Wind Turbines
NASA Astrophysics Data System (ADS)
Wang, Kai; Riziotis, Vasilis A.; Voutsinas, Spyros G.
2016-09-01
Wind turbine rotors in idling operation mode can experience high angles of attack, within the post stall region that are capable of triggering stall-induced vibrations. In the present paper rotor stability in slow idling operation is assessed on the basis of non-linear time domain and linear eigenvalue analysis. Analysis is performed for a 10 MW conceptual wind turbine designed by DTU. First the flow conditions that are likely to favour stall induced instabilities are identified through non-linear time domain aeroelastic analysis. Next, for the above specified conditions, eigenvalue stability simulations are performed aiming at identifying the low damped modes of the turbine. Finally the results of the eigenvalue analysis are evaluated through computations of the work of the aerodynamic forces by imposing harmonic vibrations following the shape and frequency of the various modes. Eigenvalue analysis indicates that the asymmetric and symmetric out-of-plane modes have the lowest damping. The results of the eigenvalue analysis agree well with those of the time domain analysis.
Survey of Army/NASA rotorcraft aeroelastic stability research
NASA Technical Reports Server (NTRS)
Ormiston, Robert A.; Warmbrodt, William G.; Hodges, Dewey H.; Peters, David A.
1988-01-01
Theoretical and experimental developments in the aeroelastic and aeromechanical stability of helicopters and tilt-rotor aircraft are addressed. Included are the underlying nonlinear structural mechanics of slender rotating beams, necessary for accurate modeling of elastic cantilever rotor blades, and the development of dynamic inflow, an unsteady aerodynamic theory for low frequency aeroelastic stability applications. Analytical treatment of isolated rotor stability in hover and forward flight, coupled rotor-fuselage stability are considered. Results of parametric investigations of system behavior are presented, and correlations between theoretical results and experimental data from small- and large-scale wind tunnel and flight testing are discussed.
NASA Technical Reports Server (NTRS)
Ashley, H.
1984-01-01
Graduate research activity in the following areas is reported: the divergence of laminated composite lifting surfaces, subsonic propeller theory and aeroelastic analysis, and cross sectional resonances in wind tunnels.
NASA Technical Reports Server (NTRS)
Reddy, T. S. R.; Srivastava, R.
1996-01-01
This guide describes the input data required for using MSAP2D (Multi Stage Aeroelastic analysis Program - Two Dimensional) computer code. MSAP2D can be used for steady, unsteady aerodynamic, and aeroelastic (flutter and forced response) analysis of bladed disks arranged in multiple blade rows such as those found in compressors, turbines, counter rotating propellers or propfans. The code can also be run for single blade row. MSAP2D code is an extension of the original NPHASE code for multiblade row aerodynamic and aeroelastic analysis. Euler equations are used to obtain aerodynamic forces. The structural dynamic equations are written for a rigid typical section undergoing pitching (torsion) and plunging (bending) motion. The aeroelastic equations are solved in time domain. For single blade row analysis, frequency domain analysis is also provided to obtain unsteady aerodynamic coefficients required in an eigen analysis for flutter. In this manual, sample input and output are provided for a single blade row example, two blade row example with equal and unequal number of blades in the blade rows.
A Coupled Aeroelastic Model for Launch Vehicle Stability Analysis
NASA Technical Reports Server (NTRS)
Orr, Jeb S.
2010-01-01
A technique for incorporating distributed aerodynamic normal forces and aeroelastic coupling effects into a stability analysis model of a launch vehicle is presented. The formulation augments the linear state-space launch vehicle plant dynamics that are compactly derived as a system of coupled linear differential equations representing small angular and translational perturbations of the rigid body, nozzle, and sloshing propellant coupled with normal vibration of a set of orthogonal modes. The interaction of generalized forces due to aeroelastic coupling and thrust can be expressed as a set of augmenting non-diagonal stiffness and damping matrices in modal coordinates with no penalty on system order. While the eigenvalues of the structural response in the presence of thrust and aeroelastic forcing can be predicted at a given flight condition independent of the remaining degrees of freedom, the coupled model provides confidence in closed-loop stability in the presence of rigid-body, slosh, and actuator dynamics. Simulation results are presented that characterize the coupled dynamic response of the Ares I launch vehicle and the impact of aeroelasticity on control system stability margins.
Aeroelastic, CFD, and Dynamic Computation and Optimization for Buffet and Flutter Application
NASA Technical Reports Server (NTRS)
Kandil, Osama A.
1997-01-01
The work presented in this paper include: 'Coupled and Uncoupled Bending-Torsion Responses of Twin-Tail Buffet'; 'Fluid/Structure Twin Tail Buffet Response Over a Wide Range of Angles of Attack'; 'Resent Advances in Multidisciplinary Aeronautical Problems of Fluids/Structures/Dynamics Interaction'; and'Development of a Coupled Fluid/Structure Aeroelastic Solver with Applications to Vortex Breakdown induced Twin Tail Buffeting.
Application of TURBO-AE to Flutter Prediction: Aeroelastic Code Development
NASA Technical Reports Server (NTRS)
Hoyniak, Daniel; Simons, Todd A.; Stefko, George (Technical Monitor)
2001-01-01
The TURBO-AE program has been evaluated by comparing the obtained results to cascade rig data and to prediction made from various in-house programs. A high-speed fan cascade, a turbine cascade, a turbine cascade and a fan geometry that shower flutter in torsion mode were analyzed. The steady predictions for the high-speed fan cascade showed the TURBO-AE predictions to match in-house codes. However, the predictions did not match the measured blade surface data. Other researchers also reported similar disagreement with these data set. Unsteady runs for the fan configuration were not successful using TURBO-AE .
Aeroelastic stability analysis of flexible overexpanded rocket nozzle
NASA Astrophysics Data System (ADS)
Bekka, N.; Sellam, M.; Chpoun, A.
2016-07-01
The aim of this paper is to present a new aeroelastic stability model taking into account the viscous effects for a supersonic nozzle flow in overexpanded regimes. This model is inspired by the Pekkari model which was developed initially for perfect fluid flow. The new model called the "Modified Pekkari Model" (MPM) considers a more realistic wall pressure profile for the case of a free shock separation inside the supersonic nozzle using the free interaction theory of Chapman. To reach this objective, a code for structure computation coupled with aerodynamic excitation effects is developed that allows the analysis of aeroelastic stability for the overexpanded nozzles. The main results are presented in a comparative manner using existing models (Pekkari model and its extended version) and the modified Pekkari model developed in this work.
Effect of compressive force on aeroelastic stability of a strut-braced wing
NASA Astrophysics Data System (ADS)
Sulaeman, Erwin
2002-01-01
Recent investigations of a strut-braced wing (SBW) aircraft show that, at high positive load factors, a large tensile force in the strut leads to a considerable compressive axial force in the inner wing, resulting in a reduced bending stiffness and even buckling of the wing. Studying the influence of this compressive force on the structural response of SBW is thus of paramount importance in the early stage of SBW design. The purpose of the this research is to investigate the effect of compressive force on aeroelastic stability of the SBW using efficient structural finite element and aerodynamic lifting surface methods. A procedure is developed to generate wing stiffness distribution for detailed and simplified wing models and to include the compressive force effect in the SBW aeroelastic analysis. A sensitivity study is performed to generate response surface equations for the wing flutter speed as functions of several design variables. These aeroelastic procedures and response surface equations provide a valuable tool and trend data to study the unconventional nature of SBW. In order to estimate the effect of the compressive force, the inner part of the wing structure is modeled as a beam-column. A structural finite element method is developed based on an analytical stiffness matrix formulation of a non-uniform beam element with arbitrary polynomial variations in the cross section. By using this formulation, the number of elements to model the wing structure can be reduced without degrading the accuracy. The unsteady aerodynamic prediction is based on a discrete element lifting surface method. The present formulation improves the accuracy of existing lifting surface methods by implementing a more rigorous treatment on the aerodynamic kernel integration. The singularity of the kernel function is isolated by implementing an exact expansion series to solve an incomplete cylindrical function problem. A hybrid doublet lattice/doublet point scheme is devised to reduce
NASA Astrophysics Data System (ADS)
Försching, H.; Knaack, J. M.
1993-08-01
A parametric investigation is performed of the aeroelastic flutter stability behaviour of a semi-rigid 3-D wing-with-engine nacelle model in subsonic flow. The system under investigation is a wind tunnel model that was flutter tested some years ago. It consists of a swept-back half wing with a pylon-mounted engine nacelle, representative of modern large transport aircraft, and is elastically restrained at its wing root, so that it may execute decoupled (rigid body) rolling and pitching oscillations about two orthogonal axes. For this binary aeroelastic system, first the equations of motion and then the aeroelastic stability equations are set up in terms of generalized coordinates. In addition to the basic wind tunnel model configuration, two artificial configurations with other positions of the rotation axes and corresponding mode shapes are investigated. For the computation of the motion-induced generalized airloads, a panel technique is used for both the wing and the engine nacelle that is replaced by an annular wing. Numerical results are presented for several systematic parameter variations and Mach numbers, where special emphasis is placed on the effects of the motion-induced unsteady airloads acting on the engine nacelle, the position of the rotation axes, and the frequency ratio of the two modes in roll and pitch. Moreover, a comparison is made with some wind tunnel test results.
Whirl Flutter Stability of Two-Bladed Proprotor/Pylon Systems In High Speed Flight
NASA Technical Reports Server (NTRS)
Singh, Beerinder; Chopra, Inderjit; Pototzky, A. (Technical Monitor)
2002-01-01
The lack of polar symmetry in two-bladed rotors leads to equations of motion with periodic coefficients in axial flight, which is contrary to three or more bladed rotors that result in constant coefficient equations. With periodic coefficients, the analysis becomes involved, as a result very few studies have been directed towards the analysis of two-bladed rotors. In this paper, the aeroelastic stability of two-bladed proprotor/pylon/wing combinations is examined in high speed axial flight. Several parametric studies are carried out to illustrate the special nature of two-bladed proprotors and to better understand the mechanism of whirl-flutter in such rotors. The wing beam bending mode for two-bladed rotors is found to be stable over the range of parameters examined, a behaviour very different from three-bladed rotors. Also, the wing torsion mode exhibits a new type of instability similar to a wing torsional divergence scouring at I/rev frequency. This type of behaviour is not seen in three and more bladed rotors. The interaction between wing chordwise bending and torsion modes is found to be much greater in the case of two-bladed rotors and, over the range of parameters considered, these two modes govern the stability of the system.
Analyzing Aeroelastic Stability of a Tilt-Rotor Aircraft
NASA Technical Reports Server (NTRS)
Kvaternil, Raymond G.
2006-01-01
Proprotor Aeroelastic Stability Analysis, now at version 4.5 (PASTA 4.5), is a FORTRAN computer program for analyzing the aeroelastic stability of a tiltrotor aircraft in the airplane mode of flight. The program employs a 10-degree- of-freedom (DOF), discrete-coordinate, linear mathematical model of a rotor with three or more blades and its drive system coupled to a 10-DOF modal model of an airframe. The user can select which DOFs are included in the analysis. Quasi-steady strip-theory aerodynamics is employed for the aerodynamic loads on the blades, a quasi-steady representation is employed for the aerodynamic loads acting on the vibrational modes of the airframe, and a stability-derivative approach is used for the aerodynamics associated with the rigid-body DOFs of the airframe. Blade parameters that vary with the blade collective pitch can be obtained by interpolation from a user-defined table. Stability is determined by examining the eigenvalues that are obtained by solving the coupled equations of motions as a matrix eigenvalue problem. Notwithstanding the relative simplicity of its mathematical foundation, PASTA 4.5 and its predecessors have played key roles in a number of engineering investigations over the years.
Studies in hypersonic aeroelasticity
NASA Astrophysics Data System (ADS)
Nydick, Ira Harvey
2000-11-01
This dissertation describes the aeroelastic analysis of a generic hypersonic vehicle, focusing on two specific problems: (1) hypersonic panel flutter, and (2) aeroelastic behavior of a complete unrestrained generic hypersonic vehicle operating at very high Mach numbers. The panels are modeled as shallow shells using Marguerre nonlinear shallow shell theory for orthotropic panels and the aerodynamic loads are obtained from third order piston theory. Two models of curvature, several applied temperature distributions, and the presence of a shock are also included in the model. Results indicate that the flutter speed of the panel is significantly reduced by temperature variations comparable to the buckling temperature and by the presence of a shock. A panel with initial curvature can be more stable than the flat panel but the increase in stability depends in a complex way on the material properties of the panel and the amount of curvature. At values of dynamic pressure above critical, aperiodic motion was observed. The value of dynamic pressure for which this occurs in both heated panels and curved panels is much closer to the critical dynamic pressure than for the flat, unheated panel. A comparison of piston theory aerodynamics and Euler and Navier-Stokes aerodynamics was performed for a two dimensional panel with prescribed motion and the results indicate that while 2nd or higher order piston theory agrees very well with the Euler solution for the frequencies seen in hypersonic panel flutter, it differs substantially from the Navier-Stokes solution. The aeroelastic behavior of the complete vehicle was simulated using the unrestrained equations of motion, utilizing the method of quasi-coordinates. The unrestrained mode shapes of the vehicle were obtained from an equivalent plate analysis using an available code (ELAPS). The effects of flexible trim and rigid body degrees of freedom are carefully incorporated in the mathematical model. This model was applied to a
Some observations on four current subjects related to aeroelastic stability
NASA Technical Reports Server (NTRS)
Ashley, H.
1978-01-01
After introductory comments on the literature and the purposes of this paper, a table is presented summarizing the author's views on some currently solved vs partially unsolved problems related to aeroelastic stability. The term 'solved' is used in the practical sense that engineers are able to cope confidently with that problem during the process of structural design. Selected entries in the table are reviewed, partially to motivate the topics in the rest of the paper. The 'four current subjects' are chosen both for timeliness and because they are among the ongoing interests of the Stanford group. The first involves the prediction of linearized unsteady aerodynamic loads due to arbitrary motions of streamlined shapes. Some contributions by Edwards are refined, which were motivated by the requirements of active control system design. The second subject is nonlinear unsteady aerodynamics for the transonic regime. After describing a few useful developments from locally-linear theory and computational fluid dynamics, there is suggested an empirical procedure for interim-analysis purposes. The third and fourth subjects concern recent discoveries regarding the aeroelastic stability of large-aspect-ratio wings and wind turbines. The former work is mainly that of Petre and Boyd. The latter includes some of the author's own preliminary discoveries about the performance and dynamics of vertical-axis machines.
NASA Technical Reports Server (NTRS)
Herrick, Gregory P.
2014-01-01
Concerns regarding noise, propulsive efficiency, and fuel burn are inspiring aircraft designs wherein the propulsive turbomachines are partially (or fully) embedded within the airframe; such designs present serious concerns with regard to aerodynamic and aeromechanic performance of the compression system in response to inlet distortion. Previously, a preliminary design of a forward-swept high-speed fan exhibited flutter concerns in cleaninlet flows, and the present author then studied this fan further in the presence of off-design distorted in-flows. Continuing this research, a three-dimensional, unsteady, Navier-Stokes computational fluid dynamics code is again applied to analyze and corroborate fan performance with clean inlet flow and now with a simplified, sinusoidal distortion of total pressure at the aerodynamic interface plane. This code, already validated in its application to assess aerodynamic damping of vibrating blades at various flow conditions using a one-way coupled energy-exchange approach, is modified to include a two-way coupled time-marching aeroelastic simulation capability. The two coupling methods are compared in their evaluation of flutter stability in the presence of distorted in-flows.
NASA Technical Reports Server (NTRS)
Herrick, Gregory P.
2014-01-01
Concerns regarding noise, propulsive efficiency, and fuel burn are inspiring aircraft designs wherein the propulsive turbomachines are partially (or fully)embedded within the airframe; such designs present serious concerns with regard to aerodynamic and aeromechanic performance of the compression system in response to inlet distortion. Previously, a preliminary design of a forward-swept high-speed fan exhibited flutter concerns in clean-inlet flows, and the present author then studied this fan further in the presence of off-design distorted in-flows. A three-dimensional, unsteady, Navier-Stokes computational fluid dynamics code is applied to analyze and corroborate fan performance with clean inlet flow. This code, already validated in its application to assess aerodynamic damping of vibrating blades at various flow conditions using a loosely-coupled approach, is modified to include a tightly-coupled aeroelastic simulation capability, and then loosely-coupled and tightly-coupled methods arecompared in their evaluation of flutter stability in distorted in-flows.
NASA Technical Reports Server (NTRS)
Herrick, Gregory P.
2014-01-01
Concerns regarding noise, propulsive efficiency, and fuel burn are inspiring aircraft designs wherein the propulsive turbomachines are partially (or fully) embedded within the airframe; such designs present serious concerns with regard to aerodynamic and aeromechanic performance of the compression system in response to inlet distortion. Previously, a preliminary design of a forward-swept high-speed fan exhibited flutter concerns in clean-inlet flows, and the present author then studied this fan further in the presence of off-design distorted in-flows. Continuing this research, a three-dimensional, unsteady, Navier-Stokes computational fluid dynamics code is again applied to analyze and corroborate fan performance with clean inlet flow and now with a simplified, sinusoidal distortion of total pressure at the aerodynamic interface plane. This code, already validated in its application to assess aerodynamic damping of vibrating blades at various flow conditions using a one-way coupled energy-exchange approach, is modified to include a two-way coupled timemarching aeroelastic simulation capability. The two coupling methods are compared in their evaluation of flutter stability in the presence of distorted in-flows.
NASA Technical Reports Server (NTRS)
Howard, Anna K. T.
1999-01-01
The tiltrotor offers the best mix of hovering and cruise flight of any of the current V/STOL configurations. One possible improvement on the tiltrotors of today designs would be using a soft-inplane hingeless hub. The advantages to a soft-inplane hingeless hub range from reduced weight and maintenance to reduced vibration and loads. However, soft-inplane rotor systems are inherently in danger of the aeromechanical instabilities of ground and air resonance. Furthermore tiltrotors can be subject to whirl flutter. At least in part because of the potential for air and ground resonance in a soft-inplane rotor, the Bell XV-15, the Bell-Boeing V-22 Osprey, and the new Bell Augusta 609 have stiff-inplane, gimballed rotors which do not experience these instabilities. In order to design soft-inplane V/STOL aircraft that do not experience ground or air resonance, it is important to be able to predict these instabilities accurately. Much of the research studying the stability of tiltrotors has been focused on the understanding and prediction of whirl flutter. As this instability is increasingly well understood, air and ground resonance for a tiltrotor need to be investigated. Once we understand the problems of air and ground resonance in a tiltrotor, we must look for solutions to these instabilities. Other researchers have found composite or kinematic couplings in the blades of a helicopter helpful for ground and air resonance stability. Tiltrotor research has shown composite couplings in the wing to be helpful for whirl flutter. Therefore, this project will undertake to model ground and air resonance of a soft-inplane hingeless tiltrotor to understand the mechanisms involved and to evaluate whether aeroelastic couplings in the wing or kinematic couplings in the blades would aid in stabilizing these instabilities in a tiltrotor.
Level-Set Topology Optimization with Aeroelastic Constraints
NASA Technical Reports Server (NTRS)
Dunning, Peter D.; Stanford, Bret K.; Kim, H. Alicia
2015-01-01
Level-set topology optimization is used to design a wing considering skin buckling under static aeroelastic trim loading, as well as dynamic aeroelastic stability (flutter). The level-set function is defined over the entire 3D volume of a transport aircraft wing box. Therefore, the approach is not limited by any predefined structure and can explore novel configurations. The Sequential Linear Programming (SLP) level-set method is used to solve the constrained optimization problems. The proposed method is demonstrated using three problems with mass, linear buckling and flutter objective and/or constraints. A constraint aggregation method is used to handle multiple buckling constraints in the wing skins. A continuous flutter constraint formulation is used to handle difficulties arising from discontinuities in the design space caused by a switching of the critical flutter mode.
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.
Optimal mistuning for enhanced aeroelastic stability of transonic fans
NASA Technical Reports Server (NTRS)
Hall, K. C.; Crawley, E. F.
1983-01-01
An inverse design procedure was developed for the design of a mistuned rotor. The design requirements are that the stability margin of the eigenvalues of the aeroelastic system be greater than or equal to some minimum stability margin, and that the mass added to each blade be positive. The objective was to achieve these requirements with a minimal amount of mistuning. Hence, the problem was posed as a constrained optimization problem. The constrained minimization problem was solved by the technique of mathematical programming via augmented Lagrangians. The unconstrained minimization phase of this technique was solved by the variable metric method. The bladed disk was modelled as being composed of a rigid disk mounted on a rigid shaft. Each of the blades were modelled with a single tosional degree of freedom.
Flutter Analysis of a Transonic Fan
NASA Technical Reports Server (NTRS)
Srivastava, R.; Bakhle, M. A.; Keith, T. G., Jr.; Stefko, G. L.
2002-01-01
This paper describes the calculation of flutter stability characteristics for a transonic forward swept fan configuration using a viscous aeroelastic analysis program. Unsteady Navier-Stokes equations are solved on a dynamically deforming, body fitted, grid to obtain the aeroelastic characteristics using the energy exchange method. The non-zero inter-blade phase angle is modeled using phase-lagged boundary conditions. Results obtained show good correlation with measurements. It is found that the location of shock and variation of shock strength strongly influenced stability. Also, outboard stations primarily contributed to stability characteristics. Results demonstrate that changes in blade shape impact the calculated aerodynamic damping, indicating importance of using accurate blade operating shape under centrifugal and steady aerodynamic loading for flutter prediction. It was found that the calculated aerodynamic damping was relatively insensitive to variation in natural frequency.
Aeroelastic Stability of Rotor Blades Using Finite Element Analysis
NASA Technical Reports Server (NTRS)
Chopra, I.; Sivaneri, N.
1982-01-01
The flutter stability of flap bending, lead-lag bending, and torsion of helicopter rotor blades in hover is investigated using a finite element formulation based on Hamilton's principle. The blade is divided into a number of finite elements. Quasi-steady strip theory is used to evaluate the aerodynamic loads. The nonlinear equations of motion are solved for steady-state blade deflections through an iterative procedure. The equations of motion are linearized assuming blade motion to be a small perturbation about the steady deflected shape. The normal mode method based on the coupled rotating natural modes is used to reduce the number of equations in the flutter analysis. First the formulation is applied to single-load-path blades (articulated and hingeless blades). Numerical results show very good agreement with existing results obtained using the modal approach. The second part of the application concerns multiple-load-path blades, i.e. bearingless blades. Numerical results are presented for several analytical models of the bearingless blade. Results are also obtained using an equivalent beam approach wherein a bearingless blade is modelled as a single beam with equivalent properties. Results show the equivalent beam model.
Aeroelastic stability and response of horizontal axis wind turbine blades
NASA Technical Reports Server (NTRS)
Kottapalli, S. B. R.; Friedmann, P. P.; Rosen, A.
1978-01-01
The coupled flap-lag-torsion equations of motion of an isolated horizontal axis wind turbine blade are formulated. Quasi-steady blade-element strip theory was applied to derive the aerodynamic operator which includes boundary layer type gradient winds. The final equations which have periodic coefficients were solved in order to obtain the aeroelastic response and stability of large horizontal axis wind turbine blade. A new method of generating an appropriate time-dependent equilibrium position (required for the stability analysis) has been implemented. Representative steady-state responses and stability boundaries, applicable mainly to an existing blade design (NASA/-ERDA MOD-0), are presented. The results indicate that the MOD-0 configuration is a basically stable design and that blade stability is not sensitive to offsets between blade elastic axis and aerodynamic center. Blade stability appears to be sensitive to precone. The tower shadow (or wake) has a considerable effect on the flap response but leaves blade stability unchanged. Finally, it was found that non linear terms in the equations of motion can significantly affect the linearized stability boundaries, however, these terms have a negligible effect on blade response at operating conditions.
The aeroelastic stability improvements of soft-inplane tiltrotors by active and passive approaches
NASA Astrophysics Data System (ADS)
Paik, Jinho
Soft-inplane tiltrotors in cruise mode have exhibited unacceptably low subcritical damping in the wing vertical bending mode as well as reduced critical whirl-flutter speed. However, soft-inplane rotor system is highly advantageous over stiff-inplane rotor system in terms of inplane dynamic hub loads which results in weight/performance penalties. Therefore, ensuring adequate aeroelastic stability characteristics is a prerequisite for soft-inplane rotor system to be used in future advanced tiltrotors. This dissertation constitutes fundamental studies of soft-inplane tiltrotors and appropriate methods to alleviate whirl-flutter instability. This study consists of four major investigations. The first investigation includes validation efforts of present analytical model against the recently available data for the Bell generic semi-span model in airplane mode and the SASIP model in hover mode. The second investigation addresses the approaches which have been employed to establish a physical understanding of the very low sub-critical damping phenomenon, which is consistently exhibited by soft-inplane tiltrotor configurations. Through analyses and comparison studies mainly between the Bell generic soft- and stiff-inplane semi-span models, the physics behind this phenomenon is emphasized. In the third investigation, parametric studies and design optimization of the rotor/wing design variables are performed in order to passively improve the whirl stability boundaries. For the last investigation, the effectiveness of active control through wing-flaperon and swashplate control inputs is examined in terms of stability improvement of soft-inplane tiltrotors. Scheduled gain and constant gain controllers are first compared for each actuation scheme and then output feedback controllers based on easily measurable wing states are compared with full-state feedback controllers. The baseline soft-inplane configurations used in passive and active studies are the full-scale Boeing Model
Aeroelastic Tailoring for Stability Augmentation and Performance Enhancements of Tiltrotor Aircraft
NASA Technical Reports Server (NTRS)
Nixon, Mark W.; Piatak, David J.; Corso, Lawrence M.; Popelka, David A.
1999-01-01
The requirements for increased speed and productivity for tiltrotors has spawned several investigations associated with proprotor aeroelastic stability augmentation and aerodynamic performance enhancements. Included among these investigations is a focus on passive aeroelastic tailoring concepts which exploit the anisotropic capabilities of fiber composite materials. Researchers at Langley Research Center and Bell Helicopter have devoted considerable effort to assess the potential for using these materials to obtain aeroelastic responses which are beneficial to the important stability and performance considerations of tiltrotors. Both experimental and analytical studies have been completed to examine aeroelastic tailoring concepts for the tiltrotor, applied either to the wing or to the rotor blades. This paper reviews some of the results obtained in these aeroelastic tailoring investigations and discusses the relative merits associated with these approaches.
ASTROP2 users manual: A program for aeroelastic stability analysis of propfans
NASA Technical Reports Server (NTRS)
Narayanan, G. V.; Kaza, K. R. V.
1991-01-01
A user's manual is presented for the aeroelastic stability and response of propulsion systems computer program called ASTROP2. The ASTROP2 code preforms aeroelastic stability analysis of rotating propfan blades. This analysis uses a two-dimensional, unsteady cascade aerodynamics model and a three-dimensional, normal-mode structural model. Analytical stability results from this code are compared with published experimental results of a rotating composite advanced turboprop model and of nonrotating metallic wing model.
NASA Technical Reports Server (NTRS)
Dowell, E. H.
1972-01-01
Criteria are presented for the prediction of panel flutter, determination of its occurrence, design for its prevention, and evaluation of its severity. Theoretical analyses recommended for the prediction of flutter stability boundaries, vibration amplitudes, and frequencies for several types of panels are described. Vibration tests and wind tunnel tests are recommended for certain panels and environmental flow conditions to provide information for design of verification analysis. Appropriate design margins on flutter stability boundaries are given and general criteria are presented for evaluating the severity of possible short-duration, limited-amplitude panel flutter on nonreusable vehicles.
Dynamic structural aeroelastic stability testing of the XV-15 tilt rotor research aircraft
NASA Technical Reports Server (NTRS)
Schroers, L. G.
1982-01-01
For the past 20 years, a significant effort has been made to understand and predict the structural aeroelastic stability characteristics of the tilt rotor concept. Beginning with the rotor-pylon oscillation of the XV-3 aircraft, the problem was identified and then subjected to a series of theoretical studies, plus model and full-scale wind tunnel tests. From this data base, methods were developed to predict the structural aeroelastic stability characteristics of the XV-15 Tilt Rotor Research Aircraft. The predicted aeroelastic characteristics are examined in light of the major parameters effecting rotor-pylon-wing stability. Flight test techniques used to obtain XV-15 aeroelastic stability are described. Flight test results are summarized and compared to the predicted values. Wind tunnel results are compared to flight test results and correlated with predicted values.
Aeroelastic stability of wind turbines: the problem, the methods and the issues
NASA Astrophysics Data System (ADS)
Riziotis, V. A.; Voutsinas, S. G.; Politis, E. S.; Chaviaropoulos, P. K.
2004-10-01
Aeroelastic stability is a key issue in the design process of wind turbines towards both enchanced stability and increased fatigue life. The theory and models behind the state-of-the-art aeroelastic stability tools developed for the analysis of the complete wind turbine at the Centre for Renewable Energy Sources and the National Technical University of Athens are presented in this article. Application examples of stability calculations for a pitch, variable speed and a stall-regulated wind turbine are also presented. Copyright
Unstalled flutter stability predictions and comparisons to test data for a composite prop-fan model
NASA Technical Reports Server (NTRS)
Turnberg, J. E.
1986-01-01
The aeroelastic stability analyses for three graphite/epoxy composite Prop-Fan designs and post-test stability analysis for one of the designs, the SR-3C-X2 are presented. It was shown that Prop-Fan stability can be effectively analyzed using the F203 modal aeroelastic stability analysis developed at Hamilton Standard and that first mode torsion-bending coupling has a direct effect on blade stability. Positive first mode torsion-bending coupling is a destabilizing factor and the minimization of this parameter will increase Prop-Fan stability. It was also shown that Prop-Fan stability analysis using F203 is sensitive to the blade modal data used as input. Calculated blade modal properties varied significantly with the structural analysis used, and these variations are reflected in the F203 calculations.
NASA Technical Reports Server (NTRS)
Nixon, Mark W.
1993-01-01
There is a potential for improving the performance and aeroelastic stability of tiltrotors through the use of elastically-coupled composite rotor blades. To study the characteristics of tiltrotors with these types of rotor blades it is necessary to formulate a new analysis which has the capabilities of modeling both a tiltrotor configuration and an anisotropic rotor blade. Background for these formulations is established in two preliminary investigations. In the first, the influence of several system design parameters on tiltrotor aeroelastic stability is examined for the high-speed axial flight mode using a newly-developed rigid-blade analysis with an elastic wing finite element model. The second preliminary investigation addresses the accuracy of using a one-dimensional beam analysis to predict frequencies of elastically-coupled highly-twisted rotor blades. Important aspects of the new aeroelastic formulations are the inclusion of a large steady pylon angle which controls tilt of the rotor system with respect to the airflow, the inclusion of elastic pitch-lag coupling terms related to rotor precone, the inclusion of hub-related degrees of freedom which enable modeling of a gimballed rotor system and engine drive-train dynamics, and additional elastic coupling terms which enable modeling of the anisotropic features for both the rotor blades and the tiltrotor wing. Accuracy of the new tiltrotor analysis is demonstrated by a comparison of the results produced for a baseline case with analytical and experimental results reported in the open literature. Two investigations of elastically tailored blades on a baseline tiltrotor are then conducted. One investigation shows that elastic bending-twist coupling of the rotor blade is a very effective means for increasing the flutter velocity of a tiltrotor, and the magnitude of coupling required does not have an adverse effect on performance or blade loads. The second investigation shows that passive blade twist control via
Unified Formulation of the Aeroelasticity of Swept Lifting Surfaces
NASA Technical Reports Server (NTRS)
Silva, Walter; Marzocca, Piergiovanni; Librescu, Liviu
2001-01-01
An unified approach for dealing with stability and aeroelastic response to time-dependent pressure pulses of swept wings in an incompressible flow is developed. To this end the indicial function concept in time and frequency domains, enabling one to derive the proper unsteady aerodynamic loads is used. Results regarding stability in the frequency and time domains, and subcritical aeroelastic response to arbitrary time-dependent external excitation obtained via the direct use of the unsteady aerodynamic derivatives for 3-D wings are supplied. Closed form expressions for unsteady aerodynamic derivatives using this unified approach have been derived and used to illustrate their application to flutter and aeroelastic response to blast and sonic-boom signatures. In this context, an original representation of the aeroelastic response in the phase space was presented and pertinent conclusions on the implications of some basic parameters have been outlined.
Nonlinear flutter of composite plates with damage evolution
NASA Astrophysics Data System (ADS)
Kim, Young I.; Stragnac, Thomas W.; Kurdila, Andrew J.
1993-04-01
The investigators present a study of dynamic and aeroelastic response of structures which evolve due to damage. Aeroelastic response is shown to be dependent upon the distribution and accumulation of damage. In turn, the damage is dependent upon the presence of the aerodynamic loads. Dynamic characteristics are unique to the coupled damage/aeroelastic system and are developed as part of the solution methodology. In this study, the damage is due to the natural progression of microcracking of the composite structure; yet, the control model presented is appropriate for distributed actuation systems. The stability boundary for aeroelastic flutter and divergence evolves due to damage. Control design based upon the min-max control theory is presented which addresses model uncertainties.
Aeroelastic Stability Investigations for Large-scale Vertical Axis Wind Turbines
NASA Astrophysics Data System (ADS)
Owens, B. C.; Griffith, D. T.
2014-06-01
The availability of offshore wind resources in coastal regions, along with a high concentration of load centers in these areas, makes offshore wind energy an attractive opportunity for clean renewable electricity production. High infrastructure costs such as the offshore support structure and operation and maintenance costs for offshore wind technology, however, are significant obstacles that need to be overcome to make offshore wind a more cost-effective option. A vertical-axis wind turbine (VAWT) rotor configuration offers a potential transformative technology solution that significantly lowers cost of energy for offshore wind due to its inherent advantages for the offshore market. However, several potential challenges exist for VAWTs and this paper addresses one of them with an initial investigation of dynamic aeroelastic stability for large-scale, multi-megawatt VAWTs. The aeroelastic formulation and solution method from the BLade Aeroelastic STability Tool (BLAST) for HAWT blades was employed to extend the analysis capability of a newly developed structural dynamics design tool for VAWTs. This investigation considers the effect of configuration geometry, material system choice, and number of blades on the aeroelastic stability of a VAWT, and provides an initial scoping for potential aeroelastic instabilities in large-scale VAWT designs.
NASA Technical Reports Server (NTRS)
Chattopadhyay, Aditi; Mccarthy, Thomas R.; Madden, John F., III
1992-01-01
An optimization procedure is developed for the design of high speed prop-rotors to be used in civil tiltrotor applications. The goal is to couple aerodynamic performance, aeroelastic stability, and structural design requirements inside a closed-loop optimization procedure. The objective is to minimize the gross weight and maximize the propulsive efficiency in high speed cruise. Constraints are imposed on the rotor aeroelastic stability in both hover and cruise and rotor figure of merit in hover. Both structural and aerodynamic design variables are used.
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.
Stability and Control Properties of an Aeroelastic Fixed Wing Micro Aerial Vehicle
NASA Technical Reports Server (NTRS)
Waszak, Martin R.; Jenkins, Luther N.; Ifju, Peter
2001-01-01
Micro aerial vehicles have been the subject of considerable interest and development over the last several years. The majority of current vehicle concepts rely on rigid fixed wings or rotors. An alternate design based on an aeroelastic membrane wing concept has also been developed that has exhibited desired characteristics in flight test demonstrations and competition. This paper presents results from a wind tunnel investigation that sought to quantify stability and control properties for a family of vehicles using the aeroelastic design. The results indicate that the membrane wing does exhibit potential benefits that could be exploited to enhance the design of future flight vehicles.
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.
Recent Applications of the Volterra Theory to Aeroelastic Phenomena
NASA Technical Reports Server (NTRS)
Silva, Walter A.; Haji, Muhammad R; Prazenica, Richard J.
2005-01-01
The identification of nonlinear aeroelastic systems based on the Volterra theory of nonlinear systems is presented. Recent applications of the theory to problems in experimental aeroelasticity are reviewed. These results include the identification of aerodynamic impulse responses, the application of higher-order spectra (HOS) to wind-tunnel flutter data, and the identification of nonlinear aeroelastic phenomena from flight flutter test data of the Active Aeroelastic Wing (AAW) aircraft.
NASA Technical Reports Server (NTRS)
Sevart, F. D.
1971-01-01
An analytical and mechanization study was conducted for two flutter stability augmentation systems. One concept uses only the wing trailing edge control surface. Another concept uses leading and trailing edge control surfaces operating simultaneously. The combined use of leading and trailing edge control surfaces should improve the surface coupling (controllability) with vertical bending and torsional structural modes and decrease the coupling between bending and torsional modes. The study was directed toward stability augmentation systems characteristics for the supersonic transport aircraft.
The Wing-Body Aeroelastic Analyses Using the Inverse Design Method
NASA Astrophysics Data System (ADS)
Lee, Seung Jun; Im, Dong-Kyun; Lee, In; Kwon, Jang-Hyuk
Flutter phenomenon is one of the most dangerous problems in aeroelasticity. When it occurs, the aircraft structure can fail in a few second. In recent aeroelastic research, computational fluid dynamics (CFD) techniques become important means to predict the aeroelastic unstable responses accurately. Among various flow equations like Navier-Stokes, Euler, full potential and so forth, the transonic small disturbance (TSD) theory is widely recognized as one of the most efficient theories. However, the small disturbance assumption limits the applicable range of the TSD theory to the thin wings. For a missile which usually has small aspect ratio wings, the influence of body aerodynamics on the wing surface may be significant. Thus, the flutter stability including the body effect should be verified. In this research an inverse design method is used to complement the aerodynamic deficiency derived from the fuselage. MGM (modified Garabedian-McFadden) inverse design method is used to optimize the aerodynamic field of a full aircraft model. Furthermore, the present TSD aeroelastic analyses do not require the grid regeneration process. The MGM inverse design method converges faster than other conventional aerodynamic theories. Consequently, the inverse designed aeroelastic analyses show that the flutter stability has been lowered by the body effect.
NASA Technical Reports Server (NTRS)
Acree, C. W., Jr.
1993-01-01
In pursuit of higher performance, the XV-15 Tiltrotor Research Aircraft was modified by the installation of new composite rotor blades. Initial flights with the Advanced Technology Blades (ATB's) revealed excessive rotor control loads that were traced to a dynamic mismatch between the blades and the aircraft control system. The analytical models of both the blades and the mechanical controls were extensively revised for use by the CAMRAD computer program to better predict aeroelastic stability and loads. This report documents the most important revisions and discusses their effects on aeroelastic stability predictions for airplane-mode flight. The ATB's may be flown in several different configurations for research, including changes in blade sweep and tip twist. The effects on stability of 1 deg and 0 deg sweep are illustrated, as are those of twisted and zero-twist tips. This report also discusses the effects of stiffening the rotor control system, which was done by locking out lateral cyclic swashplate motion with shims.
NASA Technical Reports Server (NTRS)
Sevart, F. D.; Patel, S. M.; Wattman, W. J.
1972-01-01
Testing and evaluation of stability augmentation systems for aircraft flight control were conducted. The flutter suppression system analysis of a scale supersonic transport wing model is described. Mechanization of the flutter suppression system is reported. The ride control synthesis for the B-52 aeroelastic model is discussed. Model analyses were conducted using equations of motion generated from generalized mass and stiffness data.
NASA Technical Reports Server (NTRS)
Bland, Samuel R.; Bennett, Robert M.
1963-01-01
Results of an experimental investigation of propeller whirl flutter are presented for a model consisting of an isolated, rigid system of propeller and simulated power plant mounted with flexibility in pitch and yaw on a rigid sting. A range of propeller blade angles, restraint stiffnesses, and restraint damping coefficients was investigated for a system symmetrical i n pitch and yaw with a windmilling propeller. Measurements of the static-stability derivatives were also made by using a simple balance and were compared with two sets of theoretical derivatives. Whirl-flutter calculations were made with the theoretical and measured derivatives. Some limited results were obtained for the whirl flutter of the model mounted on a cantilever semispan wing. The measured whirl-flutter speeds and frequencies of the isolated model were in very good agreement with those predicted by calculations in which measured derivatives and viscous damping were used. This agreement was better than that obtained by using structural damping. Predicted whirl-flutter speeds for the isolated model were lower when theoretical stability derivatives were used than when measured derivatives were used. The theoretical and experimental static-stability derivatives exhibited the same trends, but in certain instances differed appreciably in magnitude. the measured whirl-flutter boundary for the one configuration considered.
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.
Rotation in vibration, optimization, and aeroelastic stability problems. Ph.D. Thesis
NASA Technical Reports Server (NTRS)
Kaza, K. R. V.
1974-01-01
The effects of rotation in the areas of vibrations, dynamic stability, optimization, and aeroelasticity were studied. The governing equations of motion for the study of vibration and dynamic stability of a rapidly rotating deformable body were developed starting from the nonlinear theory of elasticity. Some common features such as the limitations of the classical theory of elasticity, the choice of axis system, the property of self-adjointness, the phenomenon of frequency splitting, shortcomings of stability methods as applied to gyroscopic systems, and the effect of internal and external damping on stability in gyroscopic systems are identified and discussed, and are then applied to three specific problems.
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.
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.
Rotor Design Options for Improving XV-15 Whirl-Flutter Stability Margins
NASA Technical Reports Server (NTRS)
Acree, C. W., Jr.; Peyran, R. J.; Johnson, Wayne
2004-01-01
Rotor design changes intended to improve tiltrotor whirl-flutter stability margins were analyzed. A baseline analytical model of the XV-15 was established, and then a thinner, composite wing was designed to be representative of a high-speed tiltrotor. The rotor blade design was modified to increase the stability speed margin for the thin-wing design. Small rearward offsets of the aerodynamic-center locus with respect to the blade elastic axis created large increases in the stability boundary. The effect was strongest for offsets at the outboard part of the blade, where an offset of the aerodynamic center by 10% of tip chord improved the stability margin by over 100 knots. Forward offsets of the blade center of gravity had similar but less pronounced effects. Equivalent results were seen for swept-tip blades. Appropriate combinations of sweep and pitch stiffness completely eliminated whirl flutter within the speed range examined; alternatively, they allowed large increases in pitch-flap coupling (delta-three) for a given stability margin. A limited investigation of the rotor loads in helicopter and airplane configuration showed only minor increases in loads.
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.
NASA Technical Reports Server (NTRS)
Ramsey, John K.
1989-01-01
An engineering approach was used to include the nonlinear effects of thickness and camber in an analytical aeroelastic analysis of cascades in supersonic acial flow (supersonic leading-edge locus). A hybrid code using Lighthill's nonlinear piston theory and Lanes's linear potential theory was developed to include these nonlinear effects. Lighthill's theory was used to calculate the unsteady pressures on the noninterference surface regions of the airfoils in cascade. Lane's theory was used to calculate the unsteady pressures on the remaining interference surface regions. Two airfoil profiles was investigated (a supersonic throughflow fan design and a NACA 66-206 airfoil with a sharp leading edge). Results show that compared with predictions of Lane's potential theory for flat plates, the inclusion of thickness (with or without camber) may increase or decrease the aeroelastic stability, depending on the airfoil geometry and operating conditions. When thickness effects are included in the aeroelastic analysis, inclusion of camber will influence the predicted stability in proportion to the magnitude of the added camber. The critical interblade phase angle, depending on the airfoil profile and operating conditions, may also be influenced by thickness and camber. Compared with predictions of Lane's linear potential theory, the inclusion of thickness and camber decreased the aerodynamic stifness and increased the aerodynamic damping at Mach 2 and 2.95 for a cascade of supersonic throughflow fan airfoils oscillating 180 degrees out of phase at a reduced frequency of 0.1.
NASA Technical Reports Server (NTRS)
Schuster, David M.; Edwards, John W.
2004-01-01
The motivation behind the inclusion of unsteady aerodynamics and aeroelastic effects in the computation of stability and control (S&C) derivatives will be discussed as they pertain to aeroelastic and aeroservoelastic analysis. This topic will be addressed in the context of two applications, the first being the estimation of S&C derivatives for a cable-mounted aeroservoelastic wind tunnel model tested in the NASA Langley Research Center (LaRC) Transonic Dynamics Tunnel (TDT). The second application will be the prediction of the nonlinear aeroservoelastic phenomenon known as Residual Pitch Oscillation (RPO) on the B-2 Bomber. Techniques and strategies used in these applications to compute S&C derivatives and perform flight simulations will be reviewed, and computational results will be presented.
NASA Astrophysics Data System (ADS)
Fathi Jegarkandi, M.; Nobari, A. S.; Sabzehparvar, M.; Haddadpour, H.
2009-08-01
Aeroelastic stability of a flexible supersonic flight vehicle is considered using nonlinear dynamics, nonlinear aerodynamics, and a linear structural model. Response surfaces including global multivariate orthogonal modeling functions are invoked to derive applied nonlinear aerodynamic coefficients. A modified Gram-Schmidt method is utilized to orthogonalize the produced polynomial multivariate functions, selected and ranked by predicted squared error metric. Local variation of angle-of-attack and side-slip angle is applied to the analytical model. Identification of nonlinear aerodynamic coefficients of the flight vehicle is conducted employing a CFD code and the required analytical model for simulation purposes is constructed. The method is used to determine the aeroelastic instability and response of a selected flight vehicle.
Nonlinear effects in transonic flutter with emphasis on manifestations of limit cycle oscillations
NASA Astrophysics Data System (ADS)
Schewe, G.; Mai, H.; Dietz, G.
2003-08-01
This paper presents flutter and forced oscillation experiments in a transonic wind tunnel. For an aeroelastic supercritical 2-D airfoil configuration we studied typical transonic phenomena in as pure a form as possible. Various manifestations of small-amplitude limit cycle oscillations were observed for different flow conditions as well as coexisting limit cycles. We demonstrated how very small control forces were sufficient to excite or suppress flutter oscillations. Limit cycle oscillations occurred under free and forced turbulent boundary layer transition in a perforated wall test-section. Flutter calculations based on experimental aerodynamic forces yield stability limits which show good agreement with directly measured experimental flutter values. The results indicate that flow separation at the trailing edge, and the interactions between the shock and the marginal region of separated flow beneath it, may be responsible for limiting the amplitude of the observed limit cycle oscillations.
NASA Technical Reports Server (NTRS)
Srivastava, R.; Reddy, T. S. R.
1996-01-01
This guide describes the input data required, for steady or unsteady aerodynamic and aeroelastic analysis of propellers and the output files generated, in using PROP3D. The aerodynamic forces are obtained by solving three dimensional unsteady, compressible Euler equations. A normal mode structural analysis is used to obtain the aeroelastic equations, which are solved using either time domain or frequency domain solution method. Sample input and output files are included in this guide for steady aerodynamic analysis of single and counter-rotation propellers, and aeroelastic analysis of single-rotation propeller.
Aeroelastic Tailoring of a Plate Wing with Functionally Graded Materials
NASA Technical Reports Server (NTRS)
Dunning, Peter D.; Stanford, Bret K.; Kim, H. Alicia; Jutte, Christine V.
2014-01-01
This work explores the use of functionally graded materials for the aeroelastic tailoring of a metallic cantilevered plate-like wing. Pareto trade-off curves between dynamic stability (flutter) and static aeroelastic stresses are obtained for a variety of grading strategies. A key comparison is between the effectiveness of material grading, geometric grading (i.e., plate thickness variations), and using both simultaneously. The introduction of material grading does, in some cases, improve the aeroelastic performance. This improvement, and the physical mechanism upon which it is based, depends on numerous factors: the two sets of metallic material parameters used for grading, the sweep of the plate, the aspect ratio of the plate, and whether the material is graded continuously or discretely.
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.
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.
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.
Flutter analysis of a supersonic cascade in time domain using an ADI Euler solver
NASA Technical Reports Server (NTRS)
Reddy, T. S. R.; Bakhle, M. A.; Huff, D. L.
1992-01-01
The aeroelastic stability of a two-dimensional cascade oscillating in supersonic axial flow is analyzed in the time domain. The aeroelastic model consists of a single degree of freedom typical section structural model for each blade of the cascade and an unsteady two-dimensional cascade aerodynamic model based on the Euler equations. The Euler equations are solved using a time accurate Alternating Direction Implicit (ADI) solution scheme. The aeroelastic equations are integrated in time. The effect of interblade phase angle is included in the aeroelastic analysis by an appropriate choice of initial and boundary conditions. Flutter predictions are obtained from the time response of a flat plate cascade in single degree of freedom pitching motion. The results correlate well with those obtained from a separate frequency domain flutter analysis for all values of interblade phase angles considered. Flutter results are then presented for cascades having airfoil sections representative of a supersonic throughflow fan. The validity of the time integration method for a cascade of airfoils at various interblade phase angles is demonstrated.
Probabilistic Aeroelastic Analysis Developed for Turbomachinery Components
NASA Technical Reports Server (NTRS)
Reddy, T. S. R.; Mital, Subodh K.; Stefko, George L.; Pai, Shantaram S.
2003-01-01
Aeroelastic analyses for advanced turbomachines are being developed for use at the NASA Glenn Research Center and industry. However, these analyses at present are used for turbomachinery design with uncertainties accounted for by using safety factors. This approach may lead to overly conservative designs, thereby reducing the potential of designing higher efficiency engines. An integration of the deterministic aeroelastic analysis methods with probabilistic analysis methods offers the potential to design efficient engines with fewer aeroelastic problems and to make a quantum leap toward designing safe reliable engines. In this research, probabilistic analysis is integrated with aeroelastic analysis: (1) to determine the parameters that most affect the aeroelastic characteristics (forced response and stability) of a turbomachine component such as a fan, compressor, or turbine and (2) to give the acceptable standard deviation on the design parameters for an aeroelastically stable system. The approach taken is to combine the aeroelastic analysis of the MISER (MIStuned Engine Response) code with the FPI (fast probability integration) code. The role of MISER is to provide the functional relationships that tie the structural and aerodynamic parameters (the primitive variables) to the forced response amplitudes and stability eigenvalues (the response properties). The role of FPI is to perform probabilistic analyses by utilizing the response properties generated by MISER. The results are a probability density function for the response properties. The probabilistic sensitivities of the response variables to uncertainty in primitive variables are obtained as a byproduct of the FPI technique. The combined analysis of aeroelastic and probabilistic analysis is applied to a 12-bladed cascade vibrating in bending and torsion. Out of the total 11 design parameters, 6 are considered as having probabilistic variation. The six parameters are space-to-chord ratio (SBYC), stagger angle
Aeroelastic stability and response of horizontal axis wind turbine blades
NASA Technical Reports Server (NTRS)
Kottapalli, S. B. R.; Friedmann, P. P.; Rosen, A.
1979-01-01
Coupled flap-lag-torsion equations of motion of an isolated horizontal axis wind turbine (HAWT) blade have been formulated. The analysis neglects blade-tower coupling. The final nonlinear equations have periodic coefficients. A new and convenient method of generating an appropriate time-dependent equilibrium position, required for the stability analysis, has been implemented and found to be computationally efficient. Steady-state response and stability boundaries for an existing (typical) HAWT blade are presented. Such stability boundaries have never been published in the literature. The results show that the isolated blade under study is basically stable. The tower shadow (wake) has a considerable effect on the out-of-plane response but leaves blade stability unchanged. Nonlinear terms can significantly affect linearized stability boundaries; however, they have a negligible effect on response, thus implying that a time-dependent equilibrium position (or steady-state response), based completely on the linear system, is appropriate for the type of HAWT blades under study.
A comparison of theory and experiment for aeroelastic stability of a hingeless rotor model in hover
NASA Technical Reports Server (NTRS)
Sharpe, David L.
1988-01-01
Theoretical predictions of aeroelastic stability are compared with experimental, isolated, hingeless-rotor data. The six cases selected represent a torsionally soft rotor having either a stiff or soft pitch-control system in combination with zero precone and droop, 5 degree precone, or -5 degree droop. Analyses from Bell Helicopter Textron, Boeing Vertol, Hughes Helicopters, Sikorsky Aircraft, the National Aeronautics and Space Administration, and the U.S. Army Aeromechanics Laboratory were compared with the experimental data. The correlation ranged from poor to fair.
NASA Astrophysics Data System (ADS)
Knipfer, A.
1999-02-01
The nacelles of modern aeroengines are constantly increasing in size. Thus, engine air-loads are becoming more powerful and their importance for the aeroelastic stability is becoming more significant. The principal goal of this study is to answer the question of how unsteady airloads vary while shifting to transonic Mach numbers. The investigations are carried out by applying a finite volume Euler method to a harmonically oscillating annular wing. The results show that transonic effects in the case of an annular wing are essentially weaker than in the case of an airfoil. The order of magnitude of the variations is around 10%. Possible consequences for the aeroelastic stability are examined with the example of an elastically mounted annular wing in transonic flow. The shifts of the stability curves also remain within a range of 10%. In addition, an actuator disk method, which is frequently used for the simulation of the fan jet, is expanded in such a way that unsteady flows can be treated. Some unsteady air-loads are strongly dependent on the pressure jump across the fan.
The influence of trailed vorticity on flutter speed estimations
NASA Astrophysics Data System (ADS)
Pirrung, Georg R.; Madsen, Helge Aa; Kim, Taeseong
2014-06-01
This paper briefly describes the implementation of a coupled near and far wake model for wind turbine rotor induction in the aeroelastic code HAWC2 and its application for flutter analysis of the NREL 5 MW wind turbine. The model consists of a far wake part based on Blade Element Momentum (BEM) theory, which is coupled with Beddoes' near wake model for trailed vorticity. The first part of this work outlines the implementation in HAWC2, with a focus on the interaction of the induction from the blade based near wake model with the induction from the polar grid based BEM model in HAWC2. The influence of the near wake model on the aeroelastic stability of the blades of the NREL 5 MW turbine in overspeed conditions is investigated in the second part of the paper. The analysis is based on a runaway case in which the turbine is free to speed up without generator torque and vibrations start building up at a critical rotor speed. Blades with modified torsional and flapwise stiffness are also investigated. A flutter analysis is often part of the stability investigations for new blades but is normally carried out with engineering models that do not include the influence of unsteady trailed vorticity. Including this influence results in a slightly increased safety margin against classical flutter in all simulated cases.
Full-scale wind-tunnel test of the aeroelastic stability of a bearingless main rotor
NASA Technical Reports Server (NTRS)
Warmbrodt, W.; Mccloud, J., III; Sheffler, M.; Staley, J.
1981-01-01
The rotor studied in the wind tunnel had previously been flight tested on a BO-105 helicopter. The investigation was conducted to determine the rotor's aeroelastic stability characteristics in hover and at airspeeds up to 143 knots. These characteristics are compared with those obtained from whirl-tower and flight tests and predictions from a digital computer simulation. It was found that the rotor was stable for all conditions tested. At constant tip speed, shaft angle, and airspeed, stability increases with blade collective pitch setting. No significant change in system damping occurred that was attributable to frequency coalescence between the rotor inplane regressing mode and the support modes. Stability levels determined in the wind tunnel were of the same magnitude and yielded the same trends as data obtained from whirl-tower and flight tests.
Predicting Unsteady Aeroelastic Behavior
NASA Technical Reports Server (NTRS)
Strganac, Thomas W.; Mook, Dean T.
1990-01-01
New method for predicting subsonic flutter, static deflections, and aeroelastic divergence developed. Unsteady aerodynamic loads determined by unsteady-vortex-lattice method. Accounts for aspect ratio and angle of attack. Equations for motion of wing and flow field solved iteratively and simultaneously. Used to predict transient responses to initial disturbances, and to predict steady-state static and oscillatory responses. Potential application for research in such unsteady structural/flow interactions as those in windmills, turbines, and compressors.
Field Validation of the Stability Limit of a Multi MW Turbine
NASA Astrophysics Data System (ADS)
Kallesøe, Bjarne S.; Kragh, Knud A.
2016-09-01
Long slender blades of modern multi-megawatt turbines exhibit a flutter like instability at rotor speeds above a critical rotor speed. Knowing the critical rotor speed is crucial to a safe turbine design. The flutter like instability can only be estimated using geometrically non-linear aeroelastic codes. In this study, the estimated rotor speed stability limit of a 7 MW state of the art wind turbine is validated experimentally. The stability limit is estimated using Siemens Wind Powers in-house aeroelastic code, and the results show that the predicted stability limit is within 5% of the experimentally observed limit.
Application of Aeroelastic Solvers Based on Navier Stokes Equations
NASA Technical Reports Server (NTRS)
Keith, Theo G., Jr.; Srivastava, Rakesh
2001-01-01
The propulsion element of the NASA Advanced Subsonic Technology (AST) initiative is directed towards increasing the overall efficiency of current aircraft engines. This effort requires an increase in the efficiency of various components, such as fans, compressors, turbines etc. Improvement in engine efficiency can be accomplished through the use of lighter materials, larger diameter fans and/or higher-pressure ratio compressors. However, each of these has the potential to result in aeroelastic problems such as flutter or forced response. To address the aeroelastic problems, the Structural Dynamics Branch of NASA Glenn has been involved in the development of numerical capabilities for analyzing the aeroelastic stability characteristics and forced response of wide chord fans, multi-stage compressors and turbines. In order to design an engine to safely perform a set of desired tasks, accurate information of the stresses on the blade during the entire cycle of blade motion is required. This requirement in turn demands that accurate knowledge of steady and unsteady blade loading is available. To obtain the steady and unsteady aerodynamic forces for the complex flows around the engine components, for the flow regimes encountered by the rotor, an advanced compressible Navier-Stokes solver is required. A finite volume based Navier-Stokes solver has been developed at Mississippi State University (MSU) for solving the flow field around multistage rotors. The focus of the current research effort, under NASA Cooperative Agreement NCC3- 596 was on developing an aeroelastic analysis code (entitled TURBO-AE) based on the Navier-Stokes solver developed by MSU. The TURBO-AE code has been developed for flutter analysis of turbomachine components and delivered to NASA and its industry partners. The code has been verified. validated and is being applied by NASA Glenn and by aircraft engine manufacturers to analyze the aeroelastic stability characteristics of modem fans, compressors
FLUT - A program for aeroelastic stability analysis. [of aircraft structures in subsonic flow
NASA Technical Reports Server (NTRS)
Johnson, E. H.
1977-01-01
A computer program (FLUT) that can be used to evaluate the aeroelastic stability of aircraft structures in subsonic flow is described. The algorithm synthesizes data from a structural vibration analysis with an unsteady aerodynamics analysis and then performs a complex eigenvalue analysis to assess the system stability. The theoretical basis of the program is discussed with special emphasis placed on some innovative techniques which improve the efficiency of the analysis. User information needed to efficiently and successfully utilize the program is provided. In addition to identifying the required input, the flow of the program execution and some possible sources of difficulty are included. The use of the program is demonstrated with a listing of the input and output for a simple example.
Ongoing Fixed Wing Research within the NASA Langley Aeroelasticity Branch
NASA Technical Reports Server (NTRS)
Bartels, Robert; Chwalowski, Pawel; Funk, Christie; Heeg, Jennifer; Hur, Jiyoung; Sanetrik, Mark; Scott, Robert; Silva, Walter; Stanford, Bret; Wiseman, Carol
2015-01-01
The NASA Langley Aeroelasticity Branch is involved in a number of research programs related to fixed wing aeroelasticity and aeroservoelasticity. These ongoing efforts are summarized here, and include aeroelastic tailoring of subsonic transport wing structures, experimental and numerical assessment of truss-braced wing flutter and limit cycle oscillations, and numerical modeling of high speed civil transport configurations. Efforts devoted to verification, validation, and uncertainty quantification of aeroelastic physics in a workshop setting are also discussed. The feasibility of certain future civil transport configurations will depend on the ability to understand and control complex aeroelastic phenomena, a goal that the Aeroelasticity Branch is well-positioned to contribute through these programs.
NASA Technical Reports Server (NTRS)
Perry, Boyd, III; Mukhopadhyay, Vivek; Hoadley, Sherwood Tiffany; 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 projet, and the development and successful use of a methodology for online controller performance evaluation.
Effects of leading-edge tubercles on wing flutter speeds.
Ng, B F; New, T H; Palacios, R
2016-04-12
The dynamic aeroelastic effects on wings modified with bio-inspired leading-edge (LE) tubercles are examined in this study. We adopt a state-space aeroelastic model via the coupling of unsteady vortex-lattice method and a composite beam to evaluate stability margins as a result of LE tubercles on a generic wing. The unsteady aerodynamics and spanwise mass variations due to LE tubercles have counteracting effects on stability margins with the former having dominant influence. When coupled, flutter speed is observed to be 5% higher, and this is accompanied by close to 6% decrease in reduced frequencies as an indication of lower structural stiffness requirements for wings with LE tubercles. Both tubercle amplitude and wavelength have similar influences over the change in flutter speeds, and such modifications to the LE would have minimal effect on stability margins when concentrated inboard of the wing. Lastly, when used in sweptback wings, LE tubercles are observed to have smaller impacts on stability margins as the sweep angle is increased.
Shaft flexibility effects on aeroelastic stability of a rotating bladed disk
NASA Technical Reports Server (NTRS)
Khader, Naim; Loewy, Robert
1989-01-01
A comprehensive study of Coriolis forces and shaft flexibility effects on the structural dynamics and aeroelastic stability of a rotating bladed-disk assembly attached to a cantilever, massless, flexible shaft is presented. Analyses were performed for an actual bladed-disk assembly, used as the first stage in the fan of the 'E3' engine. In the structural model, both in-plane and out-of-plane elastic deformation of the bladed-disk assembly were considered relative to their hub, in addition to rigid disk translations and rotations introduced by shaft flexibility. Besides structural coupling between blades (through the flexible disk), additional coupling is introduced through quasisteady aerodynamic loads. Rotational effects are accounted for throughout the work, and some mode shapes for the whole structure are presented at a selected rpm.
NASA Technical Reports Server (NTRS)
Reding, J. P.; Ericsson, L. E.
1976-01-01
A quasi-steady analysis of the aeroelastic stability of the lateral (antisymmetric) modes of the 747/orbiter vehicle was accomplished. The interference effect of the orbiter wake on the 747 tail furnishes an aerodynamic undamping contribution to the elastic modes. Likewise, the upstream influence of the 747 tail and aft fuselage on the orbiter beaver-tail rail fairing also is undamping. Fortunately these undamping effects cannot overpower the large damping contribution of the 747 tail and the modes are damped for the configurations analyzed. However, significant interference effects of the orbiter on the 747 tail have been observed in the pitch plane. The high response of the 747 vertical tail in the orbiter wave was also considered. Wind tunnel data points to flapping of the OMS pod wakes as the source of the wake resonance phenomenon.
NASA Technical Reports Server (NTRS)
Goldman, Benjamin D.; Dowell, Earl H.; Scott, Robert C.
2015-01-01
Conical shell theory and a supersonic potential flow aerodynamic theory are used to study the nonlinear pressure buckling and aeroelastic limit cycle behavior of the thermal protection system for NASA's Hypersonic Inflatable Aerodynamic Decelerator. The structural model of the thermal protection system consists of an orthotropic conical shell of the Donnell type, resting on several circumferential elastic supports. Classical Piston Theory is used initially for the aerodynamic pressure, but was found to be insufficient at low supersonic Mach numbers. Transform methods are applied to the convected wave equation for potential flow, and a time-dependent aerodynamic pressure correction factor is obtained. The Lagrangian of the shell system is formulated in terms of the generalized coordinates for all displacements and the Rayleigh-Ritz method is used to derive the governing differential-algebraic equations of motion. Aeroelastic limit cycle oscillations and buckling deformations are calculated in the time domain using a Runge-Kutta method in MATLAB. Three conical shell geometries were considered in the present analysis: a 3-meter diameter 70 deg. cone, a 3.7-meter 70 deg. cone, and a 6-meter diameter 70 deg. cone. The 6-meter configuration was loaded statically and the results were compared with an experimental load test of a 6-meter HIAD. Though agreement between theoretical and experimental strains was poor, the circumferential wrinkling phenomena observed during the experiments was captured by the theory and axial deformations were qualitatively similar in shape. With Piston Theory aerodynamics, the nonlinear flutter dynamic pressures of the 3-meter configuration were in agreement with the values calculated using linear theory, and the limit cycle amplitudes were generally on the order of the shell thickness. The effect of axial tension was studied for this configuration, and increasing tension was found to decrease the limit cycle amplitudes when the circumferential
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.
NASA Technical Reports Server (NTRS)
Lind, Richard C. (Inventor); Brenner, Martin J.
2001-01-01
A structured singular value (mu) analysis method of computing flutter margins has robust stability of a linear aeroelastic model with uncertainty operators (Delta). Flight data is used to update the uncertainty operators to accurately account for errors in the computed model and the observed range of aircraft dynamics of the aircraft under test caused by time-varying aircraft parameters, nonlinearities, and flight anomalies, such as test nonrepeatability. This mu-based approach computes predict flutter margins that are worst case with respect to the modeling uncertainty for use in determining when the aircraft is approaching a flutter condition and defining an expanded safe flight envelope for the aircraft that is accepted with more confidence than traditional methods that do not update the analysis algorithm with flight data by introducing mu as a flutter margin parameter that presents several advantages over tracking damping trends as a measure of a tendency to instability from available flight data.
Aeroelastic loads and stability investigation of a full-scale hingeless rotor
NASA Technical Reports Server (NTRS)
Peterson, Randall L.; Johnson, Wayne
1991-01-01
An analytical investigation was conducted to study the influence of various parameters on predicting the aeroelastic loads and stability of a full-scale hingeless rotor in hover and forward flight. The CAMRAD/JA (Comprehensive Analytical Model of Rotorcraft Aerodynamics and Dynamics, Johnson Aeronautics) analysis code is used to obtain the analytical predictions. Data are presented for rotor blade bending and torsional moments as well as inplane damping data obtained for rotor operation in hover at a constant rotor rotational speed of 425 rpm and thrust coefficients between 0.0 and 0.12. Experimental data are presented from a test in the wind tunnel. Validation of the rotor system structural model with experimental rotor blade loads data shows excellent correlation with analytical results. Using this analysis, the influence of different aerodynamic inflow models, the number of generalized blade and body degrees of freedom, and the control-system stiffness at predicted stability levels are shown. Forward flight predictions of the BO-105 rotor system for 1-G thrust conditions at advance ratios of 0.0 to 0.35 are presented. The influence of different aerodynamic inflow models, dynamic inflow models and shaft angle variations on predicted stability levels are shown as a function of advance ratio.
Three-dimensional nonlinear flutter analysis of long-span suspension bridges during erection.
Zhang, Xin-jun; Sun, Bing-nan; Xiang, Hai-fan
2003-01-01
In this work, the aerodynamic stability of the Yichang Suspension Bridge over Yangtze River during erection was determined by three-dimensional nonlinear flutter analysis, in which the nonlinearities of structural dynamic characteristics and aeroelastic forces caused by large deformation are fully considered. An interesting result obtained was that the bridge was more stable when the stiffening girders were erected in a non-symmetrical manner as opposed to the traditional symmetrical erection schedule. It was also found that the severe decrease in the aerodynamic stability was due to the nonlinear effects. Therefore, the nonlinear factors should be considered accurately in aerodynamic stability analysis of long-span suspension bridges during erection.
NASA Technical Reports Server (NTRS)
Sevart, F. D.; Patel, S. M.
1973-01-01
Testing and evaluation of a stability augmentation system for aircraft flight control were performed. The flutter suppression system and synthesis conducted on a scale model of a supersonic wing for a transport aircraft are discussed. Mechanization and testing of the leading and trailing edge surface actuation systems are described. The ride control system analyses for a 375,000 pound gross weight B-52E aircraft are presented. Analyses of the B-52E aircraft maneuver load control system are included.
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.
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
A flutter investigation of all-moveable NASP-like wings at hypersonic speeds
NASA Technical Reports Server (NTRS)
Spain, Charles V.; Zeiler, Thomas A.; Bullock, Ellen P.; Hodge, Jeffrey S.
1993-01-01
Six alternative all-moving wing configurations applicable to the NASP hypersonic/transatmospheric vehicle have undergone aeroelasticity testing in NASA-Langley's Mach-20-capable Helium Tunnel that yielded data for such parametric variations as airfoil profile and wing planform, wing-pivot flexure stiffness, and mass imbalance. While all wings fluttered at dynamic pressures lower than predicted by second-order piston-theory aerodynamics, this was of limited amplitude, suggesting nonlinear external-flow behavior. Slab airfoils were more stable than diamond-shaped ones; blunt leading edges enhance stability relative to sharp ones, and stiffer pivolts extert a stabilizing influence.
Cascade flutter analysis with transient response aerodynamics
NASA Technical Reports Server (NTRS)
Bakhle, Milind A.; Mahajan, Aparajit J.; Keith, Theo G., Jr.; Stefko, George L.
1991-01-01
Two methods for calculating linear frequency domain aerodynamic coefficients from a time marching Full Potential cascade solver are developed and verified. In the first method, the Influence Coefficient, solutions to elemental problems are superposed to obtain the solutions for a cascade in which all blades are vibrating with a constant interblade phase angle. The elemental problem consists of a single blade in the cascade oscillating while the other blades remain stationary. In the second method, the Pulse Response, the response to the transient motion of a blade is used to calculate influence coefficients. This is done by calculating the Fourier Transforms of the blade motion and the response. Both methods are validated by comparison with the Harmonic Oscillation method and give accurate results. The aerodynamic coefficients obtained from these methods are used for frequency domain flutter calculations involving a typical section blade structural model. An eigenvalue problem is solved for each interblade phase angle mode and the eigenvalues are used to determine aeroelastic stability. Flutter calculations are performed for two examples over a range of subsonic Mach numbers.
NASA Technical Reports Server (NTRS)
Kvaternik, R. G.
1973-01-01
Aeroelastic and dynamic studies which complement and extend various aspects of technology applicable to tilt-rotor VTOL aircraft are discussed. Particular attention is given to proprotor/pylon whirl instability, a precession-type instability akin to propeller/nacelle whirl flutter. The blade flapping and pitch-change freedoms of a proprotor are shown to lead to a fundamentally different situation as regards the manner in which the precession-generated aerodynamic forces and moments act on the pylon and induce whirl flutter relative to that of a propeller. The implication of these forces and moments with regard to their capacity for instigating a whirl instability is examined, demonstrating why a proprotor can exhibit whirl flutter in either the backward or forward directions in contrast to a propeller which is found to always whirl in the backward direction. Analytical trend studies delineating the effect of several system design parameters on proprotor/pylon stability and response are shown.
NASA Astrophysics Data System (ADS)
Silva, Walter A.; Chwalowski, Pawel; Perry, Boyd, III
2014-03-01
Reduced-order modelling (ROM) methods are applied to the Computational Fluid Dynamics (CFD)-based aeroelastic analysis of the AGARD 445.6 wing in order to gain insight regarding well-known discrepancies between the aeroelastic analyses and the experimental results. The results presented include aeroelastic solutions using the inviscid Computational Aeroelasticity Programme-Transonic Small Disturbance (CAP-TSD) code and the FUN3D code (Euler and Navier-Stokes). Full CFD aeroelastic solutions and ROM aeroelastic solutions, computed at several Mach numbers, are presented in the form of root locus plots in order to better reveal the aeroelastic root migrations with increasing dynamic pressure. Important conclusions are drawn from these results including the ability of the linear CAP-TSD code to accurately predict the entire experimental flutter boundary (repeat of analyses performed in the 1980s), that the Euler solutions at supersonic conditions indicate that the third mode is always unstable, and that the FUN3D Navier-Stokes solutions stabilize the unstable third mode seen in the Euler solutions.
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.
NASA Technical Reports Server (NTRS)
Reddy, T. S. R.; Srivastava, R.; Mehmed, Oral
2002-01-01
An aeroelastic analysis system for flutter and forced response analysis of turbomachines based on a two-dimensional linearized unsteady Euler solver has been developed. The ASTROP2 code, an aeroelastic stability analysis program for turbomachinery, was used as a basis for this development. The ASTROP2 code uses strip theory to couple a two dimensional aerodynamic model with a three dimensional structural model. The code was modified to include forced response capability. The formulation was also modified to include aeroelastic analysis with mistuning. A linearized unsteady Euler solver, LINFLX2D is added to model the unsteady aerodynamics in ASTROP2. By calculating the unsteady aerodynamic loads using LINFLX2D, it is possible to include the effects of transonic flow on flutter and forced response in the analysis. The stability is inferred from an eigenvalue analysis. The revised code, ASTROP2-LE for ASTROP2 code using Linearized Euler aerodynamics, is validated by comparing the predictions with those obtained using linear unsteady aerodynamic solutions.
Time simulation of flutter with large stiffness changes
NASA Technical Reports Server (NTRS)
Karpel, Mordechay; Wieseman, Carol D.
1992-01-01
Time simulation of flutter, involving large local structural changes, is formulated with a state-space model that is based on a relatively small number of generalized coordinates. Free-free vibration modes are first calculated for a nominal finite-element model with relatively large fictitious masses located at the area of structural changes. A low-frequency subset of these modes is then transformed into a set of structural modal coordinates with which the entire simulation is performed. These generalized coordinates and the associated oscillatory aerodynamic force coefficient matrices are used to construct an efficient time-domain, state-space model for a basic aeroelastic case. The time simulation can then be performed by simply changing the mass, stiffness, and damping coupling terms when structural changes occur. It is shown that the size of the aeroelastic model required for time simulation with large structural changes at a few apriori known locations is similar to that required for direct analysis of a single structural case. The method is applied to the simulation of an aeroelastic wind-tunnel model. The diverging oscillations are followed by the activation of a tip-ballast decoupling mechanism that stabilizes the system but may cause significant transient overshoots.
Aeroelasticity Benchmark Assessment: Subsonic Fixed Wing Program
NASA Technical Reports Server (NTRS)
Florance, Jennifer P.; Chwalowski, Pawel; Wieseman, Carol D.
2010-01-01
The fundamental technical challenge in computational aeroelasticity is the accurate prediction of unsteady aerodynamic phenomena and the effect on the aeroelastic response of a vehicle. Currently, a benchmarking standard for use in validating the accuracy of computational aeroelasticity codes does not exist. Many aeroelastic data sets have been obtained in wind-tunnel and flight testing throughout the world; however, none have been globally presented or accepted as an ideal data set. There are numerous reasons for this. One reason is that often, such aeroelastic data sets focus on the aeroelastic phenomena alone (flutter, for example) and do not contain associated information such as unsteady pressures and time-correlated structural dynamic deflections. Other available data sets focus solely on the unsteady pressures and do not address the aeroelastic phenomena. Other discrepancies can include omission of relevant data, such as flutter frequency and / or the acquisition of only qualitative deflection data. In addition to these content deficiencies, all of the available data sets present both experimental and computational technical challenges. Experimental issues include facility influences, nonlinearities beyond those being modeled, and data processing. From the computational perspective, technical challenges include modeling geometric complexities, coupling between the flow and the structure, grid issues, and boundary conditions. The Aeroelasticity Benchmark Assessment task seeks to examine the existing potential experimental data sets and ultimately choose the one that is viewed as the most suitable for computational benchmarking. An initial computational evaluation of that configuration will then be performed using the Langley-developed computational fluid dynamics (CFD) software FUN3D1 as part of its code validation process. In addition to the benchmarking activity, this task also includes an examination of future research directions. Researchers within the
NASA Technical Reports Server (NTRS)
Kottapalli, Sesi; Meza, Victor
1992-01-01
A rotorcraft analysis is conducted to assess tilt-rotor stability and conversion loads for the XV-15 rotor with metal blades within its specified test envelope. A 38-DOF flutter analysis based on the code by Johnson (1988) is developed to simulate a wind-tunnel test in which the rotor torque is constant and thereby study stability. The same analytical model provides the simulated loads including hub loads, blade loads, and oscillatory pitch-link loads with attention given to the nonuniform inflow through the proprotor in the presence of the wing. Tilt-rotor stability during the cruise mode is found to be sensitive to coupling effects in the control system stiffness, and a stability problem is identified in the XV-15 Advanced Technology Blades. The present analysis demonstrates that the tilt-rotor is stable within the specified test envelope of the NASA 40 x 80-ft wind tunnel.
Multi-fidelity construction of explicit boundaries: Application to aeroelasticity
NASA Astrophysics Data System (ADS)
Dribusch, Christoph
Wings, control surfaces and rotor blades subject to aerodynamic forces may exhibit aeroelastic instabilities such as flutter, divergence and limit cycle oscillations which generally reduce their life and functionality. This possibility of instability must be taken into account during the design process and numerical simulation models may be used to predict aeroelastic stability. Aeroelastic stability is a design requirement that encompasses several difficulties also found in other areas of design. For instance, the large computational time associated with stability analysis is also found in computational fluid dynamics (CFD) models. It is a major hurdle in numerical optimization and reliability analysis, which generally require large numbers of call to the simulation code. Similarly, the presence of bifurcations and discontinuities is also encountered in structural impact analysis based on nonlinear dynamic simulations and renders traditional approximation techniques such as Kriging ineffective. Finally, for a given component or system, aeroelastic instability is only one of multiple failure modes which must be accounted for during design and reliability studies. To address the above challenges, this dissertation proposes a novel algorithm to predict, over a range of parameters, the qualitative outcomes (pass/fail) of simulations based on relatively few, classified (pass/fail) simulation results. This is different from traditional approximation techniques that seek to predict simulation outcomes quantitatively, for example by fitting a response surface. The predictions of the proposed algorithm are based on the theory of support vector machines (SVM), a machine learning method originated in the field of pattern recognition. This process yields an analytical function that explicitly defines the boundary between feasible and infeasible regions of the parameter space and has the ability to reproduce nonlinear, disjoint boundaries in n dimensions. Since training the
NASA Technical Reports Server (NTRS)
Heeg, Jennifer; Gilbert, Michael G.; Pototzky, Anthony S.
1990-01-01
This work-in-progress presentation describes an ongoing research activity at the NASA Langley Research Center to develop analytical methods for the prediction of aerothermoelastic stability of hypersonic aircraft including active control systems. The objectives of this research include application of aerothermal loads to the structural finite element model, determination of the thermal effects on flutter, and assessment of active controls technology applied to overcome any potential adverse aeroelastic stability or response problems due to aerodynamic heating- namely flutter suppression and ride quality improvement. For this study, a generic hypersonic aircraft configuration was selected which incorporates wing flaps, ailerons and all-moveable fins to be used for active control purposes. The active control systems would use onboard sensors in a feedback loop through the aircraft flight control computers to move the surfaces for improved structural dynamic response as the aircraft encounters atmospheric turbulence.
An Aeroelastic Analysis of a Thin Flexible Membrane
NASA Technical Reports Server (NTRS)
Scott, Robert C.; Bartels, Robert E.; Kandil, Osama A.
2007-01-01
Studies have shown that significant vehicle mass and cost savings are possible with the use of ballutes for aero-capture. Through NASA's In-Space Propulsion program, a preliminary examination of ballute sensitivity to geometry and Reynolds number was conducted, and a single-pass coupling between an aero code and a finite element solver was used to assess the static aeroelastic effects. There remain, however, a variety of open questions regarding the dynamic aeroelastic stability of membrane structures for aero-capture, with the primary challenge being the prediction of the membrane flutter onset. The purpose of this paper is to describe and begin addressing these issues. The paper includes a review of the literature associated with the structural analysis of membranes and membrane utter. Flow/structure analysis coupling and hypersonic flow solver options are also discussed. An approach is proposed for tackling this problem that starts with a relatively simple geometry and develops and evaluates analysis methods and procedures. This preliminary study considers a computationally manageable 2-dimensional problem. The membrane structural models used in the paper include a nonlinear finite-difference model for static and dynamic analysis and a NASTRAN finite element membrane model for nonlinear static and linear normal modes analysis. Both structural models are coupled with a structured compressible flow solver for static aeroelastic analysis. For dynamic aeroelastic analyses, the NASTRAN normal modes are used in the structured compressible flow solver and 3rd order piston theories were used with the finite difference membrane model to simulate utter onset. Results from the various static and dynamic aeroelastic analyses are compared.
NASA Technical Reports Server (NTRS)
Pak, Chan-gi; Lung, Shun-fat
2009-01-01
Modern airplane design is a multidisciplinary task which combines several disciplines such as structures, aerodynamics, flight controls, and sometimes heat transfer. Historically, analytical and experimental investigations concerning the interaction of the elastic airframe with aerodynamic and in retia loads have been conducted during the design phase to determine the existence of aeroelastic instabilities, so called flutter .With the advent and increased usage of flight control systems, there is also a likelihood of instabilities caused by the interaction of the flight control system and the aeroelastic response of the airplane, known as aeroservoelastic instabilities. An in -house code MPASES (Ref. 1), modified from PASES (Ref. 2), is a general purpose digital computer program for the analysis of the closed-loop stability problem. This program used subroutines given in the International Mathematical and Statistical Library (IMSL) (Ref. 3) to compute all of the real and/or complex conjugate pairs of eigenvalues of the Hessenberg matrix. For high fidelity configuration, these aeroelastic system matrices are large and compute all eigenvalues will be time consuming. A subspace iteration method (Ref. 4) for complex eigenvalues problems with nonsymmetric matrices has been formulated and incorporated into the modified program for aeroservoelastic stability (MPASES code). Subspace iteration method only solve for the lowest p eigenvalues and corresponding eigenvectors for aeroelastic and aeroservoelastic analysis. In general, the selection of p is ranging from 10 for wing flutter analysis to 50 for an entire aircraft flutter analysis. The application of this newly incorporated code is an experiment known as the Aerostructures Test Wing (ATW) which was designed by the National Aeronautic and Space Administration (NASA) Dryden Flight Research Center, Edwards, California to research aeroelastic instabilities. Specifically, this experiment was used to study an instability
Structural resonance and mode of flutter of hummingbird tail feathers.
Clark, Christopher J; Elias, Damian O; Girard, Madeline B; Prum, Richard O
2013-09-15
Feathers can produce sound by fluttering in airflow. This flutter is hypothesized to be aeroelastic, arising from the coupling of aerodynamic forces to one or more of the feather's intrinsic structural resonance frequencies. We investigated how mode of flutter varied among a sample of hummingbird tail feathers tested in a wind tunnel. Feather vibration was measured directly at ~100 points across the surface of the feather with a scanning laser Doppler vibrometer (SLDV), as a function of airspeed, Uair. Most feathers exhibited multiple discrete modes of flutter, which we classified into types including tip, trailing vane and torsional modes. Vibratory behavior within a given mode was usually stable, but changes in independent variables such as airspeed or orientation sometimes caused feathers to abruptly 'jump' from one mode to another. We measured structural resonance frequencies and mode shapes directly by measuring the free response of 64 feathers stimulated with a shaker and recorded with the SLDV. As predicted by the aeroelastic flutter hypothesis, the mode shape (spatial distribution) of flutter corresponded to a bending or torsional structural resonance frequency of the feather. However, the match between structural resonance mode and flutter mode was better for tip or torsional mode shapes, and poorer for trailing vane modes. Often, the 3rd bending structural harmonic matched the expressed mode of flutter, rather than the fundamental. We conclude that flutter occurs when airflow excites one or more structural resonance frequencies of a feather, most akin to a vibrating violin string.
Improved Flight Test Procedures for Flutter Clearance
NASA Technical Reports Server (NTRS)
Lind, Rick C.; Brenner, Martin J.; Freudinger, Lawrence C.
1997-01-01
Flight flutter testing is an integral part of flight envelope clearance. This paper discusses advancements in several areas that are being investigated to improve efficiency and safety of flight test programs. Results are presented from recent flight testing of the F/A-18 Systems Research Aircraft. A wingtip excitation system was used to generate aeroelastic response data. This system worked well for many flight conditions but still displayed some anomalies. Wavelet processing is used to analyze the flight data. Filtered transfer functions are generated that greatly improve system identification. A flutter margin is formulated that accounts for errors between a model and flight data. Worst-case flutter margins are computed to demonstrate the flutter boundary may lie closer to the flight envelope than previously estimated. This paper concludes with developments for a distributed flight analysis environment and on-line health monitoring.
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.
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.
NASA Technical Reports Server (NTRS)
Kvaternik, Raymond G.; Piatak, David J.; Nixon, Mark W.; Langston, Chester W.; Singleton, Jeffrey D.; Bennett, Richard L.; Brown, Ross K.
2001-01-01
The results of a joint NASA/Army/Bell Helicopter Textron wind-tunnel test to assess the potential of Generalized Predictive Control (GPC) for actively controlling the swashplate of tiltrotor aircraft to enhance aeroelastic stability in the airplane mode of flight are presented. GPC is an adaptive time-domain predictive control method that uses a linear difference equation to describe the input-output relationship of the system and to design the controller. The test was conducted in the Langley Transonic Dynamics Tunnel using an unpowered 1/5-scale semispan aeroelastic model of the V-22 that was modified to incorporate a GPC-based multi-input multi-output control algorithm to individually control each of the three swashplate actuators. Wing responses were used for feedback. The GPC-based control system was highly effective in increasing the stability of the critical wing mode for all of the conditions tested, without measurable degradation of the damping in the other modes. The algorithm was also robust with respect to its performance in adjusting to rapid changes in both the rotor speed and the tunnel airspeed.
A Simple Analytical Approach to the Aeroelastic Stability Problem of Long-Span Cable-Stayed Bridges
NASA Astrophysics Data System (ADS)
Vairo, Giuseppe
2010-01-01
This paper deals with the aeroelastic stability problem of long-span cable-stayed bridges under an approaching crosswind flow. Starting from a continuous model of the fan-shaped bridge scheme with both H- or A-shaped towers, critical states of the coupled wind-structure system are identified by means of a variational formulation, accounting for torsional and flexural (vertical and lateral) bridge oscillations. The overall bridge dynamics is described by introducing simple mechanical systems with equivalent stiffness properties and, under the assumption of a prevailing truss-like bridge behavior, analytical estimates for dominant stiffness contributions are proposed. Several case studies are discussed and comparisons with experimental evidences as well as with available analytical and numerical results are presented. The proposed simplified approach proves to be consistent and effective for successfully capturing the main wind-bridge interaction mechanisms, and it could be considered as a useful engineering tool for the aeroelastic stability analysis of long-span cable-stayed bridges.
NASA Astrophysics Data System (ADS)
Goldman, Benjamin D.
The purpose of this dissertation is to study the aeroelastic stability of a proposed flexible thermal protection system (FTPS) for the NASA Hypersonic Inflatable Aerodynamic Decelerator (HIAD). A flat, square FTPS coupon exhibits violent oscillations during experimental aerothermal testing in NASA's 8 Foot High Temperature Tunnel, leading to catastrophic failure. The behavior of the structural response suggested that aeroelastic flutter may be the primary instability mechanism, prompting further experimental investigation and theoretical model development. Using Von Karman's plate theory for the panel-like structure and piston theory aerodynamics, a set of aeroelastic models were developed and limit cycle oscillations (LCOs) were calculated at the tunnel flow conditions. Similarities in frequency content of the theoretical and experimental responses indicated that the observed FTPS oscillations were likely aeroelastic in nature, specifically LCO/flutter. While the coupon models can be used for comparison with tunnel tests, they cannot predict accurately the aeroelastic behavior of the FTPS in atmospheric flight. This is because the geometry of the flight vehicle is no longer a flat plate, but rather (approximately) a conical shell. In the second phase of this work, linearized Donnell conical shell theory and piston theory aerodynamics are used to calculate natural modes of vibration and flutter dynamic pressures for various structural models composed of one or more conical shells resting on several circumferential elastic supports. When the flight vehicle is approximated as a single conical shell without elastic supports, asymmetric flutter in many circumferential waves is observed. When the elastic supports are included, the shell flutters symmetrically in zero circumferential waves. Structural damping is found to be important in this case, as "hump-mode" flutter is possible. Aeroelastic models that consider the individual FTPS layers as separate shells exhibit
NASA Technical Reports Server (NTRS)
Gilyard, G. B.; Edwards, J. W.
1983-01-01
Flight-test results of the first three flights of an aeroelastic research wing are described. The flight flutter-test technique used to obtain real-time damping estimates from fast-frequency sweep data was obtained and the open-loop flutter boundary determined. Nyquist analyses of sweep maneuvers appear to provide additional valuable information about flutter suppression system operation, both in terms of phase-margin estimates and as a means of evaluating maneuver quality. An error in implementing the flutter-suppression system required in a one-half nominal gain configuration, which caused the wing to be unstable at lower Mach numbers than anticipated, and the vehicle experienced closed-loop flutter on its third flight. Real-time flutter-testing procedures were improved.
Investigating the Transonic Flutter Boundary of the Benchmark Supercritical Wing
NASA Technical Reports Server (NTRS)
Heeg, Jennifer; Chwalowski, Pawel
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 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 focus on understanding the dip in the transonic flutter boundary at a single Mach number (0.74), exploring an angle of attack range of ??1 to 8 and dynamic pressures from wind off to beyond flutter onset. The rigid analysis results are examined for insights into the behavior of the aeroelastic system. Both static and dynamic aeroelastic simulation results are also examined.
NASTRAN level 16 programmer's manual updates for aeroelastic analysis of bladed discs
NASA Technical Reports Server (NTRS)
Gallo, A. M.; Dale, B.
1980-01-01
The programming routines for the NASTRAN Level 16program are presented. Particular emphasis is placed on its application to aeroelastic analyses, mode development, and flutter analysis for turbomachine blades.
Aeroelastic and dynamic finite element analyses of a bladder shrouded disk
NASA Technical Reports Server (NTRS)
Smith, G. C. C.; Elchuri, V.
1980-01-01
The delivery and demonstration of a computer program for the analysis of aeroelastic and dynamic properties is reported. Approaches to flutter and forced vibration of mistuned discs, and transient aerothermoelasticity are described.
Modern wing flutter analysis by computational fluid dynamics methods
NASA Technical Reports Server (NTRS)
Cunningham, Herbert J.; Batina, John T.; Bennett, Robert M.
1988-01-01
The application and assessment of the recently developed CAP-TSD transonic small-disturbance code for flutter prediction is described. The CAP-TSD code has been developed for aeroelastic analysis of complete aircraft configurations and was previously applied to the calculation of steady and unsteady pressures with favorable results. Generalized aerodynamic forces and flutter characteristics are calculated and compared with linear theory results and with experimental data for a 45 deg sweptback wing. These results are in good agreement with the experimental flutter data which is the first step toward validating CAP-TSD for general transonic aeroelastic applications. The paper presents these results and comparisons along with general remarks regarding modern wing flutter analysis by computational fluid dynamics methods.
NASA Technical Reports Server (NTRS)
Smith, Todd E.
1990-01-01
The dynamic analysis for the SSME HPOTP first stage turbine blade is presented wherein the rotor aeroelastic stability is assessed. The method employs normal modes analysis to simulate the coupled blade/fluid system. A three-dimensional finite element model of the blade is used in conjunction with a two-dimensional linearized unsteady aerodynamic theory which accounts for steady aerodynamic loading effects. This unsteady aerodynamic model is applied in stacked axisymmetric strips along the airfoil span. The blade dynamic and aerodynamic behaviors are coupled within modal space by expressing the unsteady aerodynamic forces in the frequency domain. A complex eigenvalue problem is solved to determine the stability of the rotor assuming tuned blades. The present analysis indicates that the HPOTP rotor experiences very low aerodynamic damping in the first four vibrational modes. The edgewise mode was found to be dynamically unstable. This mode of the blade became stable when the effect of mechanical damping was considered.
NASA Technical Reports Server (NTRS)
Peterson, R. L.; Warmbrodt, W.
1984-01-01
A hover test of a full-scale, hingeless rotor system was conducted in the NASA Ames 40- by 80-foot wind tunnel. The rotor was tested on the Ames rotor test apparatus. Rotor aeroelastic stability, performance, and loads at various rotational speeds and thrust coefficients were investigated. The primary objective was to determine the inplane stability characteristics of the rotor system. Rotor inplane damping data were obtained for operation between 350 and 425 rpm (design speed), and for thurst coefficients between 0.0 and 0.12. The rotor was stable for all conditions tested. At constant rotor rotational speed, a minimum inplane dampling level was obtained at a thrust coefficient approximately = 0.02. At constant rotor lift, a minimum in rotor inplane damping was measured at 400 rpm.
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.
NASA Technical Reports Server (NTRS)
Edwards, John W.
1996-01-01
A viscous-inviscid interactive coupling method is used for the computation of unsteady transonic flows involving separation and reattachment. A lag-entrainment integral boundary layer method is used with the transonic small disturbance potential equation in the CAP-TSDV (Computational Aeroelasticity Program - Transonic Small Disturbance) code. Efficient and robust computations of steady and unsteady separated flows, including steady separation bubbles and self-excited shock-induced oscillations are presented. The buffet onset boundary for the NACA 0012 airfoil is accurately predicted and shown computationally to be a Hopf bifurcation. Shock-induced oscillations are also presented for the 18 percent circular arc airfoil. The oscillation onset boundaries and frequencies are accurately predicted, as is the experimentally observed hysteresis of the oscillations with Mach number. This latter stability boundary is identified as a jump phenomenon. Transonic wing flutter boundaries are also shown for a thin swept wing and for a typical business jet wing, illustrating viscous effects on flutter and the effect of separation onset on the wing response at flutter. Calculations for both wings show limit cycle oscillations at transonic speeds in the vicinity of minimum flutter speed indices.
Predicting the aeroelastic behavior of a wind-tunnel model using transonic small disturbance theory
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 AFW wind-tunnel experiments. Dynamic results are presented in the form of root loci at different Mach numbers for a heavy gas and for air test mediums. The resultant flutter boundaries for both gases, and the effects of viscous damping and angle of attack on the flutter boundary in air, are also presented.
Recent Applications of Higher-Order Spectral Analysis to Nonlinear Aeroelastic Phenomena
NASA Technical Reports Server (NTRS)
Silva, Walter A.; Hajj, Muhammad R.; Dunn, Shane; Strganac, Thomas W.; Powers, Edward J.; Stearman, Ronald
2005-01-01
Recent applications of higher-order spectral (HOS) methods to nonlinear aeroelastic phenomena are presented. Applications include the analysis of data from a simulated nonlinear pitch and plunge apparatus and from F-18 flight flutter tests. A MATLAB model of the Texas A&MUniversity s Nonlinear Aeroelastic Testbed Apparatus (NATA) is used to generate aeroelastic transients at various conditions including limit cycle oscillations (LCO). The Gaussian or non-Gaussian nature of the transients is investigated, related to HOS methods, and used to identify levels of increasing nonlinear aeroelastic response. Royal Australian Air Force (RAAF) F/A-18 flight flutter test data is presented and analyzed. The data includes high-quality measurements of forced responses and LCO phenomena. Standard power spectral density (PSD) techniques and HOS methods are 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.
NASA Technical Reports Server (NTRS)
Silva, Walter A.; Perry, Boyd III; Chwalowski, Pawel
2014-01-01
Reduced-order modeling (ROM) methods are applied to the CFD-based aeroelastic analysis of the AGARD 445.6 wing in order to gain insight regarding well-known discrepancies between the aeroelastic analyses and the experimental results. The results presented include aeroelastic solutions using the inviscid CAP-TSD code and the FUN3D code (Euler and Navier-Stokes). Full CFD aeroelastic solutions and ROM aeroelastic solutions, computed at several Mach numbers, are presented in the form of root locus plots in order to better reveal the aeroelastic root migrations with increasing dynamic pressure. Important conclusions are drawn from these results including the ability of the linear CAP-TSD code to accurately predict the entire experimental flutter boundary (repeat of analyses performed in the 1980's), that the Euler solutions at supersonic conditions indicate that the third mode is always unstable, and that the FUN3D Navier-Stokes solutions stabilize the unstable third mode seen in the Euler solutions.
A computational transonic flutter boundary tracking procedure. M.S. Thesis
NASA Technical Reports Server (NTRS)
Gallman, J. W.; Batina, J. T.; Yang, T. Y.
1986-01-01
An automated flutter boundary tracking procedure for the efficient calculation of transonic flutter boundaries is presented. The procedure uses aeroelastic responses to march along the boundary by taking steps in speed and Mach number, thereby reducing the number of response calculations previously required to determine a transonic flutter boundary. Flutter boundary results are presented for a typical airfoil section oscillating with pitch and plunge degrees of freedom. These transonic flutter boundaries are in good agreement with exact boundaries calculated using the conventional time-marching method. The tracking procedure is extended to include static aeroelastic twist as a simulation of the static deformation of a wing and contains all of the essential features that are required to apply it to practical three-dimensional cases. The procedure is also applied to flutter boundaries as a function of structural parameters.
Aeroelastic Analysis of Counter Rotation Fans
NASA Technical Reports Server (NTRS)
Keith, Theo G., Jr.; Murthy, Durbha V.
1997-01-01
Aeroelastic problems in turbomachinery and propfans can be static or dynamic in nature. The analysis of static aeroelastic problems is involved primarily with determination: (a) of the shape of the blades and the steady aerodynamic loads on the blades (which are inter-dependent), (b) of the resultant steady stresses and (c) of the static instability (divergence) margin, if applicable. In this project, we were concerned exclusively with dynamic aeroelastic behavior. The analysis of dynamic aeroelastic problems is involved with the determination: (a) of the unsteady aerodynamic loads on blades and the dynamic motion of the blades (which are again inter-dependent), (b) of the resultant dynamic stresses and their effect on fatigue life and (c) of the dynamic instability (flutter), if applicable. There are two primary dynamic aeroelastic phenomena of interest to designers of turbomachinery and propfans: flutter and forced response. Flutter generally refers to the occurrence of rapidly growing self-excited oscillations leading to catastrophic failure of the blade. When certain nonlinear phenomena are present, flutter response may lead to a potentially dangerous limit cycle oscillation rather than an immediate catastrophic failure. Forced response generally refers to the steady-state oscillations that occur as a consequence of excitations external to the rotor in question. These excitations typically result from the presence of upstream obstructions, inflow distortions, downstream obstructions, or mechanical sources such as tip-casing contact or shaft and gear meshing. Significant forced response leads to blade fatigue, and at design conditions, generally contributes to a degradation of blade life. At other operating conditions, forced response may lead to catastrophic failure due to severe blade fatigue in a short duration of time.
NASA Technical Reports Server (NTRS)
Nguyen, Nhan; Kaul, Upender; Lebofsky, Sonia; Ting, Eric; Chaparro, Daniel; Urnes, James
2015-01-01
This paper summarizes the recent development of an adaptive aeroelastic wing shaping control technology called variable camber continuous trailing edge flap (VCCTEF). As wing flexibility increases, aeroelastic interactions with aerodynamic forces and moments become an increasingly important consideration in aircraft design and aerodynamic performance. Furthermore, aeroelastic interactions with flight dynamics can result in issues with vehicle stability and control. The initial VCCTEF concept was developed in 2010 by NASA under a NASA Innovation Fund study entitled "Elastically Shaped Future Air Vehicle Concept," which showed that highly flexible wing aerodynamic surfaces can be elastically shaped in-flight by active control of wing twist and bending deflection in order to optimize the spanwise lift distribution for drag reduction. A collaboration between NASA and Boeing Research & Technology was subsequently funded by NASA from 2012 to 2014 to further develop the VCCTEF concept. This paper summarizes some of the key research areas conducted by NASA during the collaboration with Boeing Research and Technology. These research areas include VCCTEF design concepts, aerodynamic analysis of VCCTEF camber shapes, aerodynamic optimization of lift distribution for drag minimization, wind tunnel test results for cruise and high-lift configurations, flutter analysis and suppression control of flexible wing aircraft, and multi-objective flight control for adaptive aeroelastic wing shaping control.
Stall Flutter Control of a Smart Blade Section Undergoing Asymmetric Limit Oscillations
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. Although 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. The improvement and the superiority of the proposed adaptive stall flutter controller are shown by comparison with a simple stall flutter controller.« less
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.
NASTRAN level 16 user's manual updates for aeroelastic analysis of bladed discs
NASA Technical Reports Server (NTRS)
Elchuri, V.; Gallo, A. M.
1980-01-01
The NASTRAN aeroelastic and flutter capability was extended to solve a class of problems associated with axial flow turbomachines. The capabilities of the program are briefly discussed. The aerodynamic data pertaining to the bladed disc sector, the associated aerodynamic modeling, the steady aerothermoelastic 'design/analysis' formulations, and the modal, flutter, and subcritical roots analyses are described. Sample problems and their solutions are included.
Stochastic Characterization of Flutter using Historical Wind Tunnel Data
NASA Technical Reports Server (NTRS)
Heeg, Jennifer
2007-01-01
Methods for predicting the onset of flutter during an experiment are traditionally applied treating the data as deterministic values. Uncertainty and variation in the data is often glossed over by using best-fit curves to represent the information. This paper applies stochastic treatments to wind tunnel data obtained for the Piezoelectric Aeroelastic Response Tailoring Investigation model. These methods include modal amplitude tracking, modal frequency tracking and several applications of the flutter margin method. The flutter margin method was developed by Zimmerman and Weissenburger, and extended by Poirel, Dunn and Porter to incorporate uncertainty. Much of the current work follows the future work recommendations of Poirel, Dunn and Porter.
NASA Technical Reports Server (NTRS)
Kaza, K. R. V.; Kielb, R. E.
1981-01-01
The effect of small differences between the individual blades (mistuning) on the aeroelastic stability and response of a cascade were studied. The aerodynamic, inertial, and structural coupling between the bending and torsional motions of each blade and the aerodynamic coupling between the blades was considered. A digital computer program was developed to conduct parametric studies. Results indicate that the mistuning has a beneficial effect on the coupled bending torsion and uncoupled torsion flutter. On forced response, however, the effect may be either beneficial or adverse, depending on the engine order of the forcing function. The results also illustrate that it may be feasible to utilize mistuning as a passive control to increase flutter speed while maintaining forced response at an acceptable level.
Rigid Body Stability Augmentation Studies for a Wind Tunnel Flutter Model.
1985-11-01
evident in the root locus diagrams, Figs 3 and 4. Therefore, the stability augmentation system (SAS) was required to - provide basic stability and if... augmentation system (SAS) control law for it. The RAE designed control system simulation and analysis computer package ’TSIM’ (Refs. 3, 4) was used to program...One of my activities in preparation for the Phase 2 trials was to investigate the rigid-body response of the model and to develop a stability
NASA Technical Reports Server (NTRS)
Silva, Walter A.; Bennett, Robert M.
1992-01-01
The CAP-TSD (Computational Aeroelasticity Program - Transonic Small Disturbance) code, developed at the NASA Langley Research Center, is applied to the Active Flexible Wing 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. Static aeroelastic solutions using CAP-TSD are computed. Dynamic (flutter) analyses then are performed as perturbations about the static aeroelastic deformations and 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 motions and sensitivity to the modeling of the wing tip ballast stores also are presented and compared with experimental flutter results.
Blade row interaction effects on flutter and forced response
NASA Technical Reports Server (NTRS)
Buffum, Daniel H.
1993-01-01
In the flutter or forced response analysis of a turbomachine blade row, the blade row in question is commonly treated as if it is isolated from the neigboring blade rows. Disturbances created by vibrating blades are then free to propagate away from this blade row without being disturbed. In reality, neighboring blade rows will reflect some portion of this wave energy back toward the vibrating blades, causing additional unsteady forces on them. It is of fundamental importance to determine whether or not these reflected waves can have a significant effect on the aeroelastic stability or forced response of a blade row. Therefore, a procedure to calculate intra-blade-row unsteady aerodynamic interactions was developed which relies upon results available from isolated blade row unsteady aerodynamic analyses. In addition, an unsteady aerodynamic influence coefficient technique is used to obtain a model for the vibratory response in which the neighboring blade rows are also flexible. The flutter analysis shows that interaction effects can be destabilizing, and the forced response analysis shows that interaction effects can result in a significant increase in the resonant response of a blade row.
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.
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.
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.
Viscous effects on transonic airfoil stability and response
NASA Technical Reports Server (NTRS)
Berry, H. M.; Batina, J. T.; Yang, T. Y.
1985-01-01
Viscous effects on transonic airfoil stability and response are investigated using an integral boundary layer model coupled to the inviscid XTRAN2L transonic small disturbance code. Unsteady transonic airloads required for stability analyses are computed using a pulse transfer function analysis including viscous effects. The pulse analysis provides unsteady aerodynamic forces for a wide range of reduced frequency in a single flow field computation. Nonlinear time marching aeroelastic solutions are presented which show the effects of viscosity on airfoil response behavior and flutter. Effects of amplitude on time marching responses are demonstrated. A state space aeroelastic model employing Pade approximants to describe the unsteady airloads is used to study the effects of viscosity on transonic airfoil stability. State space dynamic pressure root loci are in good overall agreement with time marching damping and frequency estimates. Parallel sets of results with and without viscous effects reveal the effects of viscosity on transonic unsteady airloads and aeroelastic characteristics of airfoils.
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.
Adaptive neural control of aeroelastic response
NASA Astrophysics Data System (ADS)
Lichtenwalner, Peter F.; Little, Gerald R.; Scott, Robert C.
1996-05-01
The Adaptive Neural Control of Aeroelastic Response (ANCAR) program is a joint research and development effort conducted by McDonnell Douglas Aerospace (MDA) and the National Aeronautics and Space Administration, Langley Research Center (NASA LaRC) under a Memorandum of Agreement (MOA). The purpose of the MOA is to cooperatively develop the smart structure technologies necessary for alleviating undesirable vibration and aeroelastic response associated with highly flexible structures. Adaptive control can reduce aeroelastic response associated with buffet and atmospheric turbulence, it can increase flutter margins, and it may be able to reduce response associated with nonlinear phenomenon like limit cycle oscillations. By reducing vibration levels and loads, aircraft structures can have lower acquisition cost, reduced maintenance, and extended lifetimes. Phase I of the ANCAR program involved development and demonstration of a neural network-based semi-adaptive flutter suppression system which used a neural network for scheduling control laws as a function of Mach number and dynamic pressure. This controller was tested along with a robust fixed-gain control law in NASA's Transonic Dynamics Tunnel (TDT) utilizing the Benchmark Active Controls Testing (BACT) wing. During Phase II, a fully adaptive on-line learning neural network control system has been developed for flutter suppression which will be tested in 1996. This paper presents the results of Phase I testing as well as the development progress of Phase II.
Elementary Applications of a Rotorcraft Dynamic Stability Analysis
NASA Technical Reports Server (NTRS)
Johnson, W.
1976-01-01
A number of applications of a rotorcraft aeroelastic analysis are presented to verify that the analysis encompasses the classical solutions of rotor dynamics, and to examine the influence of certain features of the model. Results are given for the following topics: flapping frequency response to pitch control; forward flight flapping stability; pitch/flap flutter and divergence; ground resonance instability; and the flight dynamics of several representative helicopters.
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.
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
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.
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.
Aeroelastic Tailoring via Tow Steered Composites
NASA Technical Reports Server (NTRS)
Stanford, Bret K.; Jutte, Christine V.
2014-01-01
The use of tow steered composites, where fibers follow prescribed curvilinear paths within a laminate, can improve upon existing capabilities related to aeroelastic tailoring of wing structures, though this tailoring method has received relatively little attention in the literature. This paper demonstrates the technique for both a simple cantilevered plate in low-speed flow, as well as the wing box of a full-scale high aspect ratio transport configuration. Static aeroelastic stresses and dynamic flutter boundaries are obtained for both cases. The impact of various tailoring choices upon the aeroelastic performance is quantified: curvilinear fiber steering versus straight fiber steering, certifiable versus noncertifiable stacking sequences, a single uniform laminate per wing skin versus multiple laminates, and identical upper and lower wing skins structures versus individual tailoring.
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.
Wing flutter boundary prediction using unsteady Euler aerodynamic method
NASA Technical Reports Server (NTRS)
Lee-Rausch, Elizabeth M.; Batina, John T.
1993-01-01
Modifications to an existing 3D implicit upwind Euler/Navier-Stokes code for the aeroelastic analysis of wings are described. These modifications include the incorporation of a deforming mesh algorithm and the addition of the structural equations of motion for their simultaneous time-integration with the governing flow equations. The paper gives a brief description of these modifications and presents unsteady calculations which check the modifications to the code. Euler flutter results for an isolated 45 deg swept-back wing are compared with experimental data for seven freestream Mach numbers which define the flutter boundary over a range of Mach number from 0.499 to 1.14. These comparisons show good agreement in flutter characteristics for freestream Mach numbers below unity. For freestream Mach numbers above unity, the computed aeroelastic results predict a premature rise in the flutter boundary as compared with the experimental boundary. Steady and unsteady contours of surface Mach number and pressure are included to illustrate the basic flow characteristics of the time-marching flutter calculations and to aid in identifying possible causes for the premature rise in the computational flutter boundary.
Wing flutter boundary prediction using an unsteady Euler aerodynamic method
NASA Technical Reports Server (NTRS)
Lee-Rausch, Elizabeth M.; Batina, John T.
1993-01-01
Modifications to an existing three-dimensional, implicit, upwind Euler/Navier-Stokes code (CFL3D Version 2.1) for the aeroelastic analysis of wings are described. These modifications, which were previously added to CFL3D Version 1.0, include the incorporation of a deforming mesh algorithm and the addition of the structural equations of motion for their simultaneous time-integration with the government flow equations. The paper gives a brief description of these modifications and presents unsteady calculations which check the modifications to the code. Euler flutter results for an isolated 45 degree swept-back wing are compared with experimental data for seven freestream Mach numbers which define the flutter boundary over a range of Mach number from 0.499 to 1.14. These comparisons show good agreement in flutter characteristics for freestream Mach numbers below unity. For freestream Mach numbers above unity, the computed aeroelastic results predict a premature rise in the flutter boundary as compared with the experimental boundary. Steady and unsteady contours of surface Mach number and pressure are included to illustrate the basic flow characteristics of the time-marching flutter calculations and to aid in identifying possible causes for the premature rise in the computational flutter boundary.
NASA Technical Reports Server (NTRS)
Mukhopadhyay, V.; Newsom, J. R.
1982-01-01
A stability margin evaluation method in terms of simultaneous gain and phase changes in all loops of a multiloop system is presented. A universal gain-phase margin evaluation diagram is constructed by generalizing an existing method using matrix singular value properties. Using this diagram and computing the minimum singular value of the system return difference matrix over the operating frequency range, regions of guaranteed stability margins can be obtained. Singular values are computed for a wing flutter suppression and a drone lateral attitude control problem. The numerical results indicate that this method predicts quite conservative stability margins. In the second example if the eigenvalue magnitude is used instead of the singular value, as a measure of nearness to singularity, more realistic stability margins are obtained. However, this relaxed measure generally cannot guarantee global stability.
Dynamic Aeroelastic Analysis of Wing/Store Configurations
2005-12-01
for his assistance with Gridgen as well as Jacob Freeman, John Staples, and Dr. Charles Denegri for providing F-16 data. I would also like to thank my...ure 3.5) was created using Gridgen . A calculation of the flutter point was then made using the aeroelastic program. A dynamic pressure was chosen
Inertial Force Coupling to Nonlinear Aeroelasticity of Flexible Wing Aircraft
NASA Technical Reports Server (NTRS)
Nguyen, Nhan T.; Ting, Eric
2016-01-01
This paper investigates the inertial force effect on nonlinear aeroelasticity of flexible wing aircraft. The geometric are nonlinearity due to rotational and tension stiffening. The effect of large bending deflection will also be investigated. Flutter analysis will be conducted for a truss-braced wing aircraft concept with tension stiffening and inertial force coupling.
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.
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.
Aeroelastic Optimization Study Based on the X-56A Model
NASA Technical Reports Server (NTRS)
Li, Wesley W.; Pak, Chan-Gi
2014-01-01
One way to increase the aircraft fuel efficiency is to reduce structural weight while maintaining adequate structural airworthiness, both statically and aeroelastically. A design process which incorporates the object-oriented multidisciplinary design, analysis, and optimization (MDAO) tool and the aeroelastic effects of high fidelity finite element models to characterize the design space was successfully developed and established. This paper presents two multidisciplinary design optimization studies using an object-oriented MDAO tool developed at NASA Armstrong Flight Research Center. The first study demonstrates the use of aeroelastic tailoring concepts to minimize the structural weight while meeting the design requirements including strength, buckling, and flutter. Such an approach exploits the anisotropic capabilities of the fiber composite materials chosen for this analytical exercise with ply stacking sequence. A hybrid and discretization optimization approach improves accuracy and computational efficiency of a global optimization algorithm. The second study presents a flutter mass balancing optimization study for the fabricated flexible wing of the X-56A model since a desired flutter speed band is required for the active flutter suppression demonstration during flight testing. The results of the second study provide guidance to modify the wing design and move the design flutter speeds back into the flight envelope so that the original objective of X-56A flight test can be accomplished successfully. The second case also demonstrates that the object-oriented MDAO tool can handle multiple analytical configurations in a single optimization run.
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.
Aeroelastic Optimization Study Based on X-56A Model
NASA Technical Reports Server (NTRS)
Li, Wesley; Pak, Chan-Gi
2014-01-01
A design process which incorporates the object-oriented multidisciplinary design, analysis, and optimization (MDAO) tool and the aeroelastic effects of high fidelity finite element models to characterize the design space was successfully developed and established. Two multidisciplinary design optimization studies using an object-oriented MDAO tool developed at NASA Armstrong Flight Research Center were presented. The first study demonstrates the use of aeroelastic tailoring concepts to minimize the structural weight while meeting the design requirements including strength, buckling, and flutter. A hybrid and discretization optimization approach was implemented to improve accuracy and computational efficiency of a global optimization algorithm. The second study presents a flutter mass balancing optimization study. The results provide guidance to modify the fabricated flexible wing design and move the design flutter speeds back into the flight envelope so that the original objective of X-56A flight test can be accomplished.
Volterra Series Approach for Nonlinear Aeroelastic Response of 2-D Lifting Surfaces
NASA Technical Reports Server (NTRS)
Silva, Walter A.; Marzocca, Piergiovanni; Librescu, Liviu
2001-01-01
The problem of the determination of the subcritical aeroelastic response and flutter instability of nonlinear two-dimensional lifting surfaces in an incompressible flow-field via Volterra series approach is addressed. The related aeroelastic governing equations are based upon the inclusion of structural nonlinearities, of the linear unsteady aerodynamics and consideration of an arbitrary time-dependent external pressure pulse. Unsteady aeroelastic nonlinear kernels are determined, and based on these, frequency and time histories of the subcritical aeroelastic response are obtained, and in this context the influence of geometric nonlinearities is emphasized. Conclusions and results displaying the implications of the considered effects are supplied.
Coupled nonlinear aeroelasticity and flight dynamics of fully flexible aircraft
NASA Astrophysics Data System (ADS)
Su, Weihua
This dissertation introduces an approach to effectively model and analyze the coupled nonlinear aeroelasticity and flight dynamics of highly flexible aircraft. A reduced-order, nonlinear, strain-based finite element framework is used, which is capable of assessing the fundamental impact of structural nonlinear effects in preliminary vehicle design and control synthesis. The cross-sectional stiffness and inertia properties of the wings are calculated along the wing span, and then incorporated into the one-dimensional nonlinear beam formulation. Finite-state unsteady subsonic aerodynamics is used to compute airloads along lifting surfaces. Flight dynamic equations are then introduced to complete the aeroelastic/flight dynamic system equations of motion. Instead of merely considering the flexibility of the wings, the current work allows all members of the vehicle to be flexible. Due to their characteristics of being slender structures, the wings, tail, and fuselage of highly flexible aircraft can be modeled as beams undergoing three dimensional displacements and rotations. New kinematic relationships are developed to handle the split beam systems, such that fully flexible vehicles can be effectively modeled within the existing framework. Different aircraft configurations are modeled and studied, including Single-Wing, Joined-Wing, Blended-Wing-Body, and Flying-Wing configurations. The Lagrange Multiplier Method is applied to model the nodal displacement constraints at the joint locations. Based on the proposed models, roll response and stability studies are conducted on fully flexible and rigidized models. The impacts of the flexibility of different vehicle members on flutter with rigid body motion constraints, flutter in free flight condition, and roll maneuver performance are presented. Also, the static stability of the compressive member of the Joined-Wing configuration is studied. A spatially-distributed discrete gust model is incorporated into the time simulation
Aeroelastic and Flight Dynamics Analysis of Folding Wing Systems
NASA Astrophysics Data System (ADS)
Wang, Ivan
This dissertation explores the aeroelastic stability of a folding wing using both theoretical and experimental methods. The theoretical model is based on the existing clamped-wing aeroelastic model that uses beam theory structural dynamics and strip theory aerodynamics. A higher-fidelity theoretical model was created by adding several improvements to the existing model, namely a structural model that uses ANSYS for individual wing segment modes and an unsteady vortex lattice aerodynamic model. The comparison with the lower-fidelity model shows that the higher-fidelity model typical provides better agreement between theory and experiment, but the predicted system behavior in general does not change, reinforcing the effectiveness of the low-fidelity model for preliminary design of folding wings. The present work also conducted more detailed aeroelastic analyses of three-segment folding wings, and in particular considers the Lockheed-type configurations to understand the existence of sudden changes in predicted aeroelastic behavior with varying fold angle for certain configurations. These phenomena were observed in carefully conducted experiments, and nonlinearities---structural and geometry---were shown to suppress the phenomena. Next, new experimental models with better manufacturing tolerances are designed to be tested in the Duke University Wind Tunnel. The testing focused on various configurations of three-segment folding wings in order to obtain higher quality data. Next, the theoretical model was further improved by adding aircraft longitudinal degrees of freedom such that the aeroelastic model may predict the instabilities for the entire aircraft and not just a clamped wing. The theoretical results show that the flutter instabilities typically occur at a higher air speed due to greater frequency separation between modes for the aircraft system than a clamped wing system, but the divergence instabilities occur at a lower air speed. Lastly, additional
Aeroelastic analysis of wind energy conversion systems
NASA Technical Reports Server (NTRS)
Dugundji, J.
1978-01-01
An aeroelastic investigation of horizontal axis wind turbines is described. The study is divided into two simpler areas; (1) the aeroelastic stability of a single blade on a rigid tower; and (2) the mechanical vibrations of the rotor system on a flexible tower. Some resulting instabilities and forced vibration behavior are described.
Shock-boundary layer interaction and transonic flutter
NASA Astrophysics Data System (ADS)
Tumkur Karnick, Pradeepa; Venkatraman, Kartik
2012-11-01
The transonic flutter dip of an aeroelastic system is primarily caused by compressibility of the flowing fluid. Viscous effects are not dominant in the pre-transonic dip region. In fact, an Euler solver can predict this flutter boundary with considerable accuracy. However with an increase in Mach number the shock moves towards the trailing edge causing shock induced separation. This shock-boundary layer interaction changes the flutter boundary in the transonic and post-transonic dip region significantly. We discuss the effect of viscosity in changing the flutter boundary in the post-transonic dip region using a RANS solver coupled to a two-degree of freedom model of the structural dynamics of a wing.
Stall flutter analysis of propfans
NASA Technical Reports Server (NTRS)
Reddy, T. S. R.
1988-01-01
Three semi-empirical aerodynamic stall models are compared with respect to their lift and moment hysteresis loop prediction, limit cycle behavior, easy implementation, and feasibility in developing the parameters required for stall flutter prediction of advanced turbines. For the comparison of aeroelastic response prediction including stall, a typical section model and a plate structural model are considered. The response analysis includes both plunging and pitching motions of the blades. In model A, a correction of the angle of attack is applied when the angle of attack exceeds the static stall angle. In model B, a synthesis procedure is used for angles of attack above static stall angles, and the time history effects are accounted for through the Wagner function.
Aeroelastic Computations of a Compressor Stage Using the Harmonic Balance Method
NASA Technical Reports Server (NTRS)
Reddy, T. S. R.
2010-01-01
The aeroelastic characteristics of a compressor stage were analyzed using a computational fluid dynamic (CFD) solver that uses the harmonic balance method to solve the governing equations. The three dimensional solver models the unsteady flow field due to blade vibration using the Reynolds-Averaged Navier-Stokes equations. The formulation enables the study of the effect of blade row interaction through the inclusion of coupling modes between blade rows. It also enables the study of nonlinear effects of high amplitude blade vibration by the inclusion of higher harmonics of the fundamental blade vibration frequency. In the present work, the solver is applied to study in detail the aeroelastic characteristics of a transonic compressor stage. Various parameters were included in the study: number of coupling modes, blade row axial spacing, and operating speeds. Only the first vibration mode is considered with amplitude of oscillation in the linear range. Both aeroelastic stability (flutter) of rotor blade and unsteady loading on the stator are calculated. The study showed that for the stage considered, the rotor aerodynamic damping is not influenced by the presence of the stator even when the axial spacing is reduced by nearly 25 percent. However, the study showed that blade row interaction effects become important for the unsteady loading on the stator when the axial spacing is reduced by the same amount.
Aeroelastic Calculations of Quiet High- Speed Fan Performed
NASA Technical Reports Server (NTRS)
Bakhle, Milind A.; Srivastava, Rakesh; Mehmed, Oral; Min, James B.
2002-01-01
An advanced high-speed fan was recently designed under a cooperative effort between the NASA Glenn Research Center and Honeywell Engines & Systems. The principal design goals were to improve performance and to reduce fan noise at takeoff. Scale models of the Quiet High-Speed Fan were tested for operability, performance, and acoustics. During testing, the fan showed significantly improved noise characteristics, but a self-excited aeroelastic vibration known as flutter was encountered in the operating range. Flutter calculations were carried out for the Quiet High-Speed Fan using a three-dimensional, unsteady aerodynamic, Reynolds-averaged Navier-Stokes turbomachinery code named "TURBO." The TURBO code can accurately model the viscous flow effects that can play an important role in various aeroelastic problems such as flutter with flow separation, flutter at high loading conditions near the stall line (stall flutter), and flutter in the presence of shock and boundary-layer interaction. Initially, calculations were performed with no blade vibrations. These calculations were at a constant rotational speed and a varying mass flow rate. The mass flow rate was varied by changing the backpressure at the exit boundary of the computational domain. These initial steady calculations were followed by aeroelastic calculations in which the blades were prescribed to vibrate harmonically in a natural mode, at a natural frequency, and with a fixed interblade phase angle between adjacent blades. The AE-prep preprocessor was used to interpolate the in-vacuum mode shapes from the structural dynamics mesh onto the computational fluid dynamics mesh and to smoothly propagate the grid deformations from the blade surface to the interior points of the grid. The aeroelastic calculations provided the unsteady aerodynamic forces on the blade surface due to blade vibrations. These forces were vector multiplied with the structural dynamic mode shape to calculate the work done on the blade during
Nonlinear Aeroelastic Analysis of Joined-Wing Configurations
NASA Astrophysics Data System (ADS)
Cavallaro, Rauno
Aeroelastic design of joined-wing configurations is yet a relatively unexplored topic which poses several difficulties. Due to the overconstrained nature of the system combined with structural geometric nonlinearities, the behavior of Joined Wings is often counterintuitive and presents challenges not seen in standard layouts. In particular, instability observed on detailed aircraft models but never thoroughly investigated, is here studied with the aid of a theoretical/computational framework. Snap-type of instabilities are shown for both pure structural and aeroelastic cases. The concept of snap-divergence is introduced to clearly identify the true aeroelastic instability, as opposed to the usual aeroelastic divergence evaluated through eigenvalue approach. Multi-stable regions and isola-type of bifurcations are possible characterizations of the nonlinear response of Joined Wings, and may lead to branch-jumping phenomena well below nominal critical load condition. Within this picture, sensitivity to (unavoidable) manufacturing defects could have potential catastrophic effects. The phenomena studied in this work suggest that the design process for Joined Wings needs to be revisited and should focus, when instability is concerned, on nonlinear post-critical analysis since linear methods may provide wrong trend indications and also hide potentially catastrophical situations. Dynamic aeroelastic analyses are also performed. Flutter occurrence is critically analyzed with frequency and time-domain capabilities. Sensitivity to different-fidelity aeroelastic modeling (fluid-structure interface algorithm, aerodynamic solvers) is assessed showing that, for some configurations, wake modeling (rigid versus free) has a strong impact on the results. Post-flutter regimes are also explored. Limit cycle oscillations are observed, followed, in some cases, by flip bifurcations (period doubling) and loss of periodicity of the solution. Aeroelastic analyses are then carried out on a
Investigation of aeroelastic stability phenomena of a helicopter by in-flight shake test
NASA Technical Reports Server (NTRS)
Miao, W. L.; Edwards, T.; Brandt, D. E.
1976-01-01
The analytical capability of the helicopter stability program is discussed. The parameters which are found to be critical to the air resonance characteristics of the soft in-plane hingeless rotor systems are detailed. A summary of two model test programs, a 1/13.8 Froude-scaled BO-105 model and a 1.67 meter (5.5 foot) diameter Froude-scaled YUH-61A model, are presented with emphasis on the selection of the final parameters which were incorporated in the full scale YUH-61A helicopter. Model test data for this configuration are shown. The actual test results of the YUH-61A air resonance in-flight shake test stability are presented. Included are a concise description of the test setup, which employs the Grumman Automated Telemetry System (ATS), the test technique for recording in-flight stability, and the test procedure used to demonstrate favorable stability characteristics with no in-plane damping augmentation (lag damper removed). The data illustrating the stability trend of air resonance with forward speed and the stability trend of ground resonance for percent airborne are presented.
NASA Astrophysics Data System (ADS)
Merrett, Craig G.
-partial differential equations. The spatial component of the governing equations is eliminated using a series expansion of basis functions and by applying Galerkin's method. The number of terms in the series expansion affects the convergence of the spatial component, and convergence is best determined by the von Koch rules that previously appeared for column buckling problems. After elimination of the spatial component, an ordinary integral-differential equation in time remains. The dynamic stability of elastic and viscoelastic problems is assessed using the determinant of the governing system of equations and the time component of the solution in the form exp (lambda t). The determinant is in terms of lambda where the values of lambda are the latent roots of the aero-servo-viscoelastic system. The real component of lambda dictates the stability of the system. If all the real components are negative, the system is stable. If at least one real component is zero and all others are negative, the system is neutrally stable. If one or more real components are positive, the system is unstable. In aero-servo-viscoelasticity, the neutrally stable condition is termed flutter. For an aero-servo-viscoelastic lifting surface, the unstable condition is historically termed torsional divergence. The more general aero-servo-viscoelastic theory has produced a number of important results, enumerated in the following list: 1. Subsonic panel flutter can occur before panel instability. This result overturned a long held assumption in aeroelasticity, and was produced by the novel application of the von Koch rules for convergence. Further, experimental results from the 1950s by the Air Force were retrieved to provide additional proof. 2. An expanded definition for flutter of a lifting surface. The legacy definition is that flutter is the first occurrence of simple harmonic motion of a structure, and the flight velocity at which this motion occurs is taken as the flutter speed. The expanded definition
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.
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 %.
Flight testing air-to-air missiles for flutter
NASA Technical Reports Server (NTRS)
Kutschinski, C. R.
1975-01-01
The philosophy of the design of air-to-air missiles and hence of flight testing them for flutter differs from that of manned aircraft. Primary emphasis is put on analytical and laboratory evaluation of missile susceptibility to aeroelastic and aero-servo-elastic instabilities and uses flight testing for confirmation of the absence of such instabilities. Flight testing for flutter is accomplished by using specially instrumented programmed missiles, air or ground launched with a booster to reach the extreme flight conditions of tactical use, or by using guided missiles with telemetered performance data. The instrumentation and testing techniques are discussed along with the success of recent flight tests.
Comparison of supercritical and conventional wing flutter characteristics
NASA Technical Reports Server (NTRS)
Farmer, M. G.; Hanson, P. W.; Wynne, E. C.
1976-01-01
A wind-tunnel study was undertaken to directly compare the measured flutter boundaries of two dynamically similar aeroelastic models which had the same planform, maximum thickness-to-chord ratio, and as nearly identical stiffness and mass distributions as possible, with one wing having a supercritical airfoil and the other a conventional airfoil. The considerations and problems associated with flutter testing supercritical wing models at or near design lift coefficients are discussed, and the measured transonic boundaries of the two wings are compared with boundaries calculated with a subsonic lifting surface theory.
Flutter analysis of low aspect ratio wings
NASA Technical Reports Server (NTRS)
Parnell, L. A.
1986-01-01
Several very low aspect ratio flat plate wing configurations are analyzed for their aerodynamic instability (flutter) characteristics. All of the wings investigated are delta planforms with clipped tips, made of aluminum alloy plate and cantilevered from the supporting vehicle body. Results of both subsonic and supersonic NASTRAN aeroelastic analyses as well as those from another version of the program implementing the supersonic linearized aerodynamic theory are presented. Results are selectively compared with the experimental data; however, supersonic predictions of the Mach Box method in NASTRAN are found to be erratic and erroneous, requiring the use of a separate program.
Aeroelastic Response of Nonlinear Wing Section by Functional Series Technique
NASA Technical Reports Server (NTRS)
Silva, Walter A.; Marzocca, Piergiovanni
2001-01-01
This paper addresses the problem of the determination of the subcritical aeroelastic response and flutter instability of nonlinear two-dimensional lifting surfaces in an incompressible flow-field via indicial functions and Volterra series approach. The related aeroelastic governing equations are based upon the inclusion of structural and damping nonlinearities in plunging and pitching, of the linear unsteady aerodynamics and consideration of an arbitrary time-dependent external pressure pulse. Unsteady aeroelastic nonlinear kernels are determined, and based on these, frequency and time histories of the subcritical aeroelastic response are obtained, and in this context the influence of the considered nonlinearities is emphasized. Conclusions and results displaying the implications of the considered effects are supplied.
Aeroelastic Response of Nonlinear Wing Section By Functional Series Technique
NASA Technical Reports Server (NTRS)
Marzocca, Piergiovanni; Librescu, Liviu; Silva, Walter A.
2000-01-01
This paper addresses the problem of the determination of the subcritical aeroelastic response and flutter instability of nonlinear two-dimensional lifting surfaces in an incompressible flow-field via indicial functions and Volterra series approach. The related aeroelastic governing equations are based upon the inclusion of structural and damping nonlinearities in plunging and pitching, of the linear unsteady aerodynamics and consideration of an arbitrary time-dependent external pressure pulse. Unsteady aeroelastic nonlinear kernels are determined, and based on these, frequency and time histories of the subcritical aeroelastic response are obtained, and in this context the influence of the considered nonlinearities is emphasized. Conclusions and results displaying the implications of the considered effects are supplied.
Aeroelastic Stability of A Soft-Inplane Gimballed Tiltrotor Model In Hover
NASA Technical Reports Server (NTRS)
Nixon, Mark W.; Langston, Chester W.; Singleton, Jeffrey D.; Piatak, David J.; Kvaternik, Raymond G.; Corso, Lawrence M.; Brown, Ross
2001-01-01
Soft-inplane rotor systems can significantly reduce the inplane rotor loads generated during the maneuvers of large tiltrotors, thereby reducing the strength requirements and the associated structural weight of the hub. Soft-inplane rotor systems. however, are subject to instabilities associated with ground resonance, and for tiltrotors this instability has increased complexity as compared to a conventional helicopter. Researchers at Langley Research Center and Bell Helicopter-Textron, Inc. have completed ail initial study of a soft-inplane gimballed tiltrotor model subject to ground resonance conditions in hover. Parametric variations of the rotor collective pitch and blade root damping, and their associated effects oil the model stability were examined. Also considered in the study was the effectiveness of ail active swash-plate and a generalized predictive control (GPC) algorithm for stability augmentation of the ground resonance conditions. Results of this study show that the ground resonance behavior of a gimballed soft-inplane tiltrotor can be significantly different from that of a classical soft-inplane helicopter rotor. The GPC-based active swash-plate was successfully implemented, and served to significantly augment damping of the critical modes to an acceptable value.
NASA Technical Reports Server (NTRS)
Friedmann, P.; Silverthorn, L. J.
1974-01-01
Equations for large amplitude coupled flap-lag motion of a hingeless elastic helicopter blade in forward flight are derived. Only a torsionally rigid blade excited by quasi-steady aerodynamic loads is considered. The effects of reversed flow together with some new terms due to radial flow are included. Using Galerkin's method the spatial dependence is eliminated and the equations are linearized about a suitable equilibrium position. The resulting system of homogeneous periodic equations is solved using multivariable Floquet-Liapunov theory, and the transition matrix at the end of the period is evaluated by two separate methods. Computational efficiency of the two numerical methods is compared. Results illustrating the effects of forward flight and various important blade parameters on the stability boundaries are presented.
Flutter, Postflutter, and Control of a Supersonic Wing Section
NASA Technical Reports Server (NTRS)
Marzocca, Piergiovanni; Librescu, Liviu; Silva, Walter A.
2002-01-01
A number of issues related to the flutter and postflutter of two-dimensional supersonic lifting surfaces are addressed. Among them there are the 1) investigation of the implications of the nonlinear unsteady aerodynamics and structural nonlinearities on the stable/unstable character of the limit cycle and 2) study of the implications of the incorporation of a control capability on both the flutter boundary and the postflutter behavior. To this end, a powerful methodology based on the Lyapunov first quantity is implemented. Such a treatment of the problem enables one to get a better understanding of the various factors involved in the nonlinear aeroelastic problem, including the stable and unstable limit cycle. In addition, it constitutes a first step toward a more general investigation of nonlinear aeroelastic phenomena of three-dimensional lifting surfaces.
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.
Physical Insights, Steady Aerodynamic Effects, and a Design Tool for Low-Pressure Turbine Flutter
NASA Astrophysics Data System (ADS)
Waite, Joshua Joseph
The successful, efficient, and safe turbine design requires a thorough understanding of the underlying physical phenomena. This research investigates the physical understanding and parameters highly correlated to flutter, an aeroelastic instability prevalent among low pressure turbine (LPT) blades in both aircraft engines and power turbines. The modern way of determining whether a certain cascade of LPT blades is susceptible to flutter is through time-expensive computational fluid dynamics (CFD) codes. These codes converge to solution satisfying the Eulerian conservation equations subject to the boundary conditions of a nodal domain consisting fluid and solid wall particles. Most detailed CFD codes are accompanied by cryptic turbulence models, meticulous grid constructions, and elegant boundary condition enforcements all with one goal in mind: determine the sign (and therefore stability) of the aerodynamic damping. The main question being asked by the aeroelastician, "is it positive or negative?'' This type of thought-process eventually gives rise to a black-box effect, leaving physical understanding behind. Therefore, the first part of this research aims to understand and reveal the physics behind LPT flutter in addition to several related topics including acoustic resonance effects. A percentage of this initial numerical investigation is completed using an influence coefficient approach to study the variation the work-per-cycle contributions of neighboring cascade blades to a reference airfoil. The second part of this research introduces new discoveries regarding the relationship between steady aerodynamic loading and negative aerodynamic damping. Using validated CFD codes as computational wind tunnels, a multitude of low-pressure turbine flutter parameters, such as reduced frequency, mode shape, and interblade phase angle, will be scrutinized across various airfoil geometries and steady operating conditions to reach new design guidelines regarding the influence
Data Comparisons and Summary of the Second Aeroelastic Prediction Workshop
NASA Technical Reports Server (NTRS)
Heeg, Jennifer; Wieseman, Carol D.; Chwalowski, Pawel
2016-01-01
This paper presents the computational results generated by participating teams of the second Aeroelastic Prediction Workshop and compare them with experimental data. Aeroelastic and rigid configurations of the Benchmark Supercritical Wing (BSCW) wind tunnel model served as the focus for the workshop. The comparison data sets include unforced ("steady") system responses, forced pitch oscillations and coupled fluid-structure responses. Integrated coefficients, frequency response functions, and flutter onset conditions are compared. The flow conditions studied were in the transonic range, including both attached and separated flow conditions. Some of the technical discussions that took place at the workshop are summarized.
Probabilistic Aeroelastic Analysis of Turbomachinery Components
NASA Technical Reports Server (NTRS)
Reddy, T. S. R.; Mital, S. K.; Stefko, G. L.
2004-01-01
A probabilistic approach is described for aeroelastic analysis of turbomachinery blade rows. Blade rows with subsonic flow and blade rows with supersonic flow with subsonic leading edge are considered. To demonstrate the probabilistic approach, the flutter frequency, damping and forced response of a blade row representing a compressor geometry is considered. The analysis accounts for uncertainties in structural and aerodynamic design variables. The results are presented in the form of probabilistic density function (PDF) and sensitivity factors. For subsonic flow cascade, comparisons are also made with different probabilistic distributions, probabilistic methods, and Monte-Carlo simulation. The approach shows that the probabilistic approach provides a more realistic and systematic way to assess the effect of uncertainties in design variables on the aeroelastic instabilities and response.
Some considerations on the effects of the P-derivatives on bridge deck flutter
NASA Astrophysics Data System (ADS)
Zhang, Xin; Brownjohn, James Mark William
2005-05-01
Using two degrees of freedom (dof) experimental flutter derivatives to perform three-dimensional flutter analysis for a cable-supported bridge is a widely practiced method. It is important to consider the P-derivatives effect to have more accurate analysis for a long-span bridge. Through a case example, this paper studied some of the issues relating to the P-derivatives effects on flutter. The operational condition in two-dof experiments was discussed. It was suggested that due to the strong aeroelastic coupling effect of the sectional model studied in this research, there was an inherent weakness of two-dof experiments. The effect of the P-derivatives was studied for an example bridge by comparing the flutter analysis results using two-dof and three-dof experimental flutter derivatives.
Sensitivity Analysis of Wing Aeroelastic Responses
NASA Technical Reports Server (NTRS)
Issac, Jason Cherian
1995-01-01
Design for prevention of aeroelastic instability (that is, the critical speeds leading to aeroelastic instability lie outside the operating range) is an integral part of the wing design process. Availability of the sensitivity derivatives of the various critical speeds with respect to shape parameters of the wing could be very useful to a designer in the initial design phase, when several design changes are made and the shape of the final configuration is not yet frozen. These derivatives are also indispensable for a gradient-based optimization with aeroelastic constraints. In this study, flutter characteristic of a typical section in subsonic compressible flow is examined using a state-space unsteady aerodynamic representation. The sensitivity of the flutter speed of the typical section with respect to its mass and stiffness parameters, namely, mass ratio, static unbalance, radius of gyration, bending frequency, and torsional frequency is calculated analytically. A strip theory formulation is newly developed to represent the unsteady aerodynamic forces on a wing. This is coupled with an equivalent plate structural model and solved as an eigenvalue problem to determine the critical speed of the wing. Flutter analysis of the wing is also carried out using a lifting-surface subsonic kernel function aerodynamic theory (FAST) and an equivalent plate structural model. Finite element modeling of the wing is done using NASTRAN so that wing structures made of spars and ribs and top and bottom wing skins could be analyzed. The free vibration modes of the wing obtained from NASTRAN are input into FAST to compute the flutter speed. An equivalent plate model which incorporates first-order shear deformation theory is then examined so it can be used to model thick wings, where shear deformations are important. The sensitivity of natural frequencies to changes in shape parameters is obtained using ADIFOR. A simple optimization effort is made towards obtaining a minimum weight
Nonlinear flutter analysis of stiffened composite panels in supersonic flow
NASA Astrophysics Data System (ADS)
Yuan, Kaihua; Qiu, Zhiping
2010-02-01
The flutter instability of stiffened composite panels subjected to aerodynamic forces in the supersonic flow is investigated. Based on Hamilton’s principle, the aeroelastic model of the composite panel is established by using the von Karman large deflection plate theory, piston theory aerodynamics and the quasi-steady thermal stress theory. Then, using the finite element method along with Bogner-Fox-Schmit elements and three-dimensional beam elements, the nonlinear equations of motion are derived. The effect of stiffening scheme on the flutter critical dynamic pressure is demonstrated through the numerical example, and the nonlinear flutter characteristics of stiffened composite panels are also analyzed in the time domain. This will lay the foundation for design of panel structures employed in aerospace vehicles.
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.
Computational Transonic Flutter Solutions for Cranked Wings by the Direct Eulerian-Lagrangian Method
NASA Astrophysics Data System (ADS)
Mellquist, Erik Charles
In this dissertation, a three-dimensional computational aeroelastic simulation for cranked, highly-swept wings is developed, and solutions are presented for several wing models. The computational model is a fully nonlinear coupled fluid-structure simulation based on the Direct Eulerian-Lagrangian coupling methodology. The wing is modeled using nonlinear modified von Karman plate finite elements. Large deformation is accounted for through the use of element-attached local coordinate systems referenced to a single global coordinate system. The fluid is modeled using the mixed Eulerian-Lagrangian formulation of the classical Euler equations and is discretized using a Galerkin finite element approach on an unstructured tetrahedral mesh. The fluid and structural models are coupled by the Direct Eulerian-Lagrangian method where the finite-element shape functions and the local element coordinate systems are used to describe the fluid-structure boundary without approximation. Time synchronization and spatial accuracy are maintained to ensure accurate exchange of energy between the fluid and the structure. The computational solutions exhibit multiple types of aeroelastic response including transonic limit cycle flutter at a wide range of dynamic pressures, subsonic and supersonic bending-torsion flutter at higher dynamic pressures and a wide range of Mach numbers, and limit cycle oscillation dependent on both Mach number and angle of attack. Shock motion dependent on wing deformation is shown to play a major role in determining the response of the wings, and, depending on the flow conditions, can either stabilize or destabilize the response. Results from the simulations correlate closely with observed wind tunnel test responses.
Application of an improved cell mapping method to bilinear stiffness aeroelastic systems
NASA Astrophysics Data System (ADS)
Ding, Q.; Cooper, J. E.; Leung, A. Y. T.
2005-01-01
A “mapping trajectory pursuit (MTP)” is introduced to improve the cell mapping techniques based on spatial Poincaré sections. Such an improvement enables the cell mapping method to determine the exact properties of all cells with less computer memory and computational time. For the purpose of prediction of the stability boundary as a function of initial conditions (domains of attraction), an initial condition region is defined besides the domain of interest. The proposed CM method is used to analyse the aeroelastic behaviour of an aeroelastic system with bilinear structural nonlinearity. Different types of motions including damped stable motion, limit cycle oscillation, complicated periodic motion, chaotic motion and divergent flutter are determined as a function of initial conditions (domains of attraction). The results compare well with that from stability analysis of the system. The bifurcation diagrams are also obtained using the method to reveal the influence of disturbances on the dynamical behaviour of the system over a broad range of air speed.
An Overview of Recent Developments in Computational Aeroelasticity
NASA Technical Reports Server (NTRS)
Bennett, Robert M.; Edwards, John W.
2004-01-01
The motivation for Computational Aeroelasticity (CA) and the elements of one type of the analysis or simulation process are briefly reviewed. The need for streamlining and improving the overall process to reduce elapsed time and improve overall accuracy is discussed. Further effort is needed to establish the credibility of the methodology, obtain experience, and to incorporate the experience base to simplify the method for future use. Experience with the application of a variety of Computational Aeroelasticity programs is summarized for the transonic flutter of two wings, the AGARD 445.6 wing and a typical business jet wing. There is a compelling need for a broad range of additional flutter test cases for further comparisons. Some existing data sets that may offer CA challenges are presented.
Design of a candidate flutter suppression control law for DAST ARW-2
NASA Technical Reports Server (NTRS)
Adams, W. M., Jr.; Tiffany, S. H.
1984-01-01
A control law is developed to suppress symmetric flutter for a mathematical model of an aeroelastic research vehicle. An implementable control law is attained by including modified LQC (Linear Quadratic Gaussian) design techniques, controller order reduction, and gain scheduling. An alternate (complementary) design approach is illustrated for one flight condition wherein nongradient-based constrained optimization techniques are applied to maximize controller robustness.
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).
Multifunction tests of a frequency domain based flutter suppression system
NASA Technical Reports Server (NTRS)
Christhilf, David M.; Adams, William M., Jr.
1992-01-01
The process is described of analysis, design, digital implementation, and subsonic testing of an active control flutter suppression system for a full span, free-to-roll wind tunnel model of an advanced fighter concept. The design technique uses a frequency domain representation of the plant and used optimization techniques to generate a robust multi input/multi output controller. During testing in a fixed-in-roll configuration, simultaneous suppression of both symmetric and antisymmetric flutter was successfully shown. For a free-to-roll configuration, symmetric flutter was suppressed to the limit of the tunnel test envelope. During aggressive rolling maneuvers above the open-loop flutter boundary, simultaneous flutter suppression and maneuver load control were demonstrated. Finally, the flutter damping controller was reoptimized overnight during the test using combined experimental and analytical frequency domain data, resulting in improved stability robustness.
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.
Development of an aeroelastic methodology for surface morphing rotors
NASA Astrophysics Data System (ADS)
Cook, James R.
transmission of force and deflection information to achieve an aeroelastic coupling updated at each time step. The method is validated first by comparing the integrated aerodynamic work at CFD and CSD nodes to verify work conservation across the interface. Second, the method is verified by comparing the sectional blade loads and deflections of a rotor in hover and in forward flight with experimental data. Finally, stability analyses for pitch/plunge flutter and camber flutter are performed with comprehensive CSD/low-order-aerodynamics and tightly coupled CFD/CSD simulations and compared to analytical solutions of Peters' thin airfoil theory to verify proper aeroelastic behavior. The effects of simple harmonic camber actuation are examined and compared to the response predicted by Peters' finite-state (F-S) theory. In anticipation of active rotor experiments inside enclosed facilities, computational simulations are performed to evaluate the capability of CFD for accurately simulating flow inside enclosed volumes. A computational methodology for accurately simulating a rotor inside a test chamber is developed to determine the influence of test facility components and turbulence modeling and performance predictions. A number of factors that influence the physical accuracy of the simulation, such as temporal resolution, grid resolution, and aeroelasticity are also evaluated.
Code of Federal Regulations, 2013 CFR
2013-01-01
...) Any rational analysis used to predict freedom from flutter, control reversal and divergence must cover...-stabilizer surfaces. (e) For turbopropeller-powered airplanes, the dynamic evaluation must include— (1) Whirl mode degree of freedom which takes into account the stability of the plane of rotation of the...
Code of Federal Regulations, 2014 CFR
2014-01-01
...) Any rational analysis used to predict freedom from flutter, control reversal and divergence must cover...-stabilizer surfaces. (e) For turbopropeller-powered airplanes, the dynamic evaluation must include— (1) Whirl mode degree of freedom which takes into account the stability of the plane of rotation of the...
Aeroelasticity of Nonlinear Tail / Rudder Systems with Freeplay
NASA Astrophysics Data System (ADS)
Rishel, Evan
This thesis details the development of a linear/nonlinear three degree of freedom aeroelastic system designed and manufactured at the University of Washington (UW). Describing function analysis was carried out in the frequency domain. Time domain simulations were carried out to account for all types of motion. Nonlinear aeroelastic behavior may lead to limit cycles which can be captured in the frequency domain using describing function approximation and numerically using Runga-Kutta integration. Linear and nonlinear aeroelastic tests were conducted in the UW 3x3 low-speed wind tunnel to determine the linear flutter speed and frequency of the system as well as its nonlinear behavior when freeplay is introduced. The test data is presented along with the results of the MATLAB-based simulations. The correlation between test and numerical results is very high.
Aeroelastic Analysis Of Versatile Thermal Insulation Panels For Launchers Applications
NASA Astrophysics Data System (ADS)
Carrera, E.; Zappino, E.; Augello, G.; Ferrarese, A.; Montabone, M.
2011-05-01
The aeroelastic behavior of a Versatile Thermal Insulation (VTI) has been investigated. Among the various loadings acting on the panels in this work the attention is payed to fluid structure interaction. e.g. panel flutter phenomena. Known available results from open literature, related to similar problems, permit to analyze the effect of various Mach regimes, including boundary layers thickness effects, in-plane mechanical and thermal loadings, nonlinear effect and amplitude of so called limit cycle oscillations. Dedicated finite element model is developed for the supersonic regime. The model used for coupling orthotropic layered structural model with to Piston Theory aerodynamic models allows the calculations of flutter conditions in case of curved panels supported in a dis- crete number of points. Through this approach the flutter boundaries of the VTI-panel have been investigated.
NASTRAN level 16 demonstration manual updates for aeroelastic analysis of bladed discs
NASA Technical Reports Server (NTRS)
Elchuri, V.; Gallo, A. M.
1980-01-01
A computer program based on state of the art compressor and structural technologies applied to bladed shrouded discs was developed and made operational in NASTRAN level 16. The problems encompassed include aeroelastic analyses, modes, and flutter. The demonstration manual updates are described.
NASA Technical Reports Server (NTRS)
Jutte, Christine V.; Stanford, Bret K.; Wieseman, Carol D.; Moore, James B.
2014-01-01
This work explores the use of tow steered composite laminates, functionally graded metals (FGM), thickness distributions, and curvilinear rib/spar/stringer topologies for aeroelastic tailoring. Parameterized models of the Common Research Model (CRM) wing box have been developed for passive aeroelastic tailoring trade studies. Metrics of interest include the wing weight, the onset of dynamic flutter, and the static aeroelastic stresses. Compared to a baseline structure, the lowest aggregate static wing stresses could be obtained with tow steered skins (47% improvement), and many of these designs could reduce weight as well (up to 14%). For these structures, the trade-off between flutter speed and weight is generally strong, although one case showed both a 100% flutter improvement and a 3.5% weight reduction. Material grading showed no benefit in the skins, but moderate flutter speed improvements (with no weight or stress increase) could be obtained by grading the spars (4.8%) or ribs (3.2%), where the best flutter results were obtained by grading both thickness and material. For the topology work, large weight reductions were obtained by removing an inner spar, and performance was maintained by shifting stringers forward and/or using curvilinear ribs: 5.6% weight reduction, a 13.9% improvement in flutter speed, but a 3.0% increase in stress levels. Flutter resistance was also maintained using straightrotated ribs although the design had a 4.2% lower flutter speed than the curved ribs of similar weight and stress levels were higher. These results will guide the development of a future design optimization scheme established to exploit and combine the individual attributes of these technologies.
Flutter of aircraft engine turbine blades
NASA Astrophysics Data System (ADS)
Panovsky, Josef, Jr.
1997-11-01
The goal of this research is to eliminate occurrences of flutter of low-pressure turbine blades in aircraft engines. Fundamental unsteady aerodynamic experiments in an annular cascade plus correlating analyses are conducted to improve the understanding of the flutter mechanism in these blades and to identify the key flutter parameters. The use of two- and three-dimensional linearized Euler methods for the calculation of the unsteady pressures due to the blade motion are validated through detailed comparison with the experimental data. Unexpected features of the steady and unsteady flows are also investigated using these computational tools. The validated computer codes are used to extend the range of the experimental data in a series of parametric studies, where the influence of mode shape, reduced frequency, and blade loading are investigated. Mode shape is identified as the most important contributor to determining the stability of a blade design. A new stability parameter is introduced to gain additional insight into the key contributors to flutter. This stability parameter is derived from the influence coefficient representation of the cascade, and includes only contributions from the reference blade and its immediate neighbors. This has the effect of retaining the most important contributions while filtering out terms of less significance. Design rules for the preliminary concept phase and procedures for the detailed analysis phase of the typical blade design process are defined. Utilization of these procedures will lead to blade designs which are free of flutter.
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.
Controlled Aeroelastic Response and Airfoil Shaping Using Adaptive Materials and Integrated Systems
NASA Technical Reports Server (NTRS)
Pinkerton, Jennifer L.; McGowan, Anna-Maria R.; Moses, Robert W.; Scott, Robert C.; Heeg, Jennifer
1996-01-01
This paper presents an overview of several activities of the Aeroelasticity Branch at the NASA Langley Research Center in the area of applying adaptive materials and integrated systems for controlling both aircraft aeroelastic response and airfoil shape. The experimental results of four programs are discussed: the Piezoelectric Aeroelastic Response Tailoring Investigation (PARTI); the Adaptive Neural Control of Aeroelastic Response (ANCAR) program; the Actively Controlled Response of Buffet Affected Tails (ACROBAT) program; and the Airfoil THUNDER Testing to Ascertain Characteristics (ATTACH) project. The PARTI program demonstrated active flutter control and significant rcductions in aeroelastic response at dynamic pressures below flutter using piezoelectric actuators. The ANCAR program seeks to demonstrate the effectiveness of using neural networks to schedule flutter suppression control laws. Th,e ACROBAT program studied the effectiveness of a number of candidate actuators, including a rudder and piezoelectric actuators, to alleviate vertical tail buffeting. In the ATTACH project, the feasibility of using Thin-Layer Composite-Uimorph Piezoelectric Driver and Sensor (THUNDER) wafers to control airfoil aerodynamic characteristics was investigated. Plans for future applications are also discussed.
Internal Structural Design of the Common Research Model Wing Box for Aeroelastic Tailoring
NASA Technical Reports Server (NTRS)
Jutte, Christine V.; Stanford, Bret K.; Wieseman, Carol D.
2015-01-01
This work explores the use of alternative internal structural designs within a full-scale wing box structure for aeroelastic tailoring, with a focus on curvilinear spars, ribs, and stringers. The baseline wing model is a fully-populated, cantilevered wing box structure of the Common Research Model (CRM). Metrics of interest include the wing weight, the onset of dynamic flutter, and the static aeroelastic stresses. Twelve parametric studies alter the number of internal structural members along with their location, orientation, and curvature. Additional evaluation metrics are considered to identify design trends that lead to lighter-weight, aeroelastically stable wing designs. The best designs of the individual studies are compared and discussed, with a focus on weight reduction and flutter resistance. The largest weight reductions were obtained by removing the inner spar, and performance was maintained by shifting stringers forward and/or using curvilinear ribs: 5.6% weight reduction, a 13.9% improvement in flutter speed, but a 3.0% increase in stress levels. Flutter resistance was also maintained using straight-rotated ribs although the design had a 4.2% lower flutter speed than the curved ribs of similar weight and stress levels were higher. For some configurations, the differences between curved and straight ribs were smaller, which provides motivation for future optimization-based studies to fully exploit the trade-offs.
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.
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.
NASA Technical Reports Server (NTRS)
Kroeger, R. A.
1977-01-01
A complete ground vibration and aeroelastic analysis was made of a modified version of the Grumman American Yankee. The aircraft had been modified for four empennage configurations, a wing boom was added, a spin chute installed and provisions included for large masses in the wing tip to vary the lateral and directional inertia. Other minor changes were made which have much less influence on the flutter and vibrations. Neither static divergence nor aileron reversal was considered since the wing structure was not sufficiently changed to affect its static aeroelastic qualities. The aircraft was found to be free from flutter in all of the normal modes explored in the ground shake test. The analysis demonstrated freedom from flutter up to 214 miles per hour.
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.
F/A-18 E/F flutter clearance model in the Langley TDT
NASA Technical Reports Server (NTRS)
1994-01-01
An 18 percent aeroelastically-scaled, full span F/A-18 E/F model was tested during multiple wind-tunnel entries in the Langley Transonic Dynamics Tunnel. The primary purpose of these entries was to assist in clearing the flight vehicle design of flutter within its operating envelope. The wind-tunnel model was tested on a string and on a cable-mount system (as shown). All lifting surfaces were flutter cleared up to M=1.2 with the model string mounted. The model was then flutter cleared on the cable-mount system to assess the influence of rigid-body dynamics and fuselage flexibility on flutter. Several configuration parametric studies were also completed, including many external store configurations.
The use of the Regier number in the structural design with flutter constraints
NASA Technical Reports Server (NTRS)
Dunn, H. J.; Doggett, Robert V., Jr.
1994-01-01
This preliminary investigation introduces the use of the Regier number as a flutter constraint criterion for aeroelastic structural optimization. Artificial neural network approximations are used to approximate the flutter criterion requirements as a function of the design Mach number and the parametric variables defining the aspect ratio, center of gravity, taper ratio, mass ratio, and pitch inertia of the wing. The presented approximations are simple enough to be used in the preliminary design stage without a well defined structural model. An example problem for a low-speed, high-aspect-ratio, light-aircraft wing is presented. The example problem is analyzed for the flutter Mach number using doublet lattice aerodynamics and the PK solution method. The use of the Regier number constraint criterion to optimize the example problem for minimum structural mass while maintaining a constant flutter Mach number is demonstrated.
NASA Astrophysics Data System (ADS)
Firouz-Abadi, R. D.; Askarian, A. R.; Zarifian, P.
2013-01-01
This paper aims to investigate aeroelastic stability boundary of subsonic wings under the effect of thrust of two engines. The wing structure is modeled as a tapered composite box-beam. Moreover, an indicial function based model is used to calculate the unsteady lift and moment distribution along the wing span in subsonic compressible flow. The two jet engines mounted on the wing are modeled as concentrated masses and the effect of thrust of each engine is applied as a follower force. Using Hamilton's principle along with Galerkin's method, the governing equations of motion are derived, then the obtained equations are solved in frequency domain using the K-method and the aeroelastic instability conditions are determined. The flutter analysis results of four example wings are compared with the experimental and analytical results in the literature and good agreements are achieved which validate the present model. Furthermore, based on several case studies on a reference wing, some attempts are performed to analyze the effect of thrust on the stability margin of the wing and some conclusions are outlined.
NASA Technical Reports Server (NTRS)
Reddy, T. S. R.; Bakhle, Milind A.; Huff, Dennis L.; Swafford, Timothy W.
1992-01-01
This report presents, in two parts, a dynamic aeroelastic stability (flutter) analysis of a cascade of blades in supersonic axial flow. Each blade of the cascade is modeled as a typical section having pitching and plunging degrees of freedom. Aerodynamic forces are obtained from a time accurate, unsteady, two-dimensional cascade solver based on the Euler equations. The solver uses a time marching flux-difference splitting (FDS) scheme. Flutter stability is analyzed in the frequency domain. The unsteady force coefficients required in the analysis are obtained by harmonically oscillating (HO) the blades for a given flow condition, oscillation frequency, and interblade phase angle. The calculated time history of the forces is then Fourier decomposed to give the required unsteady force coefficients. An influence coefficient (IC) method and a pulse response (PR) method are also implemented to reduce the computational time for the calculation of the unsteady force coefficients for any phase angle and oscillation frequency. Part 1, this report, presents these analysis methods and their validation by comparison with results obtained from linear theory for a selected flat plate cascade geometry. A typical calculation for a rotor airfoil is also included to show the applicability of the present solver for airfoil configurations. The predicted unsteady aerodynamic forces for a selected flat plate cascade geometry and flow conditions correlated well with those obtained from linear theory for different interblade phase angles and oscillation frequencies. All the three methods of predicting unsteady force coefficients, namely, HO, IC, and PR, showed good correlations with each other. It was established that only a single calculation with four blade passages is required to calculate the aerodynamic forces for any phase angle for a cascade consisting of any number of blades, for any value of the oscillation frequency. Flutter results, including mistuning effects, for a cascade of
Control Law Design in a Computational Aeroelasticity Environment
NASA Technical Reports Server (NTRS)
Newsom, Jerry R.; Robertshaw, Harry H.; Kapania, Rakesh K.
2003-01-01
A methodology for designing active control laws in a computational aeroelasticity environment is given. The methodology involves employing a systems identification technique to develop an explicit state-space model for control law design from the output of a computational aeroelasticity code. The particular computational aeroelasticity code employed in this paper solves the transonic small disturbance aerodynamic equation using a time-accurate, finite-difference scheme. Linear structural dynamics equations are integrated simultaneously with the computational fluid dynamics equations to determine the time responses of the structure. These structural responses are employed as the input to a modern systems identification technique that determines the Markov parameters of an "equivalent linear system". The Eigensystem Realization Algorithm is then employed to develop an explicit state-space model of the equivalent linear system. The Linear Quadratic Guassian control law design technique is employed to design a control law. The computational aeroelasticity code is modified to accept control laws and perform closed-loop simulations. Flutter control of a rectangular wing model is chosen to demonstrate the methodology. Various cases are used to illustrate the usefulness of the methodology as the nonlinearity of the aeroelastic system is increased through increased angle-of-attack changes.
Flutter of Darrieus wind turbine blades
NASA Technical Reports Server (NTRS)
Ham, N. D.
1978-01-01
The testing of Darrieus wind turbines has indicated that under certain conditions, serious vibrations of the blades can occur, involving flatwise bending, torsion, and chordwise bending. A theoretical method of predicting the aeroelastic stability of the coupled bending and torsional motion of such blades with a view to determining the cause of these vibrations, and a means of suppressing them was developed.
Code of Federal Regulations, 2012 CFR
2012-01-01
... analysis used to predict freedom from flutter, control reversal and divergence must cover all speeds up to... criteria, and (iii) Has fixed-fin and fixed-stabilizer surfaces. (e) For turbopropeller-powered airplanes, the dynamic evaluation must include— (1) Whirl mode degree of freedom which takes into account...
Panel Flutter and Sonic Fatigue Analysis for RLV
NASA Technical Reports Server (NTRS)
Mei, Chuh; Cheng, Guangfeng
2001-01-01
A methodology is presented for the flutter analysis of the seal of thermal protection system (TPS) panel of X-33 Advanced Technology Demonstrator test vehicle. The seal is simulated as a two-dimensional cantilevered panel with an elastic stopper, which is modeled as an equivalent spring. This cantilever beam-spring model under the aerodynamic pressure at supersonic speeds turns out to be an impact nonlinear dynamic system. The flutter analysis of the seal is thus carried out using, time domain numerical simulation with a displacement stability criterion. The flutter boundary of the seal is further verified with a family of three traditional and one nontraditional panel flutter models. The frequency domain method that applies eigenanalysis on the traditional panel flutter problem was used. The results showed that the critical dynamic pressure could be more than doubled with properly chosen material for the base stopper. The proposed methodology can be easily extended to three-dimensional panel seals with flow angularity.
Three-Dimensional Aeroelastic and Aerothermoelastic Behavior in Hypersonic Flow
NASA Technical Reports Server (NTRS)
McNamara, Jack J.; Friedmann, Peretz P.; Powell, Kenneth G.; Thuruthimattam, Biju J.; Bartels, Robert E.
2005-01-01
The aeroelastic and aerothermoelastic behavior of three-dimensional configurations in hypersonic flow regime are studied. The aeroelastic behavior of a low aspect ratio wing, representative of a fin or control surface on a generic hypersonic vehicle, is examined using third order piston theory, Euler and Navier-Stokes aerodynamics. The sensitivity of the aeroelastic behavior generated using Euler and Navier-Stokes aerodynamics to parameters governing temporal accuracy is also examined. Also, a refined aerothermoelastic model, which incorporates the heat transfer between the fluid and structure using CFD generated aerodynamic heating, is used to examine the aerothermoelastic behavior of the low aspect ratio wing in the hypersonic regime. Finally, the hypersonic aeroelastic behavior of a generic hypersonic vehicle with a lifting-body type fuselage and canted fins is studied using piston theory and Euler aerodynamics for the range of 2.5 less than or equal to M less than or equal to 28, at altitudes ranging from 10,000 feet to 80,000 feet. This analysis includes a study on optimal mesh selection for use with Euler aerodynamics. In addition to the aeroelastic and aerothermoelastic results presented, three time domain flutter identification techniques are compared, namely the moving block approach, the least squares curve fitting method, and a system identification technique using an Auto-Regressive model of the aeroelastic system. In general, the three methods agree well. The system identification technique, however, provided quick damping and frequency estimations with minimal response record length, and therefore o ers significant reductions in computational cost. In the present case, the computational cost was reduced by 75%. The aeroelastic and aerothermoelastic results presented illustrate the applicability of the CFL3D code for the hypersonic flight regime.
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.
Centrifugal Compressor Aeroelastic Analysis Code
NASA Astrophysics Data System (ADS)
Keith, Theo G., Jr.; Srivastava, Rakesh
2002-01-01
Centrifugal compressors are very widely used in the turbomachine industry where low mass flow rates are required. Gas turbine engines for tanks, rotorcraft and small jets rely extensively on centrifugal compressors for rugged and compact design. These compressors experience problems related with unsteadiness of flowfields, such as stall flutter, separation at the trailing edge over diffuser guide vanes, tip vortex unsteadiness, etc., leading to rotating stall and surge. Considerable interest exists in small gas turbine engine manufacturers to understand and eventually eliminate the problems related to centrifugal compressors. The geometric complexity of centrifugal compressor blades and the twisting of the blade passages makes the linear methods inapplicable. Advanced computational fluid dynamics (CFD) methods are needed for accurate unsteady aerodynamic and aeroelastic analysis of centrifugal compressors. Most of the current day industrial turbomachines and small aircraft engines are designed with a centrifugal compressor. With such a large customer base and NASA Glenn Research Center being, the lead center for turbomachines, it is important that adequate emphasis be placed on this area as well. Currently, this activity is not supported under any project at NASA Glenn.
An analytical and experimental study to investigate flutter suppression via piezoelectric actuation
NASA Astrophysics Data System (ADS)
Heeg, Jennifer
The objective was to analytically and experimentally study the capabilities of adaptive material plate actuators for suppressing flutter. The validity of analytical modeling techniques for piezoelectric materials was also investigated. 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 airfoil and a flexible mount system which permitted a translational and a rotational degree of freedom. It was designed such that flutter was encounted within the testing envelope of the wind tunnel. Actuators, made of piezoelectric material were affixed to leaf springs of the mount system. Each degree of freedom was controlled by a separate leaf spring. Command signals, applied to the piezoelectric actuators, exerted control over the damping and stiffness properties. A mathematical aeroservoelastic model was constructed 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. It demonstrates that small, carefully placed actuating plates can be used effectively to control aeroelastic response.
Non-linear aeroelastic prediction for aircraft applications
NASA Astrophysics Data System (ADS)
de C. Henshaw, M. J.; Badcock, K. J.; Vio, G. A.; Allen, C. B.; Chamberlain, J.; Kaynes, I.; Dimitriadis, G.; Cooper, J. E.; Woodgate, M. A.; Rampurawala, A. M.; Jones, D.; Fenwick, C.; Gaitonde, A. L.; Taylor, N. V.; Amor, D. S.; Eccles, T. A.; Denley, C. J.
2007-05-01
Current industrial practice for the prediction and analysis of flutter relies heavily on linear methods and this has led to overly conservative design and envelope restrictions for aircraft. Although the methods have served the industry well, it is clear that for a number of reasons the inclusion of non-linearity in the mathematical and computational aeroelastic prediction tools is highly desirable. The increase in available and affordable computational resources, together with major advances in algorithms, mean that non-linear aeroelastic tools are now viable within the aircraft design and qualification environment. The Partnership for Unsteady Methods in Aerodynamics (PUMA) Defence and Aerospace Research Partnership (DARP) was sponsored in 2002 to conduct research into non-linear aeroelastic prediction methods and an academic, industry, and government consortium collaborated to address the following objectives: To develop useable methodologies to model and predict non-linear aeroelastic behaviour of complete aircraft. To evaluate the methodologies on real aircraft problems. To investigate the effect of non-linearities on aeroelastic behaviour and to determine which have the greatest effect on the flutter qualification process. These aims have been very effectively met during the course of the programme and the research outputs include: New methods available to industry for use in the flutter prediction process, together with the appropriate coaching of industry engineers. Interesting results in both linear and non-linear aeroelastics, with comprehensive comparison of methods and approaches for challenging problems. Additional embryonic techniques that, with further research, will further improve aeroelastics capability. This paper describes the methods that have been developed and how they are deployable within the industrial environment. We present a thorough review of the PUMA aeroelastics programme together with a comprehensive review of the relevant research
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.
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.
Aeroelastic dynamic response and control of an airfoil section with control surface nonlinearities
NASA Astrophysics Data System (ADS)
Li, Daochun; Guo, Shijun; Xiang, Jinwu
2010-10-01
Nonlinearities in aircraft mechanisms are inevitable, especially in the control system. It is necessary to investigate the effects of them on the dynamic response and control performance of aeroelastic system. In this paper, based on the state-dependent Riccati equation method, a state feedback suboptimal control law is derived for aeroelastic response and flutter suppression of a three degree-of-freedom typical airfoil section. With the control law designed, nonlinear effects of freeplay in the control surface and time delay between the control input and actuator are investigated by numerical approach. A cubic nonlinearity in pitch degree is adopted to prevent the aeroelastic responses from divergence when the flow velocity exceeds the critical flutter speed. For the system with a freeplay, the responses of both open- and closed-loop systems are determined with Runge-Kutta algorithm in conjunction with Henon's method. This method is used to locate the switching points accurately and efficiently as the system moves from one subdomain into another. The simulation results show that the freeplay leads to a forward phase response and a slight increase of flutter speed of the closed-loop system. The effect of freeplay on the aeroelastic response decreases as the flow velocity increases. The time delay between the control input and actuator may impair control performance and cause high-frequency motion and quasi-periodic vibration.
Aeroelasticity of Axially Loaded Aerodynamic Structures for Truss-Braced Wing Aircraft
NASA Technical Reports Server (NTRS)
Nguyen, Nhan; Ting, Eric; Lebofsky, Sonia
2015-01-01
This paper presents an aeroelastic finite-element formulation for axially loaded aerodynamic structures. The presence of axial loading causes the bending and torsional sitffnesses to change. For aircraft with axially loaded structures such as the truss-braced wing aircraft, the aeroelastic behaviors of such structures are nonlinear and depend on the aerodynamic loading exerted on these structures. Under axial strain, a tensile force is created which can influence the stiffness of the overall aircraft structure. This tension stiffening is a geometric nonlinear effect that needs to be captured in aeroelastic analyses to better understand the behaviors of these types of aircraft structures. A frequency analysis of a rotating blade structure is performed to demonstrate the analytical method. A flutter analysis of a truss-braced wing aircraft is performed to analyze the effect of geometric nonlinear effect of tension stiffening on the flutter speed. The results show that the geometric nonlinear tension stiffening effect can have a significant impact on the flutter speed prediction. In general, increased wing loading results in an increase in the flutter speed. The study illustrates the importance of accounting for the geometric nonlinear tension stiffening effect in analyzing the truss-braced wing aircraft.
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.
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.
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
Computerized Analysis Of Helicopter-Rotor Aeroelasticity
NASA Technical Reports Server (NTRS)
Reddy, T. S. R.
1988-01-01
Analysis of aeroelastic stability of helicopter rotor automated. Symbolic-manipulation program, HESL, written in FORTRAN, used to aid in derivation of government equations of motion for elastic-bladed rotor. Operates both on expressions and matrices. By transferring some burden of algebraic manipulations from human analyst to computer, program reduces tedium analysis and conequent opportunity for errors.
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.
Design for active and passive flutter suppression and gust alleviation. Ph.D. Thesis
NASA Technical Reports Server (NTRS)
Karpel, M.
1981-01-01
Analytical design techniques for active and passive control of aeroelastic systems are based on a rational approximation of the unsteady aerodynamic loads in the entire Laplace domain, which yields matrix equations of motion with constant coefficients. Some existing schemes are reviewed, the matrix Pade approximant is modified, and a technique which yields a minimal number of augmented states for a desired accuracy is presented. The state-space aeroelastic model is used to design an active control system for simultaneous flutter suppression and gust alleviation. The design target is for a continuous controller which transfers some measurements taken on the vehicle to a control command applied to a control surface. Structural modifications are formulated in a way which enables the treatment of passive flutter suppression system with the same procedures by which active control systems are designed.
Optimal Topology of Aircraft Rib and Spar Structures under Aeroelastic Loads
NASA Technical Reports Server (NTRS)
Stanford, Bret K.; Dunning, Peter D.
2014-01-01
Several topology optimization problems are conducted within the ribs and spars of a wing box. It is desired to locate the best position of lightening holes, truss/cross-bracing, etc. A variety of aeroelastic metrics are isolated for each of these problems: elastic wing compliance under trim loads and taxi loads, stress distribution, and crushing loads. Aileron effectiveness under a constant roll rate is considered, as are dynamic metrics: natural vibration frequency and flutter. This approach helps uncover the relationship between topology and aeroelasticity in subsonic transport wings, and can therefore aid in understanding the complex aircraft design process which must eventually consider all these metrics and load cases simultaneously.
NASA Technical Reports Server (NTRS)
Stanford, Bret K.; Jutte, Christine V.
2016-01-01
A series of aeroelastic optimization problems are solved on a high aspect ratio wingbox of the Common Research Model, in an effort to minimize structural mass under coupled stress, buckling, and flutter constraints. Two technologies are of particular interest: tow steered composite laminate skins and curvilinear stiffeners. Both methods are found to afford feasible reductions in mass over their non-curvilinear structural counterparts, through both distinct and shared mechanisms for passively controlling aeroelastic performance. Some degree of diminishing returns are seen when curvilinear stiffeners and curvilinear fiber tow paths are used simultaneously.
NASA Technical Reports Server (NTRS)
Lehman, L. L.
1981-01-01
A computational technique has been developed for performing preliminary design aeroelastic analyses of large aspect ratio lifting surfaces. This technique, applicable to both fixed and rotating wing configurations, is based upon a formulation of the structural equilibrium equations in terms of a hybrid state vector containing generalized force and displacement variables. An integrating matrix is employed to solve these equations for divergence and flutter eigenvalues and steady aeroelastic deformation. Results are presented for simple examples which verify the technique and demonstrate how it can be applied to analyze lifting surfaces, including those constructed from composite materials.
The benchmark aeroelastic models program: Description and highlights of initial results
NASA Technical Reports Server (NTRS)
Bennett, Robert M.; Eckstrom, Clinton V.; Rivera, Jose A., Jr.; Dansberry, Bryan E.; Farmer, Moses G.; Durham, Michael H.
1992-01-01
An experimental effort was implemented in aeroelasticity called the Benchmark Models Program. The primary purpose of this program is to provide the necessary data to evaluate computational fluid dynamic codes for aeroelastic analysis. It also focuses on increasing the understanding of the physics of unsteady flows and providing data for empirical design. An overview is given of this program and some results obtained in the initial tests are highlighted. The tests that were completed include measurement of unsteady pressures during flutter of a rigid wing with an NACA 0012 airfoil section and dynamic response measurements of a flexible rectangular wing with a thick circular arc airfoil undergoing shock boundary layer oscillations.
The benchmark aeroelastic models program: Description and highlights of initial results
NASA Technical Reports Server (NTRS)
Bennett, Robert M.; Eckstrom, Clinton V.; Rivera, Jose A., Jr.; Dansberry, Bryan E.; Farmer, Moses G.; Durham, Michael H.
1991-01-01
An experimental effort was implemented in aeroelasticity called the Benchmark Models Program. The primary purpose of this program is to provide the necessary data to evaluate computational fluid dynamic codes for aeroelastic analysis. It also focuses on increasing the understanding of the physics of unsteady flows and providing data for empirical design. An overview is given of this program and some results obtained in the initial tests are highlighted. The tests that were completed include measurement of unsteady pressures during flutter of rigid wing with a NACA 0012 airfoil section and dynamic response measurements of a flexible rectangular wing with a thick circular arc airfoil undergoing shock boundary layer oscillations.
NASA Technical Reports Server (NTRS)
Doggett, R. V., Jr.; Abel, I.; Ruhlin, C. L.
1976-01-01
A status report and review of wind tunnel model experimental techniques that have been developed to study and validate the use of active control technology for the minimization of aeroelastic response are presented. Modeling techniques, test procedures, and data analysis methods used in three model studies are described. The studies include flutter mode suppression on a delta-wing model, flutter mode suppression and ride quality control on a 1/30-size model of the B-52 CCV airplane, and an active lift distribution control system on a 1/22 size C-5A model.
NASA Technical Reports Server (NTRS)
Livne, Eli; Mineau, David
1997-01-01
Analytical sensitivities of panel flutter constraints with respect to panel shape as well as thickness and material properties are derived and numerically tested. Cases of fixed in-plane loads and cases in which in-plane loads are variable (depending on panel and overall wing shape as well as material and sizing design variables) are considered. Accuracy of approximations and range of move limits required are studied in preparation for integration with nonlinear programming/approximation concept aeroelastic design synthesis methodology.
Nonlinear Stochastic Flutter of a Cantilever Wing with Joint Relaxation and Random Loading
2008-02-21
analytical, numerical, and experimental techniques. The influence of span-wise distribution of bending and torsion stiffness uncertainties on the flutter...perturbation method is very accurate for all levels of bending stiffness uncertainty examined, but the method loses its accuracy at upper levels of...in aeroelastic structures. The present work is an attempt to employ the K-L expansion to discretize the span-wise distribution of bending and torsion
Robust control design techniques for active flutter suppression
NASA Technical Reports Server (NTRS)
Ozbay, Hitay; Bachmann, Glen R.
1994-01-01
In this paper, an active flutter suppression problem is studied for a thin airfoil in unsteady aerodynamics. The mathematical model of this system is infinite dimensional because of Theodorsen's function which is irrational. Several second order approximations of Theodorsen's function are compared. A finite dimensional model is obtained from such an approximation. We use H infinity control techniques to find a robustly stabilizing controller for active flutter suppression.
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.
Aeroelastic analysis for propellers - mathematical formulations and program user's manual
NASA Technical Reports Server (NTRS)
Bielawa, R. L.; Johnson, S. A.; Chi, R. M.; Gangwani, S. T.
1983-01-01
Mathematical development is presented for a specialized propeller dedicated version of the G400 rotor aeroelastic analysis. The G400PROP analysis simulates aeroelastic characteristics particular to propellers such as structural sweep, aerodynamic sweep and high subsonic unsteady airloads (both stalled and unstalled). Formulations are presented for these expanded propeller related methodologies. Results of limited application of the analysis to realistic blade configurations and operating conditions which include stable and unstable stall flutter test conditions are given. Sections included for enhanced program user efficiency and expanded utilization include descriptions of: (1) the structuring of the G400PROP FORTRAN coding; (2) the required input data; and (3) the output results. General information to facilitate operation and improve efficiency is also provided.
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
NASA Astrophysics Data System (ADS)
Song, Pengchao
Recent studies of the occurrence of post-flutter limit cycle oscillations (LCO) of the F-16 have provided good support to the long-standing hypothesis that this phenomenon involves a nonlinear structural damping. A potential mechanism for the appearance of nonlinearity in the damping are the nonlinear geometric effects that arise when the deformations become large enough to exceed the linear regime. In this light, the focus of this investigation is first on extending nonlinear reduced order modeling (ROM) methods to include viscoelasticity which is introduced here through a linear Kelvin-Voigt model in the undeformed configuration. Proceeding with a Galerkin approach, the ROM governing equations of motion are obtained and are found to be of a generalized van der Pol-Duffing form with parameters depending on the structure and the chosen basis functions. An identification approach of the nonlinear damping parameters is next proposed which is applicable to structures modeled within commercial finite element software. The effects of this nonlinear damping mechanism on the post-flutter response is next analyzed on the Goland wing through time-marching of the aeroelastic equations comprising a rational fraction approximation of the linear aerodynamic forces. It is indeed found that the nonlinearity in the damping can stabilize the unstable aerodynamics and lead to finite amplitude limit cycle oscillations even when the stiffness related nonlinear geometric effects are neglected. The incorporation of these latter effects in the model is found to further decrease the amplitude of LCO even though the dominant bending motions do not seem to stiffen as the level of displacements is increased in static analyses.
Airloads, wakes, and aeroelasticity
NASA Technical Reports Server (NTRS)
Johnson, Wayne
1990-01-01
Fundamental considerations regarding the theory of modeling of rotary wing airloads, wakes, and aeroelasticity are presented. The topics covered are: airloads and wakes, including lifting-line theory, wake models and nonuniform inflow, free wake geometry, and blade-vortex interaction; aerodynamic and wake models for aeroelasticity, including two-dimensional unsteady aerodynamics and dynamic inflow; and airloads and structural dynamics, including comprehensive airload prediction programs. Results of calculations and correlations are presented.
Toward efficient aeroelastic energy harvesting through limit cycle shaping
NASA Astrophysics Data System (ADS)
Kirschmeier, Benjamin; Bryant, Matthew
2016-04-01
Increasing demand to harvest energy from renewable resources has caused significant research interest in unsteady aerodynamic and hydrodynamic phenomena. Apart from the traditional horizontal axis wind turbines, there has been significant growth in the study of bio-inspired oscillating wings for energy harvesting. These systems are being built to harvest electricity for wireless devices, as well as for large scale mega-watt power generation. Such systems can be driven by aeroelastic flutter phenomena which, beyond a critical wind speed, will cause the system to enter into limitcycle oscillations. When the airfoil enters large amplitude, high frequency motion, leading and trailing edge vortices form and, when properly synchronized with the airfoil kinematics, enhance the energy extraction efficiency of the device. A reduced order dynamic stall model is employed on a nonlinear aeroelastic structural model to investigate whether the parameters of a fully passive aeroelastic device can be tuned to produce limit cycle oscillations at desired kinematics. This process is done through an optimization technique to find the necessary structural parameters to achieve desired structural forces and moments corresponding to a target limit cycle. Structural nonlinearities are explored to determine the essential nonlinearities such that the system's limit cycle closely matches the desired kinematic trajectory. The results from this process demonstrate that it is possible to tune system parameters such that a desired limit cycle trajectory can be achieved. The simulations also demonstrate that the high efficiencies predicted by previous computational aerodynamics studies can be achieved in fully passive aeroelastic devices.
Estimation of the Hopf Bifurcation Point for Aeroelastic Systems
NASA Astrophysics Data System (ADS)
SEDAGHAT, A.; COOPER, J. E.; LEUNG, A. Y. T.; WRIGHT, J. R.
2001-11-01
The estimation of the Hopf bifurcation point is an important prerequisite for the non-linear analysis of non-linear instabilities in aircraft using the classical normal form theory. For unsteady transonic aerodynamics, the aeroelastic response is frequency-dependent and therefore a very costly trial-and-error and iterative scheme, frequency-matching, is used to determine flutter conditions. Furthermore, the standard algebraic methods have usually been used for systems not bigger than two degrees of freedom and do not appear to have been applied for frequency-dependent aerodynamics. In this study, a procedure is developed to produce and solve algebraic equations for any order aeroelastic systems, with and without frequency-dependent aerodynamics, to predict the Hopf bifurcation point. The approach performs the computation in a single step using symbolic programming and does not require trial and error and repeated calculations at various speeds required when using classical iterative methods. To investigate the validity of the approach, a Hancock two-degrees-of-freedom aeroelastic wing model and a multi-degree-of-freedom cantilever wind model were studied in depth. Hancock experimental data was used for curve fitting the unsteady aerodynamic damping term as a function of frequency. Fairly close agreement was obtained between the analytical and simulated aeroelastic solutions with and without frequency-dependent aerodynamics.
Development of an Aeroelastic Analysis Including a Viscous Flow Model
NASA Technical Reports Server (NTRS)
Keith, Theo G., Jr.; Bakhle, Milind A.
2001-01-01
Under this grant, Version 4 of the three-dimensional Navier-Stokes aeroelastic code (TURBO-AE) has been developed and verified. The TURBO-AE Version 4 aeroelastic code allows flutter calculations for a fan, compressor, or turbine blade row. This code models a vibrating three-dimensional bladed disk configuration and the associated unsteady flow (including shocks, and viscous effects) to calculate the aeroelastic instability using a work-per-cycle approach. Phase-lagged (time-shift) periodic boundary conditions are used to model the phase lag between adjacent vibrating blades. The direct-store approach is used for this purpose to reduce the computational domain to a single interblade passage. A disk storage option, implemented using direct access files, is available to reduce the large memory requirements of the direct-store approach. Other researchers have implemented 3D inlet/exit boundary conditions based on eigen-analysis. Appendix A: Aeroelastic calculations based on three-dimensional euler analysis. Appendix B: Unsteady aerodynamic modeling of blade vibration using the turbo-V3.1 code.
Aeroelastic Analysis of a Distributed Electric Propulsion Wing
NASA Technical Reports Server (NTRS)
Massey, Steven J.; Stanford, Bret K.; Wieseman, Carol D.; Heeg, Jennifer
2017-01-01
An aeroelastic analysis of a prototype distributed electric propulsion wing is presented. Results using MSC Nastran (Registered Trademark) doublet lattice aerodynamics are compared to those based on FUN3D Reynolds Averaged Navier- Stokes aerodynamics. Four levels of grid refinement were examined for the FUN3D solutions and solutions were seen to be well converged. It was found that no oscillatory instability existed, only that of divergence, which occurred in the first bending mode at a dynamic pressure of over three times the flutter clearance condition.
Uncertainty Quantification in Aeroelasticity
NASA Astrophysics Data System (ADS)
Beran, Philip; Stanford, Bret; Schrock, Christopher
2017-01-01
Physical interactions between a fluid and structure, potentially manifested as self-sustained or divergent oscillations, can be sensitive to many parameters whose values are uncertain. Of interest here are aircraft aeroelastic interactions, which must be accounted for in aircraft certification and design. Deterministic prediction of these aeroelastic behaviors can be difficult owing to physical and computational complexity. New challenges are introduced when physical parameters and elements of the modeling process are uncertain. By viewing aeroelasticity through a nondeterministic prism, where key quantities are assumed stochastic, one may gain insights into how to reduce system uncertainty, increase system robustness, and maintain aeroelastic safety. This article reviews uncertainty quantification in aeroelasticity using traditional analytical techniques not reliant on computational fluid dynamics; compares and contrasts this work with emerging methods based on computational fluid dynamics, which target richer physics; and reviews the state of the art in aeroelastic optimization under uncertainty. Barriers to continued progress, for example, the so-called curse of dimensionality, are discussed.
Aeroelastic Airworthiness Assesment of the Adaptive Compliant Trailing Edge Flaps
NASA Technical Reports Server (NTRS)
Herrera, Claudia Y.; Spivey, Natalie D.; Lung, Shun-fat; Ervin, Gregory; Flick, Peter
2015-01-01
The Adaptive Compliant Trailing Edge (ACTE) demonstrator is a joint task under the National Aeronautics and Space Administration Environmentally Responsible Aviation Project in partnership with the Air Force Research Laboratory and FlexSys, Inc. (Ann Arbor, Michigan). The project goal is to develop advanced technologies that enable environmentally friendly aircraft, such as adaptive compliant technologies. The ACTE demonstrator flight-test program encompassed replacing the Fowler flaps on the SubsoniC Aircraft Testbed, a modified Gulfstream III (Gulfstream Aerospace, Savannah, Georgia) aircraft, with control surfaces developed by FlexSys. The control surfaces developed by FlexSys are a pair of uniquely-designed unconventional flaps to be used as lifting surfaces during flight-testing to validate their structural effectiveness. The unconventional flaps required a multidisciplinary airworthiness assessment to prove they could withstand the prescribed flight envelope. Several challenges were posed due to the large deflections experienced by the structure, requiring non-linear analysis methods. The aeroelastic assessment necessitated both conventional and extensive testing and analysis methods. A series of ground vibration tests (GVTs) were conducted to provide modal characteristics to validate and update finite element models (FEMs) used for the flutter analyses for a subset of the various flight configurations. Numerous FEMs were developed using data from FlexSys and the ground tests. The flap FEMs were then attached to the aircraft model to generate a combined FEM that could be analyzed for aeroelastic instabilities. The aeroelastic analysis results showed the combined system of aircraft and flaps were predicted to have the required flutter margin to successfully demonstrate the adaptive compliant technology. This paper documents the details of the aeroelastic airworthiness assessment described, including the ground testing and analyses, and subsequent flight
MAVRIC Flutter Model Transonic Limit Cycle Oscillation Test
NASA Technical Reports Server (NTRS)
Edwards, John W.; Schuster, David M.; Spain, Charles V.; Keller, Donald F.; Moses, Robert W.
2001-01-01
The Models for Aeroelastic Validation Research Involving Computation semi-span wind-tunnel model (MAVRIC-I), a business jet wing-fuselage flutter model, was tested in NASA Langley's Transonic Dynamics Tunnel with the goal of obtaining experimental data suitable for Computational Aeroelasticity code validation at transonic separation onset conditions. This research model is notable for its inexpensive construction and instrumentation installation procedures. Unsteady pressures and wing responses were obtained for three wingtip configurations of clean, tipstore, and winglet. Traditional flutter boundaries were measured over the range of M = 0.6 to 0.9 and maps of Limit Cycle Oscillation (LCO) behavior were made in the range of M = 0.85 to 0.95. Effects of dynamic pressure and angle-of-attack were measured. Testing in both R134a heavy gas and air provided unique data on Reynolds number, transition effects, and the effect of speed of sound on LCO behavior. The data set provides excellent code validation test cases for the important class of flow conditions involving shock-induced transonic flow separation onset at low wing angles, including LCO behavior.
Flutter of wings involving a locally distributed flexible control surface
NASA Astrophysics Data System (ADS)
Mozaffari-Jovin, S.; Firouz-Abadi, R. D.; Roshanian, J.
2015-11-01
This paper undertakes to facilitate appraisal of aeroelastic interaction of a locally distributed, flap-type control surface with aircraft wings operating in a subsonic potential flow field. The extended Hamilton's principle serves as a framework to ascertain the Euler-Lagrange equations for coupled bending-torsional-flap vibration. An analytical solution to this boundary-value problem is then accomplished by assumed modes and the extended Galerkin's method. The developed aeroelastic model considers both the inherent flexibility of the control surface displaced on the wing and the inertial coupling between these two flexible bodies. The structural deformations also obey the Euler-Bernoulli beam theory, along with the Kelvin-Voigt viscoelastic constitutive law. Meanwhile, the unsteady thin-airfoil and strip theories are the tools of producing the three-dimensional airloads. The origin of aerodynamic instability undergoes analysis in light of the oscillatory loads as well as the loads owing to arbitrary motions. After successful verification of the model, a systematic flutter survey was conducted on the theoretical effects of various control surface parameters. The results obtained demonstrate that the flapping modes and parameters of the control surface can significantly impact the flutter characteristics of the wings, which leads to a series of pertinent conclusions.
NASA Technical Reports Server (NTRS)
Gardner, Kevin D.; Liu, Jong-Shang; Murthy, Durbha V.; Kruse, Marlin J.; James, Darrell
1999-01-01
AlliedSignal Engines, in cooperation with NASA GRC (National Aeronautics and Space Administration Glenn Research Center), completed an evaluation of recently-developed aeroelastic computer codes using test cases from the AlliedSignal Engines fan blisk and turbine databases. Test data included strain gage, performance, and steady-state pressure information obtained for conditions where synchronous or flutter vibratory conditions were found to occur. Aeroelastic codes evaluated included quasi 3-D UNSFLO (MIT Developed/AE Modified, Quasi 3-D Aeroelastic Computer Code), 2-D FREPS (NASA-Developed Forced Response Prediction System Aeroelastic Computer Code), and 3-D TURBO-AE (NASA/Mississippi State University Developed 3-D Aeroelastic Computer Code). Unsteady pressure predictions for the turbine test case were used to evaluate the forced response prediction capabilities of each of the three aeroelastic codes. Additionally, one of the fan flutter cases was evaluated using TURBO-AE. The UNSFLO and FREPS evaluation predictions showed good agreement with the experimental test data trends, but quantitative improvements are needed. UNSFLO over-predicted turbine blade response reductions, while FREPS under-predicted them. The inviscid TURBO-AE turbine analysis predicted no discernible blade response reduction, indicating the necessity of including viscous effects for this test case. For the TURBO-AE fan blisk test case, significant effort was expended getting the viscous version of the code to give converged steady flow solutions for the transonic flow conditions. Once converged, the steady solutions provided an excellent match with test data and the calibrated DAWES (AlliedSignal 3-D Viscous Steady Flow CFD Solver). However, efforts expended establishing quality steady-state solutions prevented exercising the unsteady portion of the TURBO-AE code during the present program. AlliedSignal recommends that unsteady pressure measurement data be obtained for both test cases examined
Computational Aeroelastic Analysis of the Ares Launch Vehicle During Ascent
NASA Technical Reports Server (NTRS)
Bartels, Robert E.; Chwalowski, Pawel; Massey, Steven J.; Vatsa, Veer N.; Heeg, Jennifer; Wieseman, Carol D.; Mineck, Raymond E.
2010-01-01
This paper presents the static and dynamic computational aeroelastic (CAE) analyses of the Ares crew launch vehicle (CLV) during atmospheric ascent. The influence of launch vehicle flexibility on the static aerodynamic loading and integrated aerodynamic force and moment coefficients is discussed. The ultimate purpose of this analysis is to assess the aeroelastic stability of the launch vehicle along the ascent trajectory. A comparison of analysis results for several versions of the Ares CLV will be made. Flexible static and dynamic analyses based on rigid computational fluid dynamic (CFD) data are compared with a fully coupled aeroelastic time marching CFD analysis of the launch vehicle.
Aeroelastic Studies of a Rectangular Wing with a Hole: Correlation of Theory and Experiment
NASA Technical Reports Server (NTRS)
Conyers, Howard J.; Dowell, Earl H.; Hall, Kenneth C.
2010-01-01
Two rectangular wing models with a hole have been designed and tested in the Duke University wind tunnel to better understand the effects of damage. A rectangular hole is used to simulate damage. The wing with a hole is modeled structurally as a thin elastic plate using the finite element method. The unsteady aerodynamics of the plate-like wing with a hole is modeled using the doublet lattice method. The aeroelastic equations of motion are derived using Lagrange's equation. The flutter boundary is found using the V-g method. The hole's location effects the wing's mass, stiffness, aerodynamics and therefore the aeroelastic behavior. Linear theoretical models were shown to be capable of predicting the critical flutter velocity and frequency as verified by wind tunnel tests.
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.
X-HALE: The Development of a Research Platform for the Validation of Nonlinear Aeroelastic Codes
2011-03-01
flight structure that changes the aircraft’s modes in all three axes [38]. Patil, Hodges and Cesnik studied the aeroelastic dynamics of HALE...wing’s natural mode that may cause the wing’s flutter velocity to fall to aircraft’s cruise velocity. Continuing the research, Patil, Hodges , and...analysis and lifting line aerodynamics coupled with a one-lag term for unsteadiness corrections. Patil and Hodges also developed NATASHA [18]. This
Aeroelastic Analysis for Aeropropulsion Applications
NASA Technical Reports Server (NTRS)
Keith, Theo G., Jr.; Bakhle, Milind A.
2002-01-01
Aeroelastic codes with advanced capabilities for modeling flow require substantial computational time. On the other hand, fast-running linear aeroelastic codes lack the capability to model three-dimensional, transonic, vortical, and viscous flows. The goal of this work was to develop an aeroelastic code with accurate modeling capabilities and small computational requirements.
NASA Technical Reports Server (NTRS)
Kroo, I. M.
1981-01-01
One-fifth-scale models of three basic ultralight glider designs were constructed to simulate the elastic properties of full scale gliders and were tested at Reynolds numbers close to full scale values. Twenty-four minor modifications were made to the basic configurations in order to evaluate the effects of twist, reflex, dihedral, and various stability enhancement devices. Longitudinal and lateral data were obtained at several speeds through an angle of attack range of -30 deg to +45 deg with sideslip angles of up to 20 deg. The importance of vertical center of gravity displacement is discussed. Lateral data indicate that effective dihedral is lost at low angles of attack for nearly all of the configurations tested. Drag data suggest that lift-dependent viscous drag is a large part of the glider's total drag as is expected for thin, cambered sections at these relatively low Reynolds numbers.
Coupled-Mode Flutter of Bending-Bending Type in Highly-Flexible Uniform Airfoils
NASA Astrophysics Data System (ADS)
Pourazarm, Pariya; Modarres-Sadeghi, Yahya
2016-11-01
We study the behavior of a highly flexible uniform airfoil placed in wind both numerically and experimentally. It is shown that for a non-rotating highly-flexible cantilevered airfoil, placed at very small angles of attack (less than 1 degree), the airfoil loses its stability by buckling. For slightly higher angles of attack (more than 1 degree) a coupled-mode flutter in which the first and the second flapwise modes coalesce toward a flutter mode is observed, and thus the observed flutter has a bending-bending nature. The flutter onset and frequency found experimentally matched the numerical predictions. If the same airfoil is forced to rotate about its fixed end, the static deflection decreases and the observed couple-mode flutter becomes of flapwise-torsional type, same as what has already been observed for flutter of rotating wind turbine blades. The support provided by the National Science Foundation, CBET-1437988, is greatly acknowledged.
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.
NASA Technical Reports Server (NTRS)
Heeg, Jennifer
1991-01-01
The objective was to analytically and experimentally study the capabilities of adaptive material plate actuators for suppressing flutter. The validity of analytical modeling techniques for piezoelectric materials was also investigated. 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 airfoil and a flexible mount system which permitted a translational and a rotational degree of freedom. It was designed such that flutter was encounted within the testing envelope of the wind tunnel. Actuators, made of piezoelectric material were affixed to leaf springs of the mount system. Each degree of freedom was controlled by a separate leaf spring. Command signals, applied to the piezoelectric actuators, exerted control over the damping and stiffness properties. A mathematical aeroservoelastic model was constructed 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. It demonstrates that small, carefully placed actuating plates can be used effectively to control aeroelastic response.
FUN3D Analyses in Support of the Second Aeroelastic Prediction Workshop
NASA Technical Reports Server (NTRS)
Chwalowski, Pawel; Heeg, Jennifer
2016-01-01
This paper presents the computational aeroelastic results generated in support of the second Aeroelastic Prediction Workshop for the Benchmark Supercritical Wing (BSCW) configurations and compares them to the experimental data. The computational results are obtained using FUN3D, an unstructured grid Reynolds- Averaged Navier-Stokes solver developed at NASA Langley Research Center. The analysis results include aerodynamic coefficients and surface pressures obtained for steady-state, static aeroelastic equilibrium, and unsteady flow due to a pitching wing or flutter prediction. Frequency response functions of the pressure coefficients with respect to the angular displacement are computed and compared with the experimental data. The effects of spatial and temporal convergence on the computational results are examined.
NASA Astrophysics Data System (ADS)
Qin, Z.; Librescu, L.
2003-08-01
An encompassing aeroelastic model developed toward investigating the influence of directionality property of advanced composite materials and non-classical effects such as transverse shear and warping restraint on the aeroelastic instability of composite aircraft wings is presented. Within the model developed herein, both divergence and flutter instabilities are simultaneously addressed. The aircraft wing is modelled as an anisotropic composite thin-walled beam featuring circumferentially asymmetric stiffness lay-up that generates, for the problem at hand, elastic coupling among plunging, pitching and transverse shear motions. The unsteady incompressible aerodynamics used here is based on the concept of indicial functions. Issues related to aeroelastic instability are discussed, the influence of warping restraint and transverse shear on the critical speed are evaluated, and pertinent conclusions are outlined.
Plans and Example Results for the 2nd AIAA Aeroelastic Prediction Workshop
NASA Technical Reports Server (NTRS)
Heeg, Jennifer; Chwalowski, Pawel; Schuster, David M.; Raveh, Daniella; Jirasek, Adam; Dalenbring, Mats
2015-01-01
This paper summarizes the plans for the second AIAA Aeroelastic Prediction Workshop. The workshop is designed to assess the state-of-the-art of computational methods for predicting unsteady flow fields and aeroelastic response. The goals are to provide an impartial forum to evaluate the effectiveness of existing computer codes and modeling techniques, and to identify computational and experimental areas needing additional research and development. This paper provides guidelines and instructions for participants including the computational aerodynamic model, the structural dynamic properties, the experimental comparison data and the expected output data from simulations. The Benchmark Supercritical Wing (BSCW) has been chosen as the configuration for this workshop. The analyses to be performed will include aeroelastic flutter solutions of the wing mounted on a pitch-and-plunge apparatus.
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.
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
Simplified aeroelastic modeling of horizontal axis wind turbines
NASA Technical Reports Server (NTRS)
Wendell, J. H.
1982-01-01
Certain aspects of the aeroelastic modeling and behavior of the horizontal axis wind turbine (HAWT) are examined. Two simple three degree of freedom models are described in this report, and tools are developed which allow other simple models to be derived. The first simple model developed is an equivalent hinge model to study the flap-lag-torsion aeroelastic stability of an isolated rotor blade. The model includes nonlinear effects, preconing, and noncoincident elastic axis, center of gravity, and aerodynamic center. A stability study is presented which examines the influence of key parameters on aeroelastic stability. Next, two general tools are developed to study the aeroelastic stability and response of a teetering rotor coupled to a flexible tower. The first of these tools is an aeroelastic model of a two-bladed rotor on a general flexible support. The second general tool is a harmonic balance solution method for the resulting second order system with periodic coefficients. The second simple model developed is a rotor-tower model which serves to demonstrate the general tools. This model includes nacelle yawing, nacelle pitching, and rotor teetering. Transient response time histories are calculated and compared to a similar model in the literature. Agreement between the two is very good, especially considering how few harmonics are used. Finally, a stability study is presented which examines the effects of support stiffness and damping, inflow angle, and preconing.
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.
Advanced Aeroelastic Technologies for Turbomachinery Application
NASA Technical Reports Server (NTRS)
DeWitt, Kenneth; Srivastava, Rakesh; Reddy, T. S. R.
2004-01-01
A summary of the work performed under the grant NCC-1068 is presented. More details can be found in the cited references. The summary is presented in two parts to represent two areas of research. In the first part, methods to analyze a high temperature ceramic guide vane subjected to cooling jets are presented, and in the second part, the effect of unsteady aerodynamic forces on aeroelastic stability as implemented into the turbo-REDUCE code are presented
Dynamics and Aeroelasticity of Composite Structures.
1987-04-22
UNCLASSIFIED/UNLIMITEO SAME AS aPT Z OTIC USERS C3UNCLASSIFIED 22a. NAME OF RESPONSIBLE INDIVIDUAL 22b. TELEPHONE NUMBER 22c. OFFICE SYMBOL flncliads A’Wa...support related dynamic instability which could be eliminated by 3roper adjustment of the sutnport stiffness. Good agreement with linear thoery was found...Aeroelastic analysis 38 2.3 Wind Tunnel Support Stability Analysis 40 Chapter 3 Experiment 50 3.1 Wind Tunnel Model, Support System, and 50
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
Time efficient aeroelastic simulations based on radial basis functions
NASA Astrophysics Data System (ADS)
Liu, Wen; Huang, ChengDe; Yang, Guowei
2017-02-01
Aeroelasticity studies the interaction between aerodynamic forces and structural responses, and is one of the fundamental problems to be considered in the design of modern aircraft. The fluid-structure interpolation (FSI) and mesh deformation are two key issues in the CFD-CSD coupling approach (the partitioned approach), which is the mainstream numerical strategy in aeroelastic simulations. In this paper, a time efficient coupling scheme is developed based on the radial basis function interpolations. During the FSI process, the positive definite system of linear equations is constructed with the introduction of pseudo structural forces. The acting forces on the structural nodes can be calculated more efficiently via the solution of the linear system, avoiding the costly computations of the aerodynamic/structural coupling matrix. The multi-layer sequential mesh motion algorithm (MSM) is proposed to improve the efficiency of the volume mesh deformations, which is adequate for large-scale time dependent applications with frequent mesh updates. Two-dimensional mesh motion cases show that the MSM algorithm can reduce the computing cost significantly compared to the standard RBF-based method. The computations of the AGARD 445.6 wing flutter and the static deflections of the three-dimensional high-aspect-ratio aircraft demonstrate that the developed coupling scheme is applicable to both dynamic and static aeroelastic problems.
NASA Technical Reports Server (NTRS)
Freudinger, Lawrence C.
1989-01-01
An F-18 aircraft was modified with wingtip instrumentation pods for use in NASA's high-angle-of-attack research program. Ground vibration and flight flutter testing were performed to clear an acceptable flight envelope for the aircraft. Flight test utilized atmospheric turbulence for structural excitation; the aircraft displayed no adverse aeroelastic trends within the envelope tested. The data presented in this report include mode shapes from the ground vibration and estimates of frequency and damping as a function of Mach number.
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.
Multidisciplinary aeroelastic analysis of a generic hypersonic vehicle
NASA Technical Reports Server (NTRS)
Gupta, K. K.; Petersen, K. L.
1993-01-01
This paper presents details of a flutter and stability analysis of aerospace structures such as hypersonic vehicles. Both structural and aerodynamic domains are discretized by the common finite element technique. A vibration analysis is first performed by the STARS code employing a block Lanczos solution scheme. This is followed by the generation of a linear aerodynamic grid for subsequent linear flutter analysis within subsonic and supersonic regimes of the flight envelope; the doublet lattice and constant pressure techniques are employed to generate the unsteady aerodynamic forces. Flutter analysis is then performed for several representative flight points. The nonlinear flutter solution is effected by first implementing a CFD solution of the entire vehicle. Thus, a 3-D unstructured grid for the entire flow domain is generated by a moving front technique. A finite element Euler solution is then implemented employing a quasi-implicit as well as an explicit solution scheme. A novel multidisciplinary analysis is next effected that employs modal and aerodynamic data to yield aerodynamic damping characteristics. Such analyses are performed for a number of flight points to yield a large set of pertinent data that define flight flutter characteristics of the vehicle. This paper outlines the finite-element-based integrated analysis procedures in detail, which is followed by the results of numerical analyses of flight flutter simulation.
Multidisciplinary aeroelastic analysis of a generic hypersonic vehicle
NASA Astrophysics Data System (ADS)
Gupta, K. K.; Petersen, K. L.
1993-10-01
This paper presents details of a flutter and stability analysis of aerospace structures such as hypersonic vehicles. Both structural and aerodynamic domains are discretized by the common finite element technique. A vibration analysis is first performed by the STARS code employing a block Lanczos solution scheme. This is followed by the generation of a linear aerodynamic grid for subsequent linear flutter analysis within subsonic and supersonic regimes of the flight envelope; the doublet lattice and constant pressure techniques are employed to generate the unsteady aerodynamic forces. Flutter analysis is then performed for several representative flight points. The nonlinear flutter solution is effected by first implementing a CFD solution of the entire vehicle. Thus, a 3-D unstructured grid for the entire flow domain is generated by a moving front technique. A finite element Euler solution is then implemented employing a quasi-implicit as well as an explicit solution scheme. A novel multidisciplinary analysis is next effected that employs modal and aerodynamic data to yield aerodynamic damping characteristics. Such analyses are performed for a number of flight points to yield a large set of pertinent data that define flight flutter characteristics of the vehicle. This paper outlines the finite-element-based integrated analysis procedures in detail, which is followed by the results of numerical analyses of flight flutter simulation.
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.
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.
Aeroelastic analysis of wings using the Euler equations with a deforming mesh
NASA Technical Reports Server (NTRS)
Robinson, Brian A.; Batina, John T.; Yang, Henry T. Y.
1990-01-01
Modifications to the CFL3D three dimensional unsteady Euler/Navier-Stokes code for the aeroelastic analysis of wings are described. The modifications involve including a deforming mesh capability which can move the mesh to continuously conform to the instantaneous shape of the aeroelastically deforming wing, and including the structural equations of motion for their simultaneous time-integration with the governing flow equations. Calculations were performed using the Euler equations to verify the modifications to the code and as a first step toward aeroelastic analysis using the Navier-Stokes equations. Results are presented for the NACA 0012 airfoil and a 45 deg sweptback wing to demonstrate applications of CFL3D for generalized force computations and aeroelastic analysis. Comparisons are made with published Euler results for the NACA 0012 airfoil and with experimental flutter data for the 45 deg sweptback wing to assess the accuracy of the present capability. These comparisons show good agreement and, thus, the CFL3D code may be used with confidence for aeroelastic analysis of wings.
Aeroelastic analysis of wings using the Euler equations with a deforming mesh
NASA Technical Reports Server (NTRS)
Robinson, Brian A.; Batina, John T.; Yang, Henry T. Y.
1990-01-01
Modifications to the CFL3D three-dimensional unsteady Euler/Navier-Stokes code for the aeroelastic analysis of wings are described. The modifications involve including a deforming mesh capability which can move the mesh to continuously conform to the instantaneous shape of the aeroelastically deforming wing, and including the structural equations of motion for their simultaneous time-integration with the governing flow equations. Calculations were performed using the Euler equations to verify the modifications to the code and as a first-step toward aeroelastic analysis using the Navier-Stokes equations. Results are presented for the NACA 0012 airfoil and a 45 deg sweptback wing to demonstrate applications of CFL3D for generalized force computations and aeroelastic analysis. Comparisons are made with published Euler results for the NACA 0012 airfoil and with experimental flutter data for the 45 deg sweptback wing to assess the accuracy of the present capability. These comparisons show good agreement and, thus, the CFL3D code may be used with confidence for aeroelastic analysis of wings. The paper describes the modifications that were made to the code and presents results and comparisons which assess the capability.
NASA Technical Reports Server (NTRS)
Bradley, Marty K.; Allen, Timothy J.; Droney, Christopher
2014-01-01
This Test Report summarizes the Truss Braced Wing (TBW) Aeroelastic Test (Task 3.1) work accomplished by the Boeing Subsonic Ultra Green Aircraft Research (SUGAR) team, which includes the time period of February 2012 through June 2014. The team consisted of Boeing Research and Technology, Boeing Commercial Airplanes, Virginia Tech, and NextGen Aeronautics. The model was fabricated by NextGen Aeronautics and designed to meet dynamically scaled requirements from the sized full scale TBW FEM. The test of the dynamically scaled SUGAR TBW half model was broken up into open loop testing in December 2013 and closed loop testing from January 2014 to April 2014. Results showed the flutter mechanism to primarily be a coalescence of 2nd bending mode and 1st torsion mode around 10 Hz, as predicted by analysis. Results also showed significant change in flutter speed as angle of attack was varied. This nonlinear behavior can be explained by including preload and large displacement changes to the structural stiffness and mass matrices in the flutter analysis. Control laws derived from both test system ID and FEM19 state space models were successful in suppressing flutter. The control laws were robust and suppressed flutter for a variety of Mach, dynamic pressures, and angle of attacks investigated.
A thermal model for nonlinear panel flutter
NASA Astrophysics Data System (ADS)
Gee, David John
The subject of this research is the nonlinear panel flutter behavior for high Mach number, viscous compressible flow over one side of an isotropic, elastic panel. The fluid/structure interaction is treated as an aerothermoelastic problem in the sense that in addition to aeroelastic coupling, we consider thermal coupling between the fluid and structure. Specifically, at least two distinct heat transfer mechanisms for thermal interaction between the fluid and structure may be important in the panel flutter problem considered here. The primary contribution to thermal stress in the panel is aerodynamic heating. The temperature rise is obtained from a known solution of the compressible Navier-Stokes equations for flow over a flat plate. A secondary source of thermal heating may be due to the panel profile shape and/or any fluttering motion. An unsteady temperature component is obtained by assuming that the unsteady pressure and temperature above the panel are related through an isentropic flow relation. The equation of motion for the panel transverse deflection is based on von Karman large deflection plate theory. The unsteady pressure is computed using aerodynamic piston theory. The PDE is reduced to a system of nonlinear, coupled ordinary differential equations via Galerkin's method and is solved numerically using the 4th-order Runge-Kutta method. However, due to large variation in magnitude of the thermal compressive stress, the stepsize requirement for the numerical integrations is critical and, in fact, limits the usefulness of this procedure. Therefore, bifurcation diagrams are traced out using the pseudo-arclength continuation method (AUTO94). Representative results are given for several combinations of cruise Mach number and altitude. Several stable attractors are found as functions of the in-plane load and dynamic pressure parameters. These two parameters are coupled to the flight conditions and are, therefore, not wholly independent. A direct consequence of
Perturbation and harmonic balance methods for nonlinear panel flutter.
NASA Technical Reports Server (NTRS)
Kuo, C.-C.; Morino, L.; Dugundji, J.
1972-01-01
A systematic way of applying both perturbation methods and harmonic balance methods to nonlinear panel flutter problems is developed here. Results obtained by both these methods for two-dimensional simply supported and three-dimensional clamped-clamped plates with six modes agree well with those obtained by the straightforward direct integration method, yet require less computer time and provide better insight into the solutions. Effects of viscoelastic structural damping on the flutter stability boundary are generally found to be destabilizing and the postflutter behavior becomes more explosive. The methods developed here may be of interest in related vibration problems.
Wavelet Applications for Flight Flutter Testing
NASA Technical Reports Server (NTRS)
Lind, Rick; Brenner, Marty; Freudinger, Lawrence C.
1999-01-01
Wavelets present a method for signal processing that may be useful for analyzing responses of dynamical systems. This paper describes several wavelet-based tools that have been developed to improve the efficiency of flight flutter testing. One of the tools uses correlation filtering to identify properties of several modes throughout a flight test for envelope expansion. Another tool uses features in time-frequency representations of responses to characterize nonlinearities in the system dynamics. A third tool uses modulus and phase information from a wavelet transform to estimate modal parameters that can be used to update a linear model and reduce conservatism in robust stability margins.
A methodology for aeroelastic constraint analysis in a conceptual design environment
NASA Astrophysics Data System (ADS)
de Baets, Peter Wilfried Gaston
The objective of this study is the infusion of aeroelastic constraint knowledge into the design space. The mapping of such aeroelastic information in the conceptual design space has long been a desire of the design community. The conceptual design phase of an aircraft is a multidisciplinary environment and has the most influence on the future design of the vehicle. However, sufficient results cannot he obtained in a timely enough manner to materially contribute to early design decisions. Furthermore, the natural division of the engineering team into specialty groups is not well supported by the monolithic aerodynamic-structures codes typically used in modern aeroelastic analysis. The research examines how the Bi-Level Integrated System Synthesis decomposition technique can be adapted to perform as the conceptual aeroelastic design tool. The study describes a comprehensive solution of the aeroelastic coupled problem cast in this decomposition format and implemented in an integrated framework. The method is supported by application details of a proof of concept high speed vehicle. Physics-based codes such as finite element and an aerodynamic panel method are used to model the high-definition geometric characteristics of the vehicle. A synthesis and sizing code was added to referee the conflicts that arise between the two disciplines. This research's novelty lies in four points. First is the use of physics-based tools at the conceptual design phase to calculate the aeroelastic properties. Second is the projection of flutter and divergence velocity constraint lines in a power loading versus wing loading graph. Third is the aeroelastic assessment time reduction, which has moved from a matter of years to months. Lastly, this assessment allowed verification of the impact of changing velocity, altitude, and angle of attack on the aeroelastic properties. This then allowed identification of robust design space with respect to these three mission properties. The method
Application of the Finite Element Method to Rotary Wing Aeroelasticity
NASA Technical Reports Server (NTRS)
Straub, F. K.; Friedmann, P. P.
1982-01-01
A finite element method for the spatial discretization of the dynamic equations of equilibrium governing rotary-wing aeroelastic problems is presented. Formulation of the finite element equations is based on weighted Galerkin residuals. This Galerkin finite element method reduces algebraic manipulative labor significantly, when compared to the application of the global Galerkin method in similar problems. The coupled flap-lag aeroelastic stability boundaries of hingeless helicopter rotor blades in hover are calculated. The linearized dynamic equations are reduced to the standard eigenvalue problem from which the aeroelastic stability boundaries are obtained. The convergence properties of the Galerkin finite element method are studied numerically by refining the discretization process. Results indicate that four or five elements suffice to capture the dynamics of the blade with the same accuracy as the global Galerkin method.
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.
NASA Astrophysics Data System (ADS)
Navalkar, S. T.; Bernhammer, L. O.; Sodja, J.; Slinkman, C. J.; van Wingerden, J. W.; van Kuik, G. A. M.
2016-09-01
Trailing edge flaps located outboard on wind turbine blades have recently shown considerable potential in the alleviation of turbine lifetime dynamic loads. The concept of the free-floating flap is specifically interesting for wind turbines, on account of its modularity and enhanced control authority. Such a flap is free to rotate about its axis; camberline control of the free-floating flap allows for aeroelastic control of blade loads. This paper describes the design of a scaled wind turbine blade instrumented with free-floating flaps, intended for use in wind tunnel experiments. The nature of the flap introduces a coupled form of flutter due to the aeroelastic coupling of flap rigid-body and blade out-of-plane modes; for maximal control authority it is desired to operate close to the flutter limit. Analytical and numerical methods are used to perform a flutter analysis of the turbine blade. It is shown that the potential flow aeroelastic model can be recast as a continuous-time Linear-Parameter-Varying (LPV) state space model of a low order, for which formal controller design methodologies are readily available.
LINFLUX-AE: A Turbomachinery Aeroelastic Code Based on a 3-D Linearized Euler Solver
NASA Technical Reports Server (NTRS)
Reddy, T. S. R.; Bakhle, M. A.; Trudell, J. J.; Mehmed, O.; Stefko, G. L.
2004-01-01
This report describes the development and validation of LINFLUX-AE, a turbomachinery aeroelastic code based on the linearized unsteady 3-D Euler solver, LINFLUX. A helical fan with flat plate geometry is selected as the test case for numerical validation. The steady solution required by LINFLUX is obtained from the nonlinear Euler/Navier Stokes solver TURBO-AE. The report briefly describes the salient features of LINFLUX and the details of the aeroelastic extension. The aeroelastic formulation is based on a modal approach. An eigenvalue formulation is used for flutter analysis. The unsteady aerodynamic forces required for flutter are obtained by running LINFLUX for each mode, interblade phase angle and frequency of interest. The unsteady aerodynamic forces for forced response analysis are obtained from LINFLUX for the prescribed excitation, interblade phase angle, and frequency. The forced response amplitude is calculated from the modal summation of the generalized displacements. The unsteady pressures, work done per cycle, eigenvalues and forced response amplitudes obtained from LINFLUX are compared with those obtained from LINSUB, TURBO-AE, ASTROP2, and ANSYS.
Chirality-dependent flutter of Typha blades in wind
Zhao, Zi-Long; Liu, Zong-Yuan; Feng, Xi-Qiao
2016-01-01
Cattail or Typha, an emergent aquatic macrophyte widely distributed in lakes and other shallow water areas, has slender blades with a chiral morphology. The wind-resilient Typha blades can produce distinct hydraulic resistance for ecosystem functions. However, their stem may rupture and dislodge in excessive wind drag. In this paper, we combine fluid dynamics simulations and experimental measurements to investigate the aeroelastic behavior of Typha blades in wind. It is found that the chirality-dependent flutter, including wind-induced rotation and torsion, is a crucial strategy for Typha blades to accommodate wind forces. Flow visualization demonstrates that the twisting morphology of blades provides advantages over the flat one in the context of two integrated functions: improving wind resistance and mitigating vortex-induced vibration. The unusual dynamic responses and superior mechanical properties of Typha blades are closely related to their biological/ecosystem functions and macro/micro structures. This work decodes the physical mechanisms of chirality-dependent flutter in Typha blades and holds potential applications in vortex-induced vibration suppression and the design of, e.g., bioinspired flight vehicles. PMID:27432079
NACA 0012 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 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. 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.
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.
Application of two design methods for active flutter suppression and wind-tunnel test results
NASA Technical Reports Server (NTRS)
Newsom, J. R.; Abel, I.; Dunn, H. J.
1980-01-01
The synthesis, implementation, and wind tunnel test of two flutter suppression control laws for an aeroelastic model equipped with a trailing edge control surface are presented. One control law is based on the aerodynamic energy method, and the other is based on results of optimal control theory. Analytical methods used to design the control laws and evaluate their performance are described. At Mach 0.6, 0.8, and 0.9, increases in flutter dynamic pressure were obtained but the full 44 percent increase was not achieved. However at Mach 0.95, the 44 percent increase was achieved with both control laws. Experimental results indicate that the performance of the systems is not so effective as that predicted by analysis, and that wind tunnel turbulence plays an important role in both control law synthesis and demonstration of system performance.
NASA Technical Reports Server (NTRS)
Goldman, Benjamin D.; Scott, Robert C,; Dowell, Earl H.
2014-01-01
The purpose of this work is to develop a set of theoretical and experimental techniques to characterize the aeroelasticity of the thermal protection system (TPS) on the NASA Hypersonic Inflatable Aerodynamic Decelerator (HIAD). A square TPS coupon experiences trailing edge oscillatory behavior during experimental testing in the 8' High Temperature Tunnel (HTT), which may indicate the presence of aeroelastic flutter. Several theoretical aeroelastic models have been developed, each corresponding to a different experimental test configuration. Von Karman large deflection theory is used for the plate-like components of the TPS, along with piston theory for the aerodynamics. The constraints between the individual TPS layers and the presence of a unidirectional foundation at the back of the coupon are included by developing the necessary energy expressions and using the Rayleigh Ritz method to derive the nonlinear equations of motion. Free vibrations and limit cycle oscillations are computed and the frequencies and amplitudes are compared with accelerometer and photogrammetry data from the experiments.
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.
Experimental analysis of energy harvesting from self-induced flutter of a composite beam
Zakaria, Mohamed Y. Al-Haik, Mohammad Y.; Hajj, Muhammad R.
2015-07-13
Previous attempts to harvest energy from aeroelastic vibrations have been based on attaching a beam to a moving wing or structure. Here, we exploit self-excited oscillations of a fluttering composite beam to harvest energy using piezoelectric transduction. Details of the beam properties and experimental setup are presented. The effects of preset angle of attack, wind speed, and load resistance on the levels of harvested power are determined. The results point to a complex relation between the aerodynamic loading and its impact on the static deflection and amplitudes of the limit cycle oscillations on one hand and the load resistance and level of power harvested on the other hand.
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.
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.
Investigation on transonic flutter active auppression with CFD-Based ROMs
NASA Astrophysics Data System (ADS)
Nie, XueYuan; Yang, GuoWei; Zhang, MingFeng
2015-01-01
The calculation of accurate unsteady aerodynamic forces is critical in the analysis of aeroelastic problems, however the efficiency is low because of high computational costs of the computational fluid dynamics (CFD) portion. Additionally, direct integrated CFD and computational structural dynamics (CSD) technique is unsuitable for the analysis of ASE and the flutter active suppression in state-space form. A reduced-order model (ROM) based on Volterra series was developed using CFD calculation and used to predict the flutter coupled with the structure. The closed-loop control systems designed by the sliding mode control (SMC) and linear quadratic Gaussian (LQG) control were constructed with ROM/CSD to suppress the AGARD 445.6 wing flutter. The detailed implementation of the two control approaches is presented, and the flutter suppression effectiveness is discussed and compared. The results indicate that SMC method can make the controlled object response decay to the stable equilibrium more rapidly and has better control effects than the LQG control.
Code of Federal Regulations, 2010 CFR
2010-01-01
... STANDARDS: NORMAL CATEGORY ROTORCRAFT Design and Construction General § 27.629 Flutter. Each aerodynamic surface of the rotorcraft must be free from flutter under each appropriate speed and power...
State and development of flutter calculation
NASA Technical Reports Server (NTRS)
Teichmann, Alfred
1951-01-01
This report discusses the need for considering a wide variation in certain of the basic flutter parameters in conducting a flutter analysis. Conclusions are drawn stating that design charts or simple rules may be misleading. Due to inherent difficulties, dynamic model testing may also yield misleading results. The general flutter equations and various methods of solution are discussed. Of particular interest, curves are presented showing computational effort plotted against a number of degrees of freedom used in a flutter analysis.
Nonlinear Aeroelastic Analysis of UAVs: Deterministic and Stochastic Approaches
NASA Astrophysics Data System (ADS)
Sukut, Thomas Woodrow
Aeroelastic aspects of unmanned aerial vehicles (UAVs) is analyzed by treatment of a typical section containing geometrical nonlinearities. Equations of motion are derived and numerical integration of these equations subject to quasi-steady aerodynamic forcing is performed. Model properties are tailored to a high-altitude long-endurance unmanned aircraft. Harmonic balance approximation is employed based on the steady-state oscillatory response of the aerodynamic forcing. Comparisons are made between time integration results and harmonic balance approximation. Close agreement between forcing and displacement oscillatory frequencies is found. Amplitude agreement is off by a considerable margin. Additionally, stochastic forcing effects are examined. Turbulent flow velocities generated from the von Karman spectrum are applied to the same nonlinear structural model. Similar qualitative behavior is found between quasi-steady and stochastic forcing models illustrating the importance of considering the non-steady nature of atmospheric turbulence when operating near critical flutter velocity.
Fluttering in Stratified Flows
NASA Astrophysics Data System (ADS)
Lam, Try; Vincent, Lionel; Kanso, Eva
2016-11-01
The descent motion of heavy objects under the influence of gravitational and aerodynamic forces is relevant to many branches of engineering and science. Examples range from estimating the behavior of re-entry space vehicles to studying the settlement of marine larvae and its influence on underwater ecology. The behavior of regularly shaped objects freely falling in homogeneous fluids is relatively well understood. For example, the complex interaction of a rigid coin with the surrounding fluid will cause it to either fall steadily, flutter, tumble, or be chaotic. Less is known about the effect of density stratification on the descent behavior. Here, we experimentally investigate the descent of discs in both pure water and in a linearly salt-stratified fluids where the density is varied from 1.0 to 1.14 of that of water where the Brunt-Vaisala frequency is 1.7 rad/sec and the Froude number Fr < 1. We found that stratification enhances the radial dispersion of the disc at landing, and simultaneously, decrease the descent speed and the inclination (or nutation) angle while falling. We conclude by commenting on the relevance of these results to the use of unpowered vehicles and robots for space exploration and underwater missions.
Flutter analysis of a sounding rocket fin
NASA Astrophysics Data System (ADS)
Natori, M.; Onoda, J.; Kitamura, T.
The procedures used to characterize the flutter behavior of the fin of the ISAS M-3S II launch vehicle (capable of launching 750 kg to LEO) are described. Consideration is given to supersonic flutter computations, single-point-excitation and vibration testing, construction of flutter models, and transonic wind-tunnel tests. Tables, graphs, diagrams, and photographs are provided.
NASA Technical Reports Server (NTRS)
Smith, Todd E.
1991-01-01
An aeroelastic analysis is developed which has general application to all types of axial-flow turbomachinery blades. The approach is based on linear modal analysis, where the blade's dynamic response is represented as a linear combination of contributions from each of its in-vacuum free vibrational modes. A compressible linearized unsteady potential theory is used to model the flow over the oscillating blades. The two-dimensional unsteady flow is evaluated along several stacked axisymmetric strips along the span of the airfoil. The unsteady pressures at the blade surface are integrated to result in the generalized force acting on the blade due to simple harmonic motions. The unsteady aerodynamic forces are coupled to the blade normal modes in the frequency domain using modal analysis. An iterative eigenvalue problem is solved to determine the stability of the blade when the unsteady aerodynamic forces are included in the analysis. The approach is demonstrated by applying it to a high-energy subsonic turbine blade from a rocket engine turbopump power turbine. The results indicate that this turbine could undergo flutter in an edgewise mode of vibration.
Aeroelastic airfoil smart spar
NASA Technical Reports Server (NTRS)
Greenhalgh, Skott; Pastore, Christopher M.; Garfinkle, Moishe
1993-01-01
Aircraft wings and rotor-blades are subject to undesirable bending and twisting excursions that arise from unsteady aerodynamic forces during high speed flight, abrupt maneuvers, or hard landings. These bending excursions can range in amplitude from wing-tip flutter to failure. A continuous-filament construction 'smart' laminated composite box-beam spar is described which corrects itself when subject to undesirable bending excursions or flutter. The load-bearing spar is constructed so that any tendency for the wing or rotor-blade to bend from its normal position is met by opposite twisting of the spar to restore the wing to its normal position. Experimental and theoretical characterization of these spars was made to evaluate the torsion-flexure coupling associated with symmetric lay-ups. The materials used were uniweave AS-4 graphite and a matrix comprised of Shell 8132 resin and U-40 hardener. Experimental tests were conducted on five spars to determine spar twist and bend as a function of load for 0, 17, 30, 45 and 60 deg fiber angle lay-ups. Symmetric fiber lay-ups do exhibit torsion-flexure couplings. Predictions of the twist and bend versus load were made for different fiber orientations in laminated spars using a spline function structural analysis. The analytical results were compared with experimental results for validation. Excellent correlation between experimental and analytical values was found.
The aeromechanical stability of soft-inplane tiltrotors
NASA Astrophysics Data System (ADS)
Howard, Anna Kathleen Tyler
A soft-inplane tiltrotor can be subject to the aeromechanical instabilities of ground and air resonance in addition to whirl flutter. While whirl flutter has been studied by many researchers, air and ground resonance in a tiltrotor has not been explored thoroughly. This dissertation investigates the aeromechanical instabilities of air and ground resonance in a tiltrotor and to assess some passive methods for improving the stability of a soft-inplane tiltrotor. This research is broken into two parts, addressing the semi-span model and the full-span model. In the first chapters, a semi-span analytical model consisting of a rigid blade rotor coupled to a rigid pylon and an elastic wing is developed and validated. The Boeing Model 222 which was wind-tunnel tested in 1972 is used to validate the current analytical model. This aircraft is used as the baseline for the analytical investigation of ground resonance and air resonance in hover, air resonance in transition, and air resonance and whirl flutter in airplane configuration. Neither ground resonance nor air resonance in hover are predicted for the Model 222 though air resonance in transition and cruise as well as whirl flutter at high speeds are predicted. Aeroelastic couplings in the rotor blades and wing are shown to affect the air resonance stability in cruise and transition and to be useful in augmenting the inherent stability of the aircraft. Wing vertical bending coupled to wing torsion and rotor low frequency lag coupled to blade torsion were helpful for air resonance. However, the rotor coupling was quite detrimental to whirl flutter stability and the wing coupling was unable to stabilize all the air resonance regions completely. The second part of the thesis expands the semi-span model to full-span: fuselage pitch and roll motion, anti-symmetric wing motion, and the second rotor motion are added. The full-span model determined that fuselage motion can couple with rotor low frequency lag to exhibit air
Transonic single-mode flutter and buffet of a low aspect ratio wing having a subsonic airfoil shape
NASA Technical Reports Server (NTRS)
Erickson, L. L.
1974-01-01
Transonic flutter and buffet results obtained from wind-tunnel tests of a low aspect ratio semispan wing model are presented. The tests were conducted to investigate potential transonic aeroelastic problems of vehicles having subsonic airfoil sections. The model employed NACA 00XX-64 airfoil sections in the streamwise direction and had a 14 deg leading edge sweep angle. Aspect ratio, and average thickness were 4.0, 0.35, and 8 percent, respectively. The model was tested at Mach numbers from 0.6 to 0.95 at angles of attack from 0 deg to 15 deg. Two zero lift flutter conditions were found that involved essentially single normal mode vibrations. With boundary layer trips on the model, flutter occurred in a narrow Mach number range centered at about Mach 0.90. The frequency and motion of this flutter were like that of the first normal mode vibration. With the trips removed flutter occurred at a slightly high Mach number but in a mode strongly resembling that of the second normal mode.
Nonlinear aeroelastic analysis, flight dynamics, and control of a complete aircraft
NASA Astrophysics Data System (ADS)
Patil, Mayuresh Jayawant
The focus of this research was to analyze a high-aspect-ratio wing aircraft flying at low subsonic speeds. Such aircraft are designed for high-altitude, long-endurance missions. Due to the high flexibility and associated wing deformation, accurate prediction of aircraft response requires use of nonlinear theories. Also strong interactions between flight dynamics and aeroelasticity are expected. To analyze such aircraft one needs to have an analysis tool which includes the various couplings and interactions. A theoretical basis has been established for a consistent analysis which takes into account, (i) material anisotropy, (ii) geometrical nonlinearities of the structure, (iii) rigid-body motions, (iv) unsteady flow behavior, and (v) dynamic stall. The airplane structure is modeled as a set of rigidly attached beams. Each of the beams is modeled using the geometrically exact mixed variational formulation, thus taking into account geometrical nonlinearities arising due to large displacements and rotations. The cross-sectional stiffnesses are obtained using an asymptotically exact analysis, which can model arbitrary cross sections and material properties. An aerodynamic model, consisting of a unified lift model, a consistent combination of finite-state inflow model and a modified ONERA dynamic stall model, is coupled to the structural system to determine the equations of motion. The results obtained indicate the necessity of including nonlinear effects in aeroelastic analysis. Structural geometric nonlinearities result in drastic changes in aeroelastic characteristics, especially in case of high-aspect-ratio wings. The nonlinear stall effect is the dominant factor in limiting the amplitude of oscillation for most wings. The limit cycle oscillation (LCO) phenomenon is also investigated. Post-flutter and pre-flutter LCOs are possible depending on the disturbance mode and amplitude. Finally, static output feedback (SOF) controllers are designed for flutter suppression
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.
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.
Material and Thickness Grading for Aeroelastic Tailoring of the Common Research Model Wing Box
NASA Technical Reports Server (NTRS)
Stanford, Bret K.; Jutte, Christine V.
2014-01-01
This work quantifies the potential aeroelastic benefits of tailoring a full-scale wing box structure using tailored thickness distributions, material distributions, or both simultaneously. These tailoring schemes are considered for the wing skins, the spars, and the ribs. Material grading utilizes a spatially-continuous blend of two metals: Al and Al+SiC. Thicknesses and material fraction variables are specified at the 4 corners of the wing box, and a bilinear interpolation is used to compute these parameters for the interior of the planform. Pareto fronts detailing the conflict between static aeroelastic stresses and dynamic flutter boundaries are computed with a genetic algorithm. In some cases, a true material grading is found to be superior to a single-material structure.
NASA Astrophysics Data System (ADS)
O'Donnell, K.; Schober, S.; Stolk, M.; Marzocca, P.; De Breuker, R.; Abdalla, M.; Nicolini, E.; Gürdal, Z.
2007-04-01
This paper discusses modeling, simulations and experimental aspects of active aeroelastic control on aircraft wings by using Synthetic Jet Actuators (SJAs). SJAs, a particular class of zero-net mass-flux actuators, have shown very promising results in numerous aeronautical applications, such as boundary layer control and delay of flow separation. A less recognized effect resulting from the SJAs is a momentum exchange that occurs with the flow, leading to a rearrangement of the streamlines around the airfoil modifying the aerodynamic loads. Discussions pertinent to the use of SJAs for flow and aeroelastic control and how these devices can be exploited for flutter suppression and for aerodynamic performances improvement are presented and conclusions are outlined.
Evaluation of an aeroelastic model technique for predicting airplane buffet loads
NASA Technical Reports Server (NTRS)
Hanson, P. W.
1973-01-01
A wind-tunnel technique which makes use of a dynamically scaled aeroelastic model to predict full-scale airplane buffet loads during buffet boundary penetration is evaluated. A 1/8-scale flutter model of a fighter airplane with remotely controllable variable-sweep wings and trimming surfaces was used for the evaluation. The model was flown on a cable-mount system which permitted high lift forces comparable to those in maneuvering flight. Bending moments and accelerations due to buffet were measured on the flutter model and compared with those measured on the full-scale airplane in an independent flight buffet research study. It is concluded that the technique can provide valuable information on airplane buffet load characteristics not available from any other source except flight test.
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.
Aerodynamic Indicial Functions and Their Use in Aeroelastic Formulation of Lifting Surfaces
NASA Technical Reports Server (NTRS)
Marzocca, Piergiovanni; Librescu, Liviu; Silva, Walter A.
2000-01-01
An investigation related to the use of linear indicial functions in the time and frequency domains, enabling one to derive the proper aerodynamic loads as to study the subcritical response and flutter of swept lifting surfaces, respectively, of the open/closed loop aeroelastic system is presented. The expressions of the lift and aerodynamic moment in the frequency domain are given in terms of the Theodorsen's function, while, in the time domain, these are obtained directly with the help of the Wagner's function. Closed form solutions of aerodynamic derivatives are obtained, graphical representations are supplied and conclusions and prospects for further developments are outlined.
Reduced order modeling of aeroelasticity analysis for a wing under static deformation effect
NASA Astrophysics Data System (ADS)
Tamayama, Masato
2017-01-01
The full order analysis of aeroelasticity system, which solves the Euler or Navier Stokes equations in a time domain, is usually expensive in a sense of time consumed. To improve this situation, the Reduced Order Modeling (ROM) method has been developed. If there is a pressure difference between upper and lower surfaces of a wing, the aerodynamic forces loaded on the wing cause static deformations. The ROM, therefore, should have a capability to simulate wing vibrations under the static deformation effect. To include this effect, sequential processing of ROMs for two times is proposed in this study. The 1st step ROM predicts the flutter condition for the rigid wing. The 2nd step ROM predicts the flutter condition for the statically deformed wing under the aerodynamic load caused by the 1st step ROM flutter dynamic pressure. The accuracy of this method is verified by comparing the results with those predicted only by the full order analysis. In this study, the identification of aerodynamic forces is conducted by the Eigensystem Realization Algorithm (ERA). In the ERA, reduction of singular value matrix influences the accuracy of identification. Two methods are introduced to reduce the singular value matrix, and the flutter conditions acquired by these two methods are compared each other.
Effects of Inlet Distortion on Aeromechanical Stability of a Forward-Swept High-Speed Fan
NASA Technical Reports Server (NTRS)
Herrick, Gregory P.
2011-01-01
Concerns regarding noise, propulsive efficiency, and fuel burn are inspiring aircraft designs wherein the propulsive turbomachines are partially (or fully) embedded within the airframe; such designs present serious concerns with regard to aerodynamic and aeromechanic performance of the compression system in response to inlet distortion. Separately, a forward-swept high-speed fan was developed to address noise concerns of modern podded turbofans; however this fan encounters aeroelastic instability (flutter) as it approaches stall. A three-dimensional, unsteady, Navier-Stokes computational fluid dynamics code is applied to analyze and corroborate fan performance with clean inlet flow. This code, already validated in its application to assess aerodynamic damping of vibrating blades at various flow conditions, is modified and then applied in a computational study to preliminarily assess the effects of inlet distortion on aeroelastic stability of the fan. Computational engineering application and implementation issues are discussed, followed by an investigation into the aeroelastic behavior of the fan with clean and distorted inlets.
Douglas Experience in Flight Flutter Testing
NASA Technical Reports Server (NTRS)
Philbrick, J.
1975-01-01
Douglas Aircraft Company experience in flight flutter testing is reviewed briefly, with comments on state-of-the-art excitation and instrumentation techniques used up to the present time. The limitations of previous techniques are discussed with emphasis on the problem of: (1) establishing a flutter margin of safety for predicted marginal flutter modes; (2) resolving instances of flutter not predicted by theoretical calculations in advance; and (3) delaying the airplane demonstration by time consumed in acquisition and reduction of flutter data. Current Douglas philosophy in flight flutter testing is presented and a description given of steady-state vane excitation system development, automatic data handling system, and the potential application of automatic computing methods for increasing flutter data yield.
Computational Aeroelasticity: Success, Progress, Challenge
NASA Technical Reports Server (NTRS)
Schuster, David M.; Liu, Danny D.; Huttsell, Lawrence J.
2003-01-01
The formal term Computational Aeroelasticity (CAE) has only been recently adopted to describe aeroelastic analysis methods coupling high-level computational fluid dynamics codes with structural dynamics techniques. However, the general field of aeroelastic computations has enjoyed a rich history of development and application since the first hand-calculations performed in the mid 1930 s. This paper portrays a much broader definition of Computational Aeroelasticity; one that encompasses all levels of aeroelastic computation from the simplest linear aerodynamic modeling to the highest levels of viscous unsteady aerodynamics, from the most basic linear beam structural models to state-of-the-art Finite Element Model (FEM) structural analysis. This paper is not written as a comprehensive history of CAE, but rather serves to review the development and application of aeroelastic analysis methods. It describes techniques and example applications that are viewed as relatively mature and accepted, the "successes" of CAE. Cases where CAE has been successfully applied to unique or emerging problems, but the resulting techniques have proven to be one-of-a-kind analyses or areas where the techniques have yet to evolve into a routinely applied methodology are covered as "progress" in CAE. Finally the true value of this paper is rooted in the description of problems where CAE falls short in its ability to provide relevant tools for industry, the so-called "challenges" to CAE.
NASA Astrophysics Data System (ADS)
Tan, B. H.; Lucey, A. D.; Howell, R. M.
2013-12-01
We study the effect of adding localised stiffness, via a spring support, on the stability of flexible panels subjected to axial uniform incompressible flow. Applications are considered that range from the hydro-elasticity of hull panels of high-speed ships to the aero-elasticity of glass panels in the curtain walls of high-rise buildings in very strong winds. A two-dimensional linear analysis is conducted using a hybrid of theoretical and computational methods that calculates the system eigen-states but can also be used to capture the transient behaviour that precedes these. We show that localised stiffening is a very effective means to increase the divergence-onset flow speed in both hydro- and aero-elastic applications. It is most effective when located at the mid-chord of the panel and there exists an optimum value of added stiffness beyond which further increases to the divergence-onset flow speed do not occur. For aero-elastic applications, localised stiffening can be used to replace the more destructive flutter instability that follows divergence at higher flow speeds by an extended range of divergence. The difference in eigen-solution morphology between aero- and hydro-elastic applications is highlighted, showing that for the former coalescence of two non-oscillatory divergence modes is the mechanism for flutter onset. This variation in solution morphology is mapped out in terms of a non-dimensional mass ratio. Finally, we present a short discussion of the applicability of the stabilisation strategy in a full three-dimensional system.
Aeroelastic structural acoustic control.
Clark, R L; Frampton, K D
1999-02-01
Static, constant-gain, output-feedback control compensators were designed to increase the transmission loss across a panel subjected to mean flow on one surface and a stationary, acoustic half-space on the opposite surface. The multi-input, multi-output control system was based upon the use of an array of colocated transducer pairs. The performance of the static-gain, output-feedback controller was compared to that of the full state-feedback controller using the same control actuator arrays, and was found to yield comparable levels of performance for practical limitations on control effort. Additionally, the resulting static compensators proved to be dissipative in nature, and thus the design varied little as a function of the aeroelastic coupling induced by the fluid-structure interaction under subsonic flow conditions. Several parametric studies were performed, comparing the effects of control-effort penalty as well as the number of transducer pairs used in the control system.
Flutter of buckled shape memory alloy reinforced laminates
NASA Astrophysics Data System (ADS)
Kuo, Shih-Yao; Shiau, Le-Chung; Lai, Chin-Hsin
2012-03-01
The effect of shape memory alloys (SMA) on the linear and nonlinear flutter behaviors of buckled cross-ply and angle-ply laminates was investigated in the frequency and time domains using the finite element method. In particular, this study takes the first move toward examining the effect of varying the SMA fiber spacing. Von Karman large deformation assumptions and quasi-steady aerodynamic theory were employed. The flutter boundary, stability boundary, time history response, and phase plane plots of SMA reinforced cross-ply and angle-ply laminates are presented. The numerical results show that increase in the SMA fiber volume fraction and prestrain may generate more recovery stress, and increase the stiffness of the SMA reinforced laminates. Therefore, the flutter boundary and critical load of the plate may be increased significantly. All five types of panel behavior, namely flat, buckled, limit-cycle, periodic, and chaotic motion, are clearly displayed and successively identified. This study sheds light on improving the flutter boundary efficiently by increasing the SMA fiber volume fraction to reinforce the center of the plate.
Nonlinear Time Delayed Feedback Control of Aeroelastic Systems: A Functional Approach
NASA Technical Reports Server (NTRS)
Marzocca, Piergiovanni; Librescu, Liviu; Silva, Walter A.
2003-01-01
In addition to its intrinsic practical importance, nonlinear time delayed feedback control applied to lifting surfaces can result in interesting aeroelastic behaviors. In this paper, nonlinear aeroelastic response to external time-dependent loads and stability boundary for actively controlled lifting surfaces, in an incompressible flow field, are considered. The structural model and the unsteady aerodynamics are considered linear. The implications of the presence of time delays in the linear/nonlinear feedback control and of geometrical parameters on the aeroelasticity of lifting surfaces are analyzed and conclusions on their implications are highlighted.
Improved Aerodynamic Influence Coefficients for Dynamic Aeroelastic Analyses
NASA Astrophysics Data System (ADS)
Gratton, Patrice
2011-12-01
Currently at Bombardier Aerospace, aeroelastic analyses are performed using the Doublet Lattice Method (DLM) incorporated in the NASTRAN solver. This method proves to be very reliable and fast in preliminary design stages where wind tunnel experimental results are often not available. Unfortunately, the geometric simplifications and limitations of the DLM, based on the lifting surfaces theory, reduce the ability of this method to give reliable results for all flow conditions, particularly in transonic flow. Therefore, a new method has been developed involving aerodynamic data from high-fidelity CFD codes which solve the Euler or Navier-Stokes equations. These new aerodynamic loads are transmitted to the NASTRAN aeroelastic module through improved aerodynamic influence coefficients (AIC). A cantilevered wing model is created from the Global Express structural model and a set of natural modes is calculated for a baseline configuration of the structure. The baseline mode shapes are then combined with an interpolation scheme to deform the 3-D CFD mesh necessary for Euler and Navier-Stokes analyses. An uncoupled approach is preferred to allow aerodynamic information from different CFD codes. Following the steady state CFD analyses, pressure differences ( DeltaCp), calculated between the deformed models and the original geometry, lead to aerodynamic loads which are transferred to the DLM model. A modal-based AIC method is applied to the aerodynamic matrices of NASTRAN based on a least-square approximation to evaluate aerodynamic loads of a different wing configuration which displays similar types of mode shapes. The methodology developed in this research creates weighting factors based on steady CFD analyses which have an equivalent reduced frequency of zero. These factors are applied to both the real and imaginary part of the aerodynamic matrices as well as all reduced frequencies used in the PK-Method which solves flutter problems. The modal-based AIC method
NASA Astrophysics Data System (ADS)
KIM, DONG-HYUN; LEE, IN
2000-07-01
A two-degree-of-freedom airfoil with a freeplay non-linearity in the pitch and plunge directions has been analyzed in the transonic and low-supersonic flow region, where aerodynamic non-linearities also exist. The primary purpose of this study is to show aeroelastic characteristics due to freeplay structural non-linearity in the transonic and low-supersonic regions. The unsteady aerodynamic forces on the airfoil were evaluated using two-dimensional unsteady Euler code, and the resulting aeroelastic equations are numerically integrated to obtain the aeroelastic time responses of the airfoil motions and to investigate the dynamic instability. The present model has been considered as a simple aeroelastic model, which is equivalent to the folding fin of an advanced generic missile. From the results of the present study, characteristics of important vibration responses and aeroelastic instabilities can be observed in the transonic and supersonic regions, especially considering the effect of structural non-linearity in the pitch and plunge directions. The regions of limit-cycle oscillation are shown at much lower velocities, especially in the supersonic flow region, than the divergent flutter velocities of the linear structure model. It is also shown that even small freeplay angles can lead to severe dynamic instabilities and dangerous fatigue conditions for the flight vehicle wings and control fins.
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
Aeroelastic simulation of higher harmonic control
NASA Technical Reports Server (NTRS)
Robinson, Lawson H.; Friedmann, Peretz P.
1994-01-01
This report describes the development of an aeroelastic analysis of a helicopter rotor and its application to the simulation of helicopter vibration reduction through higher harmonic control (HHC). An improved finite-state, time-domain model of unsteady aerodynamics is developed to capture high frequency aerodynamic effects. An improved trim procedure is implemented which accounts for flap, lead-lag, and torsional deformations of the blade. The effect of unsteady aerodynamics is studied and it is found that its impact on blade aeroelastic stability and low frequency response is small, but it has a significant influence on rotor hub vibrations. Several different HHC algorithms are implemented on a hingeless rotor and their effectiveness in reducing hub vibratory shears is compared. All the controllers are found to be quite effective, but very differing HHC inputs are required depending on the aerodynamic model used. Effects of HHC on rotor stability and power requirements are found to be quite small. Simulations of roughly equivalent articulated and hingeless rotors are carried out, and it is found that hingeless rotors can require considerably larger HHC inputs to reduce vibratory shears. This implies that the practical implementation of HHC on hingeless rotors might be considerably more difficult than on articulated rotors.
Response studies of rotors and rotor blades with application to aeroelastic tailoring
NASA Technical Reports Server (NTRS)
Friedmann, P. P.
1982-01-01
Various tools for the aeroelastic stability and response analysis of rotor blades in hover and forward flight were developed and incorporated in a comprehensive package capable of performing aeroelastic tailoring of rotor blades in forward flight. The results indicate that substantial vibration reductions, of order 15-40%, in the vibratory hub shears can be achieved by relatively small modifications of the initial design. Furthermore the optimized blade can be up to 20% lighter than the original design. Accomplishments are reported for the following tasks: (1) finite element modeling of rotary-wing aeroelastic problems in hover and forward flight; (2) development of numerical methods for calculating the aeroelastic response and stability of rotor blades in forward fight; (3) formulation of the helicopter air resonance problem in hover with active controls; and (4) optimum design of rotor blades for vibration reduction in forward flight.
NASA Astrophysics Data System (ADS)
Sabri, Farhad
circular cylindrical shell or truncated conical shell subjected to internal/external pressure and axial compression loading. This is a typical example of external liquid propellant tanks of space shuttles and re-entry vehicles where they may experience this kind of loading during the flight. In the current work, different end boundary conditions of a circular cylindrical shell with different filling ratios were analyzed. To the best author' knowledge this is the first study where this kind of complex loading and boundary conditions are treated together during such an analysis. Only static instability, divergence, was observed where it showed that the fluid filling ratio does not have any effect on the critical buckling pressure and axial compression. It only reduces the vibration frequencies. It also revealed that the pressurized shell loses its stability at a higher critical axial load. (ii) Aeroelastic analysis of empty or partially liquid filled circular cylindrical and conical shells. Different boundary conditions with different geometries of shells subjected to supersonic air flow are studied here. In all of cases shell loses its stability though the coupled mode flutter. The results showed that internal pressure has a stabilizing effect and increases the critical flutter speed. It is seen that the value of critical dynamic pressure changes rapidly and widely as the filling ratio increases from a low value. In addition, by increasing the length ratio the decrement of flutter speed is decreased and vanishes. This rapid change in critical dynamic pressure at low filling ratios and its almost steady behaviour at large filling ratios indicate that the fluid near the bottom of the shell is largely influenced by elastic deformation when a shell is subjected to external subsonic flow. Based on comparison with the existing numerical, analytical and experimental data and the power of capabilities of this hybrid finite element method to model different boundary conditions and
Flutter prediction for a wing with active aileron control
NASA Technical Reports Server (NTRS)
Penning, K.; Sandlin, D. R.
1983-01-01
A method for predicting the vibrational stability of an aircraft with an analog active aileron flutter suppression system (FSS) is expained. Active aileron refers to the use of an active control system connected to the aileron to damp vibrations. Wing vibrations are sensed by accelerometers and the information is used to deflect the aileron. Aerodynamic force caused by the aileron deflection oppose wing vibrations and effectively add additional damping to the system.
Aeroelastic Analyses of the SemiSpan SuperSonic Transport (S4T) Wind Tunnel Model at Mach 0.95
NASA Technical Reports Server (NTRS)
Hur, Jiyoung
2014-01-01
Detailed aeroelastic analyses of the SemiSpan SuperSonic Transport (S4T) wind tunnel model at Mach 0.95 with a 1.75deg fixed angle of attack are presented. First, a numerical procedure using the Computational Fluids Laboratory 3-Dimensional (CFL3D) Version 6.4 flow solver is investigated. The mesh update method for structured multi-block grids was successfully applied to the Navier-Stokes simulations. Second, the steady aerodynamic analyses with a rigid structure of the S4T wind tunnel model are reviewed in transonic flow. Third, the static analyses were performed for both the Euler and Navier-Stokes equations. Both the Euler and Navier-Stokes equations predicted a significant increase of lift forces, compared to the results from the rigid structure of the S4T wind-tunnel model, over various dynamic pressures. Finally, dynamic aeroelastic analyses were performed to investigate the flutter condition of the S4T wind tunnel model at the transonic Mach number. The condition of flutter was observed at a dynamic pressure of approximately 75.0-psf for the Navier-Stokes simulations. However, it was observed that the flutter condition occurred a dynamic pressure of approximately 47.27-psf for the Euler simulations. Also, the computational efficiency of the aeroelastic analyses for the S4T wind tunnel model has been assessed.
Investigating Stall Flutter using a DS model-An application for HAWTs
NASA Astrophysics Data System (ADS)
Nichols, James; Attorni, Andrea; Haans, Wouter; Witcher, David
2014-12-01
As wind turbine blades become larger there is a tendency for the blade torsional stiffness to reduce, producing the possibility of dynamic instability at moderate windspeeds. While linearised methods can assess the envelope of allowable blade properties for avoiding classical flutter with attached flow aerodynamics, wind turbine aerofoils can experience stalled flow. Therefore, it is necessary to explore the possible effects of stall-flutter on blade stability. This paper aims to address methods for judging the stability of blade designs during both attached flow and stalled flow behaviour. This paper covers the following areas: i) Attached flow model A Beddoes-Leishman indicial model is presented and the choice of coefficients is explained in the context of Theodorsen's theory for flat-plate aerofoils and experimental results by Beddoes and Leishman. Special attention is given to the differing dynamic behaviour of the pitching moment due to flapping motion, pitching motion and dynamically varying inflow. (ii) Classical flutter analysis The time domain attached flow model is verified against a linear flutter analysis by comparing time domain results for a 3D model of a representative multi-megawatt turbine blade, varying the position of the centre of mass along the chord. The results show agreement to within 6% for a range of flutter onset speeds. (iii) Dynamic stall model On entering the stalled region, damping of torsional motion of an aerofoil section can become negative. A dynamic stall model which encompasses the effects of trailing edge separation and leading edge vortex detachment is presented and validated against published experimental data. (iv) Stall flutter The resulting time domain model is used in simulations validating the prediction of reduced flutter onset for stalled aerofoils. Representative stalled conditions for a multi-megawatt wind turbine blade are investigated to assess the possible reduction in flutter speed. A maximum reduction of 17% is
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.
A root locus based flutter synthesis procedure
NASA Technical Reports Server (NTRS)
Hajela, P.
1983-01-01
An efficient generalized constraint is proposed in the context of a nonlinear mathematical programming approach for the minimum weight design of wing structures for flutter considerations. The approach is based on a root locus analysis procedure that is better suited for flutter redesign than the conventionally used V-g method. The proposed flutter constraint does not require an actual computation of the flutter speed, allows prescription of meaningful margins of safety in the optimized design and lends itself to elegant computation of sensitivity information. The approach is implemented and results presented for representative structural models.
Aeroelastic Flight Data Analysis with the Hilbert-Huang Algorithm
NASA Technical Reports Server (NTRS)
Brenner, Martin J.; Prazenica, Chad
2006-01-01
This report investigates the utility of the Hilbert Huang transform for the analysis of aeroelastic flight data. It is well known that the classical Hilbert transform can be used for time-frequency analysis of functions or signals. Unfortunately, the Hilbert transform can only be effectively applied to an extremely small class of signals, namely those that are characterized by a single frequency component at any instant in time. The recently-developed Hilbert Huang algorithm addresses the limitations of the classical Hilbert transform through a process known as empirical mode decomposition. Using this approach, the data is filtered into a series of intrinsic mode functions, each of which admits a well-behaved Hilbert transform. In this manner, the Hilbert Huang algorithm affords time-frequency analysis of a large class of signals. This powerful tool has been applied in the analysis of scientific data, structural system identification, mechanical system fault detection, and even image processing. The purpose of this report is to demonstrate the potential applications of the Hilbert Huang algorithm for the analysis of aeroelastic systems, with improvements such as localized online processing. Applications for correlations between system input and output, and amongst output sensors, are discussed to characterize the time-varying amplitude and frequency correlations present in the various components of multiple data channels. Online stability analyses and modal identification are also presented. Examples are given using aeroelastic test data from the F-18 Active Aeroelastic Wing airplane, an Aerostructures Test Wing, and pitch plunge simulation.
Plans for Aeroelastic Prediction Workshop
NASA Technical Reports Server (NTRS)
Heeg, Jennifer; Ballmann, Josef; Bhatia, Kumar; Blades, Eric; Boucke, Alexander; Chwalowski, Pawel; Dietz, Guido; Dowell, Earl; Florance, Jennifer P.; Hansen, Thorsten; Mani, Mori; Marvriplis, Dimitri; Perry, Boyd, III; Ritter, Markus; Schuster, David M.; Smith, Marilyn; Taylor, Paul; Whiting, Brent; Wieseman, Carol C.
2011-01-01
This paper summarizes the plans for the first Aeroelastic Prediction Workshop. The workshop is designed to assess the state of the art of computational methods for predicting unsteady flow fields and aeroelastic response. The goals are to provide an impartial forum to evaluate the effectiveness of existing computer codes and modeling techniques, and to identify computational and experimental areas needing additional research and development. Three subject configurations have been chosen from existing wind tunnel data sets where there is pertinent experimental data available for comparison. For each case chosen, the wind tunnel testing was conducted using forced oscillation of the model at specified frequencies
Including Aeroelastic Effects in the Calculation of X-33 Loads and Control Characteristics
NASA Technical Reports Server (NTRS)
Zeiler, Thomas A.
1998-01-01
Up until now, loads analyses of the X-33 RLV have been done at Marshall Space Flight Center (MSFC) using aerodynamic loads derived from CFD and wind tunnel models of a rigid vehicle. Control forces and moments are determined using a rigid vehicle trajectory analysis and the detailed control load distributions for achieving the desired control forces and moments, again on the rigid vehicle, are determined by Lockheed Martin Skunk Works. However, static aeroelastic effects upon the load distributions are not known. The static aeroelastic effects will generally redistribute external loads thereby affecting both the internal structural loads as well as the forces and moments generated by aerodynamic control surfaces. Therefore, predicted structural sizes as well as maneuvering requirements can be altered by consideration of static aeroelastic effects. The objective of the present work is the development of models and solutions for including static aeroelasticity in the calculation of X-33 loads and in the determination of stability and control derivatives.
Fast-Running Aeroelastic Code Based on Unsteady Linearized Aerodynamic Solver Developed
NASA Technical Reports Server (NTRS)
Reddy, T. S. R.; Bakhle, Milind A.; Keith, T., Jr.
2003-01-01
The NASA Glenn Research Center has been developing aeroelastic analyses for turbomachines for use by NASA and industry. An aeroelastic analysis consists of a structural dynamic model, an unsteady aerodynamic model, and a procedure to couple the two models. The structural models are well developed. Hence, most of the development for the aeroelastic analysis of turbomachines has involved adapting and using unsteady aerodynamic models. Two methods are used in developing unsteady aerodynamic analysis procedures for the flutter and forced response of turbomachines: (1) the time domain method and (2) the frequency domain method. Codes based on time domain methods require considerable computational time and, hence, cannot be used during the design process. Frequency domain methods eliminate the time dependence by assuming harmonic motion and, hence, require less computational time. Early frequency domain analyses methods neglected the important physics of steady loading on the analyses for simplicity. A fast-running unsteady aerodynamic code, LINFLUX, which includes steady loading and is based on the frequency domain method, has been modified for flutter and response calculations. LINFLUX, solves unsteady linearized Euler equations for calculating the unsteady aerodynamic forces on the blades, starting from a steady nonlinear aerodynamic solution. First, we obtained a steady aerodynamic solution for a given flow condition using the nonlinear unsteady aerodynamic code TURBO. A blade vibration analysis was done to determine the frequencies and mode shapes of the vibrating blades, and an interface code was used to convert the steady aerodynamic solution to a form required by LINFLUX. A preprocessor was used to interpolate the mode shapes from the structural dynamic mesh onto the computational dynamics mesh. Then, we used LINFLUX to calculate the unsteady aerodynamic forces for a given mode, frequency, and phase angle. A postprocessor read these unsteady pressures and
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.
NASA Technical Reports Server (NTRS)
Kvaternik, Raymond G.; Juang, Jer-Nan; Bennett, Richard L.
2000-01-01
The Aeroelasticity Branch at NASA Langley Research Center has a long and substantive history of tiltrotor aeroelastic research. That research has included a broad range of experimental investigations in the Langley Transonic Dynamics Tunnel (TDT) using a variety of scale models and the development of essential analyses. Since 1994, the tiltrotor research program has been using a 1/5-scale, semispan aeroelastic model of the V-22 designed and built by Bell Helicopter Textron Inc. (BHTI) in 1981. That model has been refurbished to form a tiltrotor research testbed called the Wing and Rotor Aeroelastic Test System (WRATS) for use in the TDT. In collaboration with BHTI, studies under the current tiltrotor research program are focused on aeroelastic technology areas having the potential for enhancing the commercial and military viability of tiltrotor aircraft. Among the areas being addressed, considerable emphasis is being directed to the evaluation of modern adaptive multi-input multi- output (MIMO) control techniques for active stability augmentation and vibration control of tiltrotor aircraft. As part of this investigation, a predictive control technique known as Generalized Predictive Control (GPC) is being studied to assess its potential for actively controlling the swashplate of tiltrotor aircraft to enhance aeroelastic stability in both helicopter and airplane modes of flight. This paper summarizes the exploratory numerical and experimental studies that were conducted as part of that investigation.
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.
CFD for applications to aircraft aeroelasticity
NASA Technical Reports Server (NTRS)
Guruswamy, Guru P.
1989-01-01
Strong interactions of structures and fluids are common in many engineering environments. Such interactions can give rise to physically important phenomena such as those occurring for aircraft due to aeroelasticity. Aeroelasticity can significantly influence the safe performance of aircraft. At present exact methods are available for making aeroelastic computations when flows are in either the linear subsonic or supersonic range. However, for complex flows containing shock waves, vortices and flow separations, computational methods are still under development. Several phenomena that can be dangerous and limit the performance of an aircraft occur due to the interaction of these complex flows with flexible aircraft components such as wings. For example, aircraft with highly swept wings experience vortex induced aeroelastic oscillations. Correct understanding of these complex aeroelastic phenomena requires direct coupling of fluids and structural equations. Here, a summary is presented of the development of such coupled methods and applications to aeroelasticity since about 1978 to present. The successful use of the transonic small perturbation theory (TSP) coupled with structures is discussed. This served as a major stepping stone for the current stage of aeroelasticity using computational fluid dynamics. The need for the use of more exact Euler/Navier-Stokes (ENS) equations for aeroelastic problems is explained. The current development of unsteady aerodynamic and aeroelastic procedures based on the ENS equations are discussed. Aeroelastic results computed using both TSP and ENS equations are discussed.
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)
Code of Federal Regulations, 2011 CFR
2011-01-01
... surfaces. (e) For turbopropeller-powered airplanes, the dynamic evaluation must include— (1) Whirl mode... damping as VD is approached. (c) Any rational analysis used to predict freedom from flutter, control... must be shown by analysis to be free from flutter up to VD/MD after fatigue failure, or obvious...
Code of Federal Regulations, 2010 CFR
2010-01-01
... surfaces. (e) For turbopropeller-powered airplanes, the dynamic evaluation must include— (1) Whirl mode... damping as VD is approached. (c) Any rational analysis used to predict freedom from flutter, control... must be shown by analysis to be free from flutter up to VD/MD after fatigue failure, or obvious...
Hiatal hernia squeezing the heart to flutter.
Patel, Arpan; Shah, Rushikesh; Nadavaram, Sravanthi; Aggarwal, Aakash
2014-04-01
An 80-year-old woman presented to the emergency department with failure to thrive and weakness for 14 days. Medical history was significant for polio. On admission her electrocardiogram showed atrial flutter, and cardiac enzymes were elevated. Echocardiogram revealed a high pulmonary artery pressure, but no other wall motion abnormalities or valvulopathies. Chest x-ray showed a large lucency likely representing a diaphragmatic hernia. Computed tomographic scan confirmed the hernia. Our patient remained in atrial flutter despite rate control, and thereafter surgery was consulted to evaluate the patient. She underwent hernia repair. After surgery, the patient was taken off rate control and monitored for 72 hours; she did not have any episode of atrial flutter and was discharged with follow up in a week showing no arrhythmia. Her flutter was caused directly by the mechanical effect of the large hiatal hernia pressing against her heart, as the flutter resolved after the operation.
Stability of dynamic response of suspension bridges
NASA Astrophysics Data System (ADS)
Capsoni, Antonio; Ardito, Raffaele; Guerrieri, Andrea
2017-04-01
The potential occurrence of internal parametric resonance phenomena has been recently indicated as a potential contributory cause of the appearance of critical dynamic states in long-span suspension bridges. At the same time, suspension bridges, in view of their flexibility, are prone to aeroelastic response, such as vortex shedding, torsional divergence and flutter. In this paper, a non-linear dynamic model of a suspension bridge is devised, with the purpose of providing a first attempt toward a unified framework for the study of aeroelastic and internal resonance instabilities. Inspired by the pioneering work of Herrmann and Hauger, the analyses have been based on a linearized formulation that is able to represent the main structural non-linear effects and the coupling given by aerodynamic forces. The results confirm that the interaction between aeroelastic effects and non-linear internal resonance leads to unstable conditions for wind speeds which can be lower than the critical threshold for standard aeroelastic predictions.
Aeroelastic Sizing for High-Speed Research (HSR) Longitudinal Control Alternatives Project (LCAP)
NASA Technical Reports Server (NTRS)
Walsh, Joanne L.; Dunn, H. J.; Stroud, W. Jefferson; Barthelemy, J.-F.; Weston, Robert P.; Martin, Carl J.; Bennett, Robert M.
2005-01-01
The Longitudinal Control Alternatives Project (LCAP) compared three high-speed civil transport configurations to determine potential advantages of the three associated longitudinal control concepts. The three aircraft configurations included a conventional configuration with a layout having a horizontal aft tail, a configuration with a forward canard in addition to a horizontal aft tail, and a configuration with only a forward canard. The three configurations were aeroelastically sized and were compared on the basis of operational empty weight (OEW) and longitudinal control characteristics. The sized structure consisted of composite honeycomb sandwich panels on both the wing and the fuselage. Design variables were the core depth of the sandwich and the thicknesses of the composite material which made up the face sheets of the sandwich. Each configuration was sized for minimum structural weight under linear and nonlinear aeroelastic loads subject to strain, buckling, ply-mixture, and subsonic and supersonic flutter constraints. This report describes the methods that were used and the results that were generated for the aeroelastic sizing of the three configurations.
NASA Astrophysics Data System (ADS)
Otsuka, Keisuke; Makihara, Kanjuro
2016-05-01
Morphing wings have been developed by several organizations for a variety of applications including the changing of flight ability while in the air and reducing the amount of space required to store an aircraft. One such example of morphing wings is the deployable wing that is expected to be used for Mars exploration. When designing wings, aeroelastic simulation is important to prevent the occurrence of destructive phenomena while the wing is in use. Flutter and divergence are typical issues to be addressed. However, it has been difficult to simulate the aeroelastic motion of deployable wings because of the significant differences between these deployable wings and conventional designs. The most apparent difference is the kinematic constraints of deployment, typically a hinge joint. These constraints lead not only to deformation but also to rigid body rotation. This research provides a novel method of overcoming the difficulties associated with handling these kinematic constraints. The proposed method utilizes flexible multibody dynamics and absolute nodal coordinate formulation to describe the dynamic motion of a deployable wing. This paper presents the simulation of the rigid body rotation around the kinematic constraints as induced by the aeroelasticity. The practicality of the proposed method is confirmed.
Active aeroelastic control of aircraft composite wings impacted by explosive blasts
NASA Astrophysics Data System (ADS)
Librescu, Liviu; Na, Sungsoo; Qin, Zhanming; Lee, Bokhee
2008-11-01
In this paper, the dynamic aeroelastic response and the related robust control of aircraft swept wings exposed to gust and explosive type loads are examined. The structural model of the wing is in the form of a thin/thick-walled beam and incorporates a number of non-standard effects, such as transverse shear, material anisotropy, warping inhibition, the spanwise non-uniformity of the cross-section, and the rotatory inertias. The circumferentially asymmetric stiffness lay-up configuration is implemented to generate preferred elastic couplings, and in this context, the implications of the plunging-twist elastic coupling and of warping inhibition on the aeroelastic response are investigated. The unsteady incompressible aerodynamic theory adopted in this study is that by von-Kármán and Sears, applicable to arbitrary small motion in the time domain. The considered control methodology enabling one to enhance the aeroelastic response in the subcritical flight speed range and to suppress the occurrence of the flutter instability is based on a novel control approach that is aimed to improve the robustness to modeling uncertainties and external disturbances. To this end, a combined control based on Linear Quadratic Gaussian (LQG) controller coupled with the Sliding Mode Observer (SMO) is designed and its high efficiency is put into evidence.
Model order reduction applied to a hot-bench simulation of an aeroelastic wind-tunnel model
NASA Technical Reports Server (NTRS)
Buttrill, Carey S.; Bacon, Barton J.
1991-01-01
Simulations of an aeroelastically scaled wind-tunnel model were developed for hot-bench testing of a digital controller. The digital controller provided active flutter-suppression, rolling-maneuver-load alleviation, and plant estimation. To achieve an acceptable time scale for the hot-bench application, the mathematical model of the wind-tunnel model was reduced from 220 states to approximately 130 states while assuring that the required accuracy was preserved for all combinations of 10 inputs and 56 outputs. The reduction was achieved by focussing on a linear, aeroelastic submodel of the full mathematical model and by applying a method based on the internally balanced realization of a dynamic system. The error-bound properties of the internally balanced realization significantly contribute to its utility in the model reduction process. The reduction method and the results achieved are described.
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.
NASA Technical Reports Server (NTRS)
Mourey, D. J.
1979-01-01
The aspects of flight testing an aeroelastically tailored forward swept research wing on a BQM-34F drone vehicle are examined. The geometry of a forward swept wing, which is incorporated into the BQM-34F to maintain satisfactory flight performance, stability, and control is defined. A preliminary design of the aeroelastically tailored forward swept wing is presented.
Airfoil flutter model suspension system
NASA Technical Reports Server (NTRS)
Reed, Wilmer H. (Inventor)
1987-01-01
A wind tunnel suspension system for testing flutter models under various loads and at various angles of attack is described. The invention comprises a mounting bracket assembly affixing the suspension system to the wind tunnel, a drag-link assembly and a compound spring arrangement comprises a plunge spring working in opposition to a compressive spring so as to provide a high stiffness to trim out steady state loads and simultaneously a low stiffness to dynamic loads. By this arrangement an airfoil may be tested for oscillatory response in both plunge and pitch modes while being held under high lifting loads in a wind tunnel.
NASA Technical Reports Server (NTRS)
Radovcich, N. A.
1984-01-01
The design experience associated with a benchmark aeroelastic design of an out of production transport aircraft is discussed. Current work being performed on a high aspect ratio wing design is reported. The Preliminary Aeroelastic Design of Structures (PADS) system is briefly summarized and some operational aspects of generating the design in an automated aeroelastic design environment are discussed.
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.
1983-09-30
cycle S0 t time s V velocity m/s 7i 7 Vre f reference velocity for reduced frequency and m/sStrouhal number: Vref V for compresor cascade Vref V2 for...dimensional cascades. Such interesting phenomena as rotor-stator interact ions, stalled flutter and fully three-dimensional effects will thus be excluded...aeroelastic phenomena under in~est gation instabili- ties due to stall , choke, shockwa\\es, coupling effects between the stead’ and unstead> flow fields...). The
Experimental Results from the Active Aeroelastic Wing Wind Tunnel Test Program
NASA Technical Reports Server (NTRS)
Heeg, Jennifer; Spain, Charles V.; Florance, James R.; Wieseman, Carol D.; Ivanco, Thomas G.; DeMoss, Joshua; Silva, Walter A.; Panetta, Andrew; Lively, Peter; Tumwa, Vic
2005-01-01
The Active Aeroelastic Wing (AAW) program is a cooperative effort among NASA, the Air Force Research Laboratory and the Boeing Company, encompassing flight testing, wind tunnel testing and analyses. The objective of the AAW program is to investigate the improvements that can be realized by exploiting aeroelastic characteristics, rather than viewing them as a detriment to vehicle performance and stability. To meet this objective, a wind tunnel model was crafted to duplicate the static aeroelastic behavior of the AAW flight vehicle. The model was tested in the NASA Langley Transonic Dynamics Tunnel in July and August 2004. The wind tunnel investigation served the program goal in three ways. First, the wind tunnel provided a benchmark for comparison with the flight vehicle and various levels of theoretical analyses. Second, it provided detailed insight highlighting the effects of individual parameters upon the aeroelastic response of the AAW vehicle. This parameter identification can then be used for future aeroelastic vehicle design guidance. Third, it provided data to validate scaling laws and their applicability with respect to statically scaled aeroelastic models.
Subsonic-transonic stall flutter study
NASA Technical Reports Server (NTRS)
Stardter, H.
1979-01-01
The objective of the Subsonic/Transonic Stall Flutter Program was to obtain detailed measurements of both the steady and unsteady flow field surrounding a rotor and the mechanical state of the rotor while it was operating in both steady and flutter modes to provide a basis for future analysis and for development of theories describing the flutter phenomenon. The program revealed that while all blades flutter at the same frequency, they do not flutter at the same amplitude, and their interblade phase angles are not equal. Such a pattern represents the superposition of a number of rotating nodal diameter patterns, each characterized by a different amplitude and different phase indexing, but each rotating at a speed that results in the same flutter frequency as seen in the rotor system. Review of the steady pressure contours indicated that flutter may alter the blade passage pressure distribution. The unsteady pressure amplitude contour maps reveal regions of high unsteady pressure amplitudes near the leading edge, lower amplitudes near the trailing.
Supersonic flutter of composite sandwich panels
NASA Astrophysics Data System (ADS)
Shiau, Le-Chung
1992-12-01
A flutter-motion equation is presently derived for a 2D composite sandwich panel considering the total lateral displacement of the plate as the sum of the displacement due to bending of the plate, and that which is due to shear deformation at the core. The effects of core thickness and stacking sequence of the faces on the flutter boundary of the plate are discussed; it is shown that the sandwich panel greatly improves the flutter boundary over that of a composite laminate panel, provided it has sufficient core thickness.
Static Aeroelasticity in Combat Aircraft.
1986-01-01
aircraft design. Fuselage flexibility is, in general , a secondary consideration. The relatively high density of this structural component, designed to...representation of the structure. An effective beam representation of the total panel stiffness is generally applicable and appropriate for these needs and...loading effect Is to produce zero wing lift, but a large leading-edge-up wing torque. Aeroelastically, a significant wing lift is generated as the
Active flutter suppression - Control system design and experimental validation
NASA Technical Reports Server (NTRS)
Waszak, Martin R.; Srinathkumar, S.
1991-01-01
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 with extensive use of 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 errors in flutter dynamic pressure and flutter frequency in the mathematical model. The flutter suppression controller was also successfully operated in combination with a roll maneuver controller to perform flutter suppression during rapid rolling maneuvers.
Atrial flutter a manifestation of cardiac tamponade.
Pabón, Guillermo Mora; Ramírez, John A
2012-04-01
Atrial flutter (AFL) is a common arrhythmia that is associated with postpericardiotomy and pericarditis. The relationship of AFL with tamponade has rarely been reported. A case of AFL with acute pericarditis and cardiac tamponade is thus presented here.
Aeroelastic Modeling of X-56A Stiff-Wing Configuration Flight Test Data
NASA Technical Reports Server (NTRS)
Grauer, Jared A.; Boucher, Matthew J.
2017-01-01
Aeroelastic stability and control derivatives for the X-56A Multi-Utility Technology Testbed (MUTT), in the stiff-wing configuration, were estimated from flight test data using the output-error method. Practical aspects of the analysis are discussed. The orthogonal phase-optimized multisine inputs provided excellent data information for aeroelastic modeling. Consistent parameter estimates were determined using output error in both the frequency and time domains. The frequency domain analysis converged faster and was less sensitive to starting values for the model parameters, which was useful for determining the aeroelastic model structure and obtaining starting values for the time domain analysis. Including a modal description of the structure from a finite element model reduced the complexity of the estimation problem and improved the modeling results. Effects of reducing the model order on the short period stability and control derivatives were investigated.
Flight flutter testing the B-58 airplane
NASA Technical Reports Server (NTRS)
Mahaffey, P. T.
1975-01-01
The flight flutter tests on the B-58 airplane are described, and the philosophy of flight flutter testing is discussed. The instrumentation used in the airplane and in the telemetering receiving station on the ground is described along with the methods used for exciting the airplane and the flight test procedure. Also described is the type of data obtained and its reduction. An evaluation of the procedure and instrumentation is given with a discussion of desirable improvements for future testing.
Active Feedback Control of a Web Flutter Using Flow Control Devices
NASA Astrophysics Data System (ADS)
Hayashi, Yusuke; Watanabe, Masahiro; Hara, Kensuke
This paper develops a non-contact active feedback control of web flutter in a narrow passage by using movable plates set at inlet and outlet of the passage. The strategy of this active feedback control is based on the flow-control which cancels the exciting fluid force acting on the web, i.e., cancels the self-excited feedback mechanism. In this paper, suppression of the web flutter by the active feedback control is demonstrated experimentally. In the experiments, a web (film), as a controlled object, is subjected to air flow in a narrow passage. The web flutter occurs to the web in the translational motion over the critical flow velocity. And the web flutter is actively controlled and suppressed by the movable plate motion which changes the air flow in the passage. The critical flow velocity under controlled condition is examined with changing the controller gain and phase-shift between the web motion and the movable plate motion. As a result, it is indicated that the active feedback control increases the critical flow velocity, and suppress the web flutter effectively. Moreover, the control performance is examined experimentally, and stabilization mechanism by the active feedback control is discussed.
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.
NASA Technical Reports Server (NTRS)
Schlosser, D. C.; Dominik, D. F.
1983-01-01
The static aeroelastic effects on the longitudinal stability and elevon/aileron effectiveness of the space transportation system (STS) Space Shuttle orbiter were estimated by a simplified approach called the elevon torsional stiffness (ETS) method. This method employs rigid model wind tunnel test results to predict aeroelastic effects. Lateral/directional stability and rudder effectiveness were based on results of a wind tunnel test in which a flexible tail model was used. Comparisons with selective flight data are made in this paper. Results of correlations with flight data (although limited at the present time) verify the predicted aeroelastic effects for the orbiter. The orbiter's structural characteristics are such that the effects of aeroelasticity, whether estimated using analytical techniques or simplified methods, do not appear to affect the vehicle performance to any great extent. The large amount of scatter in the flight-extracted data made verification of the aeroelastic corrections very difficult. Generally, the simplified elevon torsional stiffness method provided better correlation with flight test results than he analytical method and reduced the verification effort and cost.
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.
The application of digital computers to near-real-time processing of flutter test data
NASA Technical Reports Server (NTRS)
Hurley, S. R.
1976-01-01
Procedures used in monitoring, analyzing, and displaying flight and ground flutter test data are presented. These procedures include three digital computer programs developed to process structural response data in near real time. Qualitative and quantitative modal stability data are derived from time history response data resulting from rapid sinusoidal frequency sweep forcing functions, tuned-mode quick stops, and pilot induced control pulses. The techniques have been applied to both fixed and rotary wing aircraft, during flight, whirl tower rotor systems tests, and wind tunnel flutter model tests. An hydraulically driven oscillatory aerodynamic vane excitation system utilized during the flight flutter test programs accomplished during Lockheed L-1011 and S-3A development is described.
Time-Delayed Feedback Control for Flutter of Supersonic Aircraft Wing
NASA Astrophysics Data System (ADS)
Zhang, Shu; Huang, Yu; Xu, Jian
An active control technique called servo delayed feedback control is proposed to control the flutter of supersonic aircraft wing. It's motivated to increase the critical flow velocity. Firstly, the servo delayed feedback control is designed based on a two-dimensional airfoil so that delayed differential equations are modelled for the controlled system under consideration. Then, the stability of the system without time delay and with time delayed feedback control are considered analytically and flutter boundary of the parameters in the delayed feedback control system is predicted when time delay varies. Finally, numerical simulation for time domain with MATLAB/SIMULINK software is made to demonstrate the effectiveness of the theoretical result. The results show that, critical flow velocity can be increased by regulating the quantity of time delay and the provided strategy of delayed feedback to control the flutter in supersonic aircraft wing system is not only valid but also easily applied to engineering structures.
Fan Stall Flutter Flow Mechanism Studied
NASA Technical Reports Server (NTRS)
Lepicovsky, Jan
2002-01-01
Modern turbofan engines employ a highly loaded fan stage with transonic or low-supersonic velocities in the blade-tip region. The fan blades are often prone to flutter at off-design conditions. Flutter is a highly undesirable and dangerous self-excited mode of blade oscillations that can result in high-cycle fatigue blade failure. The origins of blade flutter are not fully understood yet. Experimental data that can be used to clarify the origins of blade flutter in modern transonic fan designs are very limited. The Transonic Flutter Cascade Facility at the NASA Glenn Research Center was developed to experimentally study the details of flow mechanisms associated with fan flutter. The cascade airfoils are instrumented to measure high-frequency unsteady flow variations in addition to the steady flow data normally recorded in cascade tests. The test program measures the variation in surface pressure in response to the oscillation of one or more of the cascade airfoils. However, during the initial phases of the program when all airfoils were in fixed positions, conditions were found where significant time variations in the pressures near the airfoil leading edges could be observed.
Method for experimental determination of flutter speed by parameter identification
NASA Technical Reports Server (NTRS)
Nissim, E.; Gilyard, Glenn B.
1989-01-01
A method for flight flutter testing is proposed which enables one to determine the flutter dynamic pressure from flights flown far below the flutter dynamic pressure. The method is based on the identification of the coefficients of the equations of motion at low dynamic pressures, followed by the solution of these equations to compute the flutter dynamic pressure. The initial results of simulated data reported in the present work indicate that the method can accurately predict the flutter dynamic pressure, as described. If no insurmountable difficulties arise in the implementation of this method, it may significantly improve the procedures for flight flutter testing.
Aeroelastic Flight Data Analysis with the Hilbert-Huang Algorithm
NASA Technical Reports Server (NTRS)
Brenner, Marty; Prazenica, Chad
2005-01-01
This paper investigates the utility of the Hilbert-Huang transform for the analysis of aeroelastic flight data. It is well known that the classical Hilbert transform can be used for time-frequency analysis of functions or signals. Unfortunately, the Hilbert transform can only be effectively applied to an extremely small class of signals, namely those that are characterized by a single frequency component at any instant in time. The recently-developed Hilbert-Huang algorithm addresses the limitations of the classical Hilbert transform through a process known as empirical mode decomposition. Using this approach, the data is filtered into a series of intrinsic mode functions, each of which admits a well-behaved Hilbert transform. In this manner, the Hilbert-Huang algorithm affords time-frequency analysis of a large class of signals. This powerful tool has been applied in the analysis of scientific data, structural system identification, mechanical system fault detection, and even image processing. The purpose of this paper is to demonstrate the potential applications of the Hilbert-Huang algorithm for the analysis of aeroelastic systems, with improvements such as localized/online processing. Applications for correlations between system input and output, and amongst output sensors, are discussed to characterize the time-varying amplitude and frequency correlations present in the various components of multiple data channels. Online stability analyses and modal identification are also presented. Examples are given using aeroelastic test data from the F/A-18 Active Aeroelastic Wing aircraft, an Aerostructures Test Wing, and pitch-plunge simulation.
NASA Technical Reports Server (NTRS)
Curtiss, H. C., Jr.; Komatsuzaki, T.; Traybar, J. J.
1979-01-01
The influence of single loop feedbacks to improve the stability of the system are considered. Reduced order dynamic models are employed where appropriate to promote physical insight. The influence of fuselage freedom on the aeroelastic stability, and the influence of the airframe flexibility on the low frequency modes of motion relevant to the stability and control characteristics of the vehicle were examined.
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.
Design for coupled-mode flutter and non-synchronous vibration in turbomachinery
NASA Astrophysics Data System (ADS)
Clark, Stephen Thomas
This research presents the detailed investigation of coupled-mode flutter and non-synchronous vibration in turbomachinery. Coupled-mode flutter and non-synchronous vibration are two aeromechanical challenges in designing turbomachinery that, when present, can cause engine blade failure. Regarding flutter, current industry design practices calculate the aerodynamic loads on a blade due to a single mode. In response to these design standards, a quasi three-dimensional, reduced-order modeling tool was developed for identifying the aeroelastic conditions that cause multi-mode flutter. This tool predicts the onset of coupled-mode flutter reasonable well for four different configurations, though certain parameters were tuned to agree with experimentation. Additionally, the results of this research indicate that mass ratio, frequency separation, and solidity have an effect on critical rotor speed for flutter. Higher mass-ratio blades require larger rotational velocities before they experience coupled-mode flutter. Similarly, increasing the frequency separation between modes and raising the solidity increases the critical rotor speed. Finally, and most importantly, design guidelines were generated for defining when a multi-mode flutter analysis is required in practical turbomachinery design. Previous work has shown that industry computational fluid dynamics can approximately predict non-synchronous vibration (NSV), but no real understanding of frequency lock-in and blade limit-cycle amplitude exists. Therefore, to understand the causes of NSV, two different reduced-order modeling approaches were used. The first approach uses a van der Pol oscillator to model a non-linear fluid instability. The van der Pol model is then coupled to a structural degree of freedom. This coupled system exhibits the two chief properties seen in experimental and computational non-synchronous vibration. Under various conditions, the fluid instability and the natural structural frequency will lock
Aeroelastic Control of a Segmented Trailing Edge Using Fiber Optic Strain Sensing Technology
NASA Technical Reports Server (NTRS)
Graham, Corbin Jay; Martins, Benjamin; Suppanade, Nathan
2014-01-01
Currently, design of aircraft structures incorporate a safety factor which is essentially an over design to mitigate the risk of structure failure during operation. Typically this safety factor is to design the structure to withstand loads much greater than what is expected to be experienced during flight. NASA Dryden Flight Research Centers has developed a Fiber Optic Strain Sensing (FOSS) system which can measure strain values in real-time. The Aeroelastics Lab at the AERO Institute is developing a segmented trailing edged wing with multiple control surfaces that can utilize the data from the FOSS system, in conjunction with an adaptive controller to redistribute the lift across a wing. This redistribution can decrease the amount of strain experienced by the wing as well as be used to dampen vibration and reduce flutter.
NASA Technical Reports Server (NTRS)
Nagaraja, K. S.; Kraft, R. H.
1999-01-01
The HSCT Flight Controls Group has developed longitudinal control laws, utilizing PTC aeroelastic flexible models to minimize aeroservoelastic interaction effects, for a number of flight conditions. The control law design process resulted in a higher order controller and utilized a large number of sensors distributed along the body for minimizing the flexibility effects. Processes were developed to implement these higher order control laws for performing the dynamic gust loads and flutter analyses. The processes and its validation were documented in Reference 2, for selected flight condition. The analytical results for additional flight conditions are presented in this document for further validation.
14 CFR 29.629 - Flutter and divergence.
Code of Federal Regulations, 2010 CFR
2010-01-01
... AIRWORTHINESS STANDARDS: TRANSPORT CATEGORY ROTORCRAFT Design and Construction General § 29.629 Flutter and divergence. Each aerodynamic surface of the rotorcraft must be free from flutter and divergence under...
Analytic simulation of higher harmonic control using a new aeroelastic model
NASA Technical Reports Server (NTRS)
Friedmann, P. P.; Robinson, L. H.
1989-01-01
This paper describes a higher harmonic control (HHC) study of a four bladed hingeless rotor using a coupled flap-lag-torsional aeroelastic stability and response analysis which incorporates finite-state, time-domain aerodynamics. The rotor trim condition is determined using a coupled trim-aeroelastic analysis. Deterministic and cautious controllers based on local and global HHC models are implemented with different combinations of input parameters identified using a Kalman filter. The effects of unsteady versus quasisteady aerodynamic modeling on HHC simulations are investigated, including the effectiveness of the local and global HHC models and the advantages of different identification schemes.
Flight Dynamics of Flexible Aircraft with Aeroelastic and Inertial Force Interactions
NASA Technical Reports Server (NTRS)
Nguyen, Nhan T.; Tuzcu, Ilhan
2009-01-01
This paper presents an integrated flight dynamic modeling method for flexible aircraft that captures coupled physics effects due to inertial forces, aeroelasticity, and propulsive forces that are normally present in flight. The present approach formulates the coupled flight dynamics using a structural dynamic modeling method that describes the elasticity of a flexible, twisted, swept wing using an equivalent beam-rod model. The structural dynamic model allows for three types of wing elastic motion: flapwise bending, chordwise bending, and torsion. Inertial force coupling with the wing elasticity is formulated to account for aircraft acceleration. The structural deflections create an effective aeroelastic angle of attack that affects the rigid-body motion of flexible aircraft. The aeroelastic effect contributes to aerodynamic damping forces that can influence aerodynamic stability. For wing-mounted engines, wing flexibility can cause the propulsive forces and moments to couple with the wing elastic motion. The integrated flight dynamics for a flexible aircraft are formulated by including generalized coordinate variables associated with the aeroelastic-propulsive forces and moments in the standard state-space form for six degree-of-freedom flight dynamics. A computational structural model for a generic transport aircraft has been created. The eigenvalue analysis is performed to compute aeroelastic frequencies and aerodynamic damping. The results will be used to construct an integrated flight dynamic model of a flexible generic transport aircraft.
Dynamic response and aeroelastic analysis of a propeller blade of a prop-fan engine
NASA Astrophysics Data System (ADS)
Joo, Gene; Lee, Hae-Kyung
Blades are modeled as cantilevered sandwich plates with Gr/Ep composite faces and orthotropic cores and also as curved twisted beams for the aeroelastic analysis. A free vibration analysis for the cantilevered sandwich plate model is performed using Rayleigh-Ritz method. Calculated results are compared with FEM codes and free vibration test results. A free vibration equation for the aeroelastic analysis is obtained by small linear perturbation about the nonlinear static equilibrium position of the curved and twisted beam model. An aeroelastic stability is analyzed along with unsteady aerodynamic analysis results with 2-D cascade effects. For analyzing dynamic response of the real prop-fan blade mode, F.E.M. codes are used. In order to verify computed results, SR-3 composite prop-fan blades with various stacking sequencies are manufactured. Natural frequencies of prop-fan specimen are obtained by modal testing method using impact hammer and FFT analyzer.
Simulation of Non-Linear Transonic Aeroelastic Behavior on the B-2
NASA Technical Reports Server (NTRS)
Dreim, D. R.; Jacobson, S. B.; Britt, R. T.
1999-01-01
At high subsonic flight speeds, large flexible aircraft begin to encounter unsteady airloads which are not predicted by most currently available aerodynamic analysis and design methods. With increasing speed and the development of transonic flow and shocks, viscous effects quickly become very important, and flow separation can occur. The Northrop Grumman USAF B-2 Bomber encountered a nonlinear aeroelastic Residual Pitch Oscillation (RPO) under these conditions. Simulation studies were performed with the Computational Aeroelasticity Program-Transonic Small Disturbance, Viscous (CAPTSDv) computer program to evaluate its ability to predict these nonlinear aeroelastic responses. Open and closed loop simulations were performed to assess the participation of the flight control system. Control, system actuator hysteresis characteristics were modeled and found to be a significant participant in the RPO phenomenon. Simulations were also performed for varying Mach numbers and altitudes to establish the stability boundaries and compare with flight test data. These CAPTSDv simulations compared well with flight data and revealed many potential further modeling enhancements.
Prediction of wing aeroelastic effects on aircraft life and pitching moment characteristics
NASA Technical Reports Server (NTRS)
Eckstrom, Clinton V.
1987-01-01
The distribution of flight loads on an aircraft structure determine the lift and pitching moment characteristics of the aircraft. When the load distribution changes due to the aeroelastic response of the structure, the lift and pitching moment characteristics also change. An estimate of the effect of aeroelasticity on stability and control characteristics is often required for the development of aircraft simulation models of evaluation of flight characteristics. This presentation outlines a procedure for incorporating calculated linear aeroelastic effects into measured nonlinear lift and pitching moment data from wind tunnel tests. Results are presented which were obtained from applying this procedure to data for an aircraft with a very flexible transport type research wing. The procedure described is generally applicable to all types of aircraft.
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)
Blade mistuning coupled with shaft flexibility effects in rotor aeroelasticity
NASA Technical Reports Server (NTRS)
Khader, Naim; Loewy, Robert G.
1989-01-01
The effect of bladed-disk polar dissymmetry, resulting from variations in mass from one blade to another, on aeroelastic stability boundaries for a fan stage is presented. In addition to both in-plane and out-of-plane deformations of the bladed-disk, bending of the supporting shaft in two planes is considered, and the resulting Coriolis forces and gyroscopic moments are included in the analysis. A quasi-steady aerodynamics approach is combined with the Lagrangian method to develop the governing equations of motion for the flexible bladed-disk-shaft assembly. Calculations are performed for an actual fan stage.
Transonic Aeroelasticity Analysis For Helicopter Rotor Blade
NASA Technical Reports Server (NTRS)
Chang, I-Chung; Gea, Lie-Mine; Chow, Chuen-Yen
1991-01-01
Numerical-simulation method for aeroelasticity analysis of helicopter rotor blade combines established techniques for analysis of aerodynamics and vibrations of blade. Application of method clearly shows elasticity of blade modifies flow and, consequently, aerodynamic loads on blade.
Wing-Body Aeroelasticity on Parallel Computers
NASA Technical Reports Server (NTRS)
Guruswamy, Guru P.; Byun, Chansup
1996-01-01
This article presents a procedure for computing the aeroelasticity of wing-body configurations on multiple-instruction, multiple-data parallel computers. In this procedure, fluids are modeled using Euler equations discretized by a finite difference method, and structures are modeled using finite element equations. The procedure is designed in such a way that each discipline can be developed and maintained independently by using a domain decomposition approach. A parallel integration scheme is used to compute aeroelastic responses by solving the coupled fluid and structural equations concurrently while keeping modularity of each discipline. The present procedure is validated by computing the aeroelastic response of a wing and comparing with experiment. Aeroelastic computations are illustrated for a high speed civil transport type wing-body configuration.
Flutter of laminated plates in supersonic flow
NASA Technical Reports Server (NTRS)
Sawyer, J. W.
1975-01-01
A solution procedure was developed using linear small deflection theory for the flutter of simply supported laminated plates. For such plates, the bending and extensional governing equations are coupled and have cross-stiffness terms which do not appear in classical plate theory. An extended Galerkin method is used to obtain approximate solutions to the governing equations, and the aerodynamic pressure loading used in the analysis is that given by linear piston theory with flow at arbitrary cross-flow angle. A limited parametric study was conducted for typical laminated composite plates. The calculations show that both the bending-extensional coupling and the cross-stiffness terms have a large destabilizing effect on flutter. Since classical plate theory does not consider bending-extensional coupling and cross stiffness terms, it usually gives inaccurate and nonconservative flutter boundaries for laminated plates.
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.
Aeroelastic analysis of versatile thermal insulation (VTI) panels with pinched boundary conditions
NASA Astrophysics Data System (ADS)
Carrera, Erasmo; Zappino, Enrico; Patočka, Karel; Komarek, Martin; Ferrarese, Adriano; Montabone, Mauro; Kotzias, Bernhard; Huermann, Brian; Schwane, Richard
2014-03-01
Launch vehicle design and analysis is a crucial problem in space engineering. The large range of external conditions and the complexity of space vehicles make the solution of the problem really challenging. The problem considered in the present work deals with the versatile thermal insulation (VTI) panel. This thermal protection system is designed to reduce heat fluxes on the LH2 tank during the long coasting phases. Because of the unconventional boundary conditions and the large-scale geometry of the panel, the aeroelastic behaviour of VTI is investigated in the present work. Known available results from literature related to similar problem, are reviewed by considering the effect of various Mach regimes, including boundary layer thickness effects, in-plane mechanical and thermal loads, non-linear effects and amplitude of limit cycle oscillations. A dedicated finite element model is developed for the supersonic regime. The models used for coupling the orthotropic layered structural model with Piston Theory aerodynamic models allow the calculations of flutter conditions in case of curved panels supported in a discrete number of points. An advanced computational aeroelasticity tool is developed using various dedicated commercial softwares (CFX, ZAERO, EDGE). A wind tunnel test campaign is carried out to assess the computational tool in the analysis of this type of problem.
Parallel Nonlinear Aeroelastic Computation for Fighter Wings in the Transonic Region
NASA Astrophysics Data System (ADS)
Larsen, Bradley Robert
In this dissertation, a parallel three-dimensional aeroelastic simulation is applied to current and next generation fighter aircraft wings. The computational model is a nonlinear fluid and structural mesh coupled using the Direct Eulerian-Langrangian method. This method attaches unique local coordinates to each node and connects the fluid mesh to the structure in such a way that a transformation preserved to the global coordinates. This allows the fluid and structure to be updated in the same time step and maintains spatial accuracy at their interface. The structural mesh is modeled using modified nonlinear von Karman finite elements and is discretized using the Galerkin finite element method. The fluid mesh also used the Galerkin finite element method to discretize the unsteady Euler equations. Computational results over a large range of Mach numbers and densities are presented for two candidate fighter wing models for transonic wing tunnel testing. The FX-35 is a trapezoidal wing based on the F-35A, and the F-Wing is a truncated delta wing similar to the F-16. Both wings exhibit a variety of flutter behaviors including strong bending-torsion flutter, limit-cycle oscillations, and essentially single degree-of-freedom responses.
Decoupler pylon: wing/store flutter suppressor
NASA Technical Reports Server (NTRS)
Reed, W. H., III (Inventor)
1982-01-01
A device for suspending a store from a support such as an aircraft wing and more specifically for increasing the flutter speed of an aircraft flying with attached store and reducing the sensitivity of flutter to changes in the pitch inertia and center of gravity location of the store is described. It comprises softspring where the store pitch mode is decoupled from support modes and a low frequency active control mechanism which maintains store alignment. A pneumatic suspension system both isolates the store in pitch and, under conditions of changing mean load, aligns the store with the wing to which it is attached.
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.
Transonic flutter calculations using the Euler equations
NASA Technical Reports Server (NTRS)
Bendiksen, Oddvar O.; Kousen, Kenneth A.
1989-01-01
In transonic flutter problems where shock motion plays an important part, it is believed that accurate predictions of the flutter boundaries will require the use of codes based on the Euler equations. Only Euler codes can obtain the correct shock location and shock strength, and the crucially important shock excursion amplitude and phase lag. The present study is based on the finite volume scheme developed by Jameson and Venkatakrishnan for the 2-D unsteady Euler equations. The equations are solved in integral form on a moving grid. The variable are pressure, density, Cartesian velocity components, and total energy.
Flight flutter testing using pulse techniques
NASA Technical Reports Server (NTRS)
Stringham, R. H., Jr.; Lenk, E. J.
1975-01-01
A case of flutter developed at a speed lower than had been flown previously. This incident precipitated the routine procedure of pulsing control surfaces as well as the firing of explosive charges during speed build-ups. In the interest of rapid evaluation of results, simple methods of data reduction were used. A case history is presented where in the pulse technique predicted flutter by extrapolating decay rates obtained at subcritical speeds; in addition, a case is presented where no valid extrapolation could be made.
NASA Technical Reports Server (NTRS)
Fowler, Samuel B.
2000-01-01
Flutter analysis performed in support of the X33 Advanced Technology Demonstrator is described. Analysis was conducted over a range of flow regimes using several different analysis codes. The finite element and aerodynamic models used in the analysis have undergone several years of development and refinement resulting in a high degree of model detail. The flutter analysis focuses on the area of three critical points within the vehicle's design trajectory at which full sets of external loads have previously been developed. A comparison between several different aerodynamic models is also made for the selected trajectory points.
Overview of the Aeroelastic Prediction Workshop
NASA Technical Reports Server (NTRS)
Heeg, Jennifer; Chwalowski, Pawel; Schuster, David M.; Dalenbring, Mats
2013-01-01
The AIAA Aeroelastic Prediction Workshop (AePW) was held in April, 2012, bringing together communities of aeroelasticians and computational fluid dynamicists. The objective in conducting this workshop on aeroelastic prediction was to assess state-of-the-art computational aeroelasticity methods as practical tools for the prediction of static and dynamic aeroelastic phenomena. No comprehensive aeroelastic benchmarking validation standard currently exists, greatly hindering validation and state-of-the-art assessment objectives. The workshop was a step towards assessing the state of the art in computational aeroelasticity. This was an opportunity to discuss and evaluate the effectiveness of existing computer codes and modeling techniques for unsteady flow, and to identify computational and experimental areas needing additional research and development. Three configurations served as the basis for the workshop, providing different levels of geometric and flow field complexity. All cases considered involved supercritical airfoils at transonic conditions. The flow fields contained oscillating shocks and in some cases, regions of separation. The computational tools principally employed Reynolds-Averaged Navier Stokes solutions. The successes and failures of the computations and the experiments are examined in this paper.
Subsonic/transonic stall flutter investigation of a rotating rig
NASA Technical Reports Server (NTRS)
Jutras, R. R.; Fost, R. B.; Chi, R. M.; Beacher, B. F.
1981-01-01
Stall flutter is investigated by obtaining detailed quantitative steady and aerodynamic and aeromechanical measurements in a typical fan rotor. The experimental investigation is made with a 31.3 percent scale model of the Quiet Engine Program Fan C rotor system. Both subsonic/transonic (torsional mode) flutter and supersonic (flexural) flutter are investigated. Extensive steady and unsteady data on the blade deformations and aerodynamic properties surrounding the rotor are acquired while operating in both the steady and flutter modes. Analysis of this data shows that while there may be more than one traveling wave present during flutter, they are all forward traveling waves.
Panel flutter in a low-density atmosphere
NASA Technical Reports Server (NTRS)
Resende, Hugo B.
1991-01-01
The problem of panel flutter is considered in the case of a free-molecule flow, the effect of the aerodynamic shear stress being incorporated into the structural model through distributed longitudinal and bending moment loads. This kind of analysis is relevant in the case of hypersonic flight vehicles, like the NASP, especially because in these conditions the shear stress can be expected to be considerably larger that the pressure at a given point. The aerodynamic loading is derived assuming a quasi-steady approximation. Two important parameters are given by alpha-m, the 'momentum accommodation coefficient', and Theta, the temperature ratio between the panel temperature and the temperature of the undisturbed flow. For high enough values of Theta the variation of the linear flutter parameter with alpha-m is close to linear. Comparison with continuum theory, given by linear piston theory, depends on the values of Theta and alpha-m considered. Finally, it is shown that buckling with respect to a uniform distributed longitudinal load is stabilizing.
Supersonic axial-flow fan flutter
NASA Technical Reports Server (NTRS)
Ramsey, John K.
1988-01-01
Lane's (1957) analytical formulation of the unsteady pressure distribution on an oscillating two-dimensional flat plate cascade in supersonic axial flow has been developed into a computer code. This unsteady aerodynamic code has shown good agreement with other published data. This code has also been incorporated into an existing aeroelastic code to analyze the NASA Lewis supersonic through-flow fan design.
NASA Technical Reports Server (NTRS)
Roskam, J.; Lan, C.; Mehrotra, S.
1972-01-01
The computer program used to determine the rigid and elastic stability derivatives presented in the summary report is listed in this appendix along with instructions for its use, sample input data and answers. This program represents the airplane at subsonic and supersonic speeds as (a) thin surface(s) (without dihedral) composed of discrete panels of constant pressure according to the method of Woodward for the aerodynamic effects and slender beam(s) for the structural effects. Given a set of input data, the computer program calculates an aerodynamic influence coefficient matrix and a structural influence coefficient matrix.
Ground Vibration and Flight Flutter Tests of the Single-seat F-16XL Aircraft with a Modified Wing
NASA Technical Reports Server (NTRS)
Voracek, David F.
1993-01-01
The NASA single-seat F-16XL aircraft was modified by the addition of a glove to the left wing. Vibration tests were conducted on the ground to assess the changes to the aircraft caused by the glove. Flight Luther testing was conducted on the aircraft with the glove installed to ensure that the flight envelope was free of aeroelastic or aeroservoelastic instabilities. The ground vibration tests showed that above 20 Hz, several modes that involved the control surfaces were significantly changed. Flight test data showed that modal damping levels and trends were satisfactory where obtainable. The data presented in this report include estimated modal parameters from the ground vibration and flight flutter test.
NASA Astrophysics Data System (ADS)
Saitoh, Kenichi; Tamayama, Masato; Kikuchi, Takao; Machida, Shigeru; Nakamichi, Jiro
This paper reports a wind-tunnel experiment and analysis that have been conducted under the National Experimental Airplane for Supersonic Transports (NEXST-1) project of JAXA. In order to perform the flight experiment, the design of the vehicle was examined from the stand point of aeroelasticity. The aileron buzz as well as flutter was of much concern for its aileron system on the main wing. Therefore, both wind-tunnel test and analysis were carried out by using a semi-span model with fuselage. Although the buzz was not observed in the test, damping responses of the aileron rotation mode were obtained. Critical damping was observed in supersonic flow, that meant a buzz could occur in ``region C'' of Lambourne's classification. Linear unsteady aerodynamic analysis is applicable to this type of buzz and the characteristics of the buzz of the model is discussed.
Uncertainty models and associated trade-offs for wing/store flutter suppression
NASA Astrophysics Data System (ADS)
Gade, Prasad V. N.; Inman, Daniel J.
1998-04-01
An active decoupler pylon approach for wing/store flutter suppression is proposed which involves the use of a piezoceramic wafer strut as an actuator for isolating wing torsion modes from store pitch inertia effects. A two degree-of-freedom typical section of an airfoil is used to represent the structural model of the wing, while the circulatory incompressible aerodynamic loads are modeled using Jones' approximation to the Theodorsen function. The analytical model developed neglects store aerodynamics, aileron degree-of-freedom and other flutter critical flexible and rigid body moves. This paper presents some typical perturbation models used to represent such uncertainties and compares their robust stability margins obtained using controllers designed with Loop Transfer Recovery and H(infinity ) control techniques. Singular value Bode plots are used to analyze the robust stability and nominal performance characteristics.
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 Phenomena of Flight Vehicles in Transonic Region
NASA Astrophysics Data System (ADS)
Lee, In; Kim, Jong-Yun; Kim, Kyung-Seok; Lim, In-Gyu
Flight vehicles experience aeroelastic problems due to the interaction between structures and aerodynamic forces. Aeroelastic instability is usually a critical problem in transonic and lower supersonic regions. In present study, the aeroelastic analyses of several flight vehicles have been performed using the coupled techniques of computational fluid dynamics (CFD) and computational structural dynamics (CSD). The aeroelastic characteristics based on several aircraft models are investigated using the developed aeroelastic analysis system. On the other hand, structural nonlinearities always exist in flight vehicles. Structural nonlinearities such as freeplay and large deformation effects are considered in the present aeroelastic analysis system. Finally, aeroelastic characteristics of several flight vehicles will be explained considering both aerodynamic and structural nonlinearities.
Design procedures for flutter-free surface panels
NASA Technical Reports Server (NTRS)
Laurenson, R. M.; Mcpherson, J. I.
1977-01-01
An approach for the design of lightweight external surface panel configurations to preclude panel flutter was developed. Design procedures were developed for flat orthotropic panels under the interacting influence of parameters such as support flexibility, inplane loads, pressure differential, and flow angularity. The basic relationships required to define these design procedures were based on theoretical panel flutter analyses. Where possible, the design procedures were verified through comparison with available experimental panel flutter data.
Aeroelasticity and structural optimization of composite helicopter rotor blades with swept tips
NASA Technical Reports Server (NTRS)
Yuan, K. A.; Friedmann, P. P.
1995-01-01
This report describes the development of an aeroelastic analysis capability for composite helicopter rotor blades with straight and swept tips, and its application to the simulation of helicopter vibration reduction through structural optimization. A new aeroelastic model is developed in this study which is suitable for composite rotor blades with swept tips in hover and in forward flight. The hingeless blade is modeled by beam type finite elements. A single finite element is used to model the swept tip. Arbitrary cross-sectional shape, generally anisotropic material behavior, transverse shears and out-of-plane warping are included in the blade model. The nonlinear equations of motion, derived using Hamilton's principle, are based on a moderate deflection theory. Composite blade cross-sectbnal properties are calculated by a separate linear, two-dimensional cross section analysis. The aerodynamic loads are obtained from quasi-steady, incompressible aerodynamics, based on an implicit formulation. The trim and steady state blade aeroelastic response are solved in a fully coupled manner. In forward flight, where the blade equations of motion are periodic, the coupled trim-aeroelastic response solution is obtained from the harmonic balance method. Subsequently, the periodic system is linearized about the steady state response, and its stability is determined from Floquet theory.
Role of HPC in Advancing Computational Aeroelasticity
NASA Technical Reports Server (NTRS)
Guruswamy, Guru P.
2004-01-01
On behalf of the High Performance Computing and Modernization Program (HPCMP) and NASA Advanced Supercomputing Division (NAS) a study is conducted to assess the role of supercomputers on computational aeroelasticity of aerospace vehicles. The study is mostly based on the responses to a web based questionnaire that was designed to capture the nuances of high performance computational aeroelasticity, particularly on parallel computers. A procedure is presented to assign a fidelity-complexity index to each application. Case studies based on major applications using HPCMP resources are presented.
Method of performing computational aeroelastic analyses
NASA Technical Reports Server (NTRS)
Silva, Walter A. (Inventor)
2011-01-01
Computational aeroelastic analyses typically use a mathematical model for the structural modes of a flexible structure and a nonlinear aerodynamic model that can generate a plurality of unsteady aerodynamic responses based on the structural modes for conditions defining an aerodynamic condition of the flexible structure. In the present invention, a linear state-space model is generated using a single execution of the nonlinear aerodynamic model for all of the structural modes where a family of orthogonal functions is used as the inputs. Then, static and dynamic aeroelastic solutions are generated using computational interaction between the mathematical model and the linear state-space model for a plurality of periodic points in time.
Renaissance of Aeroelasticity and Its Future
NASA Technical Reports Server (NTRS)
Friedmann, Peretz P.
1999-01-01
The primary objective of this paper is to demonstrate that the field of aeroelasticity continues to play a critical role in the design of modern aerospace vehicles, and several important problems are still far from being well understood. Furthermore, the emergence of new technologies, such as the use of adaptive materials (sometimes denoted as smart structures technology), providing new actuator and sensor capabilities, has invigorated aeroelasticity, and generated a host of new and challenging research topics that can have a major impact on the design of a new generation of aerospace vehicles.
Aeroelastic tailoring for oblique wing lateral trim
NASA Technical Reports Server (NTRS)
Bohlmann, Jonathan D.; Weisshaar, Terrence A.; Eckstrom, Clinton V.
1988-01-01
Composite material aeroelastic tailoring is presently explored as a means for the correction of the roll trim imbalance of oblique-wing aircraft configurations. The concept is demonstrated through the analysis of a realistic oblique wing by a static aeroelastic computational procedure encompassing the full potential transonic aerodynamic code FLO22 and a Ritz structural plate program that models the stiffness due to symmetrical-but-unbalanced composite wing skins. Results indicate that asymetric composite tailoring reduces the aileron deflection needed for roll equilibrium, and reduces control surface hinge moment and drag. Wing skin stresses are, however, very high.
Aeroelastic character of a National Aerospace Plane demonstrator concept
NASA Technical Reports Server (NTRS)
Spain, Charles V.; Zeiler, Thomas A.; Gibbons, Michael D.; Soistmann, David L.; Pozefsky, Peter; Dejesus, Rafael O.; Brannon, Cyprian P.
1993-01-01
The paper provides an analytical assessment of the flutter character of an unclassified National Aerospace Plane configuration known as the demonstrator. Linear subsonic, supersonic, and hypersonic analysis indicate that the vehicle is prone to body-freedom flutter resulting from the decrease in vibration frequency of the all-moveable wing at high flight dynamic pressures. As the wing-pivot frequency decreases, it couples with the vehicle short-period mode resulting in dynamic instability. A similar instability sometimes occurs when the pivot mode couples with the fuselage-bending mode. Also assessed, for supersonic flight conditions, are configuration variations that include relocation of the wing further aft on the lifting-body fuselage, and the addition of body flaps to the rear of the vehicle. These changes are destabilizing because they result in severe wing-pivot/fuselage-bending instabilities at dynamic pressures lower than the instabilities indicated for the original demonstrator. Finally, a two-point wing support and actuation system concept is proposed for the National Aerospace Plane, which if developed may (according to cursory analysis) enhance overall stability.
A Presentation on Robust Flutter Margin Analysis and a Flutterometer
NASA Technical Reports Server (NTRS)
Lind, Rick C.
1997-01-01
This paper documents an invited presentation given to The Boeing Company, Seattle, Washington, on September 9, 1997. The audience consisted of structural dynamic and flight test engineers from the Boeing Commercial Airplane Group who were interested in discussing research which may be applied to future flight flutter test programs. A method to compute robust flutter margins is described which is a significant departure from traditional methods. This method uses the structured singular value, mu, to compute a flutter margin which directly accounts for modeling errors such that a worst-case flutter margin is computed with respect to those errors. This method may be applied in several ways. A post-flight application uses data sets from multiple test points to compute worst-case flutter margins and a worst-case flight envelope. An on-line implementation computes flutter margins at each test point to track the flutter margins during a flight test. This on-line implementation is the basis for a flutterometer flight test tool that displays the distance to flutter at a given test point. Such a tool was not previously possible using traditional flutter flight test analysis methods. The F/A-18 System Research Aircraft was used to demonstrate these applications using flight data recorded from test points throughout the flight envelope.
SCAR arrow-wing active flutter suppression system
NASA Technical Reports Server (NTRS)
Gordon, C. K.; Visor, O. E.
1977-01-01
The potential performance and direct operating cost benefits of an active flutter suppression system (FSS) for the NASA arrow-wing supersonic cruise configuration were determined. A FSS designed to increase the flutter speed of the baseline airplane 20 percent. A comparison was made of the performance and direct operating cost between the FSS equipped aircraft and a previously defined configuration with structural modifications to provide the same flutter speed. Control system synthesis and evaluation indicated that a FSS could provide the increase in flutter speed without degrading airplane reliability, safety, handling qualities, or ride quality, and without increasing repeated loads or hydraulic and electrical power capacity requirements.
Flutter Analysis of Annular Cascades in Counter Rotation
NASA Astrophysics Data System (ADS)
Nishino, Ryohei; Namba, Masanobu
The paper studies the effect of neighboring blade rows on flutter characteristics of cascading blades. For this purpose the computation program to calculate the unsteady blade loading based on the unsteady lifting surface theory for contra-rotating annular cascades was formulated and coded. Then a computation program to solve the coupled bending-torsion flutter equation for the contra-rotating annular cascades was also developed. Some results of the flutter analysis are presented. The presence of the neighboring blade row gives rise to significant change in the critical flutter condition when the main acoustic duct mode is of cut-on state.
Flutter spectral measurements using stationary pressure transducers
NASA Technical Reports Server (NTRS)
Kurkov, A. P.
1980-01-01
Engine-order sampling was used to eliminate the integral harmonics from the flutter spectra corresponding to a case-mounted static pressure transducer. Using the optical displacement data, it was demonstrated that the blade-order sampling of pressure data may yield erroneous results due to the interference caused by blade vibration. Two methods are presented which effectively eliminate this interference yielding the blade-pressure-difference spectra. The phase difference between the differential-pressure and the displacement spectra was evaluated.
Model mount system for testing flutter
NASA Technical Reports Server (NTRS)
Farmer, M. G. (Inventor)
1984-01-01
A wind tunnel model mount system is disclosed for effectively and accurately determining the effects of attack and airstream velocity on a model airfoil or aircraft. The model mount system includes a rigid model attached to a splitter plate which is supported away from the wind tunnel wall several of flexible rods. Conventional instrumentation is employed to effect model rotation through a turntable and to record model flutter data as a function of the angle of attack versus dynamic pressure.
A historical overview of flight flutter testing
NASA Technical Reports Server (NTRS)
Kehoe, Michael W.
1995-01-01
This paper reviews the test techniques developed over the last several decades for flight flutter testing of aircraft. Structural excitation systems, instrumentation systems, digital data preprocessing, and parameter identification algorithms (for frequency and damping estimates from the response data) are described. Practical experiences and example test programs illustrate the combined, integrated effectiveness of the various approaches used. Finally, comments regarding the direction of future developments and needs are presented.
Supersonic Chordwise Bending Flutter in Cascades
1975-05-31
such a flutter boundary can be made by utilizing the trend lines predicted from a supersonic analysis based on supersonic cascade theory (Appendix I...bonding agent was injected via hypodermic needles after the blade tabs were properly inserted, The integrity and repeatability of the mounting of the indi...in conjunction with NASTRAN predictions and supersonic cascade aerodynamic computa- tions. Comparisons between theory and experiment are discussed. DD
YF-16 flight flutter test procedures
NASA Technical Reports Server (NTRS)
Brignac, W. J.; Ness, H. B.; Johnson, M. K.; Smith, L. M.
1976-01-01
The Random Decrement technique (Randomdec) was incorporated in procedures for flight testing of the YF-16 lightweight fighter prototype. Damping values obtained substantiate the adequacy of the flutter margin of safety. To confirm the structural modes which were being excited, a spectral analysis of each channel was performed using the AFFTC time/data 1923/50 time series analyzer. Inflight test procedure included the careful monitoring of strip charts, three axis pulses, rolls, and pullups.
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.
An Aeroelastic Perspective of Floating Offshore Wind Turbine Wake Formation and Instability
NASA Astrophysics Data System (ADS)
Rodriguez, Steven N.; Jaworski, Justin W.
2015-11-01
The wake formation and wake stability of floating offshore wind turbines are investigated from an aeroelastic perspective. The aeroelastic model is composed of the Sebastian-Lackner free-vortex wake aerodynamic model coupled to the nonlinear Hodges-Dowell beam equations, which are extended to include the effects of blade profile asymmetry, higher-order torsional effects, and kinetic energy components associated with periodic rigid-body motions of floating platforms. Rigid-body platform motions are also assigned to the aerodynamic model as varying inflow conditions to emulate operational rotor-wake interactions. Careful attention is given to the wake formation within operational states where the ratio of inflow velocity to induced velocity is over 50%. These states are most susceptible to aerodynamic instabilities, and provide a range of states about which a wake stability analysis can be performed. In addition, the stability analysis used for the numerical framework is implemented into a standalone free-vortex wake aerodynamic model. Both aeroelastic and standalone aerodynamic results are compared to evaluate the level of impact that flexible blades have on the wake formation and wake stability.
NASA Technical Reports Server (NTRS)
Bielawa, Richard L.
1988-01-01
In response to a systematic methodology assessment program directed to the aeroelastic stability of hingeless helicopter rotor blades, improved basic aeroelastic reformulations and new formulations relating to structural sweep were achieved. Correlational results are presented showing the substantially improved performance of the G400 aeroelastic analysis incorporating these new formulations. The formulations pertain partly to sundry solutions to classic problem areas, relating to dynamic inflow with vortex-ring state operation and basic blade kinematics, but mostly to improved physical modeling of elastic axis offset (structural sweep) in the presence of nonlinear structural twist. Specific issues addressed are an alternate modeling of the delta EI torsional excitation due to compound bending using a force integration approach, and the detailed kinematic representation of an elastically deflected point mass of a beam with both structural sweep and nonlinear twist.
System and Method for Dynamic Aeroelastic Control
NASA Technical Reports Server (NTRS)
Suh, Peter M. (Inventor)
2015-01-01
The present invention proposes a hardware and software architecture for dynamic modal structural monitoring that uses a robust modal filter to monitor a potentially very large-scale array of sensors in real time, and tolerant of asymmetric sensor noise and sensor failures, to achieve aircraft performance optimization such as minimizing aircraft flutter, drag and maximizing fuel efficiency.
NASA Technical Reports Server (NTRS)
Gardner, J. E.; Dixon, S. C.
1984-01-01
Research was done in the following areas: development and validation of solution algorithms, modeling techniques, integrated finite elements for flow-thermal-structural analysis and design, optimization of aircraft and spacecraft for the best performance, reduction of loads and increase in the dynamic structural stability of flexible airframes by the use of active control, methods for predicting steady and unsteady aerodynamic loads and aeroelastic characteristics of flight vehicles with emphasis on the transonic range, and methods for predicting and reducing helicoper vibrations.
Patient-induced complications of a Heimlich flutter valve.
Crocker, H L; Ruffin, R E
1998-03-01
Heimlich flutter valves have gained widespread acceptance in the treatment of pneumothorax. However, some features of their design may predispose them to inadvertent misuse. A case of tension pneumothorax is described which resulted from the insertion of a drinking straw into the Heimlich flutter valve assembly.
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.
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.
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.
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.
Experimental transonic flutter characteristics of supersonic cruise configurations
NASA Technical Reports Server (NTRS)
Durham, Michael H.; Cole, Stanley R.; Cazier, F. W., Jr.; Keller, Donald F.; Parker, Ellen C.; Wilkie, W. Keats
1990-01-01
The flutter characteristics of a generic arrow-wing supersonic transport configuration are studied. The wing configuration has a 3 percent biconvex airfoil and a leading-edge sweep of 73 deg out to a cranked tip with a 60 deg leading-edge sweep. The ground vibration tests and flutter test procedure are described. The effects of flutter on engine nacelles, fuel loading, wing-mounted vertical fin, wing angle-of-attack, and wing tip mass and stiffness distributions are analyzed. The data reveal that engine nacelles reduce the transonic flutter dynamic pressure by 25-30 percent; fuel loadings decrease dynamic pressures by 25 percent; 4-6 deg wing angles-of-attack cause steep transonic boundaries; and 5-10 percent changes in flutter dynamic pressures are the result of the wing-mounted vertical fin and wing-tip mass and stiffness distributions.
Unsteady aerodynamics of fluttering and tumbling plates
NASA Astrophysics Data System (ADS)
Andersen, A.; Pesavento, U.; Wang, Z. Jane
2005-10-01
We investigate the aerodynamics of freely falling plates in a quasi-two-dimensional flow at Reynolds number of 10(3) , which is typical for a leaf or business card falling in air. We quantify the trajectories experimentally using high-speed digital video at sufficient resolution to determine the instantaneous plate accelerations and thus to deduce the instantaneous fluid forces. We compare the measurements with direct numerical solutions of the two-dimensional Navier Stokes equation. Using inviscid theory as a guide, we decompose the fluid forces into contributions due to acceleration, translation, and rotation of the plate. For both fluttering and tumbling we find that the fluid circulation is dominated by a rotational term proportional to the angular velocity of the plate, as opposed to the translational velocity for a glider with fixed angle of attack. We find that the torque on a freely falling plate is small, i.e. the torque is one to two orders of magnitude smaller than the torque on a glider with fixed angle of attack. Based on these results we revise the existing ODE models of freely falling plates. We get access to different kinds of dynamics by exploring the phase diagram spanned by the Reynolds number, the dimensionless moment of inertia, and the thickness-to-width ratio. In agreement with previous experiments, we find fluttering, tumbling, and apparently chaotic motion. We further investigate the dependence on initial conditions and find brief transients followed by periodic fluttering described by simple harmonics and tumbling with a pronounced period-two structure. Near the cusp-like turning points, the plates elevate, a feature which would be absent if the lift depended on the translational velocity alone.
Global Nonlinear Analysis of Piezoelectric Energy Harvesting from Ambient and Aeroelastic Vibrations
NASA Astrophysics Data System (ADS)
Abdelkefi, Abdessattar
effects of the linear spring coefficients and electrical load resistance on the flutter speed. Then, the normal form of the Hopf bifurcation ( utter) is derived to characterize the type of instability and determine the effects of the aerodynamic nonlinearities and the nonlinear coefficients of the springs on the system's stability near the bifurcation. This is useful to characterize the effects of different parameters on the system's output and ensure that subcritical or "catastrophic" bifurcation does not take place. Both linear and nonlinear analyses are then used to design and enhance the performance of these harvesters. In the last part, the concept of energy harvesting from vortex-induced vibrations of a circular cylinder is investigated. The power levels that can be generated from these vibrations and the variations of these levels with the freestream velocity are determined. A mathematical model that accounts for the coupled lift force, cylinder motion and generated voltage is presented. Linear analysis of the electromechanical model is performed to determine the effects of the electrical load resistance on the natural frequency of the rigid cylinder and the onset of the synchronization region. The impacts of the nonlinearities on the cylinder's response and energy harvesting are then investigated.
A Taguchi study of the aeroelastic tailoring design process
NASA Technical Reports Server (NTRS)
Bohlmann, Jonathan D.; Scott, Robert C.
1991-01-01
A Taguchi study was performed to determine the important players in the aeroelastic tailoring design process and to find the best composition of the optimization's objective function. The Wing Aeroelastic Synthesis Procedure (TSO) was used to ascertain the effects that factors such as composite laminate constraints, roll effectiveness constraints, and built-in wing twist and camber have on the optimum, aeroelastically tailored wing skin design. The results show the Taguchi method to be a viable engineering tool for computational inquiries, and provide some valuable lessons about the practice of aeroelastic tailoring.
An analytical study of effects of aeroelasticity on control effectiveness
NASA Technical Reports Server (NTRS)
Mehrotra, S. C.
1975-01-01
Structural influence coefficients were calculated for various wing planforms using the KU Aeroelastic and NASTRAN programs. The resulting matrices are compared with experimental results. Conclusions are given.
On fluttering modes for aircraft wing model in subsonic air flow.
Shubov, Marianna A
2014-12-08
The paper deals with unstable aeroelastic modes for aircraft wing model in subsonic, incompressible, inviscid air flow. In recent author's papers asymptotic, spectral and stability analysis of the model has been carried out. The model is governed by a system of two coupled integrodifferential equations and a two-parameter family of boundary conditions modelling action of self-straining actuators. The Laplace transform of the solution is given in terms of the 'generalized resolvent operator', which is a meromorphic operator-valued function of the spectral parameter λ, whose poles are called the aeroelastic modes. The residues at these poles are constructed from the corresponding mode shapes. The spectral characteristics of the model are asymptotically close to the ones of a simpler system, which is called the reduced model. For the reduced model, the following result is shown: for each value of subsonic speed, there exists a radius such that all aeroelastic modes located outside the circle of this radius centred at zero are stable. Unstable modes, whose number is always finite, can occur only inside this 'circle of instability'. Explicit estimate of the 'instability radius' in terms of model parameters is given.
On fluttering modes for aircraft wing model in subsonic air flow
Shubov, Marianna A.
2014-01-01
The paper deals with unstable aeroelastic modes for aircraft wing model in subsonic, incompressible, inviscid air flow. In recent author’s papers asymptotic, spectral and stability analysis of the model has been carried out. The model is governed by a system of two coupled integrodifferential equations and a two-parameter family of boundary conditions modelling action of self-straining actuators. The Laplace transform of the solution is given in terms of the ‘generalized resolvent operator’, which is a meromorphic operator-valued function of the spectral parameter λ, whose poles are called the aeroelastic modes. The residues at these poles are constructed from the corresponding mode shapes. The spectral characteristics of the model are asymptotically close to the ones of a simpler system, which is called the reduced model. For the reduced model, the following result is shown: for each value of subsonic speed, there exists a radius such that all aeroelastic modes located outside the circle of this radius centred at zero are stable. Unstable modes, whose number is always finite, can occur only inside this ‘circle of instability’. Explicit estimate of the ‘instability radius’ in terms of model parameters is given. PMID:25484610
Flutter and divergence instability of the multi-cracked cantilever beam-column
NASA Astrophysics Data System (ADS)
Caddemi, S.; Caliò, I.; Cannizzaro, F.
2014-03-01
For conservative systems instability can occur only by divergence and the presence of damage can produce both a reduction of the buckling loads and modification of the corresponding mode shapes, depending on the positions and intensities of the damage distribution. For nonconservative systems instability is found to occur by divergence, flutter, or both, characterised by multiple stable and unstable ranges of the loads whose boundary can be altered by the damage distribution. This paper focuses on the stability behaviour of multi-cracked cantilever Euler beam-column subjected to conservative or nonconservative axial loads. The exact flutter and divergence critical loads are obtained by means of the exact closed form solution of the multi-cracked beam-column, derived by the authors in a previous paper. The extensive numerical applications, reported in the paper, aimed at evaluating the influence of several damage scenarios for different values of the degree of nonconservativeness. It is shown how the presence of damage can strongly modify the ranges of divergence and flutter critical loads of the corresponding undamaged cantilever column, which has been the subject of several papers starting from the Pflüger paradoxical results.
Aeroelastic optimization of a helicopter rotor using an efficient sensitivity analysis
NASA Technical Reports Server (NTRS)
Lim, Joon W.; Chopra, Inderjit
1990-01-01
To reduce oscillatory hub loads in forward flight, a structural optimization analysis of a hingeless helicopter rotor has been developed and applied. The aeroelastic analysis of the rotor is based on a finite element method in space and time, and linked with automated optimization algorithms. For the optimization analysis two types of structural representation are used: a generic stiffness-distribution and a single-cell thin-walled beam. For the first type, the design variables are nonstructural mass and its placement, chordwise center of gravity offset from the elastic axis, and stiffness. For the second type, width, height and thickness of spar are used as design variables. For the behavior constraints, frequency placement, autorotational inertia and aeroelastic stability of the blade are included. The required sensitivity derivatives are obtained using a direct analytical approach. An optimum oscillatory hub load shows a 25-77 percent reduction for the generic blade, and 30-50 percent reduction for the box-beam.
NASA Technical Reports Server (NTRS)
Friedmann, Peretz P.
1992-01-01
This paper presents a review of the state-of-the-art in the field of structural optimization when applied to vibration reduction of helicopters in forward flight with aeroelastic and multidisciplinary constraints. It emphasizes the application of the modern approach where the optimization is formulated as a mathematical programming problem and the objective function consists of the vibration levels at the hub and behavior constraints are imposed on the blade frequencies, aeroelastic stability margins as well as on a number of additional ingredients which can have a significant effect on the overall performance and flight mechanics of the helicopter. It is shown that the integrated multidisciplinary optimization of rotorcraft offers the potential for substantial improvements which can be achieved by careful preliminary design and analysis without requiring additional hardware such as rotor vibration absorbers or isolation systems.
NASA Technical Reports Server (NTRS)
Friedmann, Peretz P.
1991-01-01
This paper presents a survey of the state-of-the-art in the field of structural optimization when applied to vibration reduction of helicopters in forward flight with aeroelastic and multidisciplinary constraints. It emphasizes the application of the modern approach where the optimization is formulated as a mathematical programming problem, the objective function consists of the vibration levels at the hub, and behavior constraints are imposed on the blade frequencies and aeroelastic stability margins, as well as on a number of additional ingredients that can have a significant effect on the overall performance and flight mechanics of the helicopter. It is shown that the integrated multidisciplinary optimization of rotorcraft offers the potential for substantial improvements, which can be achieved by careful preliminary design and analysis without requiring additional hardware such as rotor vibration absorbers of isolation systems.
Gust response of aeroelastically tailored wind turbines
NASA Astrophysics Data System (ADS)
Scott, S.; Capuzzi, M.; Langston, D.; Bossanyi, E.; McCann, G.; Weaver, PM; Pirrera, A.
2016-09-01
Some interesting challenges arise from the drive to build larger, more durable rotors that produce cheaper energy. The rationale is that, with current wind turbine designs, the power generated is theoretically proportional to the square of blade length. One enabling technology is aeroelastic tailoring that offers enhanced combined energy capture and system durability. The design of two adaptive, aeroelastically tailored blade configurations is considered here. One uses material bend-twist coupling; the other combines both material and geometric coupling. Each structural design meets a predefined coupling distribution, whilst approximately matching the stiffness of an uncoupled baseline blade. A gust analysis shows beneficial flapwise load alleviation for both adaptive blades, with the additional benefits of smoothing variations in electrical power and rotational speed.
The predicted effect of aerodynamic detuning on coupled bending-torsion unstalled supersonic flutter
NASA Technical Reports Server (NTRS)
Hoyniak, D.; Fleeter, S.
1986-01-01
A mathematical model is developed to predict the enhanced coupled bending-torsion unstalled supersonic flutter stability due to alternate circumferential spacing aerodynamic detuning of a turbomachine rotor. The translational and torsional unsteady aerodynamic coefficients are developed in terms of influence coefficients, with the coupled bending-torsion stability analysis developed by considering the coupled equations of motion together with the unsteady aerodynamic loading. The effect of this aerodynamic detuning on coupled bending-torsion unstalled supersonic flutter as well as the verification of the modeling are then demonstrated by considering an unstable 12 bladed rotor, with Verdon's uniformly spaced Cascade B flow geometry as a baseline. However, with the elastic axis and center of gravity at 60 percent of the chord, this type of aerodynamic detuning has a minimal effect on stability. For both uniform and nonuniform circumferentially space rotors, a single degree of freedom torsion mode analysis was shown to be appropriate for values of the bending-torsion natural frequency ratio lower than 0.6 and higher 1.2. When the elastic axis and center of gravity are not coincident, the effect of detuning on cascade stability was found to be very sensitive to the location of the center of gravity with respect to the elastic axis. In addition, it was determined that when the center of gravity was forward of an elastic axis located at midchord, a single degree of freedom torsion model did not accurately predict cascade stability.
Experimental Classical Flutter Reesults of a Composite Advanced Turboprop Model
NASA Technical Reports Server (NTRS)
Mehmed, O.; Kaza, K. R. V.
1986-01-01
Experimental results are presented that show the effects of blade pitch angle and number of blades on classical flutter of a composite advanced turboprop (propfan) model. An increase in the number of blades on the rotor or the blade pitch angle is destablizing which shows an aerodynamic coupling or cascade effect between blades. The flutter came in suddenly and all blades vibrated at the same frequency but at different amplitudes and with a common predominant phase angle between consecutive blades. This further indicates aerodynamic coupling between blades. The flutter frequency was between the first two blade normal modes, signifying an aerodynamic coupling between the normal modes. Flutter was observed at all blade pitch angles from small to large angles-of-attack of the blades. A strong blade response occurred, for four blades at the two-per-revolution (2P) frequency, when the rotor speed was near the crossing of the flutter mode frequency and the 2P order line. This is because the damping is low near the flutter condition and the interblade phase angle of the flutter mode and the 2P response are the same.
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.
Overview of the Aeroelastic Prediction Workshop
NASA Technical Reports Server (NTRS)
Heeg, Jennifer; Chwalowski, Pawel; Florance, Jennifer P.; Wieseman, Carol D.; Schuster, David M.; Perry, Raleigh B.
2013-01-01
The Aeroelastic Prediction Workshop brought together an international community of computational fluid dynamicists as a step in defining the state of the art in computational aeroelasticity. This workshop's technical focus was prediction of unsteady pressure distributions resulting from forced motion, benchmarking the results first using unforced system data. The most challenging aspects of the physics were identified as capturing oscillatory shock behavior, dynamic shock-induced separated flow and tunnel wall boundary layer influences. The majority of the participants used unsteady Reynolds-averaged Navier Stokes codes. These codes were exercised at transonic Mach numbers for three configurations and comparisons were made with existing experimental data. Substantial variations were observed among the computational solutions as well as differences relative to the experimental data. Contributing issues to these differences include wall effects and wall modeling, non-standardized convergence criteria, inclusion of static aeroelastic deflection, methodology for oscillatory solutions, post-processing methods. Contributing issues pertaining principally to the experimental data sets include the position of the model relative to the tunnel wall, splitter plate size, wind tunnel expansion slot configuration, spacing and location of pressure instrumentation, and data processing methods.
Flutter estimation of S-520 sounding rocket fins
NASA Astrophysics Data System (ADS)
Honda, Masahisa; Onoda, Junjiro; Onojima, Noboru; Nakada, Atsushi; Isogai, Koji
The flutter margin of the fins, which has been improved for S-520 sounding rocket of ISAS, is estimated by wind tunnel tests and numerical flutter analysis. Both methods require stiffness distribution of the fin. In this paper, a new approach of fitting of FEM model to the measured matrix of influence coefficient is applied in order to eliminate the errors in the measured stiffness data, which may otherwise make the mathematical model nonpositive definite. The results of the wind tunnel tests and numerical flutter analysis using the above fitted FEM model stiffness distribution are compared and discussed.
Aeroelastic Tailoring with Composites Applied to Forward Swept Wings
1981-11-01
wings a viable configo.-tion option for high perfotmance aircraft. Forward swept wings have an inherent -.endency to encounter a static aeroelastic...configuration option for high performance aircraft. Forward swept wings have an inherent tendency to encounter a static aeroelastic instability ialled divergence...conventional and super- critical airfoils. ....... ..................... 19 12 Static methods for subcritical divergence dynamic pressure projection. (a
NASA Technical Reports Server (NTRS)
Gardner, J. E.
1983-01-01
Accomplishments of the past year and plans for the coming year are highlighted as they relate to five year plans and the objectives of the following technical areas: aerothermal loads; multidisciplinary analysis and optimization; unsteady aerodynamics; and configuration aeroelasticity. Areas of interest include thermal protection system concepts, active control, nonlinear aeroelastic analysis, aircraft aeroelasticity, and rotorcraft aeroelasticity and vibrations.
Reduced-Order Models for the Aeroelastic Analysis of Ares Launch Vehicles
NASA Technical Reports Server (NTRS)
Silva, Walter A.; Vatsa, Veer N.; Biedron, Robert T.
2010-01-01
This document presents the development and application of unsteady aerodynamic, structural dynamic, and aeroelastic reduced-order models (ROMs) for the ascent aeroelastic analysis of the Ares I-X flight test and Ares I crew launch vehicles using the unstructured-grid, aeroelastic FUN3D computational fluid dynamics (CFD) code. The purpose of this work is to perform computationally-efficient aeroelastic response calculations that would be prohibitively expensive via computation of multiple full-order aeroelastic FUN3D solutions. These efficient aeroelastic ROM solutions provide valuable insight regarding the aeroelastic sensitivity of the vehicles to various parameters over a range of dynamic pressures.
APPLE - An aeroelastic analysis system for turbomachines and propfans
NASA Technical Reports Server (NTRS)
Reddy, T. S. R.; Bakhle, Milind A.; Srivastava, R.; Mehmed, Oral
1992-01-01
This paper reviews aeroelastic analysis methods for propulsion elements (advanced propellers, compressors and turbines) being developed and used at NASA Lewis Research Center. These aeroelastic models include both structural and aerodynamic components. The structural models include the typical section model, the beam model with and without disk flexibility, and the finite element blade model with plate bending elements. The aerodynamic models are based on the solution of equations ranging from the two-dimensional linear potential equation for a cascade to the three-dimensional Euler equations for multi-blade configurations. Typical results are presented for each aeroelastic model. Suggestions for further research are indicated. All the available aeroelastic models and analysis methods are being incorporated into a unified computer program named APPLE (Aeroelasticity Program for Propulsion at LEwis).
Unsteady Aerodynamic Validation Experiences From the Aeroelastic Prediction Workshop
NASA Technical Reports Server (NTRS)
Heeg, Jennifer; Chawlowski, Pawel
2014-01-01
The AIAA Aeroelastic Prediction Workshop (AePW) was held in April 2012, bringing together communities of aeroelasticians, computational fluid dynamicists and experimentalists. The extended objective was to assess the state of the art in computational aeroelastic methods as practical tools for the prediction of static and dynamic aeroelastic phenomena. As a step in this process, workshop participants analyzed unsteady aerodynamic and weakly-coupled aeroelastic cases. Forced oscillation and unforced system experiments and computations have been compared for three configurations. This paper emphasizes interpretation of the experimental data, computational results and their comparisons from the perspective of validation of unsteady system predictions. The issues examined in detail are variability introduced by input choices for the computations, post-processing, and static aeroelastic modeling. The final issue addressed is interpreting unsteady information that is present in experimental data that is assumed to be steady, and the resulting consequences on the comparison data sets.
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.
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.
Wing/store flutter with nonlinear pylon stiffness
NASA Technical Reports Server (NTRS)
Desmarais, R. N.; Reed, W. H., III
1980-01-01
Recent wind tunnel tests and analytical studies show that a store mounted on a pylon with 'soft' pitch stiffness provides substantial increase in flutter speed of fighter aircraft and reduces dependency of flutter on mass and inertia of the store. This concept, termed the decoupler pylon, utilizes a low-frequency control system to maintain pitch alignment of the store during maneuvers and changing flight conditions. Under rapidly changing transient loads, however, the alignment control system may allow the store to momentarily bottom against a relatively stiff backup structure in which case the pylon stiffness acts as a hardening nonlinear spring. Such structural nonlinearities are known to affect not only the flutter speed but also the basic behavior of the instability. This paper examines the influence of pylon stiffness nonlinearities on the flutter characteristics of wing-mounted external stores.
Wing/store flutter with nonlinear pylon stiffness
NASA Technical Reports Server (NTRS)
Desmarais, R. N.; Reed, W. H., III
1980-01-01
Recent wind tunnel tests and analytical studies show that a store mounted on a pylon with soft pitch stiffness provides substantial increase in flutter speed of fighter aircraft and reduces dependency of flutter on mass and inertia of the store. This concept, termed the decoupler pylon, utilizes a low frequency control system to maintain pitch alignment of the store during maneuvers and changing flight conditions. Under rapidly changing transient loads, however, the alignment control system may allow the store to momentarily bottom against a relatively stiff backup structure in which case the pylon stiffness acts as a hardening nonlinear spring. Such structural nonlinearities are known to affect not only the flutter speed but also the basic behavior of the instability. The influence of pylon stiffness nonlinearities or the flutter characteristics of wing mounted external stores is examined.