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.
Advanced Models for Aeroelastic Analysis of Propulsion Systems
NASA Technical Reports Server (NTRS)
Keith, Theo G., Jr.; Mahajan, Aparajit
1996-01-01
This report describes an integrated, multidisciplinary simulation capability for aeroelastic analysis and optimization of advanced propulsion systems. This research is intended to improve engine development, acquisition, and maintenance costs. One of the proposed simulations is aeroelasticity of blades, cowls, and struts in an ultra-high bypass fan. These ducted fans are expected to have significant performance, fuel, and noise improvements over existing engines. An interface program was written to use modal information from COBSTAN and NASTRAN blade models in aeroelastic analysis with a single rotation ducted fan aerodynamic code.
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).
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
Proposed Wind Turbine Aeroelasticity Studies Using Helicopter Systems Analysis
NASA Technical Reports Server (NTRS)
Ladkany, Samaan G.
1998-01-01
Advanced systems for the analysis of rotary wing aeroelastic structures (helicopters) are being developed at NASA Ames by the Rotorcraft Aeromechanics Branch, ARA. The research has recently been extended to the study of wind turbines, used for electric power generation Wind turbines play an important role in Europe, Japan & many other countries because they are non polluting & use a renewable source of energy. European countries such as Holland, Norway & France have been the world leaders in the design & manufacture of wind turbines due to their historical experience of several centuries, in building complex wind mill structures, which were used in water pumping, grain grinding & for lumbering. Fossil fuel cost in Japan & in Europe is two to three times higher than in the USA due to very high import taxes. High fuel cost combined with substantial governmental subsidies, allow wind generated power to be competitive with the more traditional sources of power generation. In the USA, the use of wind energy has been limited mainly because power production from wind is twice as expensive as from other traditional sources. Studies conducted at the National Renewable Energy Laboratories (NREL) indicate that the main cost in the production of wind turbines is due to the materials & the labor intensive processes used in the construction of turbine structures. Thus, for the US to assume world leadership in wind power generation, new lightweight & consequently very flexible wind turbines, that could be economically mass produced, would have to be developed [4,5]. This effort, if successful, would result in great benefit to the US & the developing nations that suffer from overpopulation & a very high cost of energy.
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.
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 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.
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.
Centrifugal Compressor Aeroelastic Analysis Code
NASA Technical Reports Server (NTRS)
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.
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
Analysis of the Hessian for aeroelastic optimization
NASA Technical Reports Server (NTRS)
Arian, Eyal
1995-01-01
The symbol of the Hessian for an aeroelastic optimization model problem is analyzed. The flow is modeled by the small-disturbance full potential equation and the structure is modeled by an isotropic (von Karman) plate equation. The cost function consists of both aerodynamic and structural terms. In the new analysis the symbol of the cost function Hessian near the minimum is computed. The result indicates that under some conditions, which are likely fulfilled in most applications, the system is decoupled for the non-smooth components. The result also shows that the structure part in the Hessian is well-conditioned while the aerodynamic part is ill-conditioned. Applications of the result to optimization strategies are discussed.
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 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.
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.
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.
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.
Aeroelastic Analysis of Modern Complex Wings
NASA Technical Reports Server (NTRS)
Kapania, Rakesh K.; Bhardwaj, Manoj K.; Reichenbach, Eric; Guruswamy, Guru P.
1996-01-01
A process is presented by which aeroelastic analysis is performed by using an advanced computational fluid dynamics (CFD) code coupled with an advanced computational structural dynamics (CSD) code. The process is demonstrated on an F/A-18 Stabilator using NASTD (an in-house McDonnell Douglas Aerospace East CFD code) coupled with NASTRAN. The process is also demonstrated on an aeroelastic research wing (ARW-2) using ENSAERO (an in-house NASA Ames Research Center CFD code) coupled with a finite element wing-box structures code. Good results have been obtained for the F/A-18 Stabilator while results for the ARW-2 supercritical wing are still being obtained.
Aeroelastic System Development Using Proper Orthogonal Decomposition and Volterra Theory
NASA Technical Reports Server (NTRS)
Lucia, David J.; Beran, Philip S.; Silva, Walter A.
2003-01-01
This research combines Volterra theory and proper orthogonal decomposition (POD) into a hybrid methodology for reduced-order modeling of aeroelastic systems. The out-come of the method is a set of linear ordinary differential equations (ODEs) describing the modal amplitudes associated with both the structural modes and the POD basis functions for the uid. For this research, the structural modes are sine waves of varying frequency, and the Volterra-POD approach is applied to the fluid dynamics equations. The structural modes are treated as forcing terms which are impulsed as part of the uid model realization. Using this approach, structural and uid operators are coupled into a single aeroelastic operator. This coupling converts a free boundary uid problem into an initial value problem, while preserving the parameter (or parameters) of interest for sensitivity analysis. The approach is applied to an elastic panel in supersonic cross ow. The hybrid Volterra-POD approach provides a low-order uid model in state-space form. The linear uid model is tightly coupled with a nonlinear panel model using an implicit integration scheme. The resulting aeroelastic model provides correct limit-cycle oscillation prediction over a wide range of panel dynamic pressure values. Time integration of the reduced-order aeroelastic model is four orders of magnitude faster than the high-order solution procedure developed for this research using traditional uid and structural solvers.
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.
Rotorcraft Technology for HALE Aeroelastic Analysis
NASA Technical Reports Server (NTRS)
Young, Larry; Johnson, Wayne
2008-01-01
Much of technology needed for analysis of HALE nonlinear aeroelastic problems is available from rotorcraft methodologies. Consequence of similarities in operating environment and aerodynamic surface configuration. Technology available - theory developed, validated by comparison with test data, incorporated into rotorcraft codes. High subsonic to transonic rotor speed, low to moderate Reynolds number. Structural and aerodynamic models for high aspect-ratio wings and propeller blades. Dynamic and aerodynamic interaction of wing/airframe and propellers. Large deflections, arbitrary planform. Steady state flight, maneuvers and response to turbulence. Linearized state space models. This technology has not been extensively applied to HALE configurations. Correlation with measured HALE performance and behavior required before can rely on tools.
Static Aeroelastic Analysis with an Inviscid Cartesian Method
NASA Technical Reports Server (NTRS)
Rodriguez, David L.; Aftosmis, Michael J.; Nemec, Marian; Smith, Stephen C.
2014-01-01
An embedded-boundary, Cartesian-mesh flow solver is coupled with a three degree-of-freedom structural model to perform static, aeroelastic analysis of complex aircraft geometries. The approach solves a nonlinear, aerostructural system of equations using a loosely-coupled strategy. An open-source, 3-D discrete-geometry engine is utilized to deform a triangulated surface geometry according to the shape predicted by the structural model under the computed aerodynamic loads. The deformation scheme is capable of modeling large deflections and is applicable to the design of modern, very-flexible transport wings. The coupling interface is modular so that aerodynamic or structural analysis methods can be easily swapped or enhanced. After verifying the structural model with comparisons to Euler beam theory, two applications of the analysis method are presented as validation. The first is a relatively stiff, transport wing model which was a subject of a recent workshop on aeroelasticity. The second is a very flexible model recently tested in a low speed wind tunnel. Both cases show that the aeroelastic analysis method produces results in excellent agreement with experimental data.
Aeroelastic-Acoustics Simulation of Flight Systems
NASA Technical Reports Server (NTRS)
Gupta, kajal K.; Choi, S.; Ibrahim, A.
2009-01-01
This paper describes the details of a numerical finite element (FE) based analysis procedure and a resulting code for the simulation of the acoustics phenomenon arising from aeroelastic interactions. Both CFD and structural simulations are based on FE discretization employing unstructured grids. The sound pressure level (SPL) on structural surfaces is calculated from the root mean square (RMS) of the unsteady pressure and the acoustic wave frequencies are computed from a fast Fourier transform (FFT) of the unsteady pressure distribution as a function of time. The resulting tool proves to be unique as it is designed to analyze complex practical problems, involving large scale computations, in a routine fashion.
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.
Sensitivity analysis of a wing aeroelastic response
NASA Technical Reports Server (NTRS)
Kapania, Rakesh K.; Eldred, Lloyd B.; Barthelemy, Jean-Francois M.
1991-01-01
A variation of Sobieski's Global Sensitivity Equations (GSE) approach is implemented to obtain the sensitivity of the static aeroelastic response of a three-dimensional wing model. The formulation is quite general and accepts any aerodynamics and structural analysis capability. An interface code is written to convert one analysis's output to the other's input, and visa versa. Local sensitivity derivatives are calculated by either analytic methods or finite difference techniques. A program to combine the local sensitivities, such as the sensitivity of the stiffness matrix or the aerodynamic kernel matrix, into global sensitivity derivatives is developed. The aerodynamic analysis package FAST, using a lifting surface theory, and a structural package, ELAPS, implementing Giles' equivalent plate model are used.
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.
NASA Technical Reports Server (NTRS)
Kukreja, Sunil L.; Brenner, martin J.
2006-01-01
This viewgraph presentation reviews the 1. Motivation for the study 2. Nonlinear Model Form 3. Structure Detection 4. Least Absolute Shrinkage and Selection Operator (LASSO) 5. Objectives 6. Results 7. Assess LASSO as a Structure Detection Tool: Simulated Nonlinear Models 8. Applicability to Complex Systems: F/A-18 Active Aeroelastic Wing Flight Test Data. The authors conclude that 1. this is a novel approach for detecting the structure of highly over-parameterised nonlinear models in situations where other methods may be inadequate 2. that it is a practical significance in the analysis of aircraft dynamics during envelope expansion and could lead to more efficient control strategies and 3. this could allow greater insight into the functionality of various systems dynamics, by providing a quantitative model which is easily interpretable
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
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.
Vibration and aeroelastic analysis of highly flexible HALE aircraft
NASA Astrophysics Data System (ADS)
Chang, Chong-Seok
The highly flexible HALE (High Altitude Long Endurance) aircraft analysis methodology is of interest because early studies indicated that HALE aircraft might have different vibration and aeroelastic characteristics from those of conventional aircraft. Recently the computer code Nonlinear Aeroelastic Trim And Stability of HALE Aircraft (NATASHA) was developed under NASA sponsorship. NATASHA can predict the flight dynamics and aeroelastic behavior for HALE aircraft with a flying wing configuration. Further analysis improvements for NATASHA were required to extend its capability to the ground vibration test (GVT) environment and to both GVT and aeroelastic behavior of HALE aircraft with other configurations. First, the analysis methodology, based on geometrically exact fully intrinsic beam theory, was extended to treat other aircraft cofigurations. Conventional aircraft with flexible fuselage and tail can now be modeled by treating the aircraft as an assembly of beam elements. NATASHA is now applicable to any aircraft cofiguration that can be modeled this way. The intrinsic beam formulation, which is a fundamental structural modeling approach, is now capable of being applying to a structure consisting of multiple beams by relating the virtual displacements and rotations at points where two or more beam elements are connected to each other. Additional aspects are also considered in the analysis such as auxiliary elevator input in the horizontal tail and fuselage aerodynamics. Second, the modeling approach was extended to treat the GVT environment for HALE aircraft, which have highly flexible wings. GVT has its main purpose to provide modal characteristics for model validation. A bungee formulation was developed by the augmented Lagrangian method and coupled to the intrinsic beam formulation for the GVT modeling. After the coupling procedure, the whole formulation cannot be fully intrinsic because the geometric constraint by bungee cords makes the system statically
Propulsion Aeroelastic Analysis Developed for Flutter and Forced Response
NASA Technical Reports Server (NTRS)
Bakhle, Milind A.
2000-01-01
The NASA Glenn Research Center at Lewis Field develops new technologies to increase the fuel efficiency of aircraft engines, improve the safety of engine operation, reduce emissions, and reduce engine noise. With the development of new designs for fans, compressors, and turbines to achieve these goals, the basic aeroelastic requirements are that there should be no flutter (self-excited vibrations) or high resonant blade stresses (due to forced response) in the operating regime. Therefore, an accurate prediction and analysis capability is required to verify the aeroelastic soundness of the designs. Such a three-dimensional viscous propulsion aeroelastic analysis capability has been developed at Glenn with support from the Advanced Subsonic Technology (AST) program. This newly developed aeroelastic analysis capability is based on TURBO, a threedimensional unsteady aerodynamic Reynolds-averaged Navier-Stokes turbomachinery code developed previously under a grant from Glenn. TURBO can model the viscous flow effects that play an important role in certain aeroelastic problems such as flutter with flow separation, flutter at high loading conditions near the stall line (stall flutter), flutter in the presence of shock and boundary-layer interaction, and forced response due to wakes and shock impingement. In aeroelastic analysis, the structural dynamics representation of the blades is based on normal modes. A finite-element analysis code is used to calculate these in-vacuum vibration modes and the associated natural frequencies. In an aeroelastic analysis using the TURBO code, flutter and forced response are modeled as being uncoupled. To calculate if a blade row will flutter, one prescribes the motion of the blade to be a harmonic vibration in a specified in-vacuum normal mode. An aeroelastic analysis preprocessor is used to generate the displacement field required for the analysis. The work done by aerodynamic forces on the vibrating blade during a cycle of vibration is
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 Analysis of Modern Complex Wings Using ENSAERO and NASTRAN
NASA Technical Reports Server (NTRS)
Bhardwaj, Manoj
1995-01-01
A process is presented by which static aeroelastic analysis is performed using Euler flow equations in conjunction with an advanced structural analysis tool, NASTRAN. The process deals with the interfacing of two separate codes in the fields of computational fluid dynamics (CFD) and computational structural dynamics (CSD). The process is demonstrated successfully on an F/A-18 Stabilator (horizontal tail).
Computational aeroelastic analysis of aircraft wings including geometry nonlinearity
NASA Astrophysics Data System (ADS)
Tian, Binyu
The objective of the present study is to show the ability of solving fluid structural interaction problems more realistically by including the geometric nonlinearity of the structure so that the aeroelastic analysis can be extended into the onset of flutter, or in the post flutter regime. A nonlinear Finite Element Analysis software is developed based on second Piola-Kirchhoff stress and Green-Lagrange strain. The second Piola-Kirchhoff stress and Green-Lagrange strain is a pair of energetically conjugated tensors that can accommodate arbitrary large structural deformations and deflection, to study the flutter phenomenon. Since both of these tensors are objective tensors, i.e., the rigid-body motion has no contribution to their components, the movement of the body, including maneuvers and deformation, can be included. The nonlinear Finite Element Analysis software developed in this study is verified with ANSYS, NASTRAN, ABAQUS, and IDEAS for the linear static, nonlinear static, linear dynamic and nonlinear dynamic structural solutions. To solve the flow problems by Euler/Navier equations, the current nonlinear structural software is then embedded into ENSAERO, which is an aeroelastic analysis software package developed at NASA Ames Research Center. The coupling of the two software, both nonlinear in their own field, is achieved by domain decomposition method first proposed by Guruswamy. A procedure has been set for the aeroelastic analysis process. The aeroelastic analysis results have been obtained for fight wing in the transonic regime for various cases. The influence dynamic pressure on flutter has been checked for a range of Mach number. Even though the current analysis matches the general aeroelastic characteristic, the numerical value not match very well with previous studies and needs farther investigations. The flutter aeroelastic analysis results have also been plotted at several time points. The influences of the deforming wing geometry can be well seen
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.
Static aeroelastic analysis of a three-dimensional generic wing
NASA Technical Reports Server (NTRS)
Green, John A.; Lee, IN; Miura, Hirokazu
1990-01-01
A continuation of research on the static aeroelastic analysis of a generic wing configuration is presented. Results of the study of the asymmetric oblique wing model developed by Rockwell International, in conjunction with the NASA Oblique Wing Research Aircraft Program, are reported. The capability to perform static aeroelastic analyses of an oblique wing at arbitrary skew positions is demonstrated by applying the MSC/NASTRAN static analysis scheme modified by the aerodynamic influence coefficient matrix created by the NASA Ames aerodynamic panel codes. The oblique wing is studied at two skew angles, and in particular, the capability to calculate 3-D thickness effects on the aerodynamic properties of the wing is investigated. The ability to model asymmetric wings in both subsonic and supersonic Mach numbers is shown. The aerodynamic influence coefficient matrix computed by the external programs is inserted in MSC/NASTRAN static aeroelasticity analysis run stream to compute the aeroelastic deformation and internal forces. Various aerodynamic coefficients of the oblique wing were computed for two Mach numbers, 0.7 and 1.4, and the angle of attach -5 through 15 deg.
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.
Static aeroelastic analysis of a three-dimensional oblique wing
NASA Technical Reports Server (NTRS)
Lee, I.; Miura, H.; Chargin, M. K.
1990-01-01
A capability to perform static aeroelastic analyses of an oblique wing at arbitrary skew positions was developed based on the framework of the MSC/NASTRAN static aeroelastic analysis. By means of DMAP alterations, a portion of the subsonic static aeroelastic analysis scheme was modified to insert an aerodynamic influence coefficient matrix created externally by the NASA-Ames aerodynamic panel codes. The modified scheme can cover the subsonic as well as the supersonic range for both symmetric and asymmetric configurations. Static aeroelastic responses of the oblique wing are studied at two skew angles and, in particular, the capability to calculate 3D camber effects on the aerodynamic properties of the wing is investigated. Various aerodynamic coefficients of the rigid oblique wing are computed for two Mach numbers, 0.7 and 1.4, and the angle of attack is varied from -5 through 15 deg. Also, the wing flexibility effects on the aerodynamic coefficients and the displacement are examined at a Mach number of 0.7 for a 45-deg swept wing.
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.
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.
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.
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.
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
NASA Technical Reports Server (NTRS)
Kukreja, Sunil L.; Brenner, Martin J.
2006-01-01
This viewgraph presentation reviews the applicability of NARMAX structure detection to aeroelastic systems. In conclusion, the simulation results demonstrate bootstrap approach for structure computation of aircraft structural stiffness provided a high rate of true model selection: 1. T-test and stepwise regression methods had difficulty providing accurate results 2. Work contributes to understanding of the use of structure detection for modelling and identification of aerospace systems. 3. Limitation of model complexity that can be studied with these structure computation techniques 4. Result of the large number of candidate terms, for a given model order, and the data length required to guarantee convergence 5. Another approach to structure computation problem uses a least absolute shrinkage and selection operator (LASSO)
Static Aeroelastic Analysis with an Inviscid Cartesian Method
NASA Technical Reports Server (NTRS)
Rodriguez, David L.; Aftosmis, Michael J.; Nemec, Marian; Smith, Stephen C.
2014-01-01
An embedded-boundary Cartesian-mesh flow solver is coupled with a three degree-offreedom structural model to perform static, aeroelastic analysis of complex aircraft geometries. The approach solves the complete system of aero-structural equations using a modular, loosely-coupled strategy which allows the lower-fidelity structural model to deform the highfidelity CFD model. The approach uses an open-source, 3-D discrete-geometry engine to deform a triangulated surface geometry according to the shape predicted by the structural model under the computed aerodynamic loads. The deformation scheme is capable of modeling large deflections and is applicable to the design of modern, very-flexible transport wings. The interface is modular so that aerodynamic or structural analysis methods can be easily swapped or enhanced. This extended abstract includes a brief description of the architecture, along with some preliminary validation of underlying assumptions and early results on a generic 3D transport model. The final paper will present more concrete cases and validation of the approach. Preliminary results demonstrate convergence of the complete aero-structural system and investigate the accuracy of the approximations used in the formulation of the structural model.
CFD and Aeroelastic Analysis of the MEXICO Wind Turbine
NASA Astrophysics Data System (ADS)
Carrión, M.; Woodgate, M.; Steijl, R.; Barakos, G.; Gómez-Iradi, S.; Munduate, X.
2014-12-01
This paper presents an aerodynamic and aeroelastic analysis of the MEXICO wind turbine, using the compressible HMB solver of Liverpool. The aeroelasticity of the blade, as well as the effect of a low-Mach scheme were studied for the zero-yaw 15m/s wind case and steady- state computations. The wake developed behind the rotor was also extracted and compared with the experimental data, using the compressible solver and a low-Mach scheme. It was found that the loads were not sensitive to the Mach number effects, although the low-Mach scheme improved the wake predictions. The sensitivity of the results to the blade structural properties was also highlighted.
Aeroelastic Analysis of the NASA/ARMY/MIT Active Twist Rotor
NASA Technical Reports Server (NTRS)
Wilkie, W. Keats; Wilbur, Matthew L.; Mirick, Paul H.; Cesnik, Carlos E. S.; Shin, Sangloon
1999-01-01
Aeroelastic modeling procedures used in the design of a piezoelectric controllable twist helicopter rotor wind tunnel model are described. Two aeroelastic analysis methods developed for active twist rotor studies, and used in the design of the model blade, are described in this paper. The first procedure uses a simple flap-torsion dynamic representation of the active twist blade, and is intended for rapid and efficient control law and design optimization studies. The second technique employs a commercially available comprehensive rotor analysis package, and is used for more detailed analytical studies. Analytical predictions of hovering flight twist actuation frequency responses are presented for both techniques. Forward flight fixed system nP vibration suppression capabilities of the model active twist rotor system are also presented. Frequency responses predicted using both analytical procedures agree qualitatively for all design cases considered, with best correlation for cases where uniform blade properties are assumed.
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.
NRT Rotor Structural / Aeroelastic Analysis for the Preliminary Design Review
Ennis, Brandon Lee; Paquette, Joshua A.
2015-10-01
This document describes the initial structural design for the National Rotor Testbed blade as presented during the preliminary design review at Sandia National Laboratories on October 28- 29, 2015. The document summarizes the structural and aeroelastic requirements placed on the NRT rotor for satisfactory deployment at the DOE/SNL SWiFT experimental facility to produce high-quality datasets for wind turbine model validation. The method and result of the NRT blade structural optimization is also presented within this report, along with analysis of its satisfaction of the design requirements.
Aeroelastic Analysis of Aircraft: Wing and Wing/Fuselage Configurations
NASA Technical Reports Server (NTRS)
Chen, H. H.; Chang, K. C.; Tzong, T.; Cebeci, T.
1997-01-01
A previously developed interface method for coupling aerodynamics and structures is used to evaluate the aeroelastic effects for an advanced transport wing at cruise and under-cruise conditions. The calculated results are compared with wind tunnel test data. The capability of the interface method is also investigated for an MD-90 wing/fuselage configuration. In addition, an aircraft trim analysis is described and applied to wing configurations. The accuracy of turbulence models based on the algebraic eddy viscosity formulation of Cebeci and Smith is studied for airfoil flows at low Mach numbers by using methods based on the solutions of the boundary-layer and Navier-Stokes equations.
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.
A Review of Recent Aeroelastic Analysis Methods for Propulsion at NASA Lewis Research Center
NASA Technical Reports Server (NTRS)
Reddy, T. S. R.; Bakhle, Milind A.; Srivastava, R.; Mehmed, Oral; Stefko, George L.
1993-01-01
This report reviews aeroelastic analyses for propulsion components (propfans, compressors and turbines) being developed and used at NASA LeRC. These aeroelastic analyses include both structural and aerodynamic models. The structural models include a typical section, a beam (with and without disk flexibility), and a 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 to the three-dimensional Euler equations for multibladed configurations. Typical calculated results are presented for each aeroelastic model. Suggestions for further research are made. Many of the currently available aeroelastic models and analysis methods are being incorporated in a unified computer program, APPLE (Aeroelasticity Program for Propulsion at LEwis).
NASA Technical Reports Server (NTRS)
Pak, Chan-gi; Lung, Shu
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
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.
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.
A General Interface Method for Aeroelastic Analysis of Aircraft
NASA Technical Reports Server (NTRS)
Tzong, T.; Chen, H. H.; Chang, K. C.; Wu, T.; Cebeci, T.
1996-01-01
The aeroelastic analysis of an aircraft requires an accurate and efficient procedure to couple aerodynamics and structures. The procedure needs an interface method to bridge the gap between the aerodynamic and structural models in order to transform loads and displacements. Such an interface method is described in this report. This interface method transforms loads computed by any aerodynamic code to a structural finite element (FE) model and converts the displacements from the FE model to the aerodynamic model. The approach is based on FE technology in which virtual work is employed to transform the aerodynamic pressures into FE nodal forces. The displacements at the FE nodes are then converted back to aerodynamic grid points on the aircraft surface through the reciprocal theorem in structural engineering. The method allows both high and crude fidelities of both models and does not require an intermediate modeling. In addition, the method performs the conversion of loads and displacements directly between individual aerodynamic grid point and its corresponding structural finite element and, hence, is very efficient for large aircraft models. This report also describes the application of this aero-structure interface method to a simple wing and an MD-90 wing. The results show that the aeroelastic effect is very important. For the simple wing, both linear and nonlinear approaches are used. In the linear approach, the deformation of the structural model is considered small, and the loads from the deformed aerodynamic model are applied to the original geometry of the structure. In the nonlinear approach, the geometry of the structure and its stiffness matrix are updated in every iteration and the increments of loads from the previous iteration are applied to the new structural geometry in order to compute the displacement increments. Additional studies to apply the aero-structure interaction procedure to more complicated geometry will be conducted in the second phase
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.
Maximized Gust Loads of a Closed-Loop, Nonlinear Aeroelastic System Using Nonlinear Systems Theory
NASA Technical Reports Server (NTRS)
Silva, Walter A.
1999-01-01
The problem of computing the maximized gust load for a nonlinear, closed-loop aeroelastic aircraft is discusses. The Volterra theory of nonlinear systems is applied in order to define a linearized system that provides a bounds on the response of the nonlinear system of interest. The method is applied to a simplified model of an Airbus A310.
Sensitivity analysis for aeroacoustic and aeroelastic design of turbomachinery blades
NASA Technical Reports Server (NTRS)
Lorence, Christopher B.; Hall, Kenneth C.
1995-01-01
A new method for computing the effect that small changes in the airfoil shape and cascade geometry have on the aeroacoustic and aeroelastic behavior of turbomachinery cascades is presented. The nonlinear unsteady flow is assumed to be composed of a nonlinear steady flow plus a small perturbation unsteady flow that is harmonic in time. First, the full potential equation is used to describe the behavior of the nonlinear mean (steady) flow through a two-dimensional cascade. The small disturbance unsteady flow through the cascade is described by the linearized Euler equations. Using rapid distortion theory, the unsteady velocity is split into a rotational part that contains the vorticity and an irrotational part described by a scalar potential. The unsteady vorticity transport is described analytically in terms of the drift and stream functions computed from the steady flow. Hence, the solution of the linearized Euler equations may be reduced to a single inhomogeneous equation for the unsteady potential. The steady flow and small disturbance unsteady flow equations are discretized using bilinear quadrilateral isoparametric finite elements. The nonlinear mean flow solution and streamline computational grid are computed simultaneously using Newton iteration. At each step of the Newton iteration, LU decomposition is used to solve the resulting set of linear equations. The unsteady flow problem is linear, and is also solved using LU decomposition. Next, a sensitivity analysis is performed to determine the effect small changes in cascade and airfoil geometry have on the mean and unsteady flow fields. The sensitivity analysis makes use of the nominal steady and unsteady flow LU decompositions so that no additional matrices need to be factored. Hence, the present method is computationally very efficient. To demonstrate how the sensitivity analysis may be used to redesign cascades, a compressor is redesigned for improved aeroelastic stability and two different fan exit guide
Non-Linear System Identification for Aeroelastic Systems with Application to Experimental Data
NASA Technical Reports Server (NTRS)
Kukreja, Sunil L.
2008-01-01
Representation and identification of a non-linear aeroelastic pitch-plunge system as a model of the NARMAX class is considered. A non-linear difference equation describing this aircraft model is derived theoretically and shown to be of the NARMAX form. Identification methods for NARMAX models are applied to aeroelastic dynamics and its properties demonstrated via continuous-time simulations of experimental conditions. Simulation results show that (i) the outputs of the NARMAX model match closely those generated using continuous-time methods and (ii) NARMAX identification methods applied to aeroelastic dynamics provide accurate discrete-time parameter estimates. Application of NARMAX identification to experimental pitch-plunge dynamics data gives a high percent fit for cross-validated data.
Nonlinear System Identification for Aeroelastic Systems with Application to Experimental Data
NASA Technical Reports Server (NTRS)
Kukreja, Sunil L.
2008-01-01
Representation and identification of a nonlinear aeroelastic pitch-plunge system as a model of the Nonlinear AutoRegressive, Moving Average eXogenous (NARMAX) class is considered. A nonlinear difference equation describing this aircraft model is derived theoretically and shown to be of the NARMAX form. Identification methods for NARMAX models are applied to aeroelastic dynamics and its properties demonstrated via continuous-time simulations of experimental conditions. Simulation results show that (1) the outputs of the NARMAX model closely match those generated using continuous-time methods, and (2) NARMAX identification methods applied to aeroelastic dynamics provide accurate discrete-time parameter estimates. Application of NARMAX identification to experimental pitch-plunge dynamics data gives a high percent fit for cross-validated data.
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.
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.
Static aeroelastic analysis and tailoring of missile control fins
NASA Technical Reports Server (NTRS)
Mcintosh, S. C., Jr.; Dillenius, M. F. E.
1989-01-01
A concept for enhancing the design of control fins for supersonic tactical missiles is described. The concept makes use of aeroelastic tailoring to create fin designs (for given planforms) that limit the variations in hinge moments that can occur during maneuvers involving high load factors and high angles of attack. It combines supersonic nonlinear aerodynamic load calculations with finite-element structural modeling, static and dynamic structural analysis, and optimization. The problem definition is illustrated. The fin is at least partly made up of a composite material. The layup is fixed, and the orientations of the material principal axes are allowed to vary; these are the design variables. The objective is the magnitude of the difference between the chordwise location of the center of pressure and its desired location, calculated for a given flight condition. Three types of constraints can be imposed: upper bounds on static displacements for a given set of load conditions, lower bounds on specified natural frequencies, and upper bounds on the critical flutter damping parameter at a given set of flight speeds and altitudes. The idea is to seek designs that reduce variations in hinge moments that would otherwise occur. The block diagram describes the operation of the computer program that accomplishes these tasks. There is an option for a single analysis in addition to the optimization.
Design and Analysis of AN Static Aeroelastic Experiment
NASA Astrophysics Data System (ADS)
Hou, Ying-Yu; Yuan, Kai-Hua; Lv, Ji-Nan; Liu, Zi-Qiang
2016-06-01
Static aeroelastic experiments are very common in the United States and Russia. The objective of static aeroelastic experiments is to investigate deformation and loads of elastic structure in flow field. Generally speaking, prerequisite of this experiment is that the stiffness distribution of structure is known. This paper describes a method for designing experimental models, in the case where the stiffness distribution and boundary condition of a real aircraft are both uncertain. The stiffness distribution form of the structure can be calculated via finite element modeling and simulation calculation and F141 steels and rigid foam are used to make elastic model. In this paper, the design and manufacturing process of static aeroelastic models is presented and a set of experiment model was designed to simulate the stiffness of the designed wings, a set of experiments was designed to check the results. The test results show that the experimental method can effectively complete the design work of elastic model. This paper introduces the whole process of the static aeroelastic experiment, and the experimental results are analyzed. This paper developed a static aeroelasticity experiment technique and established an experiment model targeting at the swept wing of a certain kind of large aspect ratio aircraft.
Sensitivity Analysis of the Static Aeroelastic Response of a Wing
NASA Technical Reports Server (NTRS)
Eldred, Lloyd B.
1993-01-01
A technique to obtain the sensitivity of the static aeroelastic response of a three dimensional wing model is designed and implemented. The formulation is quite general and accepts any aerodynamic and structural analysis capability. A program to combine the discipline level, or local, sensitivities into global sensitivity derivatives is developed. A variety of representations of the wing pressure field are developed and tested to determine the most accurate and efficient scheme for representing the field outside of the aerodynamic code. Chebyshev polynomials are used to globally fit the pressure field. This approach had some difficulties in representing local variations in the field, so a variety of local interpolation polynomial pressure representations are also implemented. These panel based representations use a constant pressure value, a bilinearly interpolated value. or a biquadraticallv interpolated value. The interpolation polynomial approaches do an excellent job of reducing the numerical problems of the global approach for comparable computational effort. Regardless of the pressure representation used. sensitivity and response results with excellent accuracy have been produced for large integrated quantities such as wing tip deflection and trim angle of attack. The sensitivities of such things as individual generalized displacements have been found with fair accuracy. In general, accuracy is found to be proportional to the relative size of the derivatives to the quantity itself.
Aeroservoelastic Model Validation and Test Data Analysis of the F/A-18 Active Aeroelastic Wing
NASA Technical Reports Server (NTRS)
Brenner, Martin J.; Prazenica, Richard J.
2003-01-01
Model validation and flight test data analysis require careful consideration of the effects of uncertainty, noise, and nonlinearity. Uncertainty prevails in the data analysis techniques and results in a composite model uncertainty from unmodeled dynamics, assumptions and mechanics of the estimation procedures, noise, and nonlinearity. A fundamental requirement for reliable and robust model development is an attempt to account for each of these sources of error, in particular, for model validation, robust stability prediction, and flight control system development. This paper is concerned with data processing procedures for uncertainty reduction in model validation for stability estimation and nonlinear identification. F/A-18 Active Aeroelastic Wing (AAW) aircraft data is used to demonstrate signal representation effects on uncertain model development, stability estimation, and nonlinear identification. Data is decomposed using adaptive orthonormal best-basis and wavelet-basis signal decompositions for signal denoising into linear and nonlinear identification algorithms. Nonlinear identification from a wavelet-based Volterra kernel procedure is used to extract nonlinear dynamics from aeroelastic responses, and to assist model development and uncertainty reduction for model validation and stability prediction by removing a class of nonlinearity from the uncertainty.
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.
NASA Astrophysics Data System (ADS)
Chassaing, J.-C.; Lucor, D.; Trégon, J.
2012-01-01
An adaptive stochastic spectral projection method is deployed for the uncertainty quantification in limit-cycle oscillations of an elastically mounted two-dimensional lifting surface in a supersonic flow field. Variabilities in the structural parameters are propagated in the aeroelastic system which accounts for nonlinear restoring force and moment by means of hardening cubic springs. The physical nonlinearities promote sharp and sudden flutter onset for small change of the reduced velocity. In a stochastic context, this behavior translates to steep solution gradients developing in the parametric space. A remedy is to expand the stochastic response of the airfoil on a piecewise generalized polynomial chaos basis. Accurate approximation andaffordable computational costs are obtained using sensitivity-based adaptivity for various types of supersonic stochastic responses depending on the selected values of the Mach number on the bifurcation map. Sensitivity analysis via Sobol' indices shows how the probability density function of the peak pitch amplitude responds to combined uncertainties: e.g. the elastic axis location, torsional stiffness and flap angle. We believe that this work demonstrates the capability and flexibility of the approach for more reliable predictions of realistic aeroelastic systems subject to a moderate number of uncertainties.
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.
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.
Aeroelastic Analysis for Rotorcraft in Flight or in a Wind Tunnel
NASA Technical Reports Server (NTRS)
Johnson, W.
1977-01-01
An analytical model is developed for the aeroelastic behavior of a rotorcraft in flight or in a wind tunnel. A unified development is presented for a wide class of rotors, helicopters, and operating conditions. The equations of motion for the rotor are derived using an integral Newtonian method, which gives considerable physical insight into the blade inertial and aerodynamic forces. The rotor model includes coupled flap-lag bending and blade torsion degrees of freedom, and is applicable to articulated, hingeless, gimballed, and teetering rotors with an arbitrary number of blades. The aerodynamic model is valid for both high and low inflow, and for axial and nonaxial flight. The rotor rotational speed dynamics, including engine inertia and damping, and the perturbation inflow dynamics are included. For a rotor on a wind-tunnel support, a normal mode representation of the test module, strut, and balance system is used. The aeroelastic analysis for the rotorcraft in flight is applicable to a general two-rotor aircraft, including single main-rotor and tandem helicopter configurations, and side-by-side or tilting proprotor aircraft configurations.
Aeroelastic analysis of a troposkien-type wind turbine blade
NASA Technical Reports Server (NTRS)
Nitzsche, F.
1981-01-01
The linear aeroelastic equations for one curved blade of a vertical axis wind turbine in state vector form are presented. The method is based on a simple integrating matrix scheme together with the transfer matrix idea. The method is proposed as a convenient way of solving the associated eigenvalue problem for general support conditions.
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.
Controlled aeroelastic response and airfoil shaping using adaptive materials and integrated systems
NASA Astrophysics Data System (ADS)
Pinkerton, Jennifer L.; McGowan, Anna-Maria R.; Moses, Robert W.; Scott, Robert C.; Heeg, Jennifer
1996-05-01
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 charcteristics (ATTACH) project. The PARTI program demonstrated active flutter control and significant reductions 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. The 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-unimorph piezoelectric driver and sensor (THUNDER) wafers to control airfoil aerodynamic characteristics was investigated. Plans for future applications are also discussed.
Aeroelastic characteristics of composite bearingless rotor blades
NASA Technical Reports Server (NTRS)
Bielawa, R. L.
1976-01-01
Owing to the inherent unique structural features of composite bearingless rotors, various assumptions upon which conventional rotor aeroelastic analyses are formulated, are violated. Three such features identified are highly nonlinear and time-varying structural twist, structural redundancy in bending and torsion, and for certain configurations a strongly coupled low frequency bending-torsion mode. An examination of these aeroelastic considerations and appropriate formulations required for accurate analyses of such rotor systems is presented. Also presented are test results from a dynamically scaled model rotor and complementary analytic results obtained with the appropriately reformulated aeroelastic analysis.
Transonic aeroelastic analysis of the B-1 wing
NASA Technical Reports Server (NTRS)
Guruswamy, G. P.; Goorjian, P. M.; Ide, H.; Miller, G. D.
1986-01-01
The flow over the B-1 wing is studied computationally, including the aeroelastic response of the wing. Computed results are compared with results from wind tunnel and flight tests for both low- and high-sweep cases, at 25.0 and 67.5 deg, respectively, for selected transonic Mach numbers. The aerodynamic and aeroelastic computations are made by using the transonic unsteady code ATRAN3S. Steady aerodynamic computations compare well with wind tunnel results for the 25.0 deg sweep case and also for small angles of attack at 67.5 deg sweep case. The aeroelastic response results show that the wing is stable at the low-sweep angle for the calculation at the Mach number at which there is a shock wave. In the higher-sweep case, for the higher angle of attack at which oscillations were observed in the flight and wind tunnel tests, the calculations do not show any shock waves. Their absence lends support to the hypothesis that the observed oscillations are due to the presence of leading-edge separation vortices and not to shock wave motion, as was previously proposed.
NASA Astrophysics Data System (ADS)
Newman, James Charles, III
1997-10-01
The first two steps in the development of an integrated multidisciplinary design optimization procedure capable of analyzing the nonlinear fluid flow about geometrically complex aeroelastic configurations have been accomplished in the present work. For the first step, a three-dimensional unstructured grid approach to aerodynamic shape sensitivity analysis and design optimization has been developed. The advantage of unstructured grids, when compared with a structured-grid approach, is their inherent ability to discretize irregularly shaped domains with greater efficiency and less effort. Hence, this approach is ideally suited for geometrically complex configurations of practical interest. In this work the time-dependent, nonlinear Euler equations are solved using an upwind, cell-centered, finite-volume scheme. The discrete, linearized systems which result from this scheme are solved iteratively by a preconditioned conjugate-gradient-like algorithm known as GMRES for the two-dimensional cases and a Gauss-Seidel algorithm for the three-dimensional; at steady-state, similar procedures are used to solve the accompanying linear aerodynamic sensitivity equations in incremental iterative form. As shown, this particular form of the sensitivity equation makes large-scale gradient-based aerodynamic optimization possible by taking advantage of memory efficient methods to construct exact Jacobian matrix-vector products. Various surface parameterization techniques have been employed in the current study to control the shape of the design surface. Once this surface has been deformed, the interior volume of the unstructured grid is adapted by considering the mesh as a system of interconnected tension springs. Grid sensitivities are obtained by differentiating the surface parameterization and the grid adaptation algorithms with ADIFOR, an advanced automatic-differentiation software tool. To demonstrate the ability of this procedure to analyze and design complex configurations of
Formulation of the aeroelastic stability and response problem of coupled rotor/support systems
NASA Technical Reports Server (NTRS)
Warmbrodt, W.; Friedmann, P.
1979-01-01
The consistent formulation of the governing nonlinear equations of motion for a coupled rotor/support system is presented. Rotor/support coupling is clearly documented by enforcing dynamic equilibrium between the rotor and the moving flexible support. The nonlinear periodic coefficient equations of motion are applicable to both coupled rotor/fuselage aeroelastic problems of helicopters in hover or forward flight and coupled rotor/tower dynamics of a large horizontal axis wind turbine (HAWT). Finally, the equations of motion are used to study the influence of flexible supports and nonlinear terms on rotor aeroelastic stability and response of a large two-bladed HAWT.
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.
NASA Technical Reports Server (NTRS)
Reding, J. P.; Ericsson, L. E.
1976-01-01
An exploratory analysis has been made of the aeroelastic stability of the Space Shuttle Launch Configuration, with the objective of defining critical flow phenomena with adverse aeroelastic effects and developing simple analytic means of describing the time-dependent flow-interference effects so that they can be incorporated into a computer program to predict the aeroelastic stability of all free-free modes of the shuttle launch configuration. Three critical flow phenomana have been identified: (1) discontinuous jump of orbiter wing shock, (2) inlet flow between orbiter and booster, and (3) H.O. tank base flow. All involve highly nonlinear and often discontinuous aerodynamics which cause limit cycle oscillations of certain critical modes. Given the appropriate static data, the dynamic effects of the wing shock jump and the HO tank bulbous base effect can be analyzed using the developed quasi-steady techniques. However, further analytic and experimental efforts are required before the dynamic effects of the inlet flow phenomenon can be predicted for the shuttle launch configuration.
NASA Astrophysics Data System (ADS)
Ganguli, R.
2002-11-01
An aeroelastic analysis based on finite elements in space and time is used to model the helicopter rotor in forward flight. The rotor blade is represented as an elastic cantilever beam undergoing flap and lag bending, elastic torsion and axial deformations. The objective of the improved design is to reduce vibratory loads at the rotor hub that are the main source of helicopter vibration. Constraints are imposed on aeroelastic stability, and move limits are imposed on the blade elastic stiffness design variables. Using the aeroelastic analysis, response surface approximations are constructed for the objective function (vibratory hub loads). It is found that second order polynomial response surfaces constructed using the central composite design of the theory of design of experiments adequately represents the aeroelastic model in the vicinity of the baseline design. Optimization results show a reduction in the objective function of about 30 per cent. A key accomplishment of this paper is the decoupling of the analysis problem and the optimization problems using response surface methods, which should encourage the use of optimization methods by the helicopter industry.
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.
Application of unsteady aeroelastic analysis techniques on the national aerospace plane
NASA Technical Reports Server (NTRS)
Pototzky, Anthony S.; Spain, Charles V.; Soistmann, David L.; Noll, Thomas E.
1988-01-01
A presentation provided at the Fourth National Aerospace Plane Technology Symposium held in Monterey, California, in February 1988 is discussed. The objective is to provide current results of ongoing investigations to develop a methodology for predicting the aerothermoelastic characteristics of NASP-type (hypersonic) flight vehicles. Several existing subsonic and supersonic unsteady aerodynamic codes applicable to the hypersonic class of flight vehicles that are generally available to the aerospace industry are described. These codes were evaluated by comparing calculated results with measured wind-tunnel aeroelastic data. The agreement was quite good in the subsonic speed range but showed mixed agreement in the supersonic range. In addition, a future endeavor to extend the aeroelastic analysis capability to hypersonic speeds is outlined. An investigation to identify the critical parameters affecting the aeroelastic characteristics of a hypersonic vehicle, to define and understand the various flutter mechanisms, and to develop trends for the important parameters using a simplified finite element model of the vehicle is summarized. This study showed the value of performing inexpensive and timely aeroelastic wind-tunnel tests to expand the experimental data base required for code validation using simple to complex models that are representative of the NASP configurations and root boundary conditions are discussed.
Static Aeroelastic Scaling and Analysis of a Sub-Scale Flexible Wing Wind Tunnel Model
NASA Technical Reports Server (NTRS)
Ting, Eric; Lebofsky, Sonia; Nguyen, Nhan; Trinh, Khanh
2014-01-01
This paper presents an approach to the development of a scaled wind tunnel model for static aeroelastic similarity with a full-scale wing model. The full-scale aircraft model is based on the NASA Generic Transport Model (GTM) with flexible wing structures referred to as the Elastically Shaped Aircraft Concept (ESAC). The baseline stiffness of the ESAC wing represents a conventionally stiff wing model. Static aeroelastic scaling is conducted on the stiff wing configuration to develop the wind tunnel model, but additional tailoring is also conducted such that the wind tunnel model achieves a 10% wing tip deflection at the wind tunnel test condition. An aeroelastic scaling procedure and analysis is conducted, and a sub-scale flexible wind tunnel model based on the full-scale's undeformed jig-shape is developed. Optimization of the flexible wind tunnel model's undeflected twist along the span, or pre-twist or wash-out, is then conducted for the design test condition. The resulting wind tunnel model is an aeroelastic model designed for the wind tunnel test condition.
Development of an Aeroelastic Modeling Capability for Transient Nozzle Side Load Analysis
NASA Technical Reports Server (NTRS)
Wang, Ten-See; Zhao, Xiang; Zhang, Sijun; Chen, Yen-Sen
2013-01-01
Lateral nozzle forces are known to cause severe structural damage to any new rocket engine in development. Currently there is no fully coupled computational tool to analyze this fluid/structure interaction process. The objective of this study was to develop a fully coupled aeroelastic modeling capability to describe the fluid/structure interaction process during the transient nozzle operations. The aeroelastic model composes of three components: the computational fluid dynamics component based on an unstructured-grid, pressure-based computational fluid dynamics formulation, the computational structural dynamics component developed in the framework of modal analysis, and the fluid-structural interface component. The developed aeroelastic model was applied to the transient nozzle startup process of the Space Shuttle Main Engine at sea level. The computed nozzle side loads and the axial nozzle wall pressure profiles from the aeroelastic nozzle are compared with those of the published rigid nozzle results, and the impact of the fluid/structure interaction on nozzle side loads is interrogated and presented.
NASA Technical Reports Server (NTRS)
Chwalowski, Pawel; Florance, Jennifer P.; Heeg, Jennifer; Wieseman, Carol D.; Perry, Boyd P.
2011-01-01
This paper presents preliminary computational aeroelastic analysis results generated in preparation for the first Aeroelastic Prediction Workshop (AePW). These results were produced using FUN3D software developed at NASA Langley and are compared against the experimental data generated during the HIgh REynolds Number Aero- Structural Dynamics (HIRENASD) Project. The HIRENASD wind-tunnel model was tested in the European Transonic Windtunnel in 2006 by Aachen University0s Department of Mechanics with funding from the German Research Foundation. The computational effort discussed here was performed (1) to obtain a preliminary assessment of the ability of the FUN3D code to accurately compute physical quantities experimentally measured on the HIRENASD model and (2) to translate the lessons learned from the FUN3D analysis of HIRENASD into a set of initial guidelines for the first AePW, which includes test cases for the HIRENASD model and its experimental data set. This paper compares the computational and experimental results obtained at Mach 0.8 for a Reynolds number of 7 million based on chord, corresponding to the HIRENASD test conditions No. 132 and No. 159. Aerodynamic loads and static aeroelastic displacements are compared at two levels of the grid resolution. Harmonic perturbation numerical results are compared with the experimental data using the magnitude and phase relationship between pressure coefficients and displacement. A dynamic aeroelastic numerical calculation is presented at one wind-tunnel condition in the form of the time history of the generalized displacements. Additional FUN3D validation results are also presented for the AGARD 445.6 wing data set. This wing was tested in the Transonic Dynamics Tunnel and is commonly used in the preliminary benchmarking of computational aeroelastic software.
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)
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.
NASA Technical Reports Server (NTRS)
Gupta, K. K.
1997-01-01
A multidisciplinary, finite element-based, highly graphics-oriented, linear and nonlinear analysis capability that includes such disciplines as structures, heat transfer, linear aerodynamics, computational fluid dynamics, and controls engineering has been achieved by integrating several new modules in the original STARS (STructural Analysis RoutineS) computer program. Each individual analysis module is general-purpose in nature and is effectively integrated to yield aeroelastic and aeroservoelastic solutions of complex engineering problems. Examples of advanced NASA Dryden Flight Research Center projects analyzed by the code in recent years include the X-29A, F-18 High Alpha Research Vehicle/Thrust Vectoring Control System, B-52/Pegasus Generic Hypersonics, National AeroSpace Plane (NASP), SR-71/Hypersonic Launch Vehicle, and High Speed Civil Transport (HSCT) projects. Extensive graphics capabilities exist for convenient model development and postprocessing of analysis results. The program is written in modular form in standard FORTRAN language to run on a variety of computers, such as the IBM RISC/6000, SGI, DEC, Cray, and personal computer; associated graphics codes use OpenGL and IBM/graPHIGS language for color depiction. This program is available from COSMIC, the NASA agency for distribution of computer programs.
NASA Astrophysics Data System (ADS)
Keller, Jonathan Allen
2001-11-01
An analysis has been developed to predict the transient aeroelastic response of a helicopter rotor system during shipboard engagement and disengagement operations. The coupled flap-lag-torsion equations of motion were developed using Hamilton's Principle and discretized spatially using the finite element method. Aerodynamics were simulated using nonlinear quasi-steady or time domain nonlinear unsteady models. The ship airwake environment was simulated with simple deterministic airwake distributions, results from experimental measurements or numerical predictions. The transient aeroelastic response of the rotor blades was then time-integrated along a specified rotor speed profile. The control of the rotor response for an analytic model of the H-46 Sea Knight rotor system was investigated with three different passive control techniques. Collective pitch scheduling was only successful in reducing the blade flapping response in a few isolated cases. In the majority of cases, the blade transient response was increased. The use of a discrete flap damper in the very low rotor speed region was also investigated. Only by raising the flap stop setting and using a flap damper four times the strength of the lag damper could the downward flap deflections be reduced. However, because the flap stop setting was raised the upward flap deflections were often increased. The use of extendable/retractable, gated leading-edge spoilers in the low rotor speed region was also investigated. Spoilers covering the outer 15% R of the rotor blade were shown to significantly reduce both the upward and downward flap response without increasing rotor torque. Previous aeroelastic analyses developed at the University of Southampton and at Penn State University were completed with flap-torsion degrees of freedom only. The addition of the lag degree of freedom was shown to significantly influence the blade response. A comparison of the two aerodynamic models showed that the nonlinear quasi
NASA Technical Reports Server (NTRS)
Bergquist, R. R.; Carlson, R. G.; Landgrebe, A. J.; Egolf, T. A.
1974-01-01
This User's Manual was prepared to provide the engineer with the information required to run the coupled mode version of the Normal Modes Rotor Aeroelastic Analysis Computer Program. The manual provides a full set of instructions for running the program, including calculation of blade modes, calculations of variable induced velocity distribution and the calculation of the time history of the response for either a single blade or a complete rotor with an airframe (the latter with constant inflow).
NASA Technical Reports Server (NTRS)
Chattopadhyay, Aditi
1996-01-01
The objective of this research is to develop analysis procedures to investigate the coupling of composite and smart materials to improve aeroelastic and vibratory response of aerospace structures. The structural modeling must account for arbitrarily thick geometries, embedded and surface bonded sensors and actuators and imperfections, such as delamination. Changes in the dynamic response due to the presence of smart materials and delaminations is investigated. Experiments are to be performed to validate the proposed mathematical model.
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)
Wilkie, W. Keats; Belvin, W. Keith; Park, K. C.
1996-01-01
A simple aeroelastic analysis of a helicopter rotor blade incorporating embedded piezoelectric fiber composite, interdigitated electrode blade twist actuators is described. The analysis consists of a linear torsion and flapwise bending model coupled with a nonlinear ONERA based unsteady aerodynamics model. A modified Galerkin procedure is performed upon the rotor blade partial differential equations of motion to develop a system of ordinary differential equations suitable for dynamics simulation using numerical integration. The twist actuation responses for three conceptual fullscale blade designs with realistic constraints on blade mass are numerically evaluated using the analysis. Numerical results indicate that useful amplitudes of nonresonant elastic twist, on the order of one to two degrees, are achievable under one-g hovering flight conditions for interdigitated electrode poling configurations. Twist actuation for the interdigitated electrode blades is also compared with the twist actuation of a conventionally poled piezoelectric fiber composite blade. Elastic twist produced using the interdigitated electrode actuators was found to be four to five times larger than that obtained with the conventionally poled actuators.
NASA Technical Reports Server (NTRS)
Wilkie, W. Keats; Park, K. C.
1996-01-01
A simple aeroelastic analysis of a helicopter rotor blade incorporating embedded piezoelectric fiber composite, interdigitated electrode blade twist actuators is described. The analysis consist of a linear torsion and flapwise bending model coupled with a nonlinear ONERA based unsteady aerodynamics model. A modified Galerkin procedure is performed upon the rotor blade partial differential equations of motion to develop a system of ordinary differential equations suitable for numerical integration. The twist actuation responses for three conceptual full-scale blade designs with realistic constraints on blade mass are numerically evaluated using the analysis. Numerical results indicate that useful amplitudes of nonresonant elastic twist, on the order of one to two degrees, are achievable under one-g hovering flight conditions for interdigitated electrode poling configurations. Twist actuation for the interdigitated electrode blades is also compared with the twist actuation of a conventionally poled piezoelectric fiber composite blade. Elastic twist produced using the interdigitated electrode actuators was found to be four to five times larger than that obtained with the conventionally poled actuators.
Development of an Aeroelastic Modeling Capability for Transient Nozzle Side Load Analysis
NASA Technical Reports Server (NTRS)
Wang, Ten-See; Zhao, Xiang; Zhang, Sijun; Chen, Yen-Sen
2013-01-01
Lateral nozzle forces are known to cause severe structural damage to any new rocket engine in development during test. While three-dimensional, transient, turbulent, chemically reacting computational fluid dynamics methodology has been demonstrated to capture major side load physics with rigid nozzles, hot-fire tests often show nozzle structure deformation during major side load events, leading to structural damages if structural strengthening measures were not taken. The modeling picture is incomplete without the capability to address the two-way responses between the structure and fluid. The objective of this study is to develop a coupled aeroelastic modeling capability by implementing the necessary structural dynamics component into an anchored computational fluid dynamics methodology. The computational fluid dynamics component is based on an unstructured-grid, pressure-based computational fluid dynamics formulation, while the computational structural dynamics component is developed in the framework of modal analysis. Transient aeroelastic nozzle startup analyses of the Block I Space Shuttle Main Engine at sea level were performed. The computed results from the aeroelastic nozzle modeling are presented.
Algorithm for Simulating Atmospheric Turbulence and Aeroelastic Effects on Simulator Motion Systems
NASA Technical Reports Server (NTRS)
Ercole, Anthony V.; Cardullo, Frank M.; Kelly, Lon C.; Houck, Jacob A.
2012-01-01
Atmospheric turbulence produces high frequency accelerations in aircraft, typically greater than the response to pilot input. Motion system equipped flight simulators must present cues representative of the aircraft response to turbulence in order to maintain the integrity of the simulation. Currently, turbulence motion cueing produced by flight simulator motion systems has been less than satisfactory because the turbulence profiles have been attenuated by the motion cueing algorithms. This report presents a new turbulence motion cueing algorithm, referred to as the augmented turbulence channel. Like the previous turbulence algorithms, the output of the channel only augments the vertical degree of freedom of motion. This algorithm employs a parallel aircraft model and an optional high bandwidth cueing filter. Simulation of aeroelastic effects is also an area where frequency content must be preserved by the cueing algorithm. The current aeroelastic implementation uses a similar secondary channel that supplements the primary motion cue. Two studies were conducted using the NASA Langley Visual Motion Simulator and Cockpit Motion Facility to evaluate the effect of the turbulence channel and aeroelastic model on pilot control input. Results indicate that the pilot is better correlated with the aircraft response, when the augmented channel is in place.
Zhang, Zhaoyan; Neubauer, Juergen; Berry, David A
2007-10-01
In an investigation of phonation onset, a linear stability analysis was performed on a two-dimensional, aeroelastic, continuum model of phonation. The model consisted of a vocal fold-shaped constriction situated in a rigid pipe coupled to a potential flow which separated at the superior edge of the vocal fold. The vocal fold constriction was modeled as a plane-strain linear elastic layer. The dominant eigenvalues and eigenmodes of the fluid-structure-interaction system were investigated as a function of glottal airflow. To investigate specific aerodynamic mechanisms of phonation onset, individual components of the glottal airflow (e.g., flow-induced stiffness, inertia, and damping) were systematically added to the driving force. The investigations suggested that flow-induced stiffness was the primary mechanism of phonation onset, involving the synchronization of two structural eigenmodes. Only under conditions of negligible structural damping and a restricted set of vocal fold geometries did flow-induced damping become the primary mechanism of phonation onset. However, for moderate to high structural damping and a more generalized set of vocal fold geometries, flow-induced stiffness remained the primary mechanism of phonation onset. PMID:17902864
NASA Astrophysics Data System (ADS)
Wilkie, William Keats
1997-12-01
. Determining the optimum tradeoff between blade torsional stiffness and piezoelectric twist actuation authority is the subject of the third study. For this investigation, a linearized hovering-flight eigenvalue analysis is developed. Linear optimal control theory is then utilized to develop an optimum active twist blade design in terms of reducing structural energy and control effort cost. The forward flight vibratory loads characteristics of the torsional stiffness optimized active twist blade are then examined using the nonlinear, forward flight aeroelastic analysis. The optimized active twist rotor blade is shown to have improved passive and active vibratory loads characteristics relative to the baseline active twist blades.
NASA Technical Reports Server (NTRS)
Whitlow, W., Jr.; Bennett, R. M.
1982-01-01
Since the aerodynamic theory is nonlinear, the method requires the coupling of two iterative processes - an aerodynamic analysis and a structural analysis. A full potential analysis code, FLO22, is combined with a linear structural analysis to yield aerodynamic load distributions on and deflections of elastic wings. This method was used to analyze an aeroelastically-scaled wind tunnel model of a proposed executive-jet transport wing and an aeroelastic research wing. The results are compared with the corresponding rigid-wing analyses, and some effects of elasticity on the aerodynamic loading are noted.
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.
Mach number effects on transonic aeroelastic forces and flutter characteristics
NASA Technical Reports Server (NTRS)
Mohr, Ross W.; Batina, John T.; Yang, Henry T. Y.
1988-01-01
Transonic aeroelastic stability analysis and flutter calculations are presented for a generic transport-type wing based on the use of the CAP-TSD (Computational Aeroelasticity Program - Transonic Small Disturbance) finite-difference code. The CAP-TSD code was recently developed for transonic unsteady aerodynamic and aeroelastic analysis of complete aircraft configurations. A binary aeroelastic system consisting of simple bending and torsion modes was used to study aeroelastic behavior at transonic speeds. Generalized aerodynamic forces are presented for a wide range of Mach number and reduced frequency. Aeroelastic characteristics are presented for variations in freestream Mach number, mass ratio, and bending-torsion frequency ratio. Flutter boundaries are presented which have two transonic dips in flutter speed. The first dip is the usual transonic dip involving a bending-dominated flutter mode. The second dip is characterized by a single degree-of-freedom torsion oscillation. These aeroelastic results are physically interpreted and shown to be related to the steady state shock location and changes in generalized aerodynamic forces due to freestream Mach number.
NASA Astrophysics Data System (ADS)
Milgram, Judah Henry
1997-08-01
A comprehensive analysis is developed to evaluate main rotor blade trailing edge flap systems for helicopter vibration reduction. The analysis incorporates a nonlinear aeroelastic rotor model, response calculations via finite element in time method, coupled wind tunnel trim to prescribed operating conditions, unsteady compressible aerodynamics of the flap, and a multicyclic flap controller. An extensive validation study was performed using experimental wind tunnel data. Correlation between predicted and measured blade natural frequencies was generally fair. Results for hover thrust and power agreed well with the experimental data. Some discrepancy was observed near the low thrust conditions. In forward flight, fair correlation was observed for the power required and trim controls. For the rotor with no trailing edge flap motion, overall correlation of blade loads was fair good, although significant descrepancies were observed in individual cases. A parametric study was conducted for a four-bladed Sikorsky S-76 main rotor. The combination of trailing edge flap and multicyclic controller is predicted to provide significant reductions in fixed system 4/rev hub loads. The flap motions could be optimized to reduce either hub shears or hub moments through variation of a single scalar weighting parameter in the control algorithm. The effects of parametric variations of design parameters such as flap length, span-wise location, and chord could largely be offset by compensating adjustments to the flap motions as determined by the multicyclic algorithm, provided the flap motions did not become too large. The results suggest that a flap with the smallest possible chord and largest possible deflections is preferred.
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
Nonlinear Aeroelastic Analysis Using a Time-Accurate Navier-Stokes Equations Solver
NASA Technical Reports Server (NTRS)
Kuruvila, Geojoe; Bartels, Robert E.; Hong, Moeljo S.; Bhatia, G.
2007-01-01
A method to simulate limit cycle oscillation (LCO) due to control surface freeplay using a modified CFL3D, a time-accurate Navier-Stokes computational fluid dynamics (CFD) analysis code with structural modeling capability, is presented. This approach can be used to analyze aeroelastic response of aircraft with structural behavior characterized by nonlinearity in the force verses displacement curve. A limited validation of the method, using very low Mach number experimental data for a three-degrees-of-freedom (pitch/plunge/flap deflection) airfoil model with flap freeplay, is also presented.
NASA Astrophysics Data System (ADS)
Bryant, Matthew; Wolff, Eric; Garcia, Ephrahim
2011-12-01
This study examines the design parameters affecting the stability characteristics of a novel fluid flow energy harvesting device powered by aeroelastic flutter vibrations. The energy harvester makes use of a modal convergence flutter instability to generate limit cycle bending oscillations of a cantilevered piezoelectric beam with a small flap connected to its free end by a revolute joint. The critical flow speed at which destabilizing aerodynamic effects cause self-excited vibrations of the structure to emerge is essential to the design of the energy harvester because it sets the lower bound on the operating wind speed and frequency range of the system. A linearized analytic model of the device that accounts for the three-way coupling between the structural, unsteady aerodynamic, and electrical aspects of the system is used to examine tuning several design parameters while the size of the system is held fixed. The effects on the aeroelastic system dynamics and relative sensitivity of the flutter stability boundary are presented and discussed. A wind tunnel experiment is performed to validate the model predictions for the most significant system parameters.
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.
NASA Technical Reports Server (NTRS)
Srivastava, R.; Reddy, T. S. R.
1997-01-01
The program DuctE3D is used for steady or unsteady aerodynamic and aeroelastic analysis of ducted fans. This guide describes the input data required and the output files generated, in using DuctE3D. The analysis solves three dimensional unsteady, compressible Euler equations to obtain the aerodynamic forces. A normal mode structural analysis is used to obtain the aeroelastic equations, which are solved using either the time domain or the frequency domain solution method. Sample input and output files are included in this guide for steady aerodynamic analysis and aeroelastic analysis of an isolated fan row.
Computational Aeroelastic Analysis of Ares Crew Launch Vehicle Bi-Modal Loading
NASA Technical Reports Server (NTRS)
Massey, Steven J.; Chwalowski, Pawel
2010-01-01
A Reynolds averaged Navier-Stokes analysis, with and without dynamic aeroelastic effects, is presented for the Ares I-X launch vehicle at transonic Mach numbers and flight Reynolds numbers for two grid resolutions and two angles of attack. The purpose of the study is to quantify the force and moment increment imparted by the sudden transition from fully separated flow around the crew module - service module junction to that of the bi-modal flow state in which only part of the flow reattaches. The bi-modal flow phenomenon is of interest to the guidance, navigation and control community because it causes a discontinuous jump in forces and moments. Computations with a rigid structure at zero zero angle of attack indicate significant increases in normal force and pitching moment. Dynamic aeroelastic computations indicate the bi-modal flow state is insensitive to vehicle flexibility due to the resulting deflections imparting only very small changes in local angle of attack. At an angle of attack of 2.5deg, the magnitude of the pitching moment increment resulting from the bi-modal state nearly triples, while occurring at a slightly lower Mach number. Significant grid induced variations between the solutions indicate that further grid refinement is warranted.
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.
Loads Model Development and Analysis for the F/A-18 Active Aeroelastic Wing Airplane
NASA Technical Reports Server (NTRS)
Allen, Michael J.; Lizotte, Andrew M.; Dibley, Ryan P.; Clarke, Robert
2005-01-01
The Active Aeroelastic Wing airplane was successfully flight-tested in March 2005. During phase 1 of the two-phase program, an onboard excitation system provided independent control surface movements that were used to develop a loads model for the wing structure and wing control surfaces. The resulting loads model, which was used to develop the control laws for phase 2, is described. The loads model was developed from flight data through the use of a multiple linear regression technique. The loads model input consisted of aircraft states and control surface positions, in addition to nonlinear inputs that were calculated from flight-measured parameters. The loads model output for each wing consisted of wing-root bending moment and torque, wing-fold bending moment and torque, inboard and outboard leading-edge flap hinge moment, trailing-edge flap hinge moment, and aileron hinge moment. The development of the Active Aeroelastic Wing loads model is described, and the ability of the model to predict loads during phase 2 research maneuvers is demonstrated. Results show a good match to phase 2 flight data for all loads except inboard and outboard leading-edge flap hinge moments at certain flight conditions. The average load prediction errors for all loads at all flight conditions are 9.1 percent for maximum stick-deflection rolls, 4.4 percent for 5-g windup turns, and 7.7 percent for 4-g rolling pullouts.
NASA Technical Reports Server (NTRS)
Bielawa, R. L.
1984-01-01
The mathematical development for the expanded capabilities of the G400 rotor aeroelastic analysis was examined. The G400PA expanded analysis simulates the dynamics of all conventional rotors, blade pendulum vibration absorbers, and the higher harmonic excitations resulting from prescribed vibratory hub motions and higher harmonic blade pitch control. The methodology for modeling the unsteady stalled airloads of two dimensional airfoils is discussed. Formulations for calculating the rotor impedance matrix appropriate to the higher harmonic blade excitations are outlined. This impedance matrix, and the associated vibratory hub loads, are the rotor dynamic characteristic elements for use in the simplified coupled rotor/fuselage vibration analysis (SIMVIB). Updates to the development of the original G400 theory, program documentation, user instructions and information are presented.
NASA Technical Reports Server (NTRS)
Guruswamy, Guru P.; MacMurdy, Dale E.; Kapania, Rakesh K.
1994-01-01
Strong interactions between flow about an aircraft wing and the wing structure can result in aeroelastic phenomena which significantly impact aircraft performance. Time-accurate methods for solving the unsteady Navier-Stokes equations have matured to the point where reliable results can be obtained with reasonable computational costs for complex non-linear flows with shock waves, vortices and separations. The ability to combine such a flow solver with a general finite element structural model is key to an aeroelastic analysis in these flows. Earlier work involved time-accurate integration of modal structural models based on plate elements. A finite element model was developed to handle three-dimensional wing boxes, and incorporated into the flow solver without the need for modal analysis. Static condensation is performed on the structural model to reduce the structural degrees of freedom for the aeroelastic analysis. Direct incorporation of the finite element wing-box structural model with the flow solver requires finding adequate methods for transferring aerodynamic pressures to the structural grid and returning deflections to the aerodynamic grid. Several schemes were explored for handling the grid-to-grid transfer of information. The complex, built-up nature of the wing-box complicated this transfer. Aeroelastic calculations for a sample wing in transonic flow comparing various simple transfer schemes are presented and discussed.
Finite element analysis for bearingless rotor blade aeroelasticity
NASA Technical Reports Server (NTRS)
Sivaneri, N. T.; Chopra, I.
1982-01-01
A conventional articulated rotor blade has mechanical flap and lag hinges, a lag damper, and a pitch bearing. In connection with an interest in designs of greater mechanical simplicity and increased maintainability, hingeless and bearingless rotors have been developed. A hingeless blade lacks the hinges and is cantilevered at the hub. It does have a pitch bearing for pitch control. A bearingless design eliminates the hinges and the pitch bearing as well. In the present investigation of bearingless rotor blade characteristics, finite element analysis has been successfully applied to determine the solutions of the nonlinear trim equations and the linearized flutter equations of multiple-load-path blades. The employed formulation is based on Hamilton's principle. The spatial dependence of the equations of motion is discretized by dividing the flexbeams, the torque tube, and the outboard into a number of elements.
An aeroelastic analysis of the Darrieus wind turbine
Meyer, E.E.; Smith, C.E.
1983-11-01
The flutter stability of a single Darrieus wind turbine blade spinning in still air is investigated. The blade is modeled as a thin, uniform beam pinned to the rotor shaft, with aerodynamic forces accounted for using strip theory. Eliminating the spatial dependence using cubic B-splines results in a system of algebraic characteristic equations from which the stability of linear motions of the blade at any rotation rate may be inferred. The two most dangerous flutter modes are characterized for a one-parameter family of blades, and the flutter mechanism is shown to be dominated by gyroscopic coupling between motions in the plane of the blade and normal to the plane of the blade.
Active Control Analysis for Aeroelastic Instabilities in Turbomachines
NASA Technical Reports Server (NTRS)
Keith, Theo G., Jr.; Srivastava, Rakesh
2002-01-01
Turbomachines onboard aircraft operate in a highly complex and harsh environment. The unsteady flowfield inherent to turbomachines leads to several problems associated with safety, stability, performance and noise. In-flight surge or flutter incidents could be catastrophic and impact the safety and reliability of the aircraft. High-Cycle-Fatigue (HCF), on the other hand, can significantly impact safety, readiness and maintenance costs. To avoid or minimize these problems generally a more conservative design method must be initiated which results in thicker blades and a loss of performance. Actively controlled turbomachines have the potential to reduce or even eliminate the instabilities by impacting the unsteady aerodynamic characteristics. By modifying the unsteady aerodynamics, active control may significantly improve the safety and performance especially at off-design conditions, reduce noise, and increase the range of operation of the turbomachine. Active control can also help improve reliability for mission critical applications such as the Mars Flyer. In recent years, HCF has become one of the major issues concerning the cost of operation for current turbomachines. HCF alone accounts for roughly 30% of maintenance cost for the United States Air-Force. Other instabilities (flutter, surge, rotating-stall, etc.) are generally identified during the design and testing phase. Usually a redesign overcomes these problems, often reducing performance and range of operation, and resulting in an increase in the development cost and time. Despite a redesign, the engines do not have the capabilities or means to cope with in-flight unforeseen vibration, stall, flutter or surge related instabilities. This could require the entire fleet worldwide to be stood down for expensive modifications. These problems can be largely overcome by incorporating active control within the turbomachine and its design. Active control can help in maintaining the integrity of the system in
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.
Identification of dominant modes in random dynamical and aeroelastic systems
NASA Astrophysics Data System (ADS)
Hossain, Md. Nurtaj; Sarkar, Soumyadipta; Ghosh, Debraj
2015-11-01
Identification of dominant modes is an important step in studying linearly vibrating systems, including flow-induced vibrations. In the presence of uncertainty, when some of the system parameters and the external excitation are modeled as random quantities, this step becomes more difficult. This work is aimed at giving a systematic treatment to this end. The ability to capture the time-averaged kinetic energy is chosen as the primary criterion for selection of modes. Accordingly, a methodology is proposed based on the overlap of probability density functions (pdf) of the natural and excitation frequencies, proximity of the natural frequencies of the mean or baseline system, modal participation factor, and stochastic variation of mode shapes in terms of the modes of the baseline system - termed here as statistical modal overlapping. The probabilistic descriptors of the natural frequencies and mode shapes are found by solving a random eigenvalue problem. Three distinct vibration scenarios are considered: (i) undamped and damped free vibrations of a bladed disk assembly, (ii) forced vibration of a building, and (iii) flutter of a bridge model. Through numerical studies, it is observed that the proposed methodology gives an accurate selection of modes.
NASA Technical Reports Server (NTRS)
Liu, G.
1985-01-01
One of the major concerns in the design of an active control system is obtaining the information needed for effective feedback. This involves the combination of sensing and estimation. A sensor location index is defined as the weighted sum of the mean square estimation errors in which the sensor locations can be regarded as estimator design parameters. The design goal is to choose these locations to minimize the sensor location index. The choice of the number of sensors is a tradeoff between the estimation quality based upon the same performance index and the total costs of installing and maintaining extra sensors. An experimental study for choosing the sensor location was conducted on an aeroelastic system. The system modeling which includes the unsteady aerodynamics model developed by Stephen Rock was improved. Experimental results verify the trend of the theoretical predictions of the sensor location index for different sensor locations at various wind speeds.
A study of aeroelastic stability for the model support system of the National Transonic Facility
NASA Technical Reports Server (NTRS)
Strganac, Thomas W.
1988-01-01
Oscillations of wind-tunnel models have been observed during testing in the National Transonic Facility. These oscillations have been the subject of an extensive investigation. As a part of this effort, a study of the aeroelastic stability of the model support structure has been performed. This structure is mathematically modelled as a wing and conventional flutter analysis is performed. The math model implemented both experimentally and numerically obtained modal characteristics. A technique for illustrating the flutter boundary for wind tunnels is demonstrated. Results indicate that the classical flutter boundary is well above the operating envelope of the facility. However, the analysis indicates a damping-dependent instability is present which is inherent in the design. One possible modification in the design has been evaluated which eliminates the predicted instability.
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.
Global Nonlinear Analysis of Piezoelectric Energy Harvesting from Ambient and Aeroelastic Vibrations
NASA Astrophysics Data System (ADS)
Abdelkefi, Abdessattar
Converting vibrations to a usable form of energy has been the topic of many recent investigations. The ultimate goal is to convert ambient or aeroelastic vibrations to operate low-power consumption devices, such as microelectromechanical systems, heath monitoring sensors, wireless sensors or replacing small batteries that have a finite life span or would require hard and expensive maintenance. The transduction mechanisms used for transforming vibrations to electric power include: electromagnetic, electrostatic, and piezoelectric mechanisms. Because it can be used to harvest energy over a wide range of frequencies and because of its ease of application, the piezoelectric option has attracted significant interest. In this work, we investigate the performance of different types of piezoelectric energy harvesters. The objective is to design and enhance the performance of these harvesters. To this end, distributed-parameter and phenomenological models of these harvesters are developed. Global analysis of these models is then performed using modern methods of nonlinear dynamics. In the first part of this Dissertation, global nonlinear distributed-parameter models for piezoelectric energy harvesters under direct and parametric excitations are developed. The method of multiple scales is then used to derive nonlinear forms of the governing equations and associated boundary conditions, which are used to evaluate their performance and determine the effects of the nonlinear piezoelectric coefficients on their behavior in terms of softening or hardening. In the second part, we assess the influence of the linear and nonlinear parameters on the dynamic behavior of a wing-based piezoaeroelastic energy harvester. The system is composed of a rigid airfoil that is constrained to pitch and plunge and supported by linear and nonlinear torsional and flexural springs with a piezoelectric coupling attached to the plunge degree of freedom. Linear analysis is performed to determine the
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.
Aeroelastic effects in the structural dynamic analysis of vertical axis wind turbines
Lobitz, D.W.; Ashwill, T.D.
1985-01-01
Aeroelastic effects impact the structural dynamic behavior of vertical axis wind turbines (VAWTs) in two major ways. First the stability phenomena of flutter and divergence are direct results of the aeroelasticity of the structure. Secondly, aerodynamic damping can be important for predicting response levels particularly near resonance but also for off resonance conditions. The inclusion of the aeroelasticity is carried out by modifying the damping and stiffness matrices in the NASTRAN finite element code. Through the use of a specially designed preprocessor which reads the usual NASTRAN input deck and adds appropriate cards to it the incorporation of the aeroelastic effects has been made relatively transparent to the user NASTRAN flutter predictions are validated using field measurements and the effect of aerodynamic damping is demonstrated through an application to the Test Bed VAWT being designed at Sandia.
Aeroelastic effects in the structural dynamic analysis of vertical axis wind turbines
Lobitz, D.W.; Ashwill, T.D.
1986-04-01
Aeroelastic effects impact the structural dynamic behavior of vertical axis wind turbines (VAWRs) in two major ways. First, the stability phenomena of flutter and divergence are direct results of the aeroelasticity of the structure. Secondly, aerodynamic damping can be important for predicting response levels, particularly near resonance, but also for off-resonance conditions. The inclusion of the aeroelasticity is carried out by modifying the damping and stiffness matrices in the NASTRAN finite element code. Through the use of a specially designed preprocessor, which reads the usual NASTRAN input deck and adds appropriate cards to it, the incorporation of the aeroelastic effects has been made relatively transparent to the user. NASTRAN flutter predictions are validated using field measurements and the effect of aerodynamic damping is demonstrated through an application to the Test Bed VAWT being designed at Sandia.
Nastran level 16 theoretical manual updates for aeroelastic analysis of bladed discs
NASA Technical Reports Server (NTRS)
Elchuri, V.; Smith, G. C. C.
1980-01-01
A computer program based on state of the art compressor and structural technologies applied to bladed shrouded disc was developed and made operational in NASTRAN Level 16. Aeroelastic analyses, modes and flutter. Theoretical manual updates are included.
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.
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.
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.
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 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 Astrophysics Data System (ADS)
Vasconcellos, R.; Abdelkefi, A.
2015-12-01
Impacts are present in real aircraft movable surfaces, such as ailerons, flaps, rudder, elevators, trim tabs among other secondary control surfaces leading to complex, dangerous, and abrupt transitions. In this research study, we investigate the effects of discontinuous nonlinear stiffness simulating regions of freeplay, linear stiffness, and stoppers in the pitch degree of freedom on the response of a two-degree of freedom aeroelastic system. This system consists of a plunging and pitching rigid airfoil supported by a linear spring in the plunge degree of freedom and a nonlinear spring which includes the simulated stoppers at high angles in the pitch degree of freedom. The unsteady representation based on the Duhamel formulation is used to model the aerodynamic lift and moment. To characterize the system's response when subjected to impacts, we use modern methods of nonlinear dynamics including phase portraits, power spectra, and Poincaré sections. The results show that five main complex transitions are observed as the freestream velocity is increased. It is demonstrated that the observed transitions can be associated with grazing and/or grazing-sliding bifurcations.
Analysis of Test Case Computations and Experiments for the First Aeroelastic Prediction Workshop
NASA Technical Reports Server (NTRS)
Schuster, David M.; Heeg, Jennifer; Wieseman, Carol D.; Chwalowski, Pawel
2013-01-01
This paper compares computational and experimental data from the Aeroelastic Prediction Workshop (AePW) held in April 2012. This workshop was designed as a series of technical interchange meetings to assess the state of the art of computational methods for predicting unsteady flowfields and static and dynamic aeroelastic response. The goals are to provide an impartial forum to evaluate the effectiveness of existing computer codes and modeling techniques to simulate aeroelastic problems and to identify computational and experimental areas needing additional research and development. Three subject configurations were chosen from existing wind-tunnel data sets where there is pertinent experimental data available for comparison. Participant researchers analyzed one or more of the subject configurations, and results from all of these computations were compared at the workshop.
Static Aeroelastic Analysis of Transonic Wind Tunnel Models Using Finite Element Methods
NASA Technical Reports Server (NTRS)
Hooker, John R.; Burner, Alpheus W.; Valla, Robert
1997-01-01
A computational method for accurately predicting the static aeroelastic deformations of typical transonic transport wind tunnel models is described. The method utilizes a finite element method (FEM) for predicting the deformations. Extensive calibration/validation of this method was carried out using a novel wind-off wind tunnel model static loading experiment and wind-on optical wing twist measurements obtained during a recent wind tunnel test in the National Transonic Facility (NTF) at NASA LaRC. Further validations were carried out using a Navier-Stokes computational fluid dynamics (CFD) flow solver to calculate wing pressure distributions about several aeroelastically deformed wings and comparing these predictions with NTF experimental data. Results from this aeroelastic deformation method are in good overall agreement with experimentally measured values. Including the predicted deformations significantly improves the correlation between CFD predicted and experimentally measured wing & pressures.
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.
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
A modal aeroelastic analysis combining structural and aerodynamic models is applied to counterrotating propfans to evaluate their structural integrity for wind-tunnel testing. The aeroelastic analysis code is an extension of the 2D analysis code called the Aeroelastic Stability and Response of Propulsion Systems. Rotational speed and freestream Mach number are the parameters for calculating the stability of the two blade designs with a modal method combining a finite-element structural model with 2D steady and unsteady cascade aerodynamic models. The model demonstrates convergence to the least stable aeroelastic mode, describes the effects of a nonuniform inflow, and permits the modification of geometry and rotation. The analysis shows that the propfan designs are suitable for the wind-tunnel test and confirms that the propfans should be flutter-free under the range of conditions of the testing.
NASA Technical Reports Server (NTRS)
Heeg, Jennifer; Dowell, Earl H.
2001-01-01
Discrete time aeroelastic models with explicitly retained aerodynamic modes have been generated employing a time marching vortex lattice aerodynamic model. This paper presents analytical results from eigenanalysis of these models. The potential of these models to calculate the behavior of modes that represent damped system motion (noncritical modes) in addition to the simple harmonic modes is explored. A typical section with only structural freedom in pitch is examined. The eigenvalues are examined and compared to experimental data. Issues regarding the convergence of the solution with regard to refining the aerodynamic discretization are investigated. Eigenvector behavior is examined; the eigenvector associated with a particular eigenvalue can be viewed as the set of modal participation factors for that particular mode. For the present formulation of the equations of motion, the vorticity for each aerodynamic element appears explicitly as an element of each eigenvector in addition to the structural dynamic generalized coordinates. Thus, modal participation of the aerodynamic degrees of freedom can be assessed in M addition to participation of structural degrees of freedom.
Analyzing Aeroelasticity in Turbomachines
NASA Technical Reports Server (NTRS)
Reddy, T. S. R.; Srivastava, R.
2003-01-01
ASTROP2-LE is a computer program that predicts flutter and forced responses of blades, vanes, and other components of such turbomachines as fans, compressors, and turbines. ASTROP2-LE is based on the ASTROP2 program, developed previously for analysis of stability of turbomachinery components. In developing ASTROP2- LE, ASTROP2 was modified to include a capability for modeling forced responses. The program was also modified to add a capability for analysis of aeroelasticity with mistuning and unsteady aerodynamic solutions from another program, LINFLX2D, that solves the linearized Euler equations of unsteady two-dimensional flow. Using LINFLX2D to calculate unsteady aerodynamic loads, it is possible to analyze effects of transonic flow on flutter and forced response. ASTROP2-LE can be used to analyze subsonic, transonic, and supersonic aerodynamics and structural mistuning for rotors with blades of differing structural properties. It calculates the aerodynamic damping of a blade system operating in airflow so that stability can be assessed. The code also predicts the magnitudes and frequencies of the unsteady aerodynamic forces on the airfoils of a blade row from incoming wakes. This information can be used in high-cycle fatigue analysis to predict the fatigue lives of the blades.
NASA Astrophysics Data System (ADS)
Sabri, Farhad
Shells of revolution, particularly cylindrical and conical shells, are one of the basic structural elements in the aerospace structures. With the advent of high speed aircrafts, these shells can show dynamic instabilities when they are exposed to a supersonic flow. Therefore, aeroelastic analysis of these elements is one of the primary design criteria which aeronautical engineers are dealing with. This analysis can be done with the help of finite element method (FEM) coupled with the computational fluid dynamic (CFD) or by experimental methods but it is time consuming and very expensive. The purpose of this dissertation is to develop such a numerical tool to do aeroelastic analysis in a fast and precise way. Meanwhile during the design stage, where the different configurations, loading and boundary conditions may need to be analyzed, this numerical method can be used very easily with the high order of reliability. In this study structural modeling is a combination of linear Sanders thin shell theory and classical finite element method. Based on this hybrid finite element method, the shell displacements are found from the exact solutions of shell theory rather than approximating by polynomial function done in traditional finite element method. This leads to a precise and fast convergence. Supersonic aerodynamic modeling is done based on the piston theory and modified piston theory with the shell curvature term. The stress stiffening due to lateral pressure and axial compression are also taken into accounts. Fluid-structure interaction in the presence of inside quiescent fluid is modeled based on the potential theory. In this method, fluid is considered as a velocity potential variable at each node of the shell element where its motion is expressed in terms of nodal elastic displacements at the fluid-structure interface. This proposed hybrid finite element has capabilities to do following analysis: (i) Buckling and vibration of an empty or partially fluid filled
Development and Analysis of a Swept Blade Aeroelastic Model for a Small Wind Turbine (Presentation)
Preus, R.; Damiani, R.; Lee, S.; Larwood, S.
2014-06-01
As part of the U.S. Department-of-Energy-funded Competitiveness Improvement Project, the National Renewable Energy Laboratory (NREL) developed new capabilities for aeroelastic modeling of precurved and preswept blades for small wind turbines. This presentation covers the quest for optimized rotors, computer-aided engineering tools, a case study, and summary of the results.
NASA Astrophysics Data System (ADS)
Pourtakdoust, Seid H.; Aliabadi, Saeed Karimain
Flapping micro air vehicle (FMAV) is considered to exhibit much better performance at low speeds and small sizes compared to fixed-wing MAVs. To maximize the potential and capabilities of FMAVs also to produce adequate design implications, a new aeroelastic model of a typical flexible FMAV is being developed utilizing Euler-Bernoulli torsion beam and quasi steady aerodynamic model. The new model accounts for all natural existing complex interactions between the mass, inertia, elastic properties, aerodynamic loading, flapping amplitude and frequency of the FMAV as well as the effects of several geometric and design parameters. To validate the proposed theoretical model, a typical FMAV as well as instrumented test stand for the online measurement of forces, flapping angle and power consumption have been constructed. The experimental results are initially utilized to validate the flight dynamic model, and several appropriate conclusions are drawn. The model is subsequently used to demonstrate the flapping propulsion characteristics of the FMAV via simulation. Using dimensionless parameters, a set of new generalized curves have been deduced. The results indicate that by proper adjustment of the wing stiffness parameter as a function of the reduced frequency, the FMAV will attain its optimum propulsive efficiency. This fact raises additional new ideas for further research in this area by utilizing intelligent variable stiffness materials and/or or active morphing technology for the sustained, high-performance flight of FMAVs. The generalized model can also be used to conduct a performance and stability analysis of FMAVs and to design and optimize flapping-wing structures.
NASA Technical Reports Server (NTRS)
Akaydin, H. Dogus; Moini-Yekta, Shayan; Housman, Jeffrey A.; Nguyen, Nhan
2015-01-01
In this paper, we present a static aeroelastic analysis of a wind tunnel test model of a wing in high-lift configuration using a viscous flow simulation code. The model wing was tailored to deform during the tests by amounts similar to a composite airliner wing in highlift conditions. This required use of a viscous flow analysis to predict the lift coefficient of the deformed wing accurately. We thus utilized an existing static aeroelastic analysis framework that involves an inviscid flow code (Cart3d) to predict the deformed shape of the wing, then utilized a viscous flow code (Overflow) to compute the aerodynamic loads on the deformed wing. This way, we reduced the cost of flow simulations needed for this analysis while still being able to predict the aerodynamic forces with reasonable accuracy. Our results suggest that the lift of the deformed wing may be higher or lower than that of the non-deformed wing, and the washout deformation of the wing is the key factor that changes the lift of the deformed wing in two distinct ways: while it decreases the lift at low to moderate angles of attack simply by lowering local angles of attack along the span, it increases the lift at high angles of attack by alleviating separation.
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.
Multivariable flight control synthesis and literal robustness analysis for an aeroelastic vehicle
NASA Technical Reports Server (NTRS)
Schmidt, David K.; Newman, Brett
1990-01-01
An integrated flight/aeroelastic control law is developed analytically for a hypothetical large supersonic transport aircraft in which the first aeroelastic mode frequency of the fuselage (6 rad/sec) is near the short-period mode (2 rad/sec). The approach employed is based on a linear-quadratic-regulator (LQR) formulation (yielding model-following state-feedback gains), followed by asymptotic loop-transfer recovery of LQR robustness (to produce an output-feedback control law). The derivation is outlined, and numerical results comparing the performance and multivariate stability robustness of the present controller with those of a classical controller are presented in graphs. The two controllers are shown to have similar characteristics, even with respect to the sources of limitations on robustness.
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.
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.
NASA Technical Reports Server (NTRS)
Nguyen, Nhan; Ting, Eric; Lebofsky, Sonia
2015-01-01
This paper presents data analysis of a flexible wing wind tunnel model with a variable camber continuous trailing edge flap (VCCTEF) design for drag minimization tested at the University of Washington Aeronautical Laboratory (UWAL). The wind tunnel test was designed to explore the relative merit of the VCCTEF concept for improved cruise efficiency through the use of low-cost aeroelastic model test techniques. The flexible wing model is a 10%-scale model of a typical transport wing and is constructed of woven fabric composites and foam core. The wing structural stiffness in bending is tailored to be half of the stiffness of a Boeing 757-era transport wing while the torsional stiffness is about the same. This stiffness reduction results in a wing tip deflection of about 10% of the wing semi-span. The VCCTEF is a multi-segment flap design having three chordwise camber segments and five spanwise flap sections for a total of 15 individual flap elements. The three chordwise camber segments can be positioned appropriately to create a desired trailing edge camber. Elastomeric material is used to cover the gaps in between the spanwise flap sections, thereby creating a continuous trailing edge. Wind tunnel data analysis conducted previously shows that the VCCTEF can achieve a drag reduction of up to 6.31% and an improvement in the lift-to-drag ratio (L=D) of up to 4.85%. A method for estimating the bending and torsional stiffnesses of the flexible wingUWAL wind tunnel model from static load test data is presented. The resulting estimation indicates that the stiffness of the flexible wing is significantly stiffer in torsion than in bending by as much as 9 to 1. The lift prediction for the flexible wing is computed by a coupled aerodynamic-structural model. The coupled model is developed by coupling a conceptual aerodynamic tool Vorlax with a finite-element model of the flexible wing via an automated geometry deformation tool. Based on the comparison of the lift curve slope
NASA Technical Reports Server (NTRS)
Bielawa, R. L.
1976-01-01
The differential equations of motion for the lateral and torsional deformations of a nonlinearly twisted rotor blade in steady flight conditions together with those additional aeroelastic features germane to composite bearingless rotors are derived. The differential equations are formulated in terms of uncoupled (zero pitch and twist) vibratory modes with exact coupling effects due to finite, time variable blade pitch and, to second order, twist. Also presented are derivations of the fully coupled inertia and aerodynamic load distributions, automatic pitch change coupling effects, structural redundancy characteristics of the composite bearingless rotor flexbeam - torque tube system in bending and torsion, and a description of the linearized equations appropriate for eigensolution analyses. Three appendixes are included presenting material appropriate to the digital computer program implementation of the analysis, program G400.
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
Strain actuated aeroelastic control
NASA Technical Reports Server (NTRS)
Lazarus, Kenneth B.
1992-01-01
Viewgraphs on strain actuated aeroelastic control are presented. Topics covered include: structural and aerodynamic modeling; control law design methodology; system block diagram; adaptive wing test article; bench-top experiments; bench-top disturbance rejection: open and closed loop response; bench-top disturbance rejection: state cost versus control cost; wind tunnel experiments; wind tunnel gust alleviation: open and closed loop response at 60 mph; wind tunnel gust alleviation: state cost versus control cost at 60 mph; wind tunnel command following: open and closed loop error at 60 mph; wind tunnel flutter suppression: open loop flutter speed; and wind tunnel flutter suppression: closed loop state cost curves.
Nonlinear aeroelastic analysis of high-aspect-ratio wings in low subsonic flow
NASA Astrophysics Data System (ADS)
Eskandary, K.; Dardel, M.; Pashaei, M. H.; Moosavi, A. K.
2012-01-01
In this study, aeroelastic characteristics of high-aspect-ratio wing models with structural nonlinearities in quasi-steady aerodynamics flows are investigated. The studied wing model is a cantilever wing with double bending and torsional vibrations and with large deflection ability in accordance with Hodges-Dowell wing model. This wing model is valid for long, straight and thin homogeneous isotropic beams. The aerodynamics model is based on quasi-steady aerodynamic which is valid for aerodynamic flows without wake, viscosity and compressibility effects. The effect of different parameters such as mass ratios and stiffness ratios on flutter and divergence velocities and limit cycle oscillation amplitudes are carefully studied.
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.
Reduced-Order Aerodynamic Model and Its Application to a Nonlinear Aeroelastic System
NASA Technical Reports Server (NTRS)
Tang, Deman; Conner, Mark D.; Dowell, Earl H.
1998-01-01
Starting from a finite state model for a two-dimensional aerodynamic flow over an airfoil, the eigenmodes of the aerodynamic flow are determined. Using a small number of these aerodynamic eigenmodes, ie., a reduced-order model, the aeroelastic model is formed by coupling them to a typical section structural model with a trailing-edge flap. A free-play nonlinearity is modeled. Results are shown from the finite state model, the reduced-order model, and previous theoretical and experimental work. All results are in good agreement.
NASA Technical Reports Server (NTRS)
Crouse, J. E.
1974-01-01
A method is presented for designing axial-flow compressor blading from blade elements defined on cones which pass through the blade-edge streamline locations. Each blade-element centerline is composed of two segments which are tangent to each other. The centerline and surfaces of each segment have constant change of angle with path distance. The stacking line for the blade elements can be leaned in both the axial and tangential directions. The output of the computer program gives coordinates for fabrication and properties for aeroelastic analysis for planar blade sections. These coordinates and properties are obtained by interpolation across conical blade elements. The program is structured to be coupled with an aerodynamic design program.
NASA Astrophysics Data System (ADS)
Marisarla, Soujanya; Ghia, Urmila; "Karman" Ghia, Kirti
2002-11-01
Towards a comprehensive aeroelastic analysis of a joined wing, fluid dynamics and structural analyses are initially performed separately. Steady flow calculations are currently performed using 3-D compressible Navier-Stokes equations. Flow analysis of M6-Onera wing served to validate the software for the fluid dynamics analysis. The complex flow field of the joined wing is analyzed and the prevailing fluid dynamic forces are computed using COBALT software. Currently, these forces are being transferred as fluid loads on the structure. For the structural analysis, several test cases were run considering the wing as a cantilever beam; these served as validation cases. A nonlinear structural analysis of the wing is being performed using ANSYS software to predict the deflections and stresses on the joined wing. Issues related to modeling, and selecting appropriate mesh for the structure were addressed by first performing a linear analysis. The frequencies and mode shapes of the deformed wing are obtained from modal analysis. Both static and dynamic analyses are carried out, and the results obtained are carefully analyzed. Loose coupling between the fluid and structural analyses is currently being examined.
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.
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
Rotary-wing aeroelasticity with application to VTOL vehicles
NASA Technical Reports Server (NTRS)
Friedmann, Peretz P.
1993-01-01
A concise assessment is presented of the state of the art in the field of rotary-wing aeroelasticity (RWE). The basic ingredients of RWE are reviewed, including structural modeling, unsteady aerodynamic modeling, formulation of the equations of motion, and solution methods. Results illustrating these methods are presented for isolated blades and coupled rotor-fuselage problems. The application of active controls to suppress aeromechanical and aeroelastic instabilities and to reduce vibration in rotorcraft is discussed. Structural optimization with aeroelastic constraints, gust response analysis of helicopters, and aeroelastic problems in special VTOL vehicles are briefly examined.
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 Analysis of Rotor Blades Using Cfd/csd Coupling in Hover Mode
NASA Astrophysics Data System (ADS)
Chen, Long; Wu, Yizhao; Xia, Jian
A computational fluid dynamics (CFD) is coupled with a computational structural dynamics (CSD) to simulate the unsteady rotor flow with aeroelasticity effects. An unstructured upwind Navier-Stokes solver was developed for this simulation, with 2nd order time-accurate dual-time stepping method for temporal discretization and low Mach number preconditioning method. For turbulent flows, both the Spalart-Allmaras and Menter's SST model are available. Mesh deformation is achieved through a fast dynamic grid method called Delaunay graph map method for unsteady flow simulation. The rotor blades are modeled as Hodges & Dowell's nonlinear beams coupled flap-lag-torsion. The rotorcraft computational structural dynamics code employs the 15-dof beam finite element formulation for modeling. The structure code was validated by comparing the natural frequencies of a rotor model with UMARC. The flow and structure codes are coupled tightly with information exchange several times at every time step. A rotor blade model's unsteady flow field in the hover mode is simulated using the coupling method. Effect of blade elasticity with aerodynamic loads was compared with rigid blade.
NASA Astrophysics Data System (ADS)
Cavagna, Luca; Ricci, Sergio; Travaglini, Lorenzo
2011-11-01
This paper presents a design framework called NeoCASS (Next generation Conceptual Aero-Structural Sizing Suite), developed at the Department of Aerospace Engineering of Politecnico di Milano in the frame of SimSAC (Simulating Aircraft Stability And Control Characteristics for Use in Conceptual Design) project, funded by EU in the context of 6th Framework Program. It enables the creation of efficient low-order, medium fidelity models particularly suitable for structural sizing, aeroelastic analysis and optimization at the conceptual design level. The whole methodology is based on the integration of geometry construction, aerodynamic and structural analysis codes that combine depictive, computational, analytical, and semi-empirical methods, validated in an aircraft design environment. The work here presented aims at including the airframe and its effect from the very beginning of the conceptual design. This aspect is usually not considered in this early phase. In most cases, very simplified formulas and datasheets are adopted, which implies a low level of detail and a poor accuracy. Through NeoCASS, a preliminar distribution of stiffness and inertias can be determined, given the initial layout. The adoption of empirical formulas is reduced to the minimum in favor of simple numerical methods. This allows to consider the aeroelastic behavior and performances, as well, improving the accuracy of the design tools during the iterative steps and lowering the development costs and reducing the time to market. The result achieved is a design tool based on computational methods for the aero-structural analysis and Multi-Disciplinary Optimization (MDO) of aircraft layouts at the conceptual design stage. A complete case study regarding the TransoniCRuiser aircraft, including validation of the results obtained using industrial standard tools like MSC/NASTRAN and a CFD (Computational Fluid Dynamics) code, is reported. As it will be shown, it is possible to improve the degree of
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
NASA Technical Reports Server (NTRS)
Bielawa, R. L.
1982-01-01
Mathematical development is presented for the expanded capabilities of the United Technologies Research Center (UTRC) G400 Rotor Aeroelastic Analysis. This expanded analysis, G400PA, simulates the dynamics of teetered rotors, blade pendulum vibration absorbers and the higher harmonic excitations resulting from prescribed vibratory hub motions and higher harmonic blade pitch control. Formulations are also presented for calculating the rotor impedance matrix appropriate to these higher harmonic blade excitations. This impedance matrix and the associated vibratory hub loads are intended as the rotor blade characteristics elements for use in the Simplified Coupled Rotor/Fuselage Vibration Analysis (SIMVIB). Sections are included presenting updates to the development of the original G400 theory, and material appropriate to the user of the G400PA computer program. This material includes: (1) a general descriptionof the tructuring of the G400PA FORTRAN coding, (2) a detaild description of the required input data and other useful information for successfully running the program, and (3) a detailed description of the output results.
NASA Technical Reports Server (NTRS)
Wang, James M.
1991-01-01
The aeroelastic stability of a shaft-fixed bearingless rotor is analyzed in wind-tunnel tests for a wide range of operating conditions in order to determine whether such a system could be made aeroelastically stable without incorporating auxiliary dampers. The model rotor and blade properties are determined and used as an input to a bearingless-rotor analysis. Theoretical predictions are compared with experimental results in hover and forward flights. The analysis predicts the lag mode damping satisfactorily for collective pitch between 5 deg and 10 deg; however, the quasi-steady linear aerodynamic modeling overpredicts the damping values for higher collective pitch settings. It is noted that soft blade pitch links improve aeroelastic stability in hover and at low advance ratio.
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.
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.
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.
A Proposed Role of Aeroelasticity in NASA's New Exploration Vision
NASA Technical Reports Server (NTRS)
Bartels, Robert E.; Moses, Robert W.; Scott, Robert C.; Templeton, Justin D.; Cheatwood, F. McNeil; Gnoffo, Peter A.; Buck, Greg M.
2005-01-01
On 14 January 2004, NASA received a mandate to return astronauts to the Moon, evolve a sustained presence there, then head out into the solar system to Mars and perhaps beyond. This new space exploration initiative directs NASA to develop human and robotic technologies that can deliver payloads larger than Apollo to the Moon, to Mars, and bring astronauts and samples safely back to Earth at costs much lower than Apollo. These challenges require creative aerospace systems. On proposed technology for safely delivering payloads to the surface of Mars and returning samples to Earth involves deployed flexible and inflatable decelerators for atmospheric entry. Because inflatable decelerators provide the entry vehicle more drag surface area at smaller mass than traditional ablative devices, this class of decelerators can potentially accomodate larger mass payloads. The flexibility of these lightweight aeroshells can pose both vehicle and aeroelastic stability problems if not properly designed for the expected flight regimes. Computational tools need to be developed for modelling the large and nonlinear deformations of these highly flexible structures. Unlike wind tunnel testing, an integrated and efficient aeroelastic analysis tool can explore the entire flight environment. This paper will provide some background on flexible deployable decelerators, survey the current state of technology and outline the proposed development of an aeroelastic analysis and capability.
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
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.
Recent advances in transonic computational aeroelasticity
NASA Technical Reports Server (NTRS)
Batina, John T.; Bennett, Robert M.; Seidel, David A.; Cunningham, Herbert J.; Bland, Samuel R.
1988-01-01
A transonic unsteady aerodynamic and aeroelasticity code called CAP-TSD was developed for application to realistic aircraft configurations. The code permits the calculation of steady and unsteady flows about complete aircraft configurations for aeroelastic analysis in the flutter critical transonic speed range. The CAP-TSD code uses a time accurate approximate factorization algorithm for solution of the unsteady transonic small disturbance potential equation. An overview is given of the CAP-TSD code development effort and results are presented which demonstrate various capabilities of the code. Calculations are presented for several configurations including the General Dynamics 1/9 scale F-16 aircraft model and the ONERA M6 wing. Calculations are also presented from a flutter analysis of a 45 deg sweptback wing which agrees well with the experimental data. Descriptions are presented of the CAP-TSD code and algorithm details along with results and comparisons which demonstrate these recent developments in transonic computational aeroelasticity.
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.
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)
Wilbur, Matthew L.
1998-01-01
At the Langley Research Center an active mount rotorcraft testbed is being developed for use in the Langley Transonic Dynamics Tunnel. This testbed, the second generation version of the Aeroelastic Rotor Experimental System (ARES-II), can impose rotor hub motions and measure the response so that rotor-body coupling phenomena may be investigated. An analytical method for coupling an aeroelastically scaled model rotor system to the ARES-II is developed in the current study. Models of the testbed and the rotor system are developed in independent analyses, and an impedance-matching approach is used to couple the rotor system to the testbed. The development of the analytical models and the coupling method is examined, and individual and coupled results are presented for the testbed and rotor system. Coupled results are presented with and without applied hub motion, and system loads and displacements are examined. The results show that a closed-loop control system is necessary to achieve desired hub motions, that proper modeling requires including the loads at the rotor hub and rotor control system, and that the strain-gauge balance placed in the rotating system of the ARES-II provided the best loads results.
NASA Technical Reports Server (NTRS)
Friedmann, P. P.
1984-01-01
An aeroelastic model suitable for the study of aeroelastic and structural dynamic effects in multirotor vehicles simulating a hybrid heavy lift vehicle was developed and applied to the study of a number of diverse problems. The analytical model developed proved capable of modeling a number of aeroelastic problems, namely: (1) isolated blade aeroelastic stability in hover and forward flight, (2) coupled rotor/fuselage aeromechanical problem in air or ground resonance, (3) tandem rotor coupled rotor/fuselage problems, and (4) the aeromechanical stability of a multirotor vehicle model representing a hybrid heavy lift airship (HHLA). The model was used to simulate the ground resonance boundaries of a three bladed hingeless rotor model, including the effect of aerodynamic loads, and the theoretical predictions compared well with experimental results. Subsequently the model was used to study the aeromechanical stability of a vehicle representing a hybrid heavy lift airship, and potential instabilities which could occur for this type of vehicle were identified. The coupling between various blade, supporting structure and rigid body modes was identified.
NASA Technical Reports Server (NTRS)
Perry, Boyd, III; Dunn, H. J.; Sandford, Maynard C.
1988-01-01
Nominal roll control laws were designed, implemented, and tested on an aeroelastically-scaled free-to-roll wind-tunnel model of an advanced fighter configuration. The tests were performed in the NASA Langley Transonic Dynamics Tunnel. A parametric study of the nominal roll control system was conducted. This parametric study determined possible control system gain variations which yielded identical closed-loop stability (roll mode pole location) and identical roll response but different maximum control-surface deflections. Comparison of analytical predictions with wind-tunnel results was generally very good.
NASA Astrophysics Data System (ADS)
Rezaee, Mousa; Jahangiri, Reza
2015-05-01
In this study, in the presence of supersonic aerodynamic loading, the nonlinear and chaotic vibrations and stability of a simply supported Functionally Graded Piezoelectric (FGP) rectangular plate with bonded piezoelectric layer have been investigated. It is assumed that the plate is simultaneously exposed to the effects of harmonic uniaxial in-plane force and transverse piezoelectric excitations and aerodynamic loading. It is considered that the potential distribution varies linearly through the piezoelectric layer thickness, and the aerodynamic load is modeled by the first order piston theory. The von-Karman nonlinear strain-displacement relations are used to consider the geometrical nonlinearity. Based on the Classical Plate Theory (CPT) and applying the Hamilton's principle, the nonlinear coupled partial differential equations of motion are derived. The Galerkin's procedure is used to reduce the equations of motion to nonlinear ordinary differential Mathieu equations. The validity of the formulation for analyzing the Limit Cycle Oscillation (LCO), aero-elastic stability boundaries is accomplished by comparing the results with those of the literature, and the convergence study of the FGP plate is performed. By applying the Multiple Scales Method, the case of 1:2 internal resonance and primary parametric resonance are taken into account and the corresponding averaged equations are derived and analyzed numerically. The results are provided to investigate the effects of the forcing/piezoelectric detuning parameter, amplitude of forcing/piezoelectric excitation and dynamic pressure, on the nonlinear dynamics and chaotic behavior of the FGP plate. It is revealed that under the certain conditions, due to the existence of bi-stable region of non-trivial solutions, system shows the hysteretic behavior. Moreover, in absence of airflow, it is observed that variation of control parameters leads to the multi periodic and chaotic motions.
Efficient Cfd/csd Coupling Methods for Aeroelastic Applications
NASA Astrophysics Data System (ADS)
Chen, Long; Xu, Tianhao; Xie, Jing
2016-06-01
A fast aeroelastic numerical simulation method using CFD/CSD coupling are developed. Generally, aeroelastic numerical simulation costs much time and significant hardware resources with CFD/CSD coupling. In this paper, dynamic grid method, full implicit scheme, parallel technology and improved coupling method are researched for efficiency simulation. An improved Delaunay graph mapping method is proposed for efficient dynamic grid deform. Hybrid grid finite volume method is used to solve unsteady flow fields. The dual time stepping method based on parallel implicit scheme is used in temporal discretization for efficiency simulation. An approximate system of linear equations is solved by the GMRES algorithm with a LU-SGS preconditioner. This method leads to a significant increase in performance over the explicit and LU-SGS implicit methods. A modification of LU-SGS is proposed to improve the parallel performance. Parallel computing overs a very effective way to improve our productivity in doing CFD/CFD coupling analysis. Improved loose coupling method is an efficiency way over the loose coupling method and tight coupling method. 3D wing's aeroelastic phenomenon is simulated by solving Reynolds-averaged Navier-Stokes equations using improved loose coupling method. The flutter boundary is calculated and agrees well with experimental data. The transonic hole is very clear in numerical simulation results.
OVERAERO-MPI: Parallel Overset Aeroelasticity Code
NASA Technical Reports Server (NTRS)
Gee, Ken; Rizk, Yehia M.
1999-01-01
An overset modal structures analysis code was integrated with a parallel overset Navier-Stokes flow solver to obtain a code capable of static aeroelastic computations. The new code was used to compute the static aeroelastic deformation of an arrow-wing-body geometry and a complex, full aircraft configuration. For the simple geometry, the results were similar to the results obtained with the ENSAERO code and the PVM version of OVERAERO. The full potential of this code suite was illustrated in the complex, full aircraft computations.
Rotor aeroelastic stability coupled with helicopter body motion
NASA Technical Reports Server (NTRS)
Miao, W. L.; Huber, H. B.
1974-01-01
A 5.5-foot-diameter, soft-in-plane, hingeless-rotor system was tested on a gimbal which allowed the helicopter rigid-body pitch and roll motions. Coupled rotor/airframe aeroelastic stability boundaries were explored and the modal damping ratios were measured. The time histories were correlated with analysis with excellent agreement. The effects of forward speed and some rotor design parameters on the coupled rotor/airframe stability were explored both by model and analysis. Some physical insights into the coupled stability phenomenon are suggested.
NASA Technical Reports Server (NTRS)
Gupta, Kajal K.
1991-01-01
The details of an integrated general-purpose finite element structural analysis computer program which is also capable of solving complex multidisciplinary problems is presented. Thus, the SOLIDS module of the program possesses an extensive finite element library suitable for modeling most practical problems and is capable of solving statics, vibration, buckling, and dynamic response problems of complex structures, including spinning ones. The aerodynamic module, AERO, enables computation of unsteady aerodynamic forces for both subsonic and supersonic flow for subsequent flutter and divergence analysis of the structure. The associated aeroservoelastic analysis module, ASE, effects aero-structural-control stability analysis yielding frequency responses as well as damping characteristics of the structure. The program is written in standard FORTRAN to run on a wide variety of computers. Extensive graphics, preprocessing, and postprocessing routines are also available pertaining to a number of terminals.
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.
Experimental aeroelasticity history, status and future in brief
NASA Technical Reports Server (NTRS)
Ricketts, Rodney H.
1990-01-01
NASA conducts wind tunnel experiments to determine and understand the aeroelastic characteristics of new and advanced flight vehicles, including fixed-wing, rotary-wing and space-launch configurations. Review and assessments are made of the state-of-the-art in experimental aeroelasticity regarding available facilities, measurement techniques, and other means and devices useful in testing. In addition, some past experimental programs are described which assisted in the development of new technology, validated new analysis codes, or provided needed information for clearing flight envelopes of unwanted aeroelastic response. Finally, needs and requirements for advances and improvements in testing capabilities for future experimental research and development programs are described.
Helicopter rotor dynamics and aeroelasticity - Some key ideas and insights
NASA Technical Reports Server (NTRS)
Friedmann, Peretz P.
1990-01-01
Four important current topics in helicopter rotor dynamics and aeroelasticity are discussed: (1) the role of geometric nonlinearities in rotary-wing aeroelasticity; (2) structural modeling, free vibration, and aeroelastic analysis of composite rotor blades; (3) modeling of coupled rotor/fuselage areomechanical problems and their active control; and (4) use of higher-harmonic control for vibration reduction in helicopter rotors in forward flight. The discussion attempts to provide an improved fundamental understanding of the current state of the art. In this way, future research can be focused on problems which remain to be solved instead of producing marginal improvements on problems which are already understood.
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.
Power extraction from aeroelastic limit cycle oscillations
NASA Astrophysics Data System (ADS)
Dunnmon, J. A.; Stanton, S. C.; Mann, B. P.; Dowell, E. H.
2011-11-01
Nonlinear limit cycle oscillations of an aeroelastic energy harvester are exploited for enhanced piezoelectric power generation from aerodynamic flows. Specifically, a flexible beam with piezoelectric laminates is excited by a uniform axial flow field in a manner analogous to a flapping flag such that the system delivers power to an electrical impedance load. Fluid-structure interaction is modeled by augmenting a system of nonlinear equations for an electroelastic beam with a discretized vortex-lattice potential flow model. Experimental results from a prototype aeroelastic energy harvester are also presented. Root mean square electrical power on the order of 2.5 mW was delivered below the flutter boundary of the test apparatus at a comparatively low wind speed of 27 m/s and a chord normalized limit cycle amplitude of 0.33. Moreover, subcritical limit cycles with chord normalized amplitudes of up to 0.46 were observed. Calculations indicate that the system tested here was able to access over 17% of the flow energy to which it was exposed. Methods for designing aeroelastic energy harvesters by exploiting nonlinear aeroelastic phenomena and potential improvements to existing relevant aerodynamic models are also discussed.
Recent Enhancements to the Development of CFD-Based Aeroelastic Reduced-Order Models
NASA Technical Reports Server (NTRS)
Silva, Walter A.
2007-01-01
Recent enhancements to the development of CFD-based unsteady aerodynamic and aeroelastic reduced-order models (ROMs) are presented. These enhancements include the simultaneous application of structural modes as CFD input, static aeroelastic analysis using a ROM, and matched-point solutions using a ROM. The simultaneous application of structural modes as CFD input enables the computation of the unsteady aerodynamic state-space matrices with a single CFD execution, independent of the number of structural modes. The responses obtained from a simultaneous excitation of the CFD-based unsteady aerodynamic system are processed using system identification techniques in order to generate an unsteady aerodynamic state-space ROM. Once the unsteady aerodynamic state-space ROM is generated, a method for computing the static aeroelastic response using this unsteady aerodynamic ROM and a state-space model of the structure, is presented. Finally, a method is presented that enables the computation of matchedpoint solutions using a single ROM that is applicable over a range of dynamic pressures and velocities for a given Mach number. These enhancements represent a significant advancement of unsteady aerodynamic and aeroelastic ROM technology.
Transonic Unsteady Aerodynamics and Aeroelasticity 1987, part 1
NASA Technical Reports Server (NTRS)
Bland, Samuel R. (Compiler)
1989-01-01
Computational fluid dynamics methods have been widely accepted for transonic aeroelastic analysis. Previously, calculations with the TSD methods were used for 2-D airfoils, but now the TSD methods are applied to the aeroelastic analysis of the complete aircraft. The Symposium papers are grouped into five subject areas, two of which are covered in this part: (1) Transonic Small Disturbance (TSD) theory for complete aircraft configurations; and (2) Full potential and Euler equation methods.
NASA Technical Reports Server (NTRS)
Pototzky, Anthony S.
2008-01-01
A simple matrix polynomial approach is introduced for approximating unsteady aerodynamics in the s-plane and ultimately, after combining matrix polynomial coefficients with matrices defining the structure, a matrix polynomial of the flutter equations of motion (EOM) is formed. A technique of recasting the matrix-polynomial form of the flutter EOM into a first order form is also presented that can be used to determine the eigenvalues near the origin and everywhere on the complex plane. An aeroservoelastic (ASE) EOM have been generalized to include the gust terms on the right-hand side. The reasons for developing the new matrix polynomial approach are also presented, which are the following: first, the "workhorse" methods such as the NASTRAN flutter analysis lack the capability to consistently find roots near the origin, along the real axis or accurately find roots farther away from the imaginary axis of the complex plane; and, second, the existing s-plane methods, such as the Roger s s-plane approximation method as implemented in ISAC, do not always give suitable fits of some tabular data of the unsteady aerodynamics. A method available in MATLAB is introduced that will accurately fit generalized aerodynamic force (GAF) coefficients in a tabular data form into the coefficients of a matrix polynomial form. The root-locus results from the NASTRAN pknl flutter analysis, the ISAC-Roger's s-plane method and the present matrix polynomial method are presented and compared for accuracy and for the number and locations of roots.
Aeroelastic problems in turbomachines
NASA Technical Reports Server (NTRS)
Bendiksen, Oddvar O.
1990-01-01
A review of the field of turbomachinery aeroelasticity is presented. Developments over the past decade are emphasized, and an assessment of possible future directions of research is offered. The paper reviews the areas of unsteady cascade flows, structural modeling, and flutter prediction methods. Representative results for unsteady flow calculations and flutter boundary predictions in subsonic, transonic, and supersonic flows are discussed, including recent calculations based on the methods of computational fluid mechanics. Results from current attempts to correlate experimental data with theoretical predictions are discussed briefly. It is recommended that future research include investigations of novel approaches to flutter calculations that can take full advantage of parallel processing supercomputers. The feasibility of using mistuning and aeroelastic tailoring as passive flutter suppression techniques should also be pursued.
NASA Astrophysics Data System (ADS)
Noor, Ahmed K.; Venneri, Samuel L.
Various papers on flight vehicle materials, structures, and dynamics are presented. Individual topics addressed include: general modeling methods, component modeling techniques, time-domain computational techniques, dynamics of articulated structures, structural dynamics in rotating systems, structural dynamics in rotorcraft, damping in structures, structural acoustics, structural design for control, structural modeling for control, control strategies for structures, system identification, overall assessment of needs and benefits in structural dynamics and controlled structures. Also discussed are: experimental aeroelasticity in wind tunnels, aeroservoelasticity, nonlinear aeroelasticity, aeroelasticity problems in turbomachines, rotary-wing aeroelasticity with application to VTOL vehicles, computational aeroelasticity, structural dynamic testing and instrumentation.
Localization of aeroelastic modes in mistuned high-energy turbines
NASA Astrophysics Data System (ADS)
Pierre, Christophe; Smith, Todd E.; Murthy, Durbha V.
1994-05-01
The effects of blade mistuning on the aeroelastic vibration characteristics of high-energy turbines are investigated, using the first stage of the oxidizer turbopump in the Space Shuttle main rocket engine as an example. A modal aeroelastic analysis procedure is used in concert with a linearized unsteady aerodynamic theory that accounts for the effects of blade thickness, camber, and steady loading. High sensitivity of the dynamic characteristics of mistuned rotors is demonstrated. In particular, the aeroelastic free vibration modes become localized to a few blades, possibly leading to rogue blade failure, and the locus of the aeroelastic eigenvalues loses its regular structure when small mistuning (of the order usually present in actual rotors) is introduced. Perturbation analyses that yield physical insights into these phenomena are presented. A powerful but easily calculated stochastic sensitivity measure that allows the global prediction of mistuning effects is developed.
Activities in Aeroelasticity at NASA Langley Research Center
NASA Technical Reports Server (NTRS)
Perry, Boyd, III; Noll, Thomas E.
1997-01-01
This paper presents the results of recently-completed research and presents status reports of current research being performed within the Aeroelasticity Branch of the NASA Langley Research Center. Within the paper this research is classified as experimental, analytical, and theoretical aeroelastic research. The paper also describes the Langley Transonic Dynamics Tunnel, its features, capabilities, a new open-architecture data acquisition system, ongoing facility modifications, and the subsequent calibration of the facility.
Aeroelastic Response of Swept Aircraft Wings in a Compressible Flow Field
NASA Technical Reports Server (NTRS)
Marzocca, Piergiovanni; Librescu, Liviu; Silva, Walter A.
2000-01-01
The present study addresses the subcritical aeroelastic response of swept wings, in various flight speed regimes, to arbitrary time-dependent external excitations. The methodology based on the concept of indicial functions is carried out in time and frequency domains. As a result of this approach, the proper unsteady aerodynamic loads necessary to study the subcritical aeroelastic response of the open/closed loop aeroelastic systems, and of flutter instability, respectively are obtained. Validation of the aeroelastic model is provided, and applications to subcritical aeroelastic response to blast pressure signatures are illustrated. In this context, an original representation of the aeroelastic response in the phase-space is displayed, and pertinent conclusions on the implications of a number of selected parameters of the system are outlined.
Static aeroelastic behavior of an adaptive laminated piezoelectric composite wing
NASA Technical Reports Server (NTRS)
Weisshaar, T. A.; Ehlers, S. M.
1990-01-01
The effect of using an adaptive material to modify the static aeroelastic behavior of a uniform wing is examined. The wing structure is idealized as a laminated sandwich structure with piezoelectric layers in the upper and lower skins. A feedback system that senses the wing root loads applies a constant electric field to the piezoelectric actuator. Modification of pure torsional deformaton behavior and pure bending deformation are investigated, as is the case of an anisotropic composite swept wing. The use of piezoelectric actuators to create an adaptive structure is found to alter static aeroelastic behavior in that the proper choice of the feedback gain can increase or decrease the aeroelastic divergence speed. This concept also may be used to actively change the lift effectiveness of a wing. The ability to modify static aeroelastic behavior is limited by physical limitations of the piezoelectric material and the manner in which it is integrated into the parent structure.
NASA Technical Reports Server (NTRS)
Venkatesan, C.; Friedman, P.
1984-01-01
This report presents a set of governing coupled differential equations for a model of a hybrid aircraft. The model consists of multiple rotor systems connected by an elastic interconnecting structure, with options to add any combination of or all of the following components; i.e., thrusters, a buoyant hull, and an underslung weight. The dynamic equations are written for the individual blade with hub motions, for the rigid body motions of the whole model, and also for the flexible modes of the interconnecting structure. One of the purposes of this study is to serve as the basis of a numerical study aimed at determining the aeroelastic stability and structural response characteristics of a Hybrid Heavy Lift Airship (HHLA). It is also expected that the formulation may be applicable to analyzing stability and responses of dual rotor helicopters such as a Heavy Lift Helicopter (HLH). Futhermore, the model is capable of representing coupled rotor/body aeromechanical problems of single rotor helicopters.
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.
Aeroelasticity - Frontiers and beyond /von Karman Lecture/
NASA Technical Reports Server (NTRS)
Garrick, I. E.
1976-01-01
The lecture aims at giving a broad survey of the current reaches of aeroelasticity with some narrower views for the specialist. After a short historical review of concepts for orientation, several topics are briefly presented. These touch on current flight vehicles having special points of aeroelastic interest; recent developments in the active control of aeroelastic response including control of flutter; remarks on the unsteady aerodynamics of arbitrary configurations; problems of the space shuttle related to aeroelasticity; and aeroelastic response in flight.
NASA Technical Reports Server (NTRS)
Gilbert, Michael G.; Silva, Walter A.
1987-01-01
A new design concept in the development of VTOL aircraft with high forward flight speed capability is that of the X-Wing, a stiff, bearingless helicopter rotor system which can be stopped in flight and the blades used as two forward-swept and two aft-swept wings. Because of the usual configuration in the fixed-wing mode, there is a high potential for aeroelastic divergence or flutter and coupling of blade vibration modes with rigid-body modes. An aeroelastic stability analysis of an X-Wing configuration aircraft was undertaken to determine if these problems could exist. This paper reports on the results of dynamic stability analyses in the lateral and longitudinal directions including the vehicle rigid-body and flexible modes. A static aeroelastic analysis using the normal vibration mode equations of motion was performed to determine the cause of a loss of longitudinal static margin with increasing airspeed. This loss of static margin was found to be due to aeroelastic washin of the forward-swept blades and washout of the aft-swept blades moving the aircraft aerodynamic center forward of the center of gravity. This phenomenon is likely to be generic to X-Wing aircraft.
NASA Technical Reports Server (NTRS)
Gilbert, Michael G.; Silva, Walter A.
1987-01-01
A new design concept in the development of vertical takeoff and landing aircraft with high forward flight speed capability is that of the X-Wing. The X-Wing is a stiff, bearingless helicopter rotor system which can be stopped in flight and the blades used as two forward-swept wings and two aft-swept wings. Because of the unusual configuration in the fixed-wing mode, there is a high potential for aeroelastic divergence or flutter and coupling of blade vibration modes with rigid-body modes. An aeroelastic stability analysis of an X-Wing configuration aircraft was undertaken to determine if these problems could exist. This paper reports on the results of dynamic stability analyses in the lateral and longitudinal directions including the vehicle rigid-body and flexible modes. A static aeroelastic analysis using the normal vibration mode equations of motion was performed to determine the cause of a loss of longitudinal static margin with increasing airspeed. This loss of static margin was found to be due to aeroelastic 'washin' of the forward-swept blades and 'washout' of the aft-swept blades moving the aircraft aerodynamic center forward of the center of gravity. This phenomenon is likely to be generic to X-Wing aircraft.
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.
NASA Astrophysics Data System (ADS)
Wu, X.; Vahdati, M.; Sayma, A.; Imregun, M.
2005-03-01
This paper describes a large-scale aeroelasticity computation for an aero-engine core compressor. The computational domain includes all 17 bladerows, resulting in a mesh with over 68 million points. The Favre-averaged Navier Stokes equations are used to represent the flow in a non-linear time-accurate fashion on unstructured meshes of mixed elements. The structural model of the first two rotor bladerows is based on a standard finite element representation. The fluid mesh is moved at each time step according to the structural motion so that changes in blade aerodynamic damping and flow unsteadiness can be accommodated automatically. An efficient domain decomposition technique, where special care was taken to balance the memory requirement across processors, was developed as part of the work. The calculation was conducted in parallel mode on 128 CPUs of an SGI Origin 3000. Ten vibration cycles were obtained using over 2.2 CPU years, though the elapsed time was a week only. Steady-state flow measurements and predictions were found to be in good agreement. A comparison of the averaged unsteady flow and the steady-state flow revealed some discrepancies. It was concluded that, in due course, the methodology would be adopted by industry to perform routine numerical simulations of the unsteady flow through entire compressor assemblies with vibrating blades not only to minimise engine and rig tests but also to improve performance predictions.
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.
Aeroelastic Deflection of NURBS Geometry
NASA Technical Reports Server (NTRS)
Samareh, Jamshid A.
1998-01-01
The purpose of this paper is to present an algorithm for using NonUniform Rational B-Spline (NURBS) representation in an aeroelastic loop. The algorithm is based on creating a least-squares NURBS surface representing the aeroelastic defection. The resulting NURBS surfaces are used to update either the original Computer- Aided Design (CAD) model, Computational Structural Mechanics (CSM) grid or the Computational Fluid Dynamics (CFD) grid. Results are presented for a generic High-Speed Civil Transport (HSCT).
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.
Development and Testing of Control Laws for the Active Aeroelastic Wing Program
NASA Technical Reports Server (NTRS)
Dibley, Ryan P.; Allen, Michael J.; Clarke, Robert; Gera, Joseph; Hodgkinson, John
2005-01-01
The Active Aeroelastic Wing research program was a joint program between the U.S. Air Force Research Laboratory and NASA established to investigate the characteristics of an aeroelastic wing and the technique of using wing twist for roll control. The flight test program employed the use of an F/A-18 aircraft modified by reducing the wing torsional stiffness and adding a custom research flight control system. The research flight control system was optimized to maximize roll rate using only wing surfaces to twist the wing while simultaneously maintaining design load limits, stability margins, and handling qualities. NASA Dryden Flight Research Center developed control laws using the software design tool called CONDUIT, which employs a multi-objective function optimization to tune selected control system design parameters. Modifications were made to the Active Aeroelastic Wing implementation in this new software design tool to incorporate the NASA Dryden Flight Research Center nonlinear F/A-18 simulation for time history analysis. This paper describes the design process, including how the control law requirements were incorporated into constraints for the optimization of this specific software design tool. Predicted performance is also compared to results from flight.
TURBO-AE: An Aeroelastic Code for Propulsion Applications
NASA Technical Reports Server (NTRS)
Bakhle, Milind A.
1997-01-01
NASA's Advanced Subsonic Technology (AST) program is developing new technologies to increase the fuel efficiency of commercial aircraft engines, improve the safety of engine operation, and reduce engine emissions and noise. With the development of new designs for ducted fans, compressors, and turbines to achieve these goals, a basic aeroelastic requirement is that there should be no flutter or high resonant blade stresses in the operating regime. To verify the aeroelastic soundness of these designs, we need an accurate prediction and analysis code. Such a two-dimensional viscous propulsion aeroelastic code, named TURBO-AE, is being developed at the NASA Lewis Research Center. The TURBO-AE aeroelastic code is based on a three-dimensional unsteady aerodynamic Euler/Navier-Stokes turbomachinery code TURBO, developed under a grant from NASA Lewis. TURBO-AE can model viscous flow effects that play an important role in certain aeroelastic problems, such as flutter with flow separation (or stall flutter) and flutter in the presence of shock and boundary-layer interaction. The structural dynamics representation of the blade in the TURBO-AE code is based on a normal mode representation. A finite element analysis code, such as NASTRAN, is used to calculate in-vacuum vibration modes and the associated natural frequency. A work-per-cycle approach is used to determine aeroelastic (flutter) stability. With this approach, the motion of the blade is prescribed to be a harmonic vibration in a specified in vacuum normal mode. The aerodynamic forces acting on the vibrating blade and the work done by these forces on the vibrating blade during a cycle of vibration are calculated. If positive work is being done on the blade by the aerodynamic forces, the blade is dynamically unstable, since it will extract energy from the flow, leading to an increase in the amplitude of the blade's oscillation. Initial calculations have been done for a configuration representative of the Energy
Finite state aeroelastic model for use in rotor design optimization
NASA Technical Reports Server (NTRS)
He, Chengjian; Peters, David A.
1993-01-01
In this article, a rotor aeroelastic model based on a newly developed finite state dynamic wake, coupled with blade finite element analysis, is described. The analysis is intended for application in rotor blade design optimization. A coupled simultaneous system of differential equations combining blade structural dynamics and aerodynamics is established in a formulation well-suited for design sensitivity computation. Each blade is assumed to be an elastic beam undergoing flap bending, lead-lag bending, elastic twist, and axial deflections. Aerodynamic loads are computed from unsteady blade element theory where the rotor three-dimensional unsteady wake is described by a generalized dynamic wake model. Correlation of results obtained from the analysis with flight test data is provided to assess model accuracy.
Wing Weight Optimization Under Aeroelastic Loads Subject to Stress Constraints
NASA Technical Reports Server (NTRS)
Kapania, Rakesh K.; Issac, J.; Macmurdy, D.; Guruswamy, Guru P.
1997-01-01
A minimum weight optimization of the wing under aeroelastic loads subject to stress constraints is carried out. The loads for the optimization are based on aeroelastic trim. The design variables are the thickness of the wing skins and planform variables. The composite plate structural model incorporates first-order shear deformation theory, the wing deflections are expressed using Chebyshev polynomials and a Rayleigh-Ritz procedure is adopted for the structural formulation. The aerodynamic pressures provided by the aerodynamic code at a discrete number of grid points is represented as a bilinear distribution on the composite plate code to solve for the deflections and stresses in the wing. The lifting-surface aerodynamic code FAST is presently being used to generate the pressure distribution over the wing. The envisioned ENSAERO/Plate is an aeroelastic analysis code which combines ENSAERO version 3.0 (for analysis of wing-body configurations) with the composite plate code.
Development of an Aeroelastic Code Based on an Euler/Navier-Stokes Aerodynamic Solver
NASA Technical Reports Server (NTRS)
Bakhle, Milind A.; Srivastava, Rakesh; Keith, Theo G., Jr.; Stefko, George L.; Janus, Mark J.
1996-01-01
This paper describes the development of an aeroelastic code (TURBO-AE) based on an Euler/Navier-Stokes unsteady aerodynamic analysis. A brief review of the relevant research in the area of propulsion aeroelasticity is presented. The paper briefly describes the original Euler/Navier-Stokes code (TURBO) and then details the development of the aeroelastic extensions. The aeroelastic formulation is described. The modeling of the dynamics of the blade using a modal approach is detailed, along with the grid deformation approach used to model the elastic deformation of the blade. The work-per-cycle approach used to evaluate aeroelastic stability is described. Representative results used to verify the code are presented. The paper concludes with an evaluation of the development thus far, and some plans for further development and validation of the TURBO-AE code.
NASA Technical Reports Server (NTRS)
Finaish, Fathi
1992-01-01
A few years ago the Structural Dynamics Division at LaRC started ambitious experimental research efforts known as the Benchmark Models Program. The primary objective of this program was to provide experimental data that may serve as a calibration source for computational fluid dynamics (CFD) efforts that deal with aeroelastic unsteady flow configurations. It also focuses on the understanding of complex flow phenomenon associated with unsteady flow developments. The overall plan for the program has been described by Bennett, including a presentation of initial test results of flutter of a rigid wing mounted on flexible supports. An example of a test model employed to measure the dynamic response along with corresponding pressure distributions is shown. This model incorporates eighty pressure transducers distributed along two spanwise stations. In addition, the models are equipped with four accelerometers and two strain gages. The data handling system for the Benchmark Model Program is under development. Several interactive computer routines designed for the user interface, dynamic memory allocation, unsteady flow measurements data extraction, and further data processing were developed. To present a few examples of measured data, the unsteady pressure distributions and the wing model dynamic response were plotted.
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.
Simplified aeroelastic modeling of horizontal axis wind turbines
NASA Astrophysics Data System (ADS)
Wendell, J. H.
1982-09-01
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.
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.
Recent rotorcraft aeroelastic testing in the Langley Transonic Dynamics Tunnel
NASA Technical Reports Server (NTRS)
Yeager, William T., Jr.; Mirick, Paul H.; Wilbur, Matthew L.; Singleton, Jeffrey D.; Wilkie, W. K.; Hamouda, M.-N. H.
1991-01-01
Wind-tunnel testing of a properly scaled aeroelastic model helicopter rotor is considered a necessary phase in the design and development of new rotor systems. For this reason, extensive testing of aeroelastically scaled model rotors is done in the Transonic Dynamics Tunnel (TDT) located at the Langley Research Center. A unique capability of this facility, which enables proper dynamic scaling, is the use of diflourodichloromethane, or Refrigerant-12 (R-12) as a test medium. The paper presents a description of the TDT and a discussion of the benefits of using R-12 as a test medium. A description of the system used to conduct model tests is provided and examples of recent rotor tests are cited to illustrate the types of aeroelastic model rotor tests conducted in the TDT.
NASA Technical Reports Server (NTRS)
Reed, W. H., III
1981-01-01
In 1955, work was started on the conversion of a subsonic wind tunnel to a 16-foot transonic tunnel with Freon-12 or air as the test medium. The new facility, designated the Transonic Dynamics Tunnel (TDT), became fully operational in 1960. A description is presented of aeroelastic testing and research performed in the TDT since 1960. It is pointed out that wind-tunnel tests of aeroelastic models require specialized experimental techniques seldom found in other types of wind-tunnel studies. Attention is given to model mount systems, launch vehicle models, aircraft models, aircraft buffet, gust response, stability derivative measurements, and subcritical testing techniques. Aspects of vehicle development testing are considered along with aeroelastic 'fixes', aeroelastic 'surprises', approaches for controlling aeroelastic effects, and unsteady pressure measurements.
Advance finite element modeling of rotor blade aeroelasticity
NASA Technical Reports Server (NTRS)
Straub, F. K.; Sangha, K. B.; Panda, B.
1994-01-01
An advanced beam finite element has been developed for modeling rotor blade dynamics and aeroelasticity. This element is part of the Element Library of the Second Generation Comprehensive Helicopter Analysis System (2GCHAS). The element allows modeling of arbitrary rotor systems, including bearingless rotors. It accounts for moderately large elastic deflections, anisotropic properties, large frame motion for maneuver simulation, and allows for variable order shape functions. The effects of gravity, mechanically applied and aerodynamic loads are included. All kinematic quantities required to compute airloads are provided. In this paper, the fundamental assumptions and derivation of the element matrices are presented. Numerical results are shown to verify the formulation and illustrate several features of the element.
Impact of Parallel Computing on Large Scale Aeroelastic Computations
NASA Technical Reports Server (NTRS)
Guruswamy, Guru P.; Kwak, Dochan (Technical Monitor)
2000-01-01
Aeroelasticity is computationally one of the most intensive fields in aerospace engineering. Though over the last three decades the computational speed of supercomputers have substantially increased, they are still inadequate for large scale aeroelastic computations using high fidelity flow and structural equations. In addition to reaching a saturation in computational speed because of changes in economics, computer manufactures are stopping the manufacturing of mainframe type supercomputers. This has led computational aeroelasticians to face the gigantic task of finding alternate approaches for fulfilling their needs. The alternate path to over come speed and availability limitations of mainframe type supercomputers is to use parallel computers. During this decade several different architectures have evolved. In FY92 the US Government started the High Performance Computing and Communication (HPCC) program. As a participant in this program NASA developed several parallel computational tools for aeroelastic applications. This talk describes the impact of those application tools on high fidelity based multidisciplinary analysis.
NASA Technical Reports Server (NTRS)
Wang, James M.; Chopra, Inderjit; Samak, D. K.; Green, Michael; Graham, Todd
1989-01-01
The aeroelastic stability of a shaft-fixed, 1/8th Froude scaled bearingless rotor was investigated in a series of wind tunnel experiments simulating a wide range of operating conditions. A finite element formulation was used to perform a parallel theoretical analysis, with the goal of determining whether a bearingless rotor system could be made aeroelastically stable without the incorporation of auxilliary dampers. A quick estimate of lag mode damping was provided by a refined moving-block analysis implemented in real time which predicted similar damping values. Model rotor and blade properties were also determined, and these properties were used as inputs for a newly refined bearingless rotor analysis. Predicted results were compared with experimental results in hover and forward flight. Results indicated that soft pitch link stiffness increases pitch-lag coupling and stabilizes lag mode stability in hover and at low advance ratios, but destabilizes at higher advance ratios.
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
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.
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.
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.
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 Model Structure Computation for Envelope Expansion
NASA Technical Reports Server (NTRS)
Kukreja, Sunil L.
2007-01-01
Structure detection is a procedure for selecting a subset of candidate terms, from a full model description, that best describes the observed output. This is a necessary procedure to compute an efficient system description which may afford greater insight into the functionality of the system or a simpler controller design. Structure computation as a tool for black-box modelling may be of critical importance in the development of robust, parsimonious models for the flight-test community. Moreover, this approach may lead to efficient strategies for rapid envelope expansion which may save significant development time and costs. In this study, a least absolute shrinkage and selection operator (LASSO) technique is investigated for computing efficient model descriptions of nonlinear aeroelastic systems. The LASSO minimises the residual sum of squares by the addition of an l(sub 1) penalty term on the parameter vector of the traditional 2 minimisation problem. Its use for structure detection is a natural extension of this constrained minimisation approach to pseudolinear regression problems which produces some model parameters that are exactly zero and, therefore, yields a parsimonious system description. Applicability of this technique for model structure computation for the F/A-18 Active Aeroelastic Wing using flight test data is shown for several flight conditions (Mach numbers) by identifying a parsimonious system description with a high percent fit for cross-validated data.
Aeroelastic Model Structure Computation for Envelope Expansion
NASA Technical Reports Server (NTRS)
Kukreja, Sunil L.
2007-01-01
Structure detection is a procedure for selecting a subset of candidate terms, from a full model description, that best describes the observed output. This is a necessary procedure to compute an efficient system description which may afford greater insight into the functionality of the system or a simpler controller design. Structure computation as a tool for black-box modeling may be of critical importance in the development of robust, parsimonious models for the flight-test community. Moreover, this approach may lead to efficient strategies for rapid envelope expansion that may save significant development time and costs. In this study, a least absolute shrinkage and selection operator (LASSO) technique is investigated for computing efficient model descriptions of non-linear aeroelastic systems. The LASSO minimises the residual sum of squares with the addition of an l(Sub 1) penalty term on the parameter vector of the traditional l(sub 2) minimisation problem. Its use for structure detection is a natural extension of this constrained minimisation approach to pseudo-linear regression problems which produces some model parameters that are exactly zero and, therefore, yields a parsimonious system description. Applicability of this technique for model structure computation for the F/A-18 (McDonnell Douglas, now The Boeing Company, Chicago, Illinois) Active Aeroelastic Wing project using flight test data is shown for several flight conditions (Mach numbers) by identifying a parsimonious system description with a high percent fit for cross-validated data.
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.
AEROELASTIC SIMULATION TOOL FOR INFLATABLE BALLUTE AEROCAPTURE
NASA Technical Reports Server (NTRS)
Liever, P. A.; Sheta, E. F.; Habchi, S. D.
2006-01-01
A multidisciplinary analysis tool is under development for predicting the impact of aeroelastic effects on the functionality of inflatable ballute aeroassist vehicles in both the continuum and rarefied flow regimes. High-fidelity modules for continuum and rarefied aerodynamics, structural dynamics, heat transfer, and computational grid deformation are coupled in an integrated multi-physics, multi-disciplinary computing environment. This flexible and extensible approach allows the integration of state-of-the-art, stand-alone NASA and industry leading continuum and rarefied flow solvers and structural analysis codes into a computing environment in which the modules can run concurrently with synchronized data transfer. Coupled fluid-structure continuum flow demonstrations were conducted on a clamped ballute configuration. The feasibility of implementing a DSMC flow solver in the simulation framework was demonstrated, and loosely coupled rarefied flow aeroelastic demonstrations were performed. A NASA and industry technology survey identified CFD, DSMC and structural analysis codes capable of modeling non-linear shape and material response of thin-film inflated aeroshells. The simulation technology will find direct and immediate applications with NASA and industry in ongoing aerocapture technology development programs.
NASA Technical Reports Server (NTRS)
Bartels, Robert E.; Scott, Robert C.; Allen, Timothy J.; Sexton, Bradley W.
2015-01-01
Considerable attention has been given in recent years to the design of highly flexible aircraft. The results of numerous studies demonstrate the significant performance benefits of strut-braced wing (SBW) and trussbraced wing (TBW) configurations. Critical aspects of the TBW configuration are its larger aspect ratio, wing span and thinner wings. These aspects increase the importance of considering fluid/structure and control system coupling. This paper presents high-fidelity Navier-Stokes simulations of the dynamic response of the flexible Boeing Subsonic Ultra Green Aircraft Research (SUGAR) truss-braced wing wind-tunnel model. The latest version of the SUGAR TBW finite element model (FEM), v.20, is used in the present simulations. Limit cycle oscillations (LCOs) of the TBW wing/strut/nacelle are simulated at angle-of-attack (AoA) values of -1, 0 and +1 degree. The modal data derived from nonlinear static aeroelastic MSC.Nastran solutions are used at AoAs of -1 and +1 degrees. The LCO amplitude is observed to be dependent on AoA. LCO amplitudes at -1 degree are larger than those at +1 degree. The LCO amplitude at zero degrees is larger than either -1 or +1 degrees. These results correlate well with both wind-tunnel data and the behavior observed in previous studies using linear aerodynamics. The LCO onset at zero degrees AoA has also been computed using unloaded v.20 FEM modes. While the v.20 model increases the dynamic pressure at which LCO onset is observed, it is found that the LCO onset at and above Mach 0.82 is much different than that produced by an earlier version of the FEM, v. 19.
Status of NASA full-scale engine aeroelasticity research
NASA Technical Reports Server (NTRS)
Lubomski, J. F.
1980-01-01
Data relevant to several types of aeroelastic instabilities were obtained using several types of turbojet and turbofan engines. In particular, data relative to separated flow (stall) flutter, choke flutter, and system mode instabilities are presented. The unique characteristics of these instabilities are discussed, and a number of correlations are presented that help identify the nature of the phenomena.
Status of NASA full-scale engine aeroelasticity research
NASA Technical Reports Server (NTRS)
Lubomski, J. F.
1980-01-01
The paper presents data relevant to several types of aeroelastic instabilities which have been obtained using several types of turbojet and turbofan engines. Special attention is given to data relative to separated flow (stall) flutter, choke flutter, and system mode instabilities. The discussion covers the characteristics of these instabilities, and a number of correlations are presented that help identify the nature of the phenomena.
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.
NASA Astrophysics Data System (ADS)
Khouli, F.
An aeroelastic phenomenon, known as blade sailing, encountered during maritime operation of helicopters is identified as being a factor that limits the tactical flexibility of helicopter operation in some sea conditions. The hazards associated with this phenomenon and its complexity, owing to the number of factors contributing to its occurrence, led previous investigators to conclude that advanced and validated simulation tools are best suited to investigate it. A research gap is identified in terms of scaled experimental investigation of this phenomenon and practical engineering solutions to alleviate its negative impact on maritime helicopter operation. The feasibility of a proposed strategy to alleviate it required addressing a gap in modelling thin-walled composite active beams/rotor blades. The modelling is performed by extending a mathematically-consistent and asymptotic reduction strategy of the 3-D elastic problem to account for embedded active materials. The derived active cross-sectional theory is validated using 2-D finite element results for closed and open cross-sections. The geometrically-exact intrinsic formulation of active maritime rotor systems is demonstrated to yield compact and symbolic governing equations. The intrinsic feature is shown to allow a classical and proven solution scheme to be successfully applied to obtain time history solutions. A Froude-scaled experimental rotor was designed, built, and tested in a scaled ship airwake environment and representative ship motion. Based on experimental and simulations data, conclusions are drawn regarding the influence of the maritime operation environment and the rotor operation parameters on the blade sailing phenomenon. The experimental data is also used to successfully validate the developed simulation tools. The feasibility of an open-loop control strategy based on the integral active twist concept to counter blade sailing is established in a Mach-scaled maritime operation environment
Prediction of aeroelastic response of a model X-wing rotor
NASA Technical Reports Server (NTRS)
Dopher, Robert; Duh, James E.
1987-01-01
The rotorcraft dynamics analysis was used to predict the aeroelastic responses of a representative X-wing model with a 10 ft diameter rotor. The aeroelastic methodology used and the tests and assumptions involved are reviewed. Results are reported on the findings concerning control power and higher harmonic control in hover, transition flight, vibratory loads at forward speed, and responses in conversion. It is concluded that the analysis can give satisfactory predictions of X-wing behavior.
Synchronized charge extraction for aeroelastic energy harvesting
NASA Astrophysics Data System (ADS)
Zhao, Liya; Tang, Lihua; Wu, Hao; Yang, Yaowen
2014-03-01
Aeroelastic instabilities have been frequently exploited for energy harvesting purpose to power standalone electronic systems, such as wireless sensors. Meanwhile, various energy harvesting interface circuits, such as synchronized charge extraction (SCE) and synchronized switching harvesting on inductor (SSHI), have been widely pursued in the literature for efficiency enhancement of energy harvesting from existing base vibrations. These interfaces, however, have not been applied for aeroelastic energy harvesting. This paper investigates the feasibility of the SCE interface in galloping-based piezoelectric energy harvesting, with a focus on its benefit for performance improvement and influence on the galloping dynamics in different electromechanical coupling regimes. A galloping-based piezoelectric energy harvester (GPEH) is prototyped with an aluminum cantilever bonded with a piezoelectric sheet. Wind tunnel test is conducted with a simple electrical interface composed of a resistive load. Circuit simulation is performed with equivalent circuit representation of the GPEH system and confirmed by experimental results. Consequently, a self-powered SCE interface is implemented with the capability of self peak-detecting and switching. Circuit simulation for various electromechanical coupling cases shows that the harvested power with SCE interface for GPEH is independent of the electrical load, similar to that for a vibration-based piezoelectric energy harvester (VPEH). The SCE interface outperforms the standard interface if the electromechanical coupling is weak, and requires much less piezoelectric material to achieve the maximum power output. Moreover, influence of electromechanical coupling on the dynamics of GPEH with SCE is found sensitive to the wind speed.
Data acquisition for aeroelastic testing at the NASA Langley Transonic Dynamics Facility
NASA Technical Reports Server (NTRS)
Doggett, Robert V., Jr.; Rosser, David C., Jr.; Bryant, Charles S.
1993-01-01
Aeroelasticity is defined and different types of aeroelastic tests are described. The instrumentation and data acquisition requirements dictated by the time varying characteristics of aeroelastic phenomena are pointed out. Instrumentation needs for various types of tests are discussed. The hardware components and operational features of the real-time, digital-computer-controlled data acquisition system used to support testing at the Langley Transonic Dynamics Facility, which includes a major wind-tunnel, a helicopter-model hover facility, and two model setup and calibration laboratories are described.
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.
Aeroelastic Modeling of a Nozzle Startup Transient
NASA Technical Reports Server (NTRS)
Wang, Ten-See; Zhao, Xiang; Zhang, Sijun; Chen, Yen-Sen
2014-01-01
Lateral nozzle forces are known to cause severe structural damage to any new rocket engine in development during test. While three-dimensional, transient, turbulent, chemically reacting computational fluid dynamics methodology has been demonstrated to capture major side load physics with rigid nozzles, hot-fire tests often show nozzle structure deformation during major side load events, leading to structural damages if structural strengthening measures were not taken. The modeling picture is incomplete without the capability to address the two-way responses between the structure and fluid. The objective of this study is to develop a tightly coupled aeroelastic modeling algorithm by implementing the necessary structural dynamics component into an anchored computational fluid dynamics methodology. The computational fluid dynamics component is based on an unstructured-grid, pressure-based computational fluid dynamics formulation, while the computational structural dynamics component is developed under the framework of modal analysis. Transient aeroelastic nozzle startup analyses at sea level were performed, and the computed transient nozzle fluid-structure interaction physics presented,
Rotary-wing aeroelasticity with application to VTOL vehicles
NASA Technical Reports Server (NTRS)
Friedmann, Peretz P.
1990-01-01
This survey presents an assessment of the state of the art in rotary-wing aeroelasticity as applied to conventional helicopters as well as other VTOL vehicles such as tilting prop-rotors, the X-wing and a hybrid heavy lift vehicle. The objective is to enable the reader to develop an awareness of what has been accomplished, what remains to be done, and where to find more comprehensive treatments of the various topics discussed. The main topics discussed are: (1) structural modeling; (2) unsteady aerodynamic modeling; (3) formulation of the equations of motion and their solutions; (4) illustrative results for isolated blades in hover and forward flight; (5) illustrative results for coupled rotor/fuselage problems; (6) active control of aeromechanical and aeroelastic problems; (7) active controls for vibration reduction; (8) structural optimization with aeroelastic constraints; (9) gust response analysis of rotors; and (10) aeroelastic problems in special VTOL vehicles. These topics are reviewed with different levels of detail and some useful observation on potentially rewarding areas of future research are made.
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.
Effect of follower forces on aeroelastic stability of flexible structures
NASA Astrophysics Data System (ADS)
Chae, Seungmook
Missile bodies and wings are typical examples of structures that can be represented by beam models. Such structures, loaded by follower forces along with aerodynamics, exhibit the vehicle's aeroelastic instabilities. The current research integrates a nonlinear beam dynamics and unsteady aerodynamics to conduct aeroelastic studies of missile bodies and wings subjected to follower forces. The structural formulations are based on a geometrically-exact, mixed finite element method. Slender-body theory and thin-airfoil theory are used for the missile aerodynamics, and two-dimensional finite-state unsteady aerodynamics is used for wing aerodynamics. The aeroelastic analyses are performed using time-marching scheme for the missile body stability, and eigenvalue analysis for the wing flutter, respectively. Results from the time-marching formulation agree with published results for dynamic stability and show the development of limit cycle oscillations for disturbed flight near and above the critical thrust. Parametric studies of the aeroelastic behavior of specific flexible missile configurations are presented, including effects of flexibility on stability, limit-cycle amplitudes, and missile loads. The results do yield a significant interaction between the thrust, which is a follower force, and the aeroelastic stability. Parametric studies based on the eigenvalue analysis for the wing flutter, show that the predicted stability boundaries are very sensitive to the ratio of bending stiffness to torsional stiffness. The effect of thrust can be either stabilizing or destabilizing, depending on the value of this parameter. An assessment whether or not the magnitude of thrust needed to influence the flutter speed is practical is made for one configuration. The flutter speed is shown to change by 11% for this specific wing configuration.
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
NASA Technical Reports Server (NTRS)
Schuster, David M.
1993-01-01
An aeroelastic analysis method, based on three-dimensional Navier-Stokes equation aerodynamics, has been applied to improve the performance of fighter wings operating at sustained maneuver flight conditions. The scheme reduces the trimmed pressure drag of wings performing high-g maneuvers through a simultaneous application of control surface deflection and aeroelastic twist. The aerodynamic and structural interactions are decoupled by assuming an aeroelastic twist mode shape and optimizing the aerodynamic performance based on this aeroelastic mode. The wing structural stiffness properties are then determined through an inverse scheme based on the aerodynamic loads and desired twist at the maneuver flight condition. The decoupled technique is verified by performing a fully coupled aeroelastic analysis using the maneuver flight conditions and the optimized structural stiffness distributions.
Linear and nonlinear dynamic analysis of redundant load path bearingless rotor systems
NASA Technical Reports Server (NTRS)
Murthy, V. R.; Shultz, Louis A.
1994-01-01
The goal of this research is to develop the transfer matrix method to treat nonlinear autonomous boundary value problems with multiple branches. The application is the complete nonlinear aeroelastic analysis of multiple-branched rotor blades. Once the development is complete, it can be incorporated into the existing transfer matrix analyses. There are several difficulties to be overcome in reaching this objective. The conventional transfer matrix method is limited in that it is applicable only to linear branch chain-like structures, but consideration of multiple branch modeling is important for bearingless rotors. Also, hingeless and bearingless rotor blade dynamic characteristics (particularly their aeroelasticity problems) are inherently nonlinear. The nonlinear equations of motion and the multiple-branched boundary value problem are treated together using a direct transfer matrix method. First, the formulation is applied to a nonlinear single-branch blade to validate the nonlinear portion of the formulation. The nonlinear system of equations is iteratively solved using a form of Newton-Raphson iteration scheme developed for differential equations of continuous systems. The formulation is then applied to determine the nonlinear steady state trim and aeroelastic stability of a rotor blade in hover with two branches at the root. A comprehensive computer program is developed and is used to obtain numerical results for the (1) free vibration, (2) nonlinearly deformed steady state, (3) free vibration about the nonlinearly deformed steady state, and (4) aeroelastic stability tasks. The numerical results obtained by the present method agree with results from other methods.
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.
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
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
Aeroelastic behavior of composite helicopter rotor blades with advanced geometry tips
Friedmann, P.P.; Yuan, K.A.
1995-12-31
A new structural and aeroelastic model capable of representing the aeroelastic stability and response of composite helicopter rotor blades with advanced geometry tips is presented. Where it is understood that advanced geometry tips are blade tips having sweep, anhedral and taper in the outboard 10% segment of the blade. The blade is modeled by beam finite elements. A single element is used to represent the swept tip. The nonlinear equations of motion are derived using the Hamilton`s principle and are based on moderate deflection theory. Thus, the nonlinearities are of the geometric type. The important structural blade attributes captured by the model are arbitrary cross-sectional shape, general anisotropic material behavior, transverse shear and out-of-plane warping. The aerodynamic loads are based on quasi-steady Greenberg theory with reverse flow effects, using an implicit formulation. The nonlinear aeroelastic response of the blade is obtained from a fully coupled propulsive trim/aeroelastic response analysis. Aeroelastic stability is obtained from linearizing the equations of motion about the steady state response of the blade and using Floquet theory. Numerical results for the aeroelastic stability and response of a hingeless composite blade with two cell type cross section are presented, together with vibratory hub shears and moments. The influence of ply orientation and tip sweep is clearly illustrated by the results.
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.