Ongoing Fixed Wing Research within the NASA Langley Aeroelasticity Branch
NASA Technical Reports Server (NTRS)
Bartels, Robert; Chwalowski, Pawel; Funk, Christie; Heeg, Jennifer; Hur, Jiyoung; Sanetrik, Mark; Scott, Robert; Silva, Walter; Stanford, Bret; Wiseman, Carol
2015-01-01
The NASA Langley Aeroelasticity Branch is involved in a number of research programs related to fixed wing aeroelasticity and aeroservoelasticity. These ongoing efforts are summarized here, and include aeroelastic tailoring of subsonic transport wing structures, experimental and numerical assessment of truss-braced wing flutter and limit cycle oscillations, and numerical modeling of high speed civil transport configurations. Efforts devoted to verification, validation, and uncertainty quantification of aeroelastic physics in a workshop setting are also discussed. The feasibility of certain future civil transport configurations will depend on the ability to understand and control complex aeroelastic phenomena, a goal that the Aeroelasticity Branch is well-positioned to contribute through these programs.
Internal Structural Design of the Common Research Model Wing Box for Aeroelastic Tailoring
NASA Technical Reports Server (NTRS)
Jutte, Christine V.; Stanford, Bret K.; Wieseman, Carol D.
2015-01-01
This work explores the use of alternative internal structural designs within a full-scale wing box structure for aeroelastic tailoring, with a focus on curvilinear spars, ribs, and stringers. The baseline wing model is a fully-populated, cantilevered wing box structure of the Common Research Model (CRM). Metrics of interest include the wing weight, the onset of dynamic flutter, and the static aeroelastic stresses. Twelve parametric studies alter the number of internal structural members along with their location, orientation, and curvature. Additional evaluation metrics are considered to identify design trends that lead to lighter-weight, aeroelastically stable wing designs. The best designs of the individual studies are compared and discussed, with a focus on weight reduction and flutter resistance. The largest weight reductions were obtained by removing the inner spar, and performance was maintained by shifting stringers forward and/or using curvilinear ribs: 5.6% weight reduction, a 13.9% improvement in flutter speed, but a 3.0% increase in stress levels. Flutter resistance was also maintained using straight-rotated ribs although the design had a 4.2% lower flutter speed than the curved ribs of similar weight and stress levels were higher. For some configurations, the differences between curved and straight ribs were smaller, which provides motivation for future optimization-based studies to fully exploit the trade-offs.
Steady pressure measurements on an Aeroelastic Research Wing (ARW-2)
NASA Technical Reports Server (NTRS)
Sandford, Maynard C.; Seidel, David A.; Eckstrom, Clinton V.
1994-01-01
Transonic steady and unsteady pressure tests have been conducted in the Langley transonic dynamics tunnel on a large elastic wing known as the DAST ARW-2. The wing has a supercritical airfoil, an aspect ratio of 10.3, a leading-edge sweep back angle of 28.8 degrees, and two inboard and one outboard trailing-edge control surfaces. Only the outboard control surface was deflected to generate steady and unsteady flow over the wing during this study. Only the steady surface pressure, control-surface hinge moment, wing-tip deflection, and wing-root bending moment measurements are presented. The results from this elastic wing test are in tabulated form to assist in calibrating advanced computational fluid dynamics (CFD) algorithms.
NASA Technical Reports Server (NTRS)
Mourey, D. J.
1979-01-01
The aspects of flight testing an aeroelastically tailored forward swept research wing on a BQM-34F drone vehicle are examined. The geometry of a forward swept wing, which is incorporated into the BQM-34F to maintain satisfactory flight performance, stability, and control is defined. A preliminary design of the aeroelastically tailored forward swept wing is presented.
Aeroelasticity Benchmark Assessment: Subsonic Fixed Wing Program
NASA Technical Reports Server (NTRS)
Florance, Jennifer P.; Chwalowski, Pawel; Wieseman, Carol D.
2010-01-01
Aeroelasticity Branch will examine other experimental efforts within the Subsonic Fixed Wing (SFW) program (such as testing of the NASA Common Research Model (CRM)) and other NASA programs and assess aeroelasticity issues and research topics.
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.
Material and Thickness Grading for Aeroelastic Tailoring of the Common Research Model Wing Box
NASA Technical Reports Server (NTRS)
Stanford, Bret K.; Jutte, Christine V.
2014-01-01
This work quantifies the potential aeroelastic benefits of tailoring a full-scale wing box structure using tailored thickness distributions, material distributions, or both simultaneously. These tailoring schemes are considered for the wing skins, the spars, and the ribs. Material grading utilizes a spatially-continuous blend of two metals: Al and Al+SiC. Thicknesses and material fraction variables are specified at the 4 corners of the wing box, and a bilinear interpolation is used to compute these parameters for the interior of the planform. Pareto fronts detailing the conflict between static aeroelastic stresses and dynamic flutter boundaries are computed with a genetic algorithm. In some cases, a true material grading is found to be superior to a single-material structure.
NASA Technical Reports Server (NTRS)
Bradley, Marty K.; Allen, Timothy J.; Droney, Christopher
2014-01-01
This Test Report summarizes the Truss Braced Wing (TBW) Aeroelastic Test (Task 3.1) work accomplished by the Boeing Subsonic Ultra Green Aircraft Research (SUGAR) team, which includes the time period of February 2012 through June 2014. The team consisted of Boeing Research and Technology, Boeing Commercial Airplanes, Virginia Tech, and NextGen Aeronautics. The model was fabricated by NextGen Aeronautics and designed to meet dynamically scaled requirements from the sized full scale TBW FEM. The test of the dynamically scaled SUGAR TBW half model was broken up into open loop testing in December 2013 and closed loop testing from January 2014 to April 2014. Results showed the flutter mechanism to primarily be a coalescence of 2nd bending mode and 1st torsion mode around 10 Hz, as predicted by analysis. Results also showed significant change in flutter speed as angle of attack was varied. This nonlinear behavior can be explained by including preload and large displacement changes to the structural stiffness and mass matrices in the flutter analysis. Control laws derived from both test system ID and FEM19 state space models were successful in suppressing flutter. The control laws were robust and suppressed flutter for a variety of Mach, dynamic pressures, and angle of attacks investigated.
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.
Wing-Body Aeroelasticity on Parallel Computers
NASA Technical Reports Server (NTRS)
Guruswamy, Guru P.; Byun, Chansup
1996-01-01
This article presents a procedure for computing the aeroelasticity of wing-body configurations on multiple-instruction, multiple-data parallel computers. In this procedure, fluids are modeled using Euler equations discretized by a finite difference method, and structures are modeled using finite element equations. The procedure is designed in such a way that each discipline can be developed and maintained independently by using a domain decomposition approach. A parallel integration scheme is used to compute aeroelastic responses by solving the coupled fluid and structural equations concurrently while keeping modularity of each discipline. The present procedure is validated by computing the aeroelastic response of a wing and comparing with experiment. Aeroelastic computations are illustrated for a high speed civil transport type wing-body configuration.
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.
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
Computational, unsteady transonic aerodynamics and aeroelasticity about airfoils and wings
NASA Technical Reports Server (NTRS)
Goorjian, Peter M.; Guruswamy, Guru P.
1987-01-01
Research in the area of computational, unsteady transonic flows about airfoils and wings, including aeroelastic effects is reviewed. In the last decade, there have been extensive developments in computational methods in response to the need for computer codes with which to study fundamental aerodynamic and aeroelastic problems in the critical transonic regime. For example, large commercial aircraft cruise most effectively in the transonic flight regime and computational fluid dynamics (CDF) provides a new tool, which can be used in combination with test facilities to reduce the costs, time, and risks of aircraft development.
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.
Loads calibrations of strain gage bridges on the DAST project Aeroelastic Research Wing (ARW-1)
NASA Technical Reports Server (NTRS)
Eckstrom, C. V.
1980-01-01
The details of and results from the procedure used to calibrate strain gage bridges for measurement of wing structural loads for the DAST project ARW-1 wing are presented. Results are in the form of loads equations and comparison of computed loads vs. actual loads for two simulated flight loading conditions.
Loads calibrations of strain gage bridges on the DAST project Aeroelastic Research Wing (ARW-2)
NASA Technical Reports Server (NTRS)
Eckstrom, C. V.
1986-01-01
Results from and details of the procedure used to calibrate strain gage bridges for measurements of wing structural loads, shear (V), bending moment (M), and torque (T), at three semispan stations on both the left and right semispans of the ARW-2 wing are presented. The ARW-2 wing has a reference area of 35 square feet, a span of 19 feet, an aspect ratio of 10.3, a midchord line sweepback angle of 25 degrees, and a taper ratio of 0.4. The ARW-2 wing was fabricated using aluminum spars and ribs covered with a fiberglass/honeycomb sandwich skin material. All strain gage bridges are mounted along with an estimate of their accuracy by means of a comparison of computed loads versus actual loads for three simulated flight conditions.
NASA Technical Reports Server (NTRS)
Murrow, H. N.
1981-01-01
Results from flight tests of the ARW-1 research wing are presented. Preliminary loads data and experiences with the active control system for flutter suppression are included along with comparative results of test and prediction for the flutter boundary of the supercritical research wing and on performance of the flutter suppression system. The status of the ARW-2 research wing is given.
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.
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
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.
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.
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.
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.
Wing Torsional Stiffness Tests of the Active Aeroelastic Wing F/A-18 Airplane
NASA Technical Reports Server (NTRS)
Lokos, William A.; Olney, Candida D.; Crawford, Natalie D.; Stauf, Rick; Reichenbach, Eric Y.
2002-01-01
The left wing of the Active Aeroelastic Wing (AAW) F/A-18 airplane has been ground-load-tested to quantify its torsional stiffness. The test has been performed at the NASA Dryden Flight Research Center in November 1996, and again in April 2001 after a wing skin modification was performed. The primary objectives of these tests were to characterize the wing behavior before the first flight, and provide a before-and-after measurement of the torsional stiffness. Two streamwise load couples have been applied. The wing skin modification is shown to have more torsional flexibility than the original configuration has. Additionally, structural hysteresis is shown to be reduced by the skin modification. Data comparisons show good repeatability between the tests.
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.
Computed Aeroelastic Motions Of Wings In Transonic Flows
NASA Technical Reports Server (NTRS)
Guruswamy, Guru P.; Obayashi, Shigeru
1995-01-01
Report describes computational simulations of aeroelastic motions of delta and swept wings in transonic flows. Study directed toward understanding aerodynamic behavior and enhancing maneuverability of fighter airplanes equipped with such wings. Also has implications for gas pumps and turbines, in which flows near tips of vanes and blades reach supersonic speeds.
NASA Technical Reports Server (NTRS)
Jutte, Christine; Stanford, Bret K.
2014-01-01
This paper provides a brief overview of the state-of-the-art for aeroelastic tailoring of subsonic transport aircraft and offers additional resources on related research efforts. Emphasis is placed on aircraft having straight or aft swept wings. The literature covers computational synthesis tools developed for aeroelastic tailoring and numerous design studies focused on discovering new methods for passive aeroelastic control. Several new structural and material technologies are presented as potential enablers of aeroelastic tailoring, including selectively reinforced materials, functionally graded materials, fiber tow steered composite laminates, and various nonconventional structural designs. In addition, smart materials and structures whose properties or configurations change in response to external stimuli are presented as potential active approaches to aeroelastic tailoring.
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.
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.
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
NASA Astrophysics Data System (ADS)
Otsuka, Keisuke; Makihara, Kanjuro
2016-05-01
Morphing wings have been developed by several organizations for a variety of applications including the changing of flight ability while in the air and reducing the amount of space required to store an aircraft. One such example of morphing wings is the deployable wing that is expected to be used for Mars exploration. When designing wings, aeroelastic simulation is important to prevent the occurrence of destructive phenomena while the wing is in use. Flutter and divergence are typical issues to be addressed. However, it has been difficult to simulate the aeroelastic motion of deployable wings because of the significant differences between these deployable wings and conventional designs. The most apparent difference is the kinematic constraints of deployment, typically a hinge joint. These constraints lead not only to deformation but also to rigid body rotation. This research provides a novel method of overcoming the difficulties associated with handling these kinematic constraints. The proposed method utilizes flexible multibody dynamics and absolute nodal coordinate formulation to describe the dynamic motion of a deployable wing. This paper presents the simulation of the rigid body rotation around the kinematic constraints as induced by the aeroelasticity. The practicality of the proposed method is confirmed.
Flight Test of the F/A-18 Active Aeroelastic Wing Airplane
NASA Technical Reports Server (NTRS)
Clarke, Robert; Allen, Michael J.; Dibley, Ryan P.; Gera, Joseph; Hodgkinson, John
2005-01-01
Successful flight-testing of the Active Aeroelastic Wing airplane was completed in March 2005. This program, which started in 1996, was a joint activity sponsored by NASA, Air Force Research Laboratory, and industry contractors. The test program contained two flight test phases conducted in early 2003 and early 2005. During the first phase of flight test, aerodynamic models and load models of the wing control surfaces and wing structure were developed. Design teams built new research control laws for the Active Aeroelastic Wing airplane using these flight-validated models; and throughout the final phase of flight test, these new control laws were demonstrated. The control laws were designed to optimize strategies for moving the wing control surfaces to maximize roll rates in the transonic and supersonic flight regimes. Control surface hinge moments and wing loads were constrained to remain within hydraulic and load limits. This paper describes briefly the flight control system architecture as well as the design approach used by Active Aeroelastic Wing project engineers to develop flight control system gains. Additionally, this paper presents flight test techniques and comparison between flight test results and predictions.
The Wing-Body Aeroelastic Analyses Using the Inverse Design Method
NASA Astrophysics Data System (ADS)
Lee, Seung Jun; Im, Dong-Kyun; Lee, In; Kwon, Jang-Hyuk
Flutter phenomenon is one of the most dangerous problems in aeroelasticity. When it occurs, the aircraft structure can fail in a few second. In recent aeroelastic research, computational fluid dynamics (CFD) techniques become important means to predict the aeroelastic unstable responses accurately. Among various flow equations like Navier-Stokes, Euler, full potential and so forth, the transonic small disturbance (TSD) theory is widely recognized as one of the most efficient theories. However, the small disturbance assumption limits the applicable range of the TSD theory to the thin wings. For a missile which usually has small aspect ratio wings, the influence of body aerodynamics on the wing surface may be significant. Thus, the flutter stability including the body effect should be verified. In this research an inverse design method is used to complement the aerodynamic deficiency derived from the fuselage. MGM (modified Garabedian-McFadden) inverse design method is used to optimize the aerodynamic field of a full aircraft model. Furthermore, the present TSD aeroelastic analyses do not require the grid regeneration process. The MGM inverse design method converges faster than other conventional aerodynamic theories. Consequently, the inverse designed aeroelastic analyses show that the flutter stability has been lowered by the body effect.
NASA Technical Reports Server (NTRS)
Schuster, David M.
1993-01-01
An inverse method has been developed to compute the structural stiffness properties of wings given a specified wing loading and aeroelastic twist distribution. The method directly solves for the bending and torsional stiffness distribution of the wing using a modal representation of these properties. An aeroelastic design problem involving the use of a computational aerodynamics method to optimize the aeroelastic twist distribution of a tighter wing operating at maneuver flight conditions is used to demonstrate the application of the method. This exercise verifies the ability of the inverse scheme to accurately compute the structural stiffness distribution required to generate a specific aeroelastic twist under a specified aeroelastic load.
NASA Astrophysics Data System (ADS)
Schuster, David M.
1993-04-01
An inverse method has been developed to compute the structural stiffness properties of wings given a specified wing loading and aeroelastic twist distribution. The method directly solves for the bending and torsional stiffness distribution of the wing using a modal representation of these properties. An aeroelastic design problem involving the use of a computational aerodynamics method to optimize the aeroelastic twist distribution of a tighter wing operating at maneuver flight conditions is used to demonstrate the application of the method. This exercise verifies the ability of the inverse scheme to accurately compute the structural stiffness distribution required to generate a specific aeroelastic twist under a specified aeroelastic load.
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.
Aeroelastic and Flight Dynamics Analysis of Folding Wing Systems
NASA Astrophysics Data System (ADS)
Wang, Ivan
This dissertation explores the aeroelastic stability of a folding wing using both theoretical and experimental methods. The theoretical model is based on the existing clamped-wing aeroelastic model that uses beam theory structural dynamics and strip theory aerodynamics. A higher-fidelity theoretical model was created by adding several improvements to the existing model, namely a structural model that uses ANSYS for individual wing segment modes and an unsteady vortex lattice aerodynamic model. The comparison with the lower-fidelity model shows that the higher-fidelity model typical provides better agreement between theory and experiment, but the predicted system behavior in general does not change, reinforcing the effectiveness of the low-fidelity model for preliminary design of folding wings. The present work also conducted more detailed aeroelastic analyses of three-segment folding wings, and in particular considers the Lockheed-type configurations to understand the existence of sudden changes in predicted aeroelastic behavior with varying fold angle for certain configurations. These phenomena were observed in carefully conducted experiments, and nonlinearities---structural and geometry---were shown to suppress the phenomena. Next, new experimental models with better manufacturing tolerances are designed to be tested in the Duke University Wind Tunnel. The testing focused on various configurations of three-segment folding wings in order to obtain higher quality data. Next, the theoretical model was further improved by adding aircraft longitudinal degrees of freedom such that the aeroelastic model may predict the instabilities for the entire aircraft and not just a clamped wing. The theoretical results show that the flutter instabilities typically occur at a higher air speed due to greater frequency separation between modes for the aircraft system than a clamped wing system, but the divergence instabilities occur at a lower air speed. Lastly, additional
Aeroelastic Tailoring of Transport Wings Including Transonic Flutter Constraints
NASA Technical Reports Server (NTRS)
Stanford, Bret K.; Wieseman, Carol D.; Jutte, Christine V.
2015-01-01
Several minimum-mass optimization problems are solved to evaluate the effectiveness of a variety of novel tailoring schemes for subsonic transport wings. Aeroelastic stress and panel buckling constraints are imposed across several trimmed static maneuver loads, in addition to a transonic flutter margin constraint, captured with aerodynamic influence coefficient-based tools. Tailoring with metallic thickness variations, functionally graded materials, balanced or unbalanced composite laminates, curvilinear tow steering, and distributed trailing edge control effectors are all found to provide reductions in structural wing mass with varying degrees of success. The question as to whether this wing mass reduction will offset the increased manufacturing cost is left unresolved for each case.
Aeroelastic stability of forward swept composite winged aircraft
NASA Technical Reports Server (NTRS)
Weisshaar, T. A.
1983-01-01
This paper reviews the author's past and present aeroelastic stability and performance studies related to forward swept, composite wing aircraft. The influence of laminate elastic bend/twist coupling upon wing divergence, lateral control, and lift effectiveness will be illustrated by means of closed-form solutions, numerical analysis and simple wind-tunnel experiments. In addition, results of analyses of a freely flying flexible FSW aircraft are discussed to indicate the possible effects of the flexible forward swept wing on aircraft dynamic stability. These studies show, both theoretically and experimentally, that, if the aircraft is not carefully designed, a phenomenon referred to as body freedom flutter may appear.
Optimal aeroelastic design of an oblique wing structure
NASA Technical Reports Server (NTRS)
Gwin, L. B.
1974-01-01
A procedure is presented for determining the optimal cover panel thickness of a wing structure to meet specified strength and static aeroelastic divergence requirements for minimum weight. Efficient reanalysis techniques using discrete structural and aerodynamic methods are used in conjunction with redesign algorithms driven by optimality criteria. The optimality conditions for the divergence constraint are established, and expressions are obtained for derivatives of the dynamic pressure at divergence with respect to design variables. The procedure is applied to an oblique wing aircraft where strength and stiffness are critical design considerations for sizing the cover thickness of the wing structure.
Twist Model Development and Results From the Active Aeroelastic Wing F/A-18 Aircraft
NASA Technical Reports Server (NTRS)
Lizotte, Andrew; Allen, Michael J.
2005-01-01
Understanding the wing twist of the active aeroelastic wing F/A-18 aircraft is a fundamental research objective for the program and offers numerous benefits. In order to clearly understand the wing flexibility characteristics, a model was created to predict real-time wing twist. A reliable twist model allows the prediction of twist for flight simulation, provides insight into aircraft performance uncertainties, and assists with computational fluid dynamic and aeroelastic issues. The left wing of the aircraft was heavily instrumented during the first phase of the active aeroelastic wing program allowing deflection data collection. Traditional data processing steps were taken to reduce flight data, and twist predictions were made using linear regression techniques. The model predictions determined a consistent linear relationship between the measured twist and aircraft parameters, such as surface positions and aircraft state variables. Error in the original model was reduced in some cases by using a dynamic pressure-based assumption and by using neural networks. These techniques produced excellent predictions for flight between the standard test points and accounted for nonlinearities in the data. This report discusses data processing techniques and twist prediction validation, and provides illustrative and quantitative results.
Twist Model Development and Results from the Active Aeroelastic Wing F/A-18 Aircraft
NASA Technical Reports Server (NTRS)
Lizotte, Andrew M.; Allen, Michael J.
2007-01-01
Understanding the wing twist of the active aeroelastic wing (AAW) F/A-18 aircraft is a fundamental research objective for the program and offers numerous benefits. In order to clearly understand the wing flexibility characteristics, a model was created to predict real-time wing twist. A reliable twist model allows the prediction of twist for flight simulation, provides insight into aircraft performance uncertainties, and assists with computational fluid dynamic and aeroelastic issues. The left wing of the aircraft was heavily instrumented during the first phase of the active aeroelastic wing program allowing deflection data collection. Traditional data processing steps were taken to reduce flight data, and twist predictions were made using linear regression techniques. The model predictions determined a consistent linear relationship between the measured twist and aircraft parameters, such as surface positions and aircraft state variables. Error in the original model was reduced in some cases by using a dynamic pressure-based assumption. This technique produced excellent predictions for flight between the standard test points and accounted for nonlinearities in the data. This report discusses data processing techniques and twist prediction validation, and provides illustrative and quantitative results.
Aeroelastic Tailoring of a Plate Wing with Functionally Graded Materials
NASA Technical Reports Server (NTRS)
Dunning, Peter D.; Stanford, Bret K.; Kim, H. Alicia; Jutte, Christine V.
2014-01-01
This work explores the use of functionally graded materials for the aeroelastic tailoring of a metallic cantilevered plate-like wing. Pareto trade-off curves between dynamic stability (flutter) and static aeroelastic stresses are obtained for a variety of grading strategies. A key comparison is between the effectiveness of material grading, geometric grading (i.e., plate thickness variations), and using both simultaneously. The introduction of material grading does, in some cases, improve the aeroelastic performance. This improvement, and the physical mechanism upon which it is based, depends on numerous factors: the two sets of metallic material parameters used for grading, the sweep of the plate, the aspect ratio of the plate, and whether the material is graded continuously or discretely.
Application of the Finite Element Method to Rotary Wing Aeroelasticity
NASA Technical Reports Server (NTRS)
Straub, F. K.; Friedmann, P. P.
1982-01-01
A finite element method for the spatial discretization of the dynamic equations of equilibrium governing rotary-wing aeroelastic problems is presented. Formulation of the finite element equations is based on weighted Galerkin residuals. This Galerkin finite element method reduces algebraic manipulative labor significantly, when compared to the application of the global Galerkin method in similar problems. The coupled flap-lag aeroelastic stability boundaries of hingeless helicopter rotor blades in hover are calculated. The linearized dynamic equations are reduced to the standard eigenvalue problem from which the aeroelastic stability boundaries are obtained. The convergence properties of the Galerkin finite element method are studied numerically by refining the discretization process. Results indicate that four or five elements suffice to capture the dynamics of the blade with the same accuracy as the global Galerkin method.
Shock Location Dominated Transonic Flight Loads on the Active Aeroelastic Wing
NASA Technical Reports Server (NTRS)
Lokos, William A.; Lizotte, Andrew; Lindsley, Ned J.; Stauf, Rick
2005-01-01
During several Active Aeroelastic Wing research flights, the shadow of the over-wing shock could be observed because of natural lighting conditions. As the plane accelerated, the shock location moved aft, and as the shadow passed the aileron and trailing-edge flap hinge lines, their associated hinge moments were substantially affected. The observation of the dominant effect of shock location on aft control surface hinge moments led to this investigation. This report investigates the effect of over-wing shock location on wing loads through flight-measured data and analytical predictions. Wing-root and wing-fold bending moment and torque and leading- and trailing-edge hinge moments have been measured in flight using calibrated strain gages. These same loads have been predicted using a computational fluid dynamics code called the Euler Navier-Stokes Three Dimensional Aeroelastic Code. The computational fluid dynamics study was based on the elastically deformed shape estimated by a twist model, which in turn was derived from in-flight-measured wing deflections provided by a flight deflection measurement system. During level transonic flight, the shock location dominated the wing trailing-edge control surface hinge moments. The computational fluid dynamics analysis based on the shape provided by the flight deflection measurement system produced very similar results and substantially correlated with the measured loads data.
Experimental Results from the Active Aeroelastic Wing Wind Tunnel Test Program
NASA Technical Reports Server (NTRS)
Heeg, Jennifer; Spain, Charles V.; Florance, James R.; Wieseman, Carol D.; Ivanco, Thomas G.; DeMoss, Joshua; Silva, Walter A.; Panetta, Andrew; Lively, Peter; Tumwa, Vic
2005-01-01
The Active Aeroelastic Wing (AAW) program is a cooperative effort among NASA, the Air Force Research Laboratory and the Boeing Company, encompassing flight testing, wind tunnel testing and analyses. The objective of the AAW program is to investigate the improvements that can be realized by exploiting aeroelastic characteristics, rather than viewing them as a detriment to vehicle performance and stability. To meet this objective, a wind tunnel model was crafted to duplicate the static aeroelastic behavior of the AAW flight vehicle. The model was tested in the NASA Langley Transonic Dynamics Tunnel in July and August 2004. The wind tunnel investigation served the program goal in three ways. First, the wind tunnel provided a benchmark for comparison with the flight vehicle and various levels of theoretical analyses. Second, it provided detailed insight highlighting the effects of individual parameters upon the aeroelastic response of the AAW vehicle. This parameter identification can then be used for future aeroelastic vehicle design guidance. Third, it provided data to validate scaling laws and their applicability with respect to statically scaled aeroelastic models.
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.
Transonic aerodynamic and aeroelastic characteristics of a variable sweep wing
NASA Technical Reports Server (NTRS)
Goorjian, P. M.; Guruswamy, G. P.; Ide, H.; Miller, G.
1985-01-01
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-sweep 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 the 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 are not due to shock wave motion as was previously proposed.
Transonic aerodynamic and aeroelastic characteristics of a variable sweep wing
NASA Technical Reports Server (NTRS)
Goorjian, P. M.; Guruswamy, G. P.; Ide, H.; Miller, G.
1985-01-01
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-sweep and high-sweep cases, at 25.0 deg. 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 the 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 are not due to shock wave motion as was previously proposed.
NASA Technical Reports Server (NTRS)
1973-01-01
The aerodynamic characteristics of transport aircraft with oblique wing flying at supersonic speeds are discussed. Aeroelastic divergence of the forward swept portion of the wing is analyzed. The effect of aspect ratio as a method for avoiding aeroelastic divergence is examined. A relatively low aspect ratio appears necessary for an oblique wing when constructed of conventional aluminum alloy materials. The aspect ratio may be increased by increasing the wing thickness ratio and by utilizing materials with higher moduli of elasticity and rigidity.
NASA Technical Reports Server (NTRS)
Adams, W. M., Jr.; Tiffany, S. H.
1983-01-01
A control law is developed to suppress symmetric flutter for a mathematical model of an aeroelastic research vehicle. An implementable control law is attained by including modified LQG (linear quadratic Gaussian) design techniques, controller order reduction, and gain scheduling. An alternate (complementary) design approach is illustrated for one flight condition wherein nongradient-based constrained optimization techniques are applied to maximize controller robustness.
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.
Prediction of wing aeroelastic effects on aircraft life and pitching moment characteristics
NASA Technical Reports Server (NTRS)
Eckstrom, Clinton V.
1987-01-01
The distribution of flight loads on an aircraft structure determine the lift and pitching moment characteristics of the aircraft. When the load distribution changes due to the aeroelastic response of the structure, the lift and pitching moment characteristics also change. An estimate of the effect of aeroelasticity on stability and control characteristics is often required for the development of aircraft simulation models of evaluation of flight characteristics. This presentation outlines a procedure for incorporating calculated linear aeroelastic effects into measured nonlinear lift and pitching moment data from wind tunnel tests. Results are presented which were obtained from applying this procedure to data for an aircraft with a very flexible transport type research wing. The procedure described is generally applicable to all types of aircraft.
Prediction of wing aeroelastic effects on aircraft lift and pitching moment characteristics
NASA Technical Reports Server (NTRS)
Eckstrom, C. V.
1985-01-01
The distribution of flight loads on an aircraft structure determines the lift and pitching moment characteristics of the aircraft. When the load distribution changes due to the aeroelastic response of the structure, the lift and pitching moment characteristics also change. Some estimate of the effect of aeroelasticity on stability and control characteristics, particularly lift and pitching moment, is required for use in aircraft simulation models for evaluation of flight characteristics. This presentation outlines a procedure to incorporate aeroelastic effects into lift and pitching moment data from wind tunnel tests. Results are presented which were obtained from applying this procedure to an aircraft with a very flexible transport-type research wing. The procedure described is generally applicable to all types of aircraft.
Prediction of wing aeroelastic effects on aircraft lift and pitching moment characteristics
NASA Technical Reports Server (NTRS)
Eckstrom, Clinton V.
1986-01-01
The distribution of flight loads on an aircraft structure determine the lift and pitching moment characteristics of the aircraft. When the load distribution changes due to the aeroelastic response of the structure, the lift and pitching moment characteristics also change. An estimate of the effect of aeroelasticity on stability and control characteristics is often required for the development of aircraft simulation models of evaluation of flight characteristics. This presentation outlines a procedure for incorporating calculated linear aeroelastic effects into measured nonlinear lift and pitching moment data from wind tunnel tests. Results are presented which were obtained from applying this procedure to data for an aircraft with a very flexible transport type research wing. The procedure described is generally applicable to all types of aircraft.
Static Aeroelastic Effects of Formation Flight for Slender Unswept Wings
NASA Technical Reports Server (NTRS)
Hanson, Curtis E.
2009-01-01
The static aeroelastic equilibrium equations for slender, straight wings are modified to incorporate the effects of aerodynamically-coupled formation flight. A system of equations is developed by applying trim constraints and is solved for component lift distribution, trim angle-of-attack, and trim aileron deflection. The trim values are then used to calculate the elastic twist distribution of the wing box. This system of equations is applied to a formation of two gliders in trimmed flight. Structural and aerodynamic properties are assumed for the gliders, and solutions are calculated for flexible and rigid wings in solo and formation flight. It is shown for a sample application of two gliders in formation flight, that formation disturbances produce greater twist in the wingtip immersed in the vortex than for either the opposing wingtip or the wings of a similar airplane in solo flight. Changes in the lift distribution, resulting from wing twist, increase the performance benefits of formation flight. A flexible wing in formation flight will require greater aileron deflection to achieve roll trim than a rigid wing.
Close-Range Photogrammetric Measurement of Static Deflections for an Aeroelastic Supercritical Wing
NASA Technical Reports Server (NTRS)
Byrdsong, Thomas A.; Adams, Richard R.; Sandford, Maynard C.
1990-01-01
Close range photogrammetric measurements were made for the lower wing surface of a full span aspect ratio 10.3 aeroelastic supercritical research wing. The measurements were made during wind tunnel tests for quasi-steady pressure distributions on the wing. The tests were conducted in the NASA Langley Transonic Dynamics Tunnel at Mach numbers up to 0.90 and dynamic pressures up to 300 pounds per square foot. Deflection data were obtained for 57 locations on the wing lower surface using dual non-metric cameras. Representative data are presented as graphical overview to show variations and trends of spar deflection with test variables. Comparative data are presented for photogrammetric and cathetometric results of measurements for the wing tip deflections. A tabulation of the basic measurements is presented in a supplement to this report.
Aeroelastic Wing Shaping Control Subject to Actuation Constraints.
NASA Technical Reports Server (NTRS)
Swei, Sean Shan-Min; Nguyen, Nhan
2014-01-01
This paper considers the control of coupled aeroelastic aircraft model which is configured with Variable Camber Continuous Trailing Edge Flap (VCCTEF) system. The relative deflection between two adjacent flaps is constrained and this actuation constraint is accounted for when designing an effective control law for suppressing the wing vibration. A simple tuned-mass damper mechanism with two attached masses is used as an example to demonstrate the effectiveness of vibration suppression with confined motion of tuned masses. In this paper, a dynamic inversion based pseudo-control hedging (PCH) and bounded control approach is investigated, and for illustration, it is applied to the NASA Generic Transport Model (GTM) configured with VCCTEF system.
NASA Technical Reports Server (NTRS)
Lind, Rick
1999-01-01
The F/A-18 Active Aeroelastic Wing research aircraft will demonstrate technologies related to aeroservoelastic effects such as wing twist and load minimization. This program presents several challenges for control design that are often not considered for traditional aircraft. This paper presents a control design based on H-infinity synthesis that simultaneously considers the multiple objectives associated with handling qualities, actuator limitations, and loads. A point design is presented to demonstrate a controller and the resulting closed-loop properties.
NASA Technical Reports Server (NTRS)
Lind, Rick
1999-01-01
The F/A-18 Active Aeroelastic Wing research aircraft will demonstrate technologies related to aeroservoelastic effects such as wing twist and load minimization. This program presents several challenges for control design that are often not considered for traditional aircraft. This paper presents a control design based on H(sub infinity) synthesis that simultaneously considers the multiple objectives associated with handling qualities, actuator limitations, and loads. A point design is presented to demonstrate a controller and the resulting closed-loop properties.
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.
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.
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.
Parallel aeroelastic computations for wing and wing-body configurations
NASA Technical Reports Server (NTRS)
Byun, Chansup
1994-01-01
The objective of this research is to develop computationally efficient methods for solving fluid-structural interaction problems by directly coupling finite difference Euler/Navier-Stokes equations for fluids and finite element dynamics equations for structures on parallel computers. This capability will significantly impact many aerospace projects of national importance such as Advanced Subsonic Civil Transport (ASCT), where the structural stability margin becomes very critical at the transonic region. This research effort will have direct impact on the High Performance Computing and Communication (HPCC) Program of NASA in the area of parallel computing.
Smithornis broadbills produce loud wing song by aeroelastic flutter of medial primary wing feathers.
Clark, Christopher J; Kirschel, Alexander N G; Hadjioannou, Louis; Prum, Richard O
2016-04-01
Broadbills in the genus Smithornis produce a loud brreeeeet during a distinctive flight display. It has been posited that this klaxon-like sound is generated non-vocally with the outer wing feathers (P9, P10), but no scientific studies have previously addressed this hypothesis. Although most birds that make non-vocal communication sounds have feathers with a shape distinctively modified for sound production, Smithornis broadbills do not. We investigated whether this song is produced vocally or with the wings in rufous-sided broadbill (S. rufolateralis) and African broad bill (S. capensis). In support of the wing song hypothesis, synchronized high-speed video and sound recordings of displays demonstrated that sound pulses were produced during the downstroke, subtle gaps sometimes appeared between the outer primary feathers P6-P10, and wing tip speed reached 16 m s(-1) Tests of a spread wing in a wind tunnel demonstrated that at a specific orientation, P6 and P7 flutter and produce sound. Wind tunnel tests on individual feathers P5-P10 from a male of each species revealed that while all of these feathers can produce sound via aeroelastic flutter, P6 and P7 produce the loudest sounds, which are similar in frequency to the wing song, at airspeeds achievable by the wing tip during display flight. Consistent with the wind tunnel experiments, field manipulations of P6, P7 and P8 changed the timbre of the wing song, and reduced its tonality, demonstrating that P6 and P7 are together the sound source, and not P9 or P10. The resultant wing song appears to have functionally replaced vocal song. PMID:27030781
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)
Stanford, Bret K.; Jutte, Christine V.
2014-01-01
Several minimum-mass aeroelastic optimization problems are solved to evaluate the effectiveness of a variety of novel tailoring schemes for subsonic transport wings. Aeroelastic strength and panel buckling constraints are imposed across a variety of trimmed maneuver loads. Tailoring with metallic thickness variations, functionally graded materials, composite laminates, tow steering, and distributed trailing edge control effectors are all found to provide reductions in structural wing mass with varying degrees of success. The question as to whether this wing mass reduction will offset the increased manufacturing cost is left unresolved for each case.
Stability and Control Properties of an Aeroelastic Fixed Wing Micro Aerial Vehicle
NASA Technical Reports Server (NTRS)
Waszak, Martin R.; Jenkins, Luther N.; Ifju, Peter
2001-01-01
Micro aerial vehicles have been the subject of considerable interest and development over the last several years. The majority of current vehicle concepts rely on rigid fixed wings or rotors. An alternate design based on an aeroelastic membrane wing concept has also been developed that has exhibited desired characteristics in flight test demonstrations and competition. This paper presents results from a wind tunnel investigation that sought to quantify stability and control properties for a family of vehicles using the aeroelastic design. The results indicate that the membrane wing does exhibit potential benefits that could be exploited to enhance the design of future flight vehicles.
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.
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.
Optimum Aeroelastic Design of Resonance Type Flapping Wing for Micro Air Vehicles
NASA Astrophysics Data System (ADS)
Isogai, Koji; Kamisawa, Yuichi; Sato, Hiroyuki
The optimum aeroelastic design method for a resonance-type flapping wing for a Micro Air Vehicle (MAV) is presented. It uses Complex Method and 3D Navier-Stokes code to determine the optimum structural and aerodynamic parameters of a 2 DOF flapping wing system. The method is used to design a dragonfly-type MAV, and numerical simulation shows that the designed flapping wings can generate sufficient lift to sustain the weight and sufficient thrust to overcome the body drag.
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.
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.
Aeroelastic stability and control of an oblique wing - Wind tunnel experiments
NASA Technical Reports Server (NTRS)
Jones, R. T.
1976-01-01
Results are presented for wind tunnel tests of an elastic wing model to verify the theoretical predictions for the aeroelastic instability of an oblique wing. The model wing has an elliptic planform of 10 to 1 axis ratio and a symmetrical airfoil section of 7-1/2% thickness/chord ratio. The wing is of wood and as may be seen in the photographs presented, slack wires are used to limit the amplitude of unstable motions. The fuselage is mounted on bearings permitting freedom of roll, but provision is made to clamp the fuselage for some of the tests. It is found that freedom in roll increases the dynamic pressure at which aeroelastic instability first appears. With the model free in roll, the effectiveness of the ailerons in maintaining trim is not noticeably affected by passage through the speed at which the wing would become unstable if clamped.
Aeroelastic Response of Nonlinear Wing Section By Functional Series Technique
NASA Technical Reports Server (NTRS)
Marzocca, Piergiovanni; Librescu, Liviu; Silva, Walter A.
2000-01-01
This paper addresses the problem of the determination of the subcritical aeroelastic response and flutter instability of nonlinear two-dimensional lifting surfaces in an incompressible flow-field via indicial functions and Volterra series approach. The related aeroelastic governing equations are based upon the inclusion of structural and damping nonlinearities in plunging and pitching, of the linear unsteady aerodynamics and consideration of an arbitrary time-dependent external pressure pulse. Unsteady aeroelastic nonlinear kernels are determined, and based on these, frequency and time histories of the subcritical aeroelastic response are obtained, and in this context the influence of the considered nonlinearities is emphasized. Conclusions and results displaying the implications of the considered effects are supplied.
Aeroelastic Response of Nonlinear Wing Section by Functional Series Technique
NASA Technical Reports Server (NTRS)
Silva, Walter A.; Marzocca, Piergiovanni
2001-01-01
This paper addresses the problem of the determination of the subcritical aeroelastic response and flutter instability of nonlinear two-dimensional lifting surfaces in an incompressible flow-field via indicial functions and Volterra series approach. The related aeroelastic governing equations are based upon the inclusion of structural and damping nonlinearities in plunging and pitching, of the linear unsteady aerodynamics and consideration of an arbitrary time-dependent external pressure pulse. Unsteady aeroelastic nonlinear kernels are determined, and based on these, frequency and time histories of the subcritical aeroelastic response are obtained, and in this context the influence of the considered nonlinearities is emphasized. Conclusions and results displaying the implications of the considered effects are supplied.
NASA Technical Reports Server (NTRS)
Erickson, Gary E.
2013-01-01
A video-based photogrammetric model deformation system was established as a dedicated optical measurement technique at supersonic speeds in the NASA Langley Research Center Unitary Plan Wind Tunnel. This system was used to measure the wing twist due to aerodynamic loads of two supersonic commercial transport airplane models with identical outer mold lines but different aeroelastic properties. One model featured wings with deflectable leading- and trailing-edge flaps and internal channels to accommodate static pressure tube instrumentation. The wings of the second model were of single-piece construction without flaps or internal channels. The testing was performed at Mach numbers from 1.6 to 2.7, unit Reynolds numbers of 1.0 million to 5.0 million, and angles of attack from -4 degrees to +10 degrees. The video model deformation system quantified the wing aeroelastic response to changes in the Mach number, Reynolds number concurrent with dynamic pressure, and angle of attack and effectively captured the differences in the wing twist characteristics between the two test articles.
Aeroelastic Studies of a Rectangular Wing with a Hole: Correlation of Theory and Experiment
NASA Technical Reports Server (NTRS)
Conyers, Howard J.; Dowell, Earl H.; Hall, Kenneth C.
2010-01-01
Two rectangular wing models with a hole have been designed and tested in the Duke University wind tunnel to better understand the effects of damage. A rectangular hole is used to simulate damage. The wing with a hole is modeled structurally as a thin elastic plate using the finite element method. The unsteady aerodynamics of the plate-like wing with a hole is modeled using the doublet lattice method. The aeroelastic equations of motion are derived using Lagrange's equation. The flutter boundary is found using the V-g method. The hole's location effects the wing's mass, stiffness, aerodynamics and therefore the aeroelastic behavior. Linear theoretical models were shown to be capable of predicting the critical flutter velocity and frequency as verified by wind tunnel tests.
NASA Technical Reports Server (NTRS)
Batina, J. T.
1985-01-01
Unsteady transonic flow calculations for aerodynamically interfering airfoil configurations are performed as a first-step toward solving the three-dimensional canard-wing interaction problem. These calculations are performed by extending the XTRAN2L two-dimensional unsteady transonic small-disturbance code to include an additional airfoil. Unsteady transonic forces due to plunge and pitch motions of a two-dimensional canard and wing are presented. Results for a variety of canard-wing separation distances reveal the effects of aerodynamic interference on unsteady transonic airloads. Aeroelastic analyses employing these unsteady airloads demonstrate the effects of aerodynamic interference on aeroelastic stability and flutter. For the configurations studied, increases in wing flutter speed result with the inclusion of the aerodynamically interfering canard.
NASA Technical Reports Server (NTRS)
Batina, J. T.
1985-01-01
Unsteady transonic flow calculations for aerodynamically interfering airfoil configurations are performed as a first step toward solving the three dimensional canard wing interaction problem. These calculations are performed by extending the XTRAN2L two dimensional unsteady transonic small disturbance code to include an additional airfoil. Unsteady transonic forces due to plunge and pitch motions of a two dimensional canard and wing are presented. Results for a variety of canard wing separation distances reveal the effects of aerodynamic interference on unsteady transonic airloads. Aeroelastic analyses employing these unsteady airloads demonstrate the effects of aerodynamic interference on aeroelastic stability and flutter. For the configurations studied, increases in wing flutter speed result with the inclusion of the aerodynamically interfering canard.
A comparative study of serial and parallel aeroelastic computations of wings
NASA Technical Reports Server (NTRS)
Byun, Chansup; Guruswamy, Guru P.
1994-01-01
A procedure for computing the aeroelasticity of wings on parallel multiple-instruction, multiple-data (MIMD) computers is presented. In this procedure, fluids are modeled using Euler equations, and structures are modeled using modal or finite element equations. The procedure is designed in such a way that each discipline can be developed and maintained independently by using a domain decomposition approach. In the present parallel procedure, each computational domain is scalable. A parallel integration scheme is used to compute aeroelastic responses by solving fluid and structural equations concurrently. The computational efficiency issues of parallel integration of both fluid and structural equations are investigated in detail. This approach, which reduces the total computational time by a factor of almost 2, is demonstrated for a typical aeroelastic wing by using various numbers of processors on the Intel iPSC/860.
Dynamic structural aeroelastic stability testing of the XV-15 tilt rotor research aircraft
NASA Technical Reports Server (NTRS)
Schroers, L. G.
1982-01-01
For the past 20 years, a significant effort has been made to understand and predict the structural aeroelastic stability characteristics of the tilt rotor concept. Beginning with the rotor-pylon oscillation of the XV-3 aircraft, the problem was identified and then subjected to a series of theoretical studies, plus model and full-scale wind tunnel tests. From this data base, methods were developed to predict the structural aeroelastic stability characteristics of the XV-15 Tilt Rotor Research Aircraft. The predicted aeroelastic characteristics are examined in light of the major parameters effecting rotor-pylon-wing stability. Flight test techniques used to obtain XV-15 aeroelastic stability are described. Flight test results are summarized and compared to the predicted values. Wind tunnel results are compared to flight test results and correlated with predicted values.
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.
Active Aeroelastic Wing Aerodynamic Model Development and Validation for a Modified F/A-18A
NASA Technical Reports Server (NTRS)
Cumming, Stephen B.; Diebler, Corey G.
2005-01-01
A new aerodynamic model has been developed and validated for a modified F/A-18A used for the Active Aeroelastic Wing (AAW) research program. The goal of the program was to demonstrate the advantages of using the inherent flexibility of an aircraft to enhance its performance. The research aircraft was an F/A-18A with wings modified to reduce stiffness and a new control system to increase control authority. There have been two flight phases. Data gathered from the first flight phase were used to create the new aerodynamic model. A maximum-likelihood output-error parameter estimation technique was used to obtain stability and control derivatives. The derivatives were incorporated into the National Aeronautics and Space Administration F-18 simulation, validated, and used to develop new AAW control laws. The second phase of flights was used to evaluate the handling qualities of the AAW aircraft and the control law design process, and to further test the accuracy of the new model. The flight test envelope covered Mach numbers between 0.85 and 1.30 and dynamic pressures from 600 to 1250 pound-force per square foot. The results presented in this report demonstrate that a thorough parameter identification analysis can be used to improve upon models that were developed using other means. This report describes the parameter estimation technique used, details the validation techniques, discusses differences between previously existing F/A-18 models, and presents results from the second phase of research flights.
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).
Simulation and Flight Control of an Aeroelastic Fixed Wing Micro Aerial Vehicle
NASA Technical Reports Server (NTRS)
Waszak, Martin; Davidson, John B.; Ifju, Peter G.
2002-01-01
Micro aerial vehicles have been the subject of continued interest and development over the last several years. The majority of current vehicle concepts rely on rigid fixed wings or rotors. An alternate design based on an aeroelastic membrane wing has also been developed that exhibits desired characteristics in flight test demonstrations, competition, and in prior aerodynamics studies. This paper presents a simulation model and an assessment of flight control characteristics of the vehicle. Linear state space models of the vehicle associated with typical trimmed level flight conditions and which are suitable for control system design are presented as well. The simulation is used as the basis for the design of a measurement based nonlinear dynamic inversion control system and outer loop guidance system. The vehicle/controller system is the subject of ongoing investigations of autonomous and collaborative control schemes. The results indicate that the design represents a good basis for further development of the micro aerial vehicle for autonomous and collaborative controls research.
Deflection-Based Structural Loads Estimation From the Active Aeroelastic Wing F/A-18 Aircraft
NASA Technical Reports Server (NTRS)
Lizotte, Andrew M.; Lokos, William A.
2005-01-01
Traditional techniques in structural load measurement entail the correlation of a known load with strain-gage output from the individual components of a structure or machine. The use of strain gages has proved successful and is considered the standard approach for load measurement. However, remotely measuring aerodynamic loads using deflection measurement systems to determine aeroelastic deformation as a substitute to strain gages may yield lower testing costs while improving aircraft performance through reduced instrumentation weight. This technique was examined using a reliable strain and structural deformation measurement system. The objective of this study was to explore the utility of a deflection-based load estimation, using the active aeroelastic wing F/A-18 aircraft. Calibration data from ground tests performed on the aircraft were used to derive left wing-root and wing-fold bending-moment and torque load equations based on strain gages, however, for this study, point deflections were used to derive deflection-based load equations. Comparisons between the strain-gage and deflection-based methods are presented. Flight data from the phase-1 active aeroelastic wing flight program were used to validate the deflection-based load estimation method. Flight validation revealed a strong bending-moment correlation and slightly weaker torque correlation. Development of current techniques, and future studies are discussed.
Aeroelasticity of Axially Loaded Aerodynamic Structures for Truss-Braced Wing Aircraft
NASA Technical Reports Server (NTRS)
Nguyen, Nhan; Ting, Eric; Lebofsky, Sonia
2015-01-01
This paper presents an aeroelastic finite-element formulation for axially loaded aerodynamic structures. The presence of axial loading causes the bending and torsional sitffnesses to change. For aircraft with axially loaded structures such as the truss-braced wing aircraft, the aeroelastic behaviors of such structures are nonlinear and depend on the aerodynamic loading exerted on these structures. Under axial strain, a tensile force is created which can influence the stiffness of the overall aircraft structure. This tension stiffening is a geometric nonlinear effect that needs to be captured in aeroelastic analyses to better understand the behaviors of these types of aircraft structures. A frequency analysis of a rotating blade structure is performed to demonstrate the analytical method. A flutter analysis of a truss-braced wing aircraft is performed to analyze the effect of geometric nonlinear effect of tension stiffening on the flutter speed. The results show that the geometric nonlinear tension stiffening effect can have a significant impact on the flutter speed prediction. In general, increased wing loading results in an increase in the flutter speed. The study illustrates the importance of accounting for the geometric nonlinear tension stiffening effect in analyzing the truss-braced wing aircraft.
Aeroelasticity of wing and wing-body configurations on parallel computers
NASA Technical Reports Server (NTRS)
Byun, Chansup
1995-01-01
The objective of this research is to develop computationally efficient methods for solving aeroelasticity problems on parallel computers. Both uncoupled and coupled methods are studied in this research. For the uncoupled approach, the conventional U-g method is used to determine the flutter boundary. The generalized aerodynamic forces required are obtained by the pulse transfer-function analysis method. For the coupled approach, the fluid-structure interaction is obtained by directly coupling finite difference Euler/Navier-Stokes equations for fluids and finite element dynamics equations for structures. This capability will significantly impact many aerospace projects of national importance such as Advanced Subsonic Civil Transport (ASCT), where the structural stability margin becomes very critical at the transonic region. This research effort will have direct impact on the High Performance Computing and Communication (HPCC) Program of NASA in the area of parallel computing.
NASA Technical Reports Server (NTRS)
Ashley, H.
1984-01-01
Graduate research activity in the following areas is reported: the divergence of laminated composite lifting surfaces, subsonic propeller theory and aeroelastic analysis, and cross sectional resonances in wind tunnels.
NASA 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.
Aeroelastic loads prediction for an arrow wing. Task 2: Evaluation of semi-empirical methods
NASA Technical Reports Server (NTRS)
Wery, A. C.; Kulfan, R. M.; Manro, M. E.
1983-01-01
The development and evaluation of a semi empirical method to predict pressure distributions on a deformed wing by using an experimental data base in addition to a linear potential flow solution is described. The experimental data accounts for the effects of aeroelasticity by relating the pressures to a parameter which is influenced by the deflected shape. Several parameters were examined before the net leading edge suction coefficient was selected as the best.
NASA Technical Reports Server (NTRS)
Byun, Chansup; Guruswamy, Guru P.
1993-01-01
This paper presents a procedure for computing the aeroelasticity of wing-body configurations on multiple-instruction, multiple-data (MIMD) parallel computers. In this procedure, fluids are modeled using Euler equations discretized by a finite difference method, and structures are modeled using finite element equations. The procedure is designed in such a way that each discipline can be developed and maintained independently by using a domain decomposition approach. A parallel integration scheme is used to compute aeroelastic responses by solving the coupled fluid and structural equations concurrently while keeping modularity of each discipline. The present procedure is validated by computing the aeroelastic response of a wing and comparing with experiment. Aeroelastic computations are illustrated for a High Speed Civil Transport type wing-body configuration.
An investigation of supersonic aeroelastic characteristics of oblique winged aircraft
NASA Technical Reports Server (NTRS)
Weisshaar, T. A.
1975-01-01
Two formulations of the oblique wing flutter problem are presented: one formulation allows wing bending deformations and the rigid body roll degree of freedom while the second formulation includes bending-torsional deformation and roll degrees of freedom. Flutter is found to occur in two basic modes. The first mode is associated with bending-roll coupling and occurs at low reduced frequency values. The other instability mode is primarily one of classical bending-torsion with negligible roll coupling; this mode occurs at much higher reduced frequencies. The occurrence of bending-roll coupling mode leads to lower flutter speeds while the bending-torsion mode is associated with higher flutter speeds. The ratio of the wing mass moment of inertial in roll to the fuselage moment of inertia evidently plays a major role in the determination of which of the two instabilities is critical.
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.
Parallel Nonlinear Aeroelastic Computation for Fighter Wings in the Transonic Region
NASA Astrophysics Data System (ADS)
Larsen, Bradley Robert
In this dissertation, a parallel three-dimensional aeroelastic simulation is applied to current and next generation fighter aircraft wings. The computational model is a nonlinear fluid and structural mesh coupled using the Direct Eulerian-Langrangian method. This method attaches unique local coordinates to each node and connects the fluid mesh to the structure in such a way that a transformation preserved to the global coordinates. This allows the fluid and structure to be updated in the same time step and maintains spatial accuracy at their interface. The structural mesh is modeled using modified nonlinear von Karman finite elements and is discretized using the Galerkin finite element method. The fluid mesh also used the Galerkin finite element method to discretize the unsteady Euler equations. Computational results over a large range of Mach numbers and densities are presented for two candidate fighter wing models for transonic wing tunnel testing. The FX-35 is a trapezoidal wing based on the F-35A, and the F-Wing is a truncated delta wing similar to the F-16. Both wings exhibit a variety of flutter behaviors including strong bending-torsion flutter, limit-cycle oscillations, and essentially single degree-of-freedom responses.
Modeling and analysis methodology for aeroelastically tailored chordwise deformable wings
NASA Technical Reports Server (NTRS)
Rehfield, Lawrence W.; Chang, Stephen; Zischka, Peter J.
1992-01-01
Structural concepts have been created which produce chordwise camber deformation that results in enhanced lift. A wing box can be tailored to utilize each of these with composites. In attempting to optimize the aerodynamic benefits, we have found there are two optimal designs that are of interest. There is a weight optimum which corresponds to the maximum lift per unit structural weight. There is also a lift optimum that corresponds to maximum absolute lift. New structural models, the basic deformation mechanisms that are utilized and typical analytical results are presented. It appears that lift enhancements of sufficient magnitude can be produced to render this type of wing tailoring of practical interest. Experiments and finite element correlations are performed which confirm the validity of the theoretical models utilized.
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.
Strain Gage Loads Calibration Testing of the Active Aeroelastic Wing F/A-18 Aircraft
NASA Technical Reports Server (NTRS)
Lokos, William A.; Olney, Candida D.; Chen, Tony; Crawford, Natalie D.; Stauf, Rick; Reichenbach, Eric Y.; Bessette, Denis (Technical Monitor)
2002-01-01
This report describes strain-gage calibration loading through the application of known loads of the Active Aeroelastic Wing F/A-18 airplane. The primary goal of this test is to produce a database suitable for deriving load equations for left and right wing root and fold shear; bending moment; torque; and all eight wing control-surface hinge moments. A secondary goal is to produce a database of wing deflections measured by string potentiometers and the onboard flight deflection measurement system. Another goal is to produce strain-gage data through both the laboratory data acquisition system and the onboard aircraft data system as a check of the aircraft system. Thirty-two hydraulic jacks have applied loads through whiffletrees to 104 tension-compression load pads bonded to the lower wing surfaces. The load pads covered approximately 60 percent of the lower wing surface. A series of 72 load cases has been performed, including single-point, double-point, and distributed load cases. Applied loads have reached 70 percent of the flight limit load. Maximum wingtip deflection has reached nearly 16 in.
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.
Real-Time Adaptive Least-Squares Drag Minimization for Performance Adaptive Aeroelastic Wing
NASA Technical Reports Server (NTRS)
Ferrier, Yvonne L.; Nguyen, Nhan T.; Ting, Eric
2016-01-01
This paper contains a simulation study of a real-time adaptive least-squares drag minimization algorithm for an aeroelastic model of a flexible wing aircraft. The aircraft model is based on the NASA Generic Transport Model (GTM). The wing structures incorporate a novel aerodynamic control surface known as the Variable Camber Continuous Trailing Edge Flap (VCCTEF). The drag minimization algorithm uses the Newton-Raphson method to find the optimal VCCTEF deflections for minimum drag in the context of an altitude-hold flight control mode at cruise conditions. The aerodynamic coefficient parameters used in this optimization method are identified in real-time using Recursive Least Squares (RLS). The results demonstrate the potential of the VCCTEF to improve aerodynamic efficiency for drag minimization for transport aircraft.
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.
2005 PathfinderPlus Aero-Elastic Research Flight
NASA Technical Reports Server (NTRS)
Navarro, Robert
2005-01-01
This viewgraph presentation describes the 2005 Pathfinder along with an investigation of its aeroelastic responses. The contents include: 1) HALE Class of Vehicles; 2) Aero-elastic Research Flights Overall Objective; 3) General Arrangement; 4) Sensor Locations; 5) NASA Ramp Operations; 6) Lakebed Operations; 7) 1st Flight Data Set; 8) Tool development / data usage; 9) HALE Tool Development & Validation; 10) Building a HALE Foundation; 11) Compelling Needs Drive HALE Efforts; and 12) Team Photo
Gain-Scheduled Approximations to H-Infinity Controllers for the F/A-18 Active Aeroelastic Wing
NASA Technical Reports Server (NTRS)
Lind, Rick
1999-01-01
The Active Aeroelastic Wing will demonstrate technologies related to aeroservoelastic effects such as wing twist and load minimization. This paper presents a control design based on H-infinity synthesis that simultaneously considers the multiple objectives associated with handling qualities, actuator limitations, and loads. The controller is realized as a filter and gain set approximation to a state-space H-infinity controller. This approximation allows scheduling of the controller over a flight envelope.
NASA Technical Reports Server (NTRS)
Diederich, Franklin W; Foss, Kenneth A
1953-01-01
Charts and approximate formulas are presented for the estimation of aeroelastic effects on the spanwise lift distribution, lift-curve slope, aerodynamic center, and damping in roll of swept and unswept wings at subsonic and supersonic speeds. Some design considerations brought out by the results of this report are discussed.
NASA Technical Reports Server (NTRS)
Ting, Eric; Nguyen, Nhan; Trinh, Khanh
2014-01-01
This paper presents a static aeroelastic model and longitudinal trim model for the analysis of a flexible wing transport aircraft. The static aeroelastic model is built using a structural model based on finite-element modeling and coupled to an aerodynamic model that uses vortex-lattice solution. An automatic geometry generation tool is used to close the loop between the structural and aerodynamic models. The aeroelastic model is extended for the development of a three degree-of-freedom longitudinal trim model for an aircraft with flexible wings. The resulting flexible aircraft longitudinal trim model is used to simultaneously compute the static aeroelastic shape for the aircraft model and the longitudinal state inputs to maintain an aircraft trim state. The framework is applied to an aircraft model based on the NASA Generic Transport Model (GTM) with wing structures allowed to flexibly deformed referred to as the Elastically Shaped Aircraft Concept (ESAC). The ESAC wing mass and stiffness properties are based on a baseline "stiff" values representative of current generation transport aircraft.
Active Aeroelastic Wing Aerodynamic Model Development and Validation for a Modified F/A-18A Airplane
NASA Technical Reports Server (NTRS)
Cumming, Stephen B.; Diebler, Corey G.
2005-01-01
A new aerodynamic model has been developed and validated for a modified F/A-18A airplane used for the Active Aeroelastic Wing (AAW) research program. The goal of the program was to demonstrate the advantages of using the inherent flexibility of an aircraft to enhance its performance. The research airplane was an F/A-18A with wings modified to reduce stiffness and a new control system to increase control authority. There have been two flight phases. Data gathered from the first flight phase were used to create the new aerodynamic model. A maximum-likelihood output-error parameter estimation technique was used to obtain stability and control derivatives. The derivatives were incorporated into the National Aeronautics and Space Administration F-18 simulation, validated, and used to develop new AAW control laws. The second phase of flights was used to evaluate the handling qualities of the AAW airplane and the control law design process, and to further test the accuracy of the new model. The flight test envelope covered Mach numbers between 0.85 and 1.30 and dynamic pressures from 600 to 1250 pound-force per square foot. The results presented in this report demonstrate that a thorough parameter identification analysis can be used to improve upon models that were developed using other means. This report describes the parameter estimation technique used, details the validation techniques, discusses differences between previously existing F/A-18 models, and presents results from the second phase of research flights.
Effect of multiple engine placement on aeroelastic trim and stability of flying wing aircraft
NASA Astrophysics Data System (ADS)
Mardanpour, Pezhman; Richards, Phillip W.; Nabipour, Omid; Hodges, Dewey H.
2014-01-01
Effects of multiple engine placement on flutter characteristics of a backswept flying wing resembling the HORTEN IV are investigated using the code NATASHA (Nonlinear Aeroelastic Trim And Stability of HALE Aircraft). Four identical engines with defined mass, inertia, and angular momentum are placed in different locations along the span with different offsets from the elastic axis while fixing the location of the aircraft c.g. The aircraft experiences body freedom flutter along with non-oscillatory instabilities that originate from flight dynamics. Multiple engine placement increases flutter speed particularly when the engines are placed in the outboard portion of the wing (60-70% span), forward of the elastic axis, while the lift to drag ratio is affected negligibly. The behavior of the sub- and supercritical eigenvalues is studied for two cases of engine placement. NATASHA captures a hump body-freedom flutter with low frequency for the clean wing case, which disappears as the engines are placed on the wings. In neither case is there any apparent coalescence between the unstable modes. NATASHA captures other non-oscillatory unstable roots with very small amplitude, apparently originating with flight dynamics. For the clean-wing case, in the absence of aerodynamic and gravitational forces, the regions of minimum kinetic energy density for the first and third bending modes are located around 60% span. For the second mode, this kinetic energy density has local minima around the 20% and 80% span. The regions of minimum kinetic energy of these modes are in agreement with calculations that show a noticeable increase in flutter speed if engines are placed forward of the elastic axis at these regions.
Control Surface Interaction Effects of the Active Aeroelastic Wing Wind Tunnel Model
NASA Technical Reports Server (NTRS)
Heeg, Jennifer
2006-01-01
This paper presents results from testing the Active Aeroelastic Wing wind tunnel model in NASA Langley s Transonic Dynamics Tunnel. The wind tunnel test provided an opportunity to study aeroelastic system behavior under combined control surface deflections, testing for control surface interaction effects. Control surface interactions were observed in both static control surface actuation testing and dynamic control surface oscillation testing. The primary method of evaluating interactions was examination of the goodness of the linear superposition assumptions. Responses produced by independently actuating single control surfaces were combined and compared with those produced by simultaneously actuating and oscillating multiple control surfaces. Adjustments to the data were required to isolate the control surface influences. Using dynamic data, the task increases, as both the amplitude and phase have to be considered in the data corrections. The goodness of static linear superposition was examined and analysis of variance was used to evaluate significant factors influencing that goodness. The dynamic data showed interaction effects in both the aerodynamic measurements and the structural measurements.
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.
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.
Effects of spoiler surfaces on the aeroelastic behavior of a low-aspect-ratio rectangular wing
NASA Technical Reports Server (NTRS)
Cole, Stanley R.
1990-01-01
An experimental research study to determine the effectiveness of spoiler surfaces in suppressing flutter onset for a low-aspect-ratio, rectangular wing was conducted in the Langley Transonic Dynamics Tunnel (TDT). The wing model used in this flutter test consisted of a rigid wing mounted to the wind-tunnel wall by a flexible, rectangular beam. The flexible beam was connected to the wing root and cantilever mounted to the wind-tunnel wall. The wing had a 1.5 aspect ratio based on wing semispan and a NACA 64A010 airfoil shape. The spoiler surfaces consisted of thin, rectangular aluminum plates that were vertically mounted to the wing surface. The spoiler surface geometry and location on the wing surface were varied to determine the effects of these parameters on the classical flutter of the wing model. Subsonically, the experiment showed that spoiler surfaces increased the flutter dynamic pressure with each successive increase in spoiler height or width. This subsonic increase in flutter dynamic pressure was approximately 15 percent for the maximum height spoiler configuration and for the maximum width spoiler configuration. At transonic Mach numbers, the flutter dynamic pressure conditions were increased even more substantially than at subsonic Mach numbers for some of the smaller spoiler surfaces. But greater than a certain spoiler size (in terms of either height or width) the spoilers forced a torsional instability in the transonic regime that was highly Mach number dependent. This detrimental torsional instability was found at dynamic pressures well below the expected flutter conditions. Variations in the spanwise location of the spoiler surfaces on the wing showed little effect on flutter. Flutter analysis was conducted for the basic configuration (clean wing with all spoiler surface mass properties included). The analysis correlated well with the clean wing experimental flutter results.
Aeroelastic prediction of the limit cycle oscillations of a cropped delta wing
NASA Astrophysics Data System (ADS)
Attar, P. J.; Gordnier, R. E.
2006-01-01
The flutter and limit cycle oscillation (LCO) behavior of a cropped delta wing are investigated using a newly developed computational aeroelastic solver. This computational model includes a well-validated Euler finite difference solver coupled to a high-fidelity finite element structural solver. The nonlinear structural model includes geometric nonlinearities which are modelled using a co-rotational formulation. The LCOs of the cropped delta wing are computed and the results are compared to previous computations and to experiment. Over the range of dynamic pressures for which experimental results are reported, the LCO magnitudes computed using the current model are comparable to those from a previous computation which used a lower-order von Karman structural model. However, for larger dynamic pressures, the current computational model and the model which used the von Karman theory start to differ significantly, with the current model predicting larger deflections for a given dynamic pressure. This results in a LCO curve which is in better qualitative agreement with experiment. Flow features which were present in the previous computational model such as a leading edge vortex and a shock wave are enhanced in the current model due to the prediction of larger deflections and rotations at the higher dynamic pressures.
NASA Technical Reports Server (NTRS)
Jutte, Christine V.; Stanford, Bret K.; Wieseman, Carol D.; Moore, James B.
2014-01-01
This work explores the use of tow steered composite laminates, functionally graded metals (FGM), thickness distributions, and curvilinear rib/spar/stringer topologies for aeroelastic tailoring. Parameterized models of the Common Research Model (CRM) wing box have been developed for passive aeroelastic tailoring trade studies. Metrics of interest include the wing weight, the onset of dynamic flutter, and the static aeroelastic stresses. Compared to a baseline structure, the lowest aggregate static wing stresses could be obtained with tow steered skins (47% improvement), and many of these designs could reduce weight as well (up to 14%). For these structures, the trade-off between flutter speed and weight is generally strong, although one case showed both a 100% flutter improvement and a 3.5% weight reduction. Material grading showed no benefit in the skins, but moderate flutter speed improvements (with no weight or stress increase) could be obtained by grading the spars (4.8%) or ribs (3.2%), where the best flutter results were obtained by grading both thickness and material. For the topology work, large weight reductions were obtained by removing an inner spar, and performance was maintained by shifting stringers forward and/or using curvilinear ribs: 5.6% weight reduction, a 13.9% improvement in flutter speed, but a 3.0% increase in stress levels. Flutter resistance was also maintained using straightrotated ribs although the design had a 4.2% lower flutter speed than the curved ribs of similar weight and stress levels were higher. These results will guide the development of a future design optimization scheme established to exploit and combine the individual attributes of these technologies.
NASA Technical Reports Server (NTRS)
Graves, Sharon S.; Burner, Alpheus W.; Edwards, John W.; Schuster, David M.
2001-01-01
The techniques used to acquire, reduce, and analyze dynamic deformation measurements of an aeroelastic semispan wind tunnel model are presented. Single-camera, single-view video photogrammetry (also referred to as videogrammetric model deformation, or VMD) was used to determine dynamic aeroelastic deformation of the semispan 'Models for Aeroelastic Validation Research Involving Computation' (MAVRIC) model in the Transonic Dynamics Tunnel at the NASA Langley Research Center. Dynamic deformation was determined from optical retroreflective tape targets at five semispan locations located on the wing from the root to the tip. Digitized video images from a charge coupled device (CCD) camera were recorded and processed to automatically determine target image plane locations that were then corrected for sensor, lens, and frame grabber spatial errors. Videogrammetric dynamic data were acquired at a 60-Hz rate for time records of up to 6 seconds during portions of this flutter/Limit Cycle Oscillation (LCO) test at Mach numbers from 0.3 to 0.96. Spectral analysis of the deformation data is used to identify dominant frequencies in the wing motion. The dynamic data will be used to separate aerodynamic and structural effects and to provide time history deflection data for Computational Aeroelasticity code evaluation and validation.
Flight Test of the F/A-18 Active Aeroelastic Wing Airplane
NASA Technical Reports Server (NTRS)
Voracek, David
2007-01-01
A viewgraph presentation of flight tests performed on the F/A active aeroelastic wing airplane is shown. The topics include: 1) F/A-18 AAW Airplane; 2) F/A-18 AAW Control Surfaces; 3) Flight Test Background; 4) Roll Control Effectiveness Regions; 5) AAW Design Test Points; 6) AAW Phase I Test Maneuvers; 7) OBES Pitch Doublets; 8) OBES Roll Doublets; 9) AAW Aileron Flexibility; 10) Phase I - Lessons Learned; 11) Control Law Development and Verification & Validation Testing; 12) AAW Phase II RFCS Envelopes; 13) AAW 1-g Phase II Flight Test; 14) Region I - Subsonic 1-g Rolls; 15) Region I - Subsonic 1-g 360 Roll; 16) Region II - Supersonic 1-g Rolls; 17) Region II - Supersonic 1-g 360 Roll; 18) Region III - Subsonic 1-g Rolls; 19) Roll Axis HOS/LOS Comparison Region II - Supersonic (open-loop); 20) Roll Axis HOS/LOS Comparison Region II - Supersonic (closed-loop); 21) AAW Phase II Elevated-g Flight Test; 22) Region I - Subsonic 4-g RPO; and 23) Phase II - Lessons Learned
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.
A methodology for robust structural design with application to active aeroelastic wings
NASA Astrophysics Data System (ADS)
Zink, Paul Scott
A new design process for Active Aeroelastic Wing (AAW) technology was developed, in which control surface gear ratios and structural design variables were treated together in the same optimization problem, acting towards the same objective of weight minimization. This is in contrast to traditional AAW design processes that treat design of the gear ratios and design of the structure as separate optimization problems, each with their own different objectives and constraints, executed in an iterative fashion. The demonstration of the new AAW design process, implemented in an efficient modal-based structural analysis and optimization code, on a lightweight fighter resulted in a 15% reduction in wing box skin weight over a more traditional AAW design process. In addition, the new process was far more streamlined than the traditional approach in that it was performed in one continuous run and did not require the exchange of data between modules. The new AAW design process was then used in the development of a methodology for the design of AAW structures that are robust to uncertainty in maneuver loads which arise from the use of linear aerodynamics. Maneuver load uncertainty was modeled probabilistically and based on typical differences between rigid loads as predicted by nonlinear and linear aerodynamic theory. These models were used to augment the linear aerodynamic loads that had been used in the AAW design process. Characteristics of the robust design methodology included: use of a criticality criterion based on a strain energy formulation to determine what loads were most critical to the structure, Latin Hypercube Sampling for the propagation of uncertainty to the criterion function, and redesign of the structure, using the new AAW design process, to the most critical loads identified. The demonstration of the methodology resulted in a wing box skin structure that was 11% heavier than an AAW structure designed only with linear aerodynamics. However, it was
NASA Technical Reports Server (NTRS)
Byun, Chansup; Guruswamy, Guru P.; Kutler, Paul (Technical Monitor)
1994-01-01
In recent years significant advances have been made for parallel computers in both hardware and software. Now parallel computers have become viable tools in computational mechanics. Many application codes developed on conventional computers have been modified to benefit from parallel computers. Significant speedups in some areas have been achieved by parallel computations. For single-discipline use of both fluid dynamics and structural dynamics, computations have been made on wing-body configurations using parallel computers. However, only a limited amount of work has been completed in combining these two disciplines for multidisciplinary applications. The prime reason is the increased level of complication associated with a multidisciplinary approach. In this work, procedures to compute aeroelasticity on parallel computers using direct coupling of fluid and structural equations will be investigated for wing-body configurations. The parallel computer selected for computations is an Intel iPSC/860 computer which is a distributed-memory, multiple-instruction, multiple data (MIMD) computer with 128 processors. In this study, the computational efficiency issues of parallel integration of both fluid and structural equations will be investigated in detail. The fluid and structural domains will be modeled using finite-difference and finite-element approaches, respectively. Results from the parallel computer will be compared with those from the conventional computers using a single processor. This study will provide an efficient computational tool for the aeroelastic analysis of wing-body structures on MIMD type parallel computers.
Real-Time Frequency Response Estimation Using Joined-Wing SensorCraft Aeroelastic Wind-Tunnel Data
NASA Technical Reports Server (NTRS)
Grauer, Jared A; Heeg, Jennifer; Morelli, Eugene A
2012-01-01
A new method is presented for estimating frequency responses and their uncertainties from wind-tunnel data in real time. The method uses orthogonal phase-optimized multi- sine excitation inputs and a recursive Fourier transform with a least-squares estimator. The method was first demonstrated with an F-16 nonlinear flight simulation and results showed that accurate short period frequency responses were obtained within 10 seconds. The method was then applied to wind-tunnel data from a previous aeroelastic test of the Joined- Wing SensorCraft. Frequency responses describing bending strains from simultaneous control surface excitations were estimated in a time-efficient manner.
NASA Technical Reports Server (NTRS)
Foss, Kenneth A; Diederich, Franklin W
1953-01-01
Charts and approximate formulas are presented for the estimation of static aeroelastic effects on the spanwise lift distribution, rolling-moment coefficient, and rate of roll due to the deflection of ailerons on swept and unswept wings at subsonic and supersonic speeds. Some design considerations brought out by the results of this report are discussed. This report treats the lateral-control case in a manner similar to that employed in NACA Report 1140 for the symmetric-flight case, and is intended to be used in conjunction with NACA Report 1140 and the charts and formulas presented therein.
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.
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)
Heeg, Jennifer; Chwalowski, Pawel; Wieseman, Carol D.; Florance, Jennifer P.; Schuster, David M.
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. The Rectangular Supercritical Wing (RSW) was chosen as the first configuration to study due to its geometric simplicity, perceived simple flow field at transonic conditions and availability of an experimental data set containing forced oscillation response data. Six teams performed analyses of the RSW; they used Reynolds-Averaged Navier-Stokes flow solvers exercised assuming that the wing had a rigid structure. Both steady-state and forced oscillation computations were performed by each team. The results of these calculations were compared with each other and with the experimental data. The steady-state results from the computations capture many of the flow features of a classical supercritical airfoil pressure distribution. The most dominant feature of the oscillatory results is the upper surface shock dynamics. Substantial variations were observed among the computational solutions as well as differences relative to the experimental data. Contributing issues to these differences include substantial wind tunnel wall effects and diverse choices in the analysis parameters.
NASA Technical Reports Server (NTRS)
Mullen, J., Jr.
1976-01-01
A comparison of program estimates of wing weight, material distribution. structural loads and elastic deformations with actual Northrop F-5A/B data is presented. Correlation coefficients obtained using data from a number of existing aircraft were computed for use in vehicle synthesis to estimate wing weights. The modifications necessary to adapt the WADES code for use in the ACSYNT program are described. Basic program flow and overlay structure is outlined. An example of the convergence of the procedure in estimating wing weights during the synthesis of a vehicle to satisfy F-5 mission requirements is given. A description of inputs required for use of the WADES program is included.
Aeroelasticity at the NASA Langley Research Center Recent progress, new challenges
NASA Technical Reports Server (NTRS)
Hanson, P. W.
1985-01-01
Recent progress in aeroelasticity, particularly at the NASA Langley Research Center is reviewed to look at the questions answered and questions raised, and to attempt to define appropriate research emphasis needed in the near future and beyond. The paper is focused primarily on the NASA Langley Research Center (LaRC) Program because Langley is the lead NASA center for aerospace structures research, and essentially is the only one working in depth in the area of aeroelasticity. Historical trends in aeroelasticity are reviewed broadly in terms of technology and staffing particularly at the LaRC. Then, selected studies of the Loads and Aeroelasticity Division at LaRC and others over the past three years are presented with attention paid to unresolved questions. Finally, based on the results of these studies and on perceptions of design trends and aircraft operational requirements, future research needs in aeroelasticity are discussed.
NASA Technical Reports Server (NTRS)
Nguyen, Nhan; Ting, Eric; Nguyen, Daniel; Dao, Tung; Trinh, Khanh
2013-01-01
This paper presents a coupled vortex-lattice flight dynamic model with an aeroelastic finite-element model to predict dynamic characteristics of a flexible wing transport aircraft. The aircraft model is based on NASA Generic Transport Model (GTM) with representative mass and stiffness properties to achieve a wing tip deflection about twice that of a conventional transport aircraft (10% versus 5%). This flexible wing transport aircraft is referred to as an Elastically Shaped Aircraft Concept (ESAC) which is equipped with a Variable Camber Continuous Trailing Edge Flap (VCCTEF) system for active wing shaping control for drag reduction. A vortex-lattice aerodynamic model of the ESAC is developed and is coupled with an aeroelastic finite-element model via an automated geometry modeler. This coupled model is used to compute static and dynamic aeroelastic solutions. The deflection information from the finite-element model and the vortex-lattice model is used to compute unsteady contributions to the aerodynamic force and moment coefficients. A coupled aeroelastic-longitudinal flight dynamic model is developed by coupling the finite-element model with the rigid-body flight dynamic model of the GTM.
NASA Technical Reports Server (NTRS)
1986-01-01
One of the most unusual experimental flight vehicles appearing at NASA's Ames-Dryden Flight Research Facility (later redesignated Dryden Flight Research Center) in the 1980s was the Rotor Systems Research Aircraft (RSRA) X-Wing aircraft, seen here on the ramp. The craft was developed originally and then modified by Sikorsky Aircraft for a joint NASA-Defense Advanced Research Projects Agency (DARPA) program and was rolled out 19 August 1986. Taxi tests and initial low-altitude flight tests without the main rotor attached were carried out at Dryden before the program was terminated in 1988. The unusual aircraft that resulted from the Ames Research Center/Army X-Wing Project was flown at the Ames-Dryden Flight Research Facility (now Dryden Flight Research Center), Edwards, California, beginning in the spring of 1984, with a follow-on program beginning in 1986. The program, was conceived to provide an efficient combination of the vertical lift characteristic of conventional helicopters and the high cruise speed of fixed-wing aircraft. It consisted of a hybrid vehicle called the NASA/Army Rotor Systems Research Aircraft (RSRA), which was equipped with advanced X-wing rotor systems. The program began in the early 1970s to investigate ways to increase the speed of rotor aircraft, as well as their performance, reliability, and safety . It also sought to reduce the noise, vibration, and maintenance costs of helicopters. Sikorsky Aircraft Division of United Technologies Laboratories built two RSRA aircraft. NASA's Langley Research Center, Hampton, Virginia, did some initial testing and transferred the program to Ames Research Center, Mountain View, California, for an extensive flight research program conducted by Ames and the Army. The purpose of the 1984 tests was to demonstrate the fixed-wing capability of the helicopter/airplane hybrid research vehicle and explore its flight envelope and flying qualities. These tests, flown by Ames pilot G. Warren Hall and Army Maj (soon
NASA Technical Reports Server (NTRS)
Heeg, Jennifer; Morelli, Eugene A.
2011-01-01
Multiple mutually orthogonal signals comprise excitation data sets for aeroservoelastic system identification. A multisine signal is a sum of harmonic sinusoid components. A set of these signals is made orthogonal by distribution of the frequency content such that each signal contains unique frequencies. This research extends the range of application of an excitation method developed for stability and control flight testing to aeroservoelastic modeling from wind tunnel testing. Wind tunnel data for the Joined Wing SensorCraft model validates this method, demonstrating that these signals applied simultaneously reproduce the frequency response estimates achieved from one-at-a-time excitation.
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.
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.
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)
Abel, Irving
1997-01-01
An overview of recently completed programs in aeroelasticity and structural dynamics research at the NASA Langley Research Center is presented. Methods used to perform flutter clearance studies in the wind-tunnel on a high performance fighter are discussed. Recent advances in the use of smart structures and controls to solve aeroelastic problems, including flutter and gust response are presented. An aeroelastic models program designed to support an advanced high speed civil transport is described. An extension to transonic small disturbance theory that better predicts flows involving separation and reattachment is presented. The results of a research study to determine the effects of flexibility on the taxi and takeoff characteristics of a high speed civil transport are presented. The use of photogrammetric methods aboard Space Shuttle to measure spacecraft dynamic response is discussed. Issues associated with the jitter response of multi-payload spacecraft are discussed. Finally a Space Shuttle flight experiment that studied the control of flexible spacecraft is described.
A historical overview of tiltrotor aeroelastic research at Langley Research Center
NASA Technical Reports Server (NTRS)
Kvaternik, Raymond G.
1992-01-01
The Bell/Boeing V-22 Osprey which is being developed for the U.S. Military is a tiltrotor aircraft combining the versatility of a helicopter with the range and speed of a turboprop airplane. The V-22 represents a tiltrotor lineage which goes back over forty years, during which time contributions to the technology base needed for its development were made by both government and industry. NASA Langley Research Center has made substantial contributions to tiltrotor technology in several areas, in particular in the area of aeroelasticity. The purpose of this talk is to present a summary of the tiltrotor aeroelastic research conducted at Langley which has contributed to that technology.
Summary Report of the Orbital X-34 Wing Static Aeroelastic Study
NASA Technical Reports Server (NTRS)
Prabhn, Ramadas K.; Weilmuenster, K. J. (Technical Monitor)
2001-01-01
This report documents the results of a computational study conducted on the Orbital Sciences X-34 vehicle to compute its inviscid aerodynamic characteristics taking into account the wing structural flexibility. This was a joint exercise between LaRC and SDRC of California. SDRC modeled the structural details of the wing, and provided the structural deformation for a given pressure distribution on its surfaces. This study was done for a Mach number of 1.35 and an angle of attack of 9 deg.; the freestream dynamic pressure was assumed to be 607 lb/sq ft. Only the wing and the body were simulated in the CFD computations. Two wing configurations were examined. The first had the elevons in the undeflected position and the second had the elevons deflected 20 deg. up. The results indicated that with elevon undeflected, the wing twists by about 1.5 deg. resulting in a reduction in the angle of attack at the wing tip to by 1.5 deg. The maximum vertical deflection of the wing is about 3.71 inches at the wing tip. For the wing with the undeflected elevons, the effect of this wing deformation is to reduce the normal force coefficient (C(sub N)) by 0.012 and introduce a noise up pitching moment coefficient (C(sub m)) of 0.042.
Unsteady-Pressure and Dynamic-Deflection Measurements on an Aeroelastic Supercritical Wing
NASA Technical Reports Server (NTRS)
Seidel, David A.; Sandford, Maynard C.; Eckstrom, Clinton V.
1991-01-01
Transonic steady and unsteady pressure tests were conducted on a large elastic wing. The wing has a supercritical airfoil, a full span aspect ratio of 10.3, a leading edge sweepback angle of 28.8 degrees, and two inboard and one outboard trailing edge control surfaces. Only the outboard control surface was deflected statically and dynamically to generate steady and unsteady flow over the wing. The unsteady surface pressure and dynamic deflection measurements of this elastic wing are presented to permit correlations of the experimental data with theoretical predictions.
Supersonic aeroelastic instability results for a NASP-like wing model
NASA Technical Reports Server (NTRS)
Cole, Stanley R.; Florance, James R.; Thomason, Lee B.; Spain, Charles V.; Bullock, Ellen P.
1993-01-01
An experimental study and an analytical study have been conducted to examine static divergence for hypersonic-vehicle wing models at supersonic conditions. A supersonic test in the Langley Unitary Plan Wind Tunnel facility was conducted for two wind-tunnel models. These models were nearly identical with the exception of airfoil shape. One model had a four-percent maximum thickness airfoil and the other model had an eight-percent maximum thickness airfoil. The wing models had low-aspect ratios and highly swept leading edges. The all-movable wing models were supported by a single-pivot mechanism along the wing root. For both of the wind-tunnel models, configuration changes could be made in the wing-pivot location along the wing root and in the wing-pivot pitch stiffness. Three divergence conditions were measured for the four-percent thick airfoil model in the Mach number range of 2.6 to 3.6 and one divergence condition was measured for the eight-percent thick airfoil model at a Mach number of 2.9. Analytical divergence calculations were made for comparison with experimental results and to evaluate the parametric effects of wing-pivot stiffness, wing-pivot location, and airfoil thickness variations. These analyses showed that decreasing airfoil thickness, moving the wing-pivot location upstream, or increasing the pitch-pivot stiffness have the beneficial effect of increasing the divergence dynamic pressures. The calculations predicted the trend of experimental divergence dynamic pressure with Mach number accurately; however, the calculations were approximately 25 percent conservative with respect to dynamic pressure.
An overview of selected NASP aeroelastic studies at the NASA Langley Research Center
NASA Technical Reports Server (NTRS)
Spain, Charles V.; Soistmann, David L.; Parker, Ellen C.; Gibbons, Michael D.; Gilbert, Michael G.
1990-01-01
Following an initial discussion of the NASP flight environment, the results of recent aeroelastic testing of NASP-type highly swept delta-wing models in Langley's Transonic Dynamics Tunnel (TDT) are summarized. Subsonic and transonic flutter characteristics of a variety of these models are described, and several analytical codes used to predict flutter of these models are evaluated. These codes generally provide good, but conservative predictions of subsonic and transonic flutter. Also, test results are presented on a nonlinear transonic phenomena known as aileron buzz which occurred in the wind tunnel on highly swept delta wings with full-span ailerons. An analytical procedure which assesses the effects of hypersonic heating on aeroelastic instabilities (aerothermoelasticity) is also described. This procedure accurately predicted flutter of a heated aluminum wing on which experimental data exists. Results are presented on the application of this method to calculate the flutter characteristics of a fine-element model of a generic NASP configuration. Finally, it is demonstrated analytically that active controls can be employed to improve the aeroelastic stability and ride quality of a generic NASP vehicle flying at hypersonic speeds.
Coupled finite-difference/finite-element approach for wing-body aeroelasticity
NASA Technical Reports Server (NTRS)
Guruswamy, Guru P.
1992-01-01
Computational methods using finite-difference approaches for fluids and finite-element approaches for structures have individually advanced to solve almost full-aircraft configurations. However, coupled approaches to solve fluid/structural interaction problems are still in their early stages of development, particularly for complex geometries using complete equations such as the Euler/Navier-Stokes equations. Earlier work demonstrated the success of coupling finite-difference and finite-element methods for simple wing configurations using the Euler/Navier-Stokes equations. In this paper, the same approach is extended for general wing-body configurations. The structural properties are represented by beam-type finite elements. The flow is modeled using the Euler/Navier-Stokes equations. A general procedure to fully couple structural finite-element boundary conditions with fluid finite-difference boundary conditions is developed for wing-body configurations. Computations are made using moving grids that adapt to wing-body structural deformations. Results are illustrated for a typical wing-body configuration.
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.
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 modeling of the active flexible wing wind-tunnel model
NASA Technical Reports Server (NTRS)
Silva, Walter A.; Heeg, Jennifer; Bennett, Robert M.
1991-01-01
The primary issues involved in the generation of linear, state-space equations of motion of a flexible wind tunnel model, the Active Flexible Wing (AFW), are discussed. The codes that were used and their inherent assumptions and limitations are also briefly discussed. The application of the CAP-TSD code to the AFW for determination of the model's transonic flutter boundary is included as well.
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.
NASA Technical Reports Server (NTRS)
Reed, W. H., III
1981-01-01
Testing of wind-tunnel aeroelastic models is a well established, widely used means of studying flutter trends, validating theory and investigating flutter margins of safety of new vehicle designs. The Langley Transonic Dynamics Tunnel was designed specifically for work on dynamics and aeroelastic problems of aircraft and space vehicles. A cross section of aeroelastic research and testing in the facility since it became operational more than two decades ago is presented. Examples selected from a large store of experience illustrate the nature and purpose of some major areas of work performed in the tunnel. These areas include: specialized experimental techniques; development testing of new aircraft and launch vehicle designs; evaluation of proposed "fixes" to solve aeroelastic problems uncovered during development testing; study of unexpected aeroelastic phenomena (i.e., "surprises"); control of aeroelastic effects by active and passive means; and, finally, fundamental research involving measurement of unsteady pressures on oscillating wings and control surface.
NASA Technical Reports Server (NTRS)
Manro, M. E.
1983-01-01
Two separated flow computer programs and a semiempirical method for incorporating the experimentally measured separated flow effects into a linear aeroelastic analysis were evaluated. The three dimensional leading edge vortex (LEV) code is evaluated. This code is an improved panel method for three dimensional inviscid flow over a wing with leading edge vortex separation. The governing equations are the linear flow differential equation with nonlinear boundary conditions. The solution is iterative; the position as well as the strength of the vortex is determined. Cases for both full and partial span vortices were executed. The predicted pressures are good and adequately reflect changes in configuration.
Aeroelastic Sizing for High-Speed Research (HSR) Longitudinal Control Alternatives Project (LCAP)
NASA Technical Reports Server (NTRS)
Walsh, Joanne L.; Dunn, H. J.; Stroud, W. Jefferson; Barthelemy, J.-F.; Weston, Robert P.; Martin, Carl J.; Bennett, Robert M.
2005-01-01
The Longitudinal Control Alternatives Project (LCAP) compared three high-speed civil transport configurations to determine potential advantages of the three associated longitudinal control concepts. The three aircraft configurations included a conventional configuration with a layout having a horizontal aft tail, a configuration with a forward canard in addition to a horizontal aft tail, and a configuration with only a forward canard. The three configurations were aeroelastically sized and were compared on the basis of operational empty weight (OEW) and longitudinal control characteristics. The sized structure consisted of composite honeycomb sandwich panels on both the wing and the fuselage. Design variables were the core depth of the sandwich and the thicknesses of the composite material which made up the face sheets of the sandwich. Each configuration was sized for minimum structural weight under linear and nonlinear aeroelastic loads subject to strain, buckling, ply-mixture, and subsonic and supersonic flutter constraints. This report describes the methods that were used and the results that were generated for the aeroelastic sizing of the three configurations.
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.
NASA Technical Reports Server (NTRS)
Gardner, J. E.; Dixon, S. C.
1986-01-01
The Langley Research Center Loads and Aeroelasticity Division's research accomplishments for FY85 and research plans for FY86 are presented. The rk under each branch (technical area) will be described in terms of highlights of accomplishments during the past year and highlights of plans for the current year as they relate to five year plans for each technical area. This information will be useful in program coordination with other government organizations and industry in areas of mutual interest.
NASA Technical Reports Server (NTRS)
Gardner, James E.; Dixon, S. C.
1987-01-01
The Loads and Aeroelasticity Division's research accomplishments for FY 86 and research plans for FY 87 are presented. The work under each Branch (technical area) is described in terms of highlights of accomplishments during the past year and highlights of plans for the current year as they relate to five year plans for each technical area. This information will be useful in program coordination with other government organizations and industry in areas of mutual interest.
NASA Technical Reports Server (NTRS)
Dixon, S. C.; Gardner, James E.
1988-01-01
The purpose of this paper is to present the Loads and Aeroelasticity Division's research accomplishments for FY87 and research plans for FY88. The work under each Branch (technical area) is described in terms of highlights of accomplishments during the past year and highlights of plans for the current year as they relate to five year plans for each technical area. This information will be useful in program coordination with other government organizations and industry in areas of mutual interest.
X-Wing Research Vehicle in Hangar
NASA Technical Reports Server (NTRS)
1987-01-01
One of the most unusual experimental flight vehicles appearing at NASA's Ames-Dryden Flight Research Facility (later redesignated Dryden Flight Research Center) in the 1980s was the Rotor Systems Research Aircraft (RSRA) X-Wing aircraft, seen here on the ramp. The craft was developed originally and then modified by Sikorsky Aircraft for a joint NASA-Defense Advanced Research Projects Agency (DARPA) program and was rolled out 19 August 1986. Taxi tests and initial low-altitude flight tests without the main rotor attached were carried out at Dryden before the program was terminated in 1988. The unusual aircraft that resulted from the Ames Research Center/Army X-Wing Project was flown at the Ames-Dryden Flight Research Facility (now Dryden Flight Research Center), Edwards, California, beginning in the spring of 1984, with a follow-on program beginning in 1986. The program, was conceived to provide an efficient combination of the vertical lift characteristic of conventional helicopters and the high cruise speed of fixed-wing aircraft. It consisted of a hybrid vehicle called the NASA/Army Rotor Systems Research Aircraft (RSRA), which was equipped with advanced X-wing rotor systems. The program began in the early 1970s to investigate ways to increase the speed of rotor aircraft, as well as their performance, reliability, and safety . It also sought to reduce the noise, vibration, and maintenance costs of helicopters. Sikorsky Aircraft Division of United Technologies Laboratories built two RSRA aircraft. NASA's Langley Research Center, Hampton, Virginia, did some initial testing and transferred the program to Ames Research Center, Mountain View, California, for an extensive flight research program conducted by Ames and the Army. The purpose of the 1984 tests was to demonstrate the fixed-wing capability of the helicopter/airplane hybrid research vehicle and explore its flight envelope and flying qualities. These tests, flown by Ames pilot G. Warren Hall and Army Maj (soon
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.
Technical activities of the configuration aeroelasticity branch
NASA Technical Reports Server (NTRS)
Cole, Stanley R. (Editor)
1991-01-01
A number of recent technical activities of the Configuration Aeroelasticity Branch of the NASA Langley Research Center are discussed in detail. The information on the research branch is compiled in twelve separate papers. The first of these topics is a summary of the purpose of the branch, including a full description of the branch and its associated projects and program efforts. The next ten papers cover specific projects and are as follows: Experimental transonic flutter characteristics of supersonic cruise configurations; Aeroelastic effects of spoiler surfaces mounted on a low aspect ratio rectangular wing; Planform curvature effects on flutter of 56 degree swept wing determined in Transonic Dynamics Tunnel (TDT); An introduction to rotorcraft testing in TDT; Rotorcraft vibration reduction research at the TDT; A preliminary study to determine the effects of tip geometry on the flutter of aft swept wings; Aeroelastic models program; NACA 0012 pressure model and test plan; Investigation of the use of extension twist coupling in composite rotor blades; and Improved finite element methods for rotorcraft structures. The final paper describes the primary facility operation by the branch, the Langley TDT.
Subsonic Ultra Green Aircraft Research. Phase II - Volume I; Truss Braced Wing Design Exploration
NASA Technical Reports Server (NTRS)
Bradley, Marty K.; Droney, Christopher K.; Allen, Timothy J.
2015-01-01
This report summarizes the Truss Braced Wing (TBW) work accomplished by the Boeing Subsonic Ultra Green Aircraft Research (SUGAR) team, consisting of Boeing Research and Technology, Boeing Commercial Airplanes, General Electric, Georgia Tech, Virginia Tech, NextGen Aeronautics, and Microcraft. A multi-disciplinary optimization (MDO) environment defined the geometry that was further refined for the updated SUGAR High TBW configuration. Airfoil shapes were tested in the NASA TCT facility, and an aeroelastic model was tested in the NASA TDT facility. Flutter suppression was successfully demonstrated using control laws derived from test system ID data and analysis models. Aeroelastic impacts for the TBW design are manageable and smaller than assumed in Phase I. Flutter analysis of TBW designs need to include pre-load and large displacement non-linear effects to obtain a reasonable match to test data. With the updated performance and sizing, fuel burn and energy use is reduced by 54% compared to the SUGAR Free current technology Baseline (Goal 60%). Use of the unducted fan version of the engine reduces fuel burn and energy by 56% compared to the Baseline. Technology development roadmaps were updated, and an airport compatibility analysis established feasibility of a folding wing aircraft at existing airports.
Aeroelastic Tailoring for Stability Augmentation and Performance Enhancements of Tiltrotor Aircraft
NASA Technical Reports Server (NTRS)
Nixon, Mark W.; Piatak, David J.; Corso, Lawrence M.; Popelka, David A.
1999-01-01
The requirements for increased speed and productivity for tiltrotors has spawned several investigations associated with proprotor aeroelastic stability augmentation and aerodynamic performance enhancements. Included among these investigations is a focus on passive aeroelastic tailoring concepts which exploit the anisotropic capabilities of fiber composite materials. Researchers at Langley Research Center and Bell Helicopter have devoted considerable effort to assess the potential for using these materials to obtain aeroelastic responses which are beneficial to the important stability and performance considerations of tiltrotors. Both experimental and analytical studies have been completed to examine aeroelastic tailoring concepts for the tiltrotor, applied either to the wing or to the rotor blades. This paper reviews some of the results obtained in these aeroelastic tailoring investigations and discusses the relative merits associated with these approaches.
NASA Technical Reports Server (NTRS)
Gardner, J. E.; Dixon, S. C.
1984-01-01
Research was done in the following areas: development and validation of solution algorithms, modeling techniques, integrated finite elements for flow-thermal-structural analysis and design, optimization of aircraft and spacecraft for the best performance, reduction of loads and increase in the dynamic structural stability of flexible airframes by the use of active control, methods for predicting steady and unsteady aerodynamic loads and aeroelastic characteristics of flight vehicles with emphasis on the transonic range, and methods for predicting and reducing helicoper vibrations.
Development of Advanced Computational Aeroelasticity Tools at NASA Langley Research Center
NASA Technical Reports Server (NTRS)
Bartels, R. E.
2008-01-01
NASA Langley Research Center has continued to develop its long standing computational tools to address new challenges in aircraft and launch vehicle design. This paper discusses the application and development of those computational aeroelastic tools. Four topic areas will be discussed: 1) Modeling structural and flow field nonlinearities; 2) Integrated and modular approaches to nonlinear multidisciplinary analysis; 3) Simulating flight dynamics of flexible vehicles; and 4) Applications that support both aeronautics and space exploration.
NASA Technical Reports Server (NTRS)
Gardner, J. E.; Dixon, S. C.
1985-01-01
The loads and aeroelasticity divisions research accomplishments are presented. The work under each branch or technical area, described in terms of highlights of accomplishments during the past year and highlights of plans for the current year as they relate to 5 year plans for each technical area. This information will be useful in program coordination with other government organizations and industry in areas of mutual interest.
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.
Investigation of the aeroelastic stability of the AFW wind-tunnel model using CAP-TSD
NASA Technical Reports Server (NTRS)
Silva, Walter A.; Bennett, Robert M.
1991-01-01
The Computational Aeroelasticity Program - Transonic Small Disturbance (CAP-TSD) code, developed at the NASA Langley Research Center, is applied to the Active Flexible Wing (AFW) wind tunnel model for prediction of the model's transonic aeroelastic behavior. A semi-span computational model is used for evaluation of symmetric motions and a full span model is used for evaluation of antisymmetric motions. Static aeroelastic solutions using CAP-TSD are computed. Dynamic (flutter) analyses are then performed as perturbations about the static aeroelastic deformations and presented as flutter boundaries in terms of Mach number and dynamic pressure. Flutter boundaries that take into account modal refinements, vorticity and entropy corrections, antisymmetric motions and sensitivity to the modeling of the wing tip ballast stores are also presented and compared with experimental flutter results.
Predicting the aeroelastic behavior of a wind-tunnel model using transonic small disturbance theory
NASA Technical Reports Server (NTRS)
Silva, Walter A.; Bennett, Robert M.
1990-01-01
The CAP-TSD (Computational Aeroelasticity Program - Transonic Small Disturbance) code, developed at the NASA-Langley Research Center, is applied to the Active Flexible Wing (AFW) wind-tunnel model for prediction of the model's transonic aeroelastic behavior. Static aeroelastic solutions using CAP-TSD are computed. Dynamic (flutter) analyses are then performed as perturbations about the static aeroelastic deformations of the AFW. The accuracy of the static aeroelastic procedure is investigated by comparing analytical results to those from AFW wind-tunnel experiments. Dynamic results are presented in the form of root loci at different Mach numbers for a heavy gas and for air test mediums. The resultant flutter boundaries for both gases, and the effects of viscous damping and angle of attack on the flutter boundary in air, are also presented.
Selected topics in experimental aeroelasticity at the NASA Langley Research Center
NASA Technical Reports Server (NTRS)
Ricketts, R. H.
1985-01-01
The results of selected studies that have been conducted by the NASA Langley Research Center in the last three years are presented. The topics presented focus primarily on the ever-important transonic flight regime and include the following: body-freedom flutter of a forward-swept-wing configuration with and without relaxed static stability; instabilities associated with a new tilt-rotor vehicle; effects of winglets, supercritical airfoils, and spanwise curvature on wing flutter; wind-tunnel investigation of a flutter-like oscillation on a high-aspect-ratio flight research wing; results of wing-tunnel demonstration of the NASA decoupler pylon concept for passive suppression of wing/store flutter; and, new flutter testing methods which include testing at cryogenic temperatures for full scale Reynolds number simulation, subcritical response techniques for predicting onset of flutter, and a two-degree-of-freedom mount system for testing side-wall-mounted models.
A CFD/CSD Interaction Methodology for Aircraft Wings
NASA Technical Reports Server (NTRS)
Bhardwaj, Manoj K.
1997-01-01
With advanced subsonic transports and military aircraft operating in the transonic regime, it is becoming important to determine the effects of the coupling between aerodynamic loads and elastic forces. Since aeroelastic effects can contribute significantly to the design of these aircraft, there is a strong need in the aerospace industry to predict these aero-structure interactions computationally. To perform static aeroelastic analysis in the transonic regime, high fidelity computational fluid dynamics (CFD) analysis tools must be used in conjunction with high fidelity computational structural fluid dynamics (CSD) analysis tools due to the nonlinear behavior of the aerodynamics in the transonic regime. There is also a need to be able to use a wide variety of CFD and CSD tools to predict these aeroelastic effects in the transonic regime. Because source codes are not always available, it is necessary to couple the CFD and CSD codes without alteration of the source codes. In this study, an aeroelastic coupling procedure is developed which will perform static aeroelastic analysis using any CFD and CSD code with little code integration. The aeroelastic coupling procedure is demonstrated on an F/A-18 Stabilator using NASTD (an in-house McDonnell Douglas CFD code) and NASTRAN. In addition, the Aeroelastic Research Wing (ARW-2) is used for demonstration of the aeroelastic coupling procedure by using ENSAERO (NASA Ames Research Center CFD code) and a finite element wing-box code (developed as part of this research).
An analytical study of effects of aeroelasticity on control effectiveness
NASA Technical Reports Server (NTRS)
Mehrotra, S. C.
1975-01-01
Structural influence coefficients were calculated for various wing planforms using the KU Aeroelastic and NASTRAN programs. The resulting matrices are compared with experimental results. Conclusions are given.
NASA Technical Reports Server (NTRS)
Kelly, G. L.; Berthold, G.; Abbott, L.
1982-01-01
A 5 MHZ single-board microprocessor system which incorporates an 8086 CPU and an 8087 Numeric Data Processor is used to implement the control laws for the NASA Drones for Aerodynamic and Structural Testing, Aeroelastic Research Wing II. The control laws program was executed in 7.02 msec, with initialization consuming 2.65 msec and the control law loop 4.38 msec. The software emulator execution times for these two tasks were 36.67 and 61.18, respectively, for a total of 97.68 msec. The space, weight and cost reductions achieved in the present, aircraft control application of this combination of a 16-bit microprocessor with an 80-bit floating point coprocessor may be obtainable in other real time control applications.
Selected topics in experimental aeroelasticity at the NASA Langley Research Center
NASA Technical Reports Server (NTRS)
Ricketts, R. H.
1985-01-01
The results of selected studies that have been conducted by the NASA Langley Research Center in the last three years are presented. The topics presented focus primarily on the ever-important transonic flight regime and include the following: body-freedom flutter of a forward-swept-wing configuration with and without relaxed static stability; instabilities associated with a new tilt-rotor vehicle; effects of winglets, supercritical airfoils, and spanwise curvature on wing flutter; wind-tunnel investigation of a flutter-like oscillation on a high-aspect-ratio flight research wing; results of wind-tunnel demonstration of the NASA decoupler pylon concept for passive suppression of wing/store flutter; and, new flutter testing methods which include testing at cryogenic temperatures for full scale Reynolds number simulation, subcritical response techniques for predicting onset of flutter, and a two-degree-of-freedom mount system for testing side-wall-mounted models.
Aeroelastic, CFD, and Dynamics Computation and Optimization for Buffet and Flutter Applications
NASA Technical Reports Server (NTRS)
Kandil, Osama A.
1997-01-01
Accomplishments achieved during the reporting period are listed. These accomplishments included 6 papers published in various journals or presented at various conferences; 1 abstract submitted to a technical conference; production of 2 animated movies; and a proposal for use of the National Aerodynamic Simulation Facility at NASA Ames Research Center for further research. The published and presented papers and animated movies addressed the following topics: aeroelasticity, computational fluid dynamics, structural dynamics, wing and tail buffet, vortical flow interactions, and delta wings.
NASA Technical Reports Server (NTRS)
Silva, Walter A.; Bennett, Robert M.
1992-01-01
The CAP-TSD (Computational Aeroelasticity Program - Transonic Small Disturbance) code, developed at the NASA Langley Research Center, is applied to the Active Flexible Wing wind-tunnel model for prediction of transonic aeroelastic behavior. A semi-span computational model is used for evaluation of symmetric motions, and a full-span model is used for evaluation of antisymmetric motions. Static aeroelastic solutions using CAP-TSD are computed. Dynamic (flutter) analyses then are performed as perturbations about the static aeroelastic deformations and presented as flutter boundaries in terms of Mach number and dynamic pressure. Flutter boundaries that take into account modal refinements, vorticity and entropy corrections, antisymmetric motions and sensitivity to the modeling of the wing tip ballast stores also are presented and compared with experimental flutter results.
NASA Technical Reports Server (NTRS)
Silva, Walter A.; Bennett, Robert M.
1990-01-01
The CAP-TSD (Computational Aeroelasticity Program - Transonic Small Disturbance) code, developed at the NASA - Langley Research Center, is applied to the Active Flexible Wing (AFW) wind tunnel model for prediction of the model's transonic aeroelastic behavior. Static aeroelastic solutions using CAP-TSD are computed. Dynamic (flutter) analyses are then performed as perturbations about the static aeroelastic deformations of the AFW. The accuracy of the static aeroelastic procedure is investigated by comparing analytical results to those from previous AFW wind tunnel experiments. Dynamic results are presented in the form of root loci at different Mach numbers for a heavy gas and air. The resultant flutter boundaries for both gases are also presented. The effects of viscous damping and angle-of-attack, on the flutter boundary in air, are presented as well.
A CFD/CSD interaction methodology for aircraft wings
NASA Astrophysics Data System (ADS)
Bhardwaj, Manoj Kumar
With advanced subsonic transports and military aircraft operating in the transonic regime, it is becoming important to determine the effects of the coupling between aerodynamic loads and elastic forces. Since aeroelastic effects can contribute significantly to the design of these aircraft, there is a strong need in the aerospace industry to predict these aero-structure interactions computationally. To perform static aeroelastic analysis in the transonic regime, high fidelity computational fluid dynamics (CFD) analysis tools must be used in conjunction with high fidelity computational structural dynamics (CSD) analysis tools due to the nonlinear behavior of the aerodynamics in the transonic regime. There is also a need to be able to use a wide variety of CFD and CSD tools to predict these aeroelastic effects in the transonic regime. Because source codes are not always available, it is necessary to couple the CFD and CSD codes without alteration of the source codes. In this study, an aeroelastic coupling procedure is developed which will perform static aeroelastic analysis using any CFD and CSD code with little code integration. The aeroelastic coupling procedure is demonstrated on an F/A-18 Stabilator using NASTD (an in-house McDonnell Douglas CFD code) and NASTRAN. In addition, the Aeroelastic Research Wing (ARW-2) is used for demonstration of the aeroelastic coupling procedure by using ENSAERO (NASA Ames Research Center CFD code) and a finite element wing-box code (developed as a part of this research). The results obtained from the present study are compared with those available from an experimental study conducted at NASA Langley Research Center and a study conducted at NASA Ames Research Center using ENSAERO and modal superposition. The results compare well with experimental data. Parallel computing power is used to investigate parallel static aeroelastic analysis because obtaining an aeroelastic solution using CFD/CSD methods is computationally intensive. A
Recent Applications of the Volterra Theory to Aeroelastic Phenomena
NASA Technical Reports Server (NTRS)
Silva, Walter A.; Haji, Muhammad R; Prazenica, Richard J.
2005-01-01
The identification of nonlinear aeroelastic systems based on the Volterra theory of nonlinear systems is presented. Recent applications of the theory to problems in experimental aeroelasticity are reviewed. These results include the identification of aerodynamic impulse responses, the application of higher-order spectra (HOS) to wind-tunnel flutter data, and the identification of nonlinear aeroelastic phenomena from flight flutter test data of the Active Aeroelastic Wing (AAW) aircraft.
NASA Technical Reports Server (NTRS)
Kandil, Osama A.
1993-01-01
Research on Navier-Stokes, dynamics, and aeroelastic computations for vortical flows, buffet, and flutter applications was performed. Progress during the period from 1 Oct. 1992 to 30 Sep. 1993 is included. Papers on the following topics are included: vertical tail buffet in vortex breakdown flows; simulation of tail buffet using delta wing-vertical tail configuration; shock-vortex interaction over a 65-degree delta wing in transonic flow; supersonic vortex breakdown over a delta wing in transonic flow; and prediction and control of slender wing rock.
A Taguchi study of the aeroelastic tailoring design process
NASA Technical Reports Server (NTRS)
Bohlmann, Jonathan D.; Scott, Robert C.
1991-01-01
A Taguchi study was performed to determine the important players in the aeroelastic tailoring design process and to find the best composition of the optimization's objective function. The Wing Aeroelastic Synthesis Procedure (TSO) was used to ascertain the effects that factors such as composite laminate constraints, roll effectiveness constraints, and built-in wing twist and camber have on the optimum, aeroelastically tailored wing skin design. The results show the Taguchi method to be a viable engineering tool for computational inquiries, and provide some valuable lessons about the practice of aeroelastic tailoring.
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.
Lift on Flexible and Rigid Cambered Wings at High Incidence
NASA Astrophysics Data System (ADS)
Jones, Anya; Mancini, Peter; Granlund, Kenneth; Ol, Michael
2014-11-01
The effects of camber and camber change due to elastic deflection of a membrane wing were investigated for wings in rectilinear translation with parameter variations in wing incidence and acceleration. Direct force and moment measurements were performed on a rigid flat plate wing, rigid cambered wings, and a membrane wing. Features in the force histories were further examined via flow visualization by planar laser illumination of fluorescent dye. Below 10 degrees of incidence, Wagner's approximation accurately predicts the time-evolution of lift for the rigid wings. At higher incidence, flow separation results in force transients, and the effect of wing camber is no longer additive. Both the rigid flat plate and rigid cambered wings reach peak lift at a 35 degree angle of attack, whereas the flexible wing experiences stall delay and reaches peak lift at 50 degrees. Due to the aeroelasticity of the flexible membrane, flow over the suction surface remains attached for much higher incidence angles than for the rigid wings. For incidence angles less than 30 degrees, the peak lift of the flexible wing is lower than that of its rigid counterparts. Beyond 30 degrees, the flexible wing experiences an aeroelastically induced stall delay that allows lift to exceed the rigid analogs. This work was supported by the Air Force Office of Scientific Research (AFOSR) Summer Faculty Fellowship Program and the U.S. Army Research Laboratory under the Micro Autonomous Systems and Technology (MAST) program.
The oblique wing-research aircraft
NASA Technical Reports Server (NTRS)
Andrews, W. H.
1980-01-01
The AD-1 airplane was designed as a low cost, low speed manned research tool to evaluate the flying qualities of the oblique wing concept. The airplane is constructed primarily of foam and fiberglass and incorporates simplicity in terms of the onboard systems. There are no hydraulics, the control system is cable and torque tube, and the electrical systems consist of engine driven generators which power the battery for engine start, cockpit gages, trim motors, and the onboard data system. The propulsion systems consist of two Microturbo TRS-18 engines sea level trust rated at 220 pounds. The airplane weighs approximately 2100 pounds and has a performance potential in the range of 200 knots and an altitude of 15,000 feet.
Plans and Example Results for the 2nd AIAA Aeroelastic Prediction Workshop
NASA Technical Reports Server (NTRS)
Heeg, Jennifer; Chwalowski, Pawel; Schuster, David M.; Raveh, Daniella; Jirasek, Adam; Dalenbring, Mats
2015-01-01
This paper summarizes the plans for the second AIAA Aeroelastic Prediction Workshop. The workshop is designed to assess the state-of-the-art of computational methods for predicting unsteady flow fields and aeroelastic response. The goals are to provide an impartial forum to evaluate the effectiveness of existing computer codes and modeling techniques, and to identify computational and experimental areas needing additional research and development. This paper provides guidelines and instructions for participants including the computational aerodynamic model, the structural dynamic properties, the experimental comparison data and the expected output data from simulations. The Benchmark Supercritical Wing (BSCW) has been chosen as the configuration for this workshop. The analyses to be performed will include aeroelastic flutter solutions of the wing mounted on a pitch-and-plunge apparatus.
FUN3D Analyses in Support of the Second Aeroelastic Prediction Workshop
NASA Technical Reports Server (NTRS)
Chwalowski, Pawel; Heeg, Jennifer
2016-01-01
This paper presents the computational aeroelastic results generated in support of the second Aeroelastic Prediction Workshop for the Benchmark Supercritical Wing (BSCW) configurations and compares them to the experimental data. The computational results are obtained using FUN3D, an unstructured grid Reynolds- Averaged Navier-Stokes solver developed at NASA Langley Research Center. The analysis results include aerodynamic coefficients and surface pressures obtained for steady-state, static aeroelastic equilibrium, and unsteady flow due to a pitching wing or flutter prediction. Frequency response functions of the pressure coefficients with respect to the angular displacement are computed and compared with the experimental data. The effects of spatial and temporal convergence on the computational results are examined.
NASA 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.
Some experiences with active control of aeroelastic response
NASA Technical Reports Server (NTRS)
Newsom, J. R.; Abel, I.
1981-01-01
Flight and wind tunnel tests were conducted and multidiscipline computer programs were developed as part of investigations of active control technology conducted at the NASA Langley Research Center. Unsteady aerodynamics approximation, optimal control theory, optimal controller design, and the Delta wing and DC-10 models are described. The drones for aerodynamics and structural testing (DAST program) for evaluating procedures for aerodynamic loads prediction and the design of active control systems on wings with significant aeroelastic effects is described as well as the DAST model used in the wind tunnel tests.
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.
Benchmark Active Controls Technology (BACT) Wing CFD Results
NASA Technical Reports Server (NTRS)
Schuster, David M.; Bartels, Robert E.
2000-01-01
The Benchmark Active Controls Technology (BACT) wing test (see chapter 8E) provides data for the validation of aerodynamic, aeroelastic, and active aeroelastic control simulation codes. These data provide a rich database for development and validation of computational aeroelastic and aeroservoelastic methods. In this vein, high-level viscous CFD analyses of the BACT wing have been performed for a subset of the test conditions available in the dataset. The computations presented in this section investigate the aerodynamic characteristics of the rigid clean wing configuration as well as simulations of the wing with a static and oscillating aileron and spoiler deflection. Two computational aeroelasticity codes extensively used at NASA Langley Research Center are implemented in this simulation. They are the ENS3DAE and CFL3DAE computational aeroelasticity programs. Both of these methods solve the three-dimensional compressible Navier-Stokes equations for both rigid and flexible vehicles, but they use significantly different approaches to the solution 6f the aerodynamic equations of motion. Detailed descriptions of both methods are presented in the following section.
Aeroelastic Tailoring via Tow Steered Composites
NASA Technical Reports Server (NTRS)
Stanford, Bret K.; Jutte, Christine V.
2014-01-01
The use of tow steered composites, where fibers follow prescribed curvilinear paths within a laminate, can improve upon existing capabilities related to aeroelastic tailoring of wing structures, though this tailoring method has received relatively little attention in the literature. This paper demonstrates the technique for both a simple cantilevered plate in low-speed flow, as well as the wing box of a full-scale high aspect ratio transport configuration. Static aeroelastic stresses and dynamic flutter boundaries are obtained for both cases. The impact of various tailoring choices upon the aeroelastic performance is quantified: curvilinear fiber steering versus straight fiber steering, certifiable versus noncertifiable stacking sequences, a single uniform laminate per wing skin versus multiple laminates, and identical upper and lower wing skins structures versus individual tailoring.
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.
A CFD/CSD interaction methodology for aircraft wings
Bhardwaj, M.K.; Kapania, R.K.; Reichenbach, E.; Guruswamy, G.P.
1998-01-01
With advanced subsonic transports and military aircraft operating in the transonic regime, it is becoming important to determine the effects of the coupling between aerodynamic loads and elastic forces. Since aeroelastic effects can significantly impact the design of these aircraft, there is a strong need in the aerospace industry to predict these interactions computationally. Such an analysis in the transonic regime requires high fidelity computational fluid dynamics (CFD) analysis tools, due to the nonlinear behavior of the aerodynamics in the transonic regime and also high fidelity computational structural dynamics (CSD) analysis tools. Also, there is a need to be able to use a wide variety of CFD and CSD methods to predict aeroelastic effects. Since source codes are not always available, it is necessary to couple the CFD and CSD codes without alteration of the source codes. In this study, an aeroelastic coupling procedure is developed to determine the static aeroelastic response of aircraft wings using any CFD and CSD code with little code integration. The aeroelastic coupling procedure is demonstrated on an F/A-18 Stabilator using NASTD (an in-house McDonnell Douglas CFD code) and NASTRAN. In addition, the Aeroelastic Research Wing (ARW-2) is used for demonstration of the aeroelastic coupling procedure by using ENSAERO (NASA Ames Research Center CFD code) and a finite element wing-box code. The results obtained from the present study are compared with those available from an experimental study conducted at NASA Langley Research Center and a study conducted at NASA Ames Research Center using ENSAERO and modal superposition. The results compare well with experimental data.
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.
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.
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
Aeroelastic Optimization Study Based on the X-56A Model
NASA Technical Reports Server (NTRS)
Li, Wesley W.; Pak, Chan-Gi
2014-01-01
One way to increase the aircraft fuel efficiency is to reduce structural weight while maintaining adequate structural airworthiness, both statically and aeroelastically. A design process which incorporates the object-oriented multidisciplinary design, analysis, and optimization (MDAO) tool and the aeroelastic effects of high fidelity finite element models to characterize the design space was successfully developed and established. This paper presents two multidisciplinary design optimization studies using an object-oriented MDAO tool developed at NASA Armstrong Flight Research Center. The first study demonstrates the use of aeroelastic tailoring concepts to minimize the structural weight while meeting the design requirements including strength, buckling, and flutter. Such an approach exploits the anisotropic capabilities of the fiber composite materials chosen for this analytical exercise with ply stacking sequence. A hybrid and discretization optimization approach improves accuracy and computational efficiency of a global optimization algorithm. The second study presents a flutter mass balancing optimization study for the fabricated flexible wing of the X-56A model since a desired flutter speed band is required for the active flutter suppression demonstration during flight testing. The results of the second study provide guidance to modify the wing design and move the design flutter speeds back into the flight envelope so that the original objective of X-56A flight test can be accomplished successfully. The second case also demonstrates that the object-oriented MDAO tool can handle multiple analytical configurations in a single optimization run.
Aeroelastic Optimization Study Based on X-56A Model
NASA Technical Reports Server (NTRS)
Li, Wesley; Pak, Chan-Gi
2014-01-01
A design process which incorporates the object-oriented multidisciplinary design, analysis, and optimization (MDAO) tool and the aeroelastic effects of high fidelity finite element models to characterize the design space was successfully developed and established. Two multidisciplinary design optimization studies using an object-oriented MDAO tool developed at NASA Armstrong Flight Research Center were presented. The first study demonstrates the use of aeroelastic tailoring concepts to minimize the structural weight while meeting the design requirements including strength, buckling, and flutter. A hybrid and discretization optimization approach was implemented to improve accuracy and computational efficiency of a global optimization algorithm. The second study presents a flutter mass balancing optimization study. The results provide guidance to modify the fabricated flexible wing design and move the design flutter speeds back into the flight envelope so that the original objective of X-56A flight test can be accomplished.
NASA Technical Reports Server (NTRS)
Edwards, John W.; Malone, John B.
1992-01-01
The current status of computational methods for unsteady aerodynamics and aeroelasticity is reviewed. The key features of challenging aeroelastic applications are discussed in terms of the flowfield state: low-angle high speed flows and high-angle vortex-dominated flows. The critical role played by viscous effects in determining aeroelastic stability for conditions of incipient flow separation is stressed. The need for a variety of flow modeling tools, from linear formulations to implementations of the Navier-Stokes equations, is emphasized. Estimates of computer run times for flutter calculations using several computational methods are given. Applications of these methods for unsteady aerodynamic and transonic flutter calculations for airfoils, wings, and configurations are summarized. Finally, recommendations are made concerning future research directions.
NASA Technical Reports Server (NTRS)
Edwards, John W.; Malone, John B.
1992-01-01
The status of computational methods for unsteady aerodynamics and aeroelasticity is reviewed. The key features of challenging aeroelastic applications is discussed in terms of the flowfield state - low angle high speed flows and high angle vortex dominated flows. The critical role played by viscous effects in determining aeroelastic stability for conditions of incipient flow separation is stressed. The need for a variety of flow modeling tools, from linear formulations to implementations of the Navier-Stokes equations, is emphasized. Estimates of computer run times for flutter calculations using several computational methods are given. Applications of these methods for unsteady aerodynamic and transonic flutter calculations for airfoils, wings, and configurations are summarized. Finally, recommendations are made concerning future research directions.
NASA Technical Reports Server (NTRS)
Edwards, J. W.; Malone, J. B.
1992-01-01
The current status of computational methods for unsteady aerodynamics and aeroelasticity is reviewed. The key features of challenging aeroelastic applications are discussed in terms of the flowfield state: low-angle high speed flows and high-angle vortex-dominated flows. The critical role played by viscous effects in determining aeroelastic stability for conditions of incipient flow separation is stressed. The need for a variety of flow modeling tools, from linear formulations to implementations of the Navier-Stokes equations, is emphasized. Estimates of computer run times for flutter calculations using several computational methods are given. Application of these methods for unsteady aerodynamic and transonic flutter calculations for airfoils, wings, and configurations are summarized. Finally, recommendations are made concerning future research directions.
FUN3D Analyses in Support of the First Aeroelastic Prediction Workshop
NASA Technical Reports Server (NTRS)
Chwalowski, Pawel; Heeg, Jennifer; Wieseman, Carol D.; Florance, Jennifer P.
2013-01-01
This paper presents the computational aeroelastic results generated in support of the first Aeroelastic Prediction Workshop for the Benchmark Supercritical Wing (BSCW) and the HIgh REynolds Number AeroStructural Dynamics (HIRENASD) configurations and compares them to the experimental data. The computational results are obtained using FUN3D, an unstructured grid Reynolds-averaged Navier-Stokes solver developed at NASA Langley Research Center. The analysis results for both configurations include aerodynamic coefficients and surface pressures obtained for steady-state or static aeroelastic equilibrium (BSCW and HIRENASD, respectively) and for unsteady flow due to a pitching wing (BSCW) or modally-excited wing (HIRENASD). Frequency response functions of the pressure coefficients with respect to displacement are computed and compared with the experimental data. For the BSCW, the shock location is computed aft of the experimentally-located shock position. The pressure distribution upstream of this shock is in excellent agreement with the experimental data, but the pressure downstream of the shock in the separated flow region does not match as well. For HIRENASD, very good agreement between the numerical results and the experimental data is observed at the mid-span wing locations.
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
Oblique Wing Remotely Piloted Research Aircraft. Volume 1: Development
NASA Technical Reports Server (NTRS)
1974-01-01
The NASA Ames/DSI oblique wing remotely piloted research aircraft is a highly unusual, variable remotely piloted vehicle whose configuration and capabilities are the result of certain initial design guidelines that, in terms of conventional aircraft structures and configurations, would be considered to be contradictory and unachievable. Accordingly, the novel design of the yawed wing RPV is at odds in many respects with conventional aircraft practice. Novelty, then, forms the first, unwritten, design guideline. This design is intended to move away from convention in geometry, structure, and materials. The specific guidelines followed in the design of the yawed wing RPV and a short discussion of the impact of each on the configuration of the vehicle are presented.
In-flight aeroelastic measurement technique development
NASA Astrophysics Data System (ADS)
Burner, Alpheus W.; Lokos, William A.; Barrows, Danny A.
2003-11-01
The initial concept and development of a low-cost, adaptable method for the measurement of static and dynamic aeroelastic deformation of aircraft during flight testing is presented. The method is adapted from a proven technique used in wind tunnel testing to measure model deformation, often referred to as the videogrammetric model deformation (or VMD) technique. The requirements for in-flight measurements are compared and contrasted with those for wind tunnel testing. The methodology for the proposed measurements and differences compared with that used for wind tunnel testing is given. Several error sources and their effects are identified. Measurement examples using the new technique, including change in wing twist and deflection as a function of time, from an F/A-18 research aircraft at NASA's Dryden Flight Research Center are presented.
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.
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.
NASA Technical Reports Server (NTRS)
Ippolito, Corey; Nguyen, Nhan; Lohn, Jason; Dolan, John
2014-01-01
The emergence of advanced lightweight materials is resulting in a new generation of lighter, flexible, more-efficient airframes that are enabling concepts for active aeroelastic wing-shape control to achieve greater flight efficiency and increased safety margins. These elastically shaped aircraft concepts require non-traditional methods for large-scale multi-objective flight control that simultaneously seek to gain aerodynamic efficiency in terms of drag reduction while performing traditional command-tracking tasks as part of a complete guidance and navigation solution. This paper presents results from a preliminary study of a notional multi-objective control law for an aeroelastic flexible-wing aircraft controlled through distributed continuous leading and trailing edge control surface actuators. This preliminary study develops and analyzes a multi-objective control law derived from optimal linear quadratic methods on a longitudinal vehicle dynamics model with coupled aeroelastic dynamics. The controller tracks commanded attack-angle while minimizing drag and controlling wing twist and bend. This paper presents an overview of the elastic aircraft concept, outlines the coupled vehicle model, presents the preliminary control law formulation and implementation, presents results from simulation, provides analysis, and concludes by identifying possible future areas for research
Wing Classification in the Virtual Research Center
NASA Technical Reports Server (NTRS)
Campbell, William H.
1999-01-01
The Virtual Research Center (VRC) is a Web site that hosts a database of documents organized to allow teams of scientists and engineers to store and maintain documents. A number of other workgroup-related capabilities are provided. My tasks as a NASA/ASEE Summer Faculty Fellow included developing a scheme for classifying the workgroups using the VRC using the various Divisions within NASA Enterprises. To this end I developed a plan to use several CGI Perl scripts to gather classification information from the leaders of the workgroups, and to display all the workgroups within a specified classification. I designed, implemented, and partially tested scripts which can be used to do the classification. I was also asked to consider directions for future development of the VRC. I think that the VRC can use XML to advantage. XML is a markup language with designer tags that can be used to build meaning into documents. An investigation as to how CORBA, an object-oriented object request broker included with JDK 1.2, might be used also seems justified.
Role of computational fluid dynamics in unsteady aerodynamics for aeroelasticity
NASA Technical Reports Server (NTRS)
Guruswamy, Guru P.; Goorjian, Peter M.
1989-01-01
In the last two decades there have been extensive developments in computational unsteady transonic aerodynamics. Such developments are essential since the transonic regime plays an important role in the design of modern aircraft. Therefore, there has been a large effort to develop computational tools with which to accurately perform flutter analysis at transonic speeds. In the area of Computational Fluid Dynamics (CFD), unsteady transonic aerodynamics are characterized by the feature of modeling the motion of shock waves over aerodynamic bodies, such as wings. This modeling requires the solution of nonlinear partial differential equations. Most advanced codes such as XTRAN3S use the transonic small perturbation equation. Currently, XTRAN3S is being used for generic research in unsteady aerodynamics and aeroelasticity of almost full aircraft configurations. Use of Euler/Navier Stokes equations for simple typical sections has just begun. A brief history of the development of CFD for aeroelastic applications is summarized. The development of unsteady transonic aerodynamics and aeroelasticity are also summarized.
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.
Experimental and theoretical studies in nonlinear aeroelasticity
NASA Astrophysics Data System (ADS)
Attar, Peter Joseph
Experimental and theoretical studies are conducted in the field of nonlinear aeroelasticity. Specifically two aeroelastic configurations, a flapping flag and a delta wing, are investigated and correlations between theory and experiment are presented. Two nonlinear structural theories are used to describe the structural behavior of the two models which are studied. The delta wing structural behavior is modeled using the nonlinear plate theory of von Karman. The nonlinearity in this model is due to the coupling between the out-of-plane and in-plane deflections and the model allows for moderately large out-of-plane plate deflections. The flapping flag structural model is a nonlinear beam theory which includes nonlinearities due to both large curvature and inertia. The axial deflection in this model is related to the out-of-plane deflection using an axially inextensible theory. The aerodynamic theory used is potential flow theory, which is applicable to low speed flows. The equation which describes potential flow is the Laplace equation, which is a linear partial differential equation. The Laplace equation is solved using a vortex lattice method. Aeroelastic solutions are found using both the classic small disturbance linearized fluid-structure interface boundary condition and the exact nonlinear boundary condition. The aeroelastic model which includes the nonlinear boundary conditions also includes a free wake solution. Several reduced order methods are explored. Normal mode solutions, both for the structural and aerodynamic models, are studied along with a proper orthogonal decomposition model for the aerodynamic flow. A brief description of a parallel implementation of the aeroelastic simulation code is also given and the parallel speedup is shown to be nearly linear for a certain class of problems. Correlation between theory and experiment is presented for both the delta wing and flapping flag model. Several steady angle of attack cases were investigated for the
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.
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.
Unified Formulation of the Aeroelasticity of Swept Lifting Surfaces
NASA Technical Reports Server (NTRS)
Silva, Walter; Marzocca, Piergiovanni; Librescu, Liviu
2001-01-01
An unified approach for dealing with stability and aeroelastic response to time-dependent pressure pulses of swept wings in an incompressible flow is developed. To this end the indicial function concept in time and frequency domains, enabling one to derive the proper unsteady aerodynamic loads is used. Results regarding stability in the frequency and time domains, and subcritical aeroelastic response to arbitrary time-dependent external excitation obtained via the direct use of the unsteady aerodynamic derivatives for 3-D wings are supplied. Closed form expressions for unsteady aerodynamic derivatives using this unified approach have been derived and used to illustrate their application to flutter and aeroelastic response to blast and sonic-boom signatures. In this context, an original representation of the aeroelastic response in the phase space was presented and pertinent conclusions on the implications of some basic parameters have been outlined.
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
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.
NASA Technical Reports Server (NTRS)
Radovcich, N. A.; Dreim, D.; Okeefe, D. A.; Linner, L.; Pathak, S. K.; Reaser, J. S.; Richardson, D.; Sweers, J.; Conner, F.
1985-01-01
Work performed in the design of a transport aircraft wing for maximum fuel efficiency is documented with emphasis on design criteria, design methodology, and three design configurations. The design database includes complete finite element model description, sizing data, geometry data, loads data, and inertial data. A design process which satisfies the economics and practical aspects of a real design is illustrated. The cooperative study relationship between the contractor and NASA during the course of the contract is also discussed.
Wing design for a civil tiltrotor transport aircraft
NASA Technical Reports Server (NTRS)
Rais-Rohani, Masoud
1994-01-01
The goal of this research is the proper tailoring of the civil tiltrotor's composite wing-box structure leading to a minimum-weight wing design. With focus on the structural design, the wing's aerodynamic shape and the rotor-pylon system are held fixed. The initial design requirement on drag reduction set the airfoil maximum thickness-to-chord ratio to 18 percent. The airfoil section is the scaled down version of the 23 percent-thick airfoil used in V-22's wing. With the project goal in mind, the research activities began with an investigation of the structural dynamic and aeroelastic characteristics of the tiltrotor configuration, and the identification of proper procedures to analyze and account for these characteristics in the wing design. This investigation led to a collection of more than thirty technical papers on the subject, some of which have been referenced here. The review of literature on the tiltrotor revealed the complexity of the system in terms of wing-rotor-pylon interactions. The aeroelastic instability or whirl flutter stemming from wing-rotor-pylon interactions is found to be the most critical mode of instability demanding careful consideration in the preliminary wing design. The placement of wing fundamental natural frequencies in bending and torsion relative to each other and relative to the rotor 1/rev frequencies is found to have a strong influence on the whirl flutter. The frequency placement guide based on a Bell Helicopter Textron study is used in the formulation of frequency constraints. The analysis and design studies are based on two different finite-element computer codes: (1) MSC/NASATRAN and (2) WIDOWAC. These programs are used in parallel with the motivation to eventually, upon necessary modifications and validation, use the simpler WIDOWAC code in the structural tailoring of the tiltrotor wing. Several test cases were studied for the preliminary comparison of the two codes. The results obtained so far indicate a good overall
Oblique wing transonic transport configuration development
NASA Technical Reports Server (NTRS)
1977-01-01
Studies of transport aircraft designed for boom-free supersonic flight show the variable sweep oblique wing to be the most efficient configuration for flight at low supersonic speeds. Use of this concept leads to a configuration that is lighter, quieter, and more fuel efficient than symmetric aircraft designed for the same mission. Aerodynamic structural, weight, aeroelastic and flight control studies show the oblique wing concept to be technically feasible. Investigations are reported for wing planform and thickness, pivot design and weight estimation, engine cycle (bypass ratio), and climb, descent and reserve fuel. Results are incorporated into a final configuration. Performance, weight, and balance characteristics are evaluated. Flight control requirements are reviewed, and areas in which further research is needed are identified.
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.
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.
In-flight gust monitoring and aeroelasticity studies
NASA Astrophysics Data System (ADS)
Alvarez-Salazar, Oscar Salvador
An in-flight gust monitoring and aeroelasticity study was conducted on board NASA Dryden's F15-B/FTF-II test platform (``FTF''). A total of four flights were completed. This study is the first in a series of flight experiments being conducted jointly by NASA Dryden Flight Research Center and UCLA's Flight Systems Research Center. The first objective of the in-flight gust- monitoring portion of the study was to demonstrate for the first time anywhere the measurability of intensity variations of a collimated Helium-Neon laser beam due to atmospheric air turbulence while having both the source and target apertures mounted outside an airborne aircraft. Intensity beam variations are the result of forward scattering of the beam by variations in the air's index of refraction, which are carried across the laser beam's path by a cross flow or air (i.e., atmospheric turbulence shifting vertically in the atmosphere). A laser beam was propagated parallel to the direction of flight for 1/2 meter outside the flight test fixture and its intensity variations due to atmospheric turbulence were successfully measured by a photo- detector. When the aircraft did not fly through a field of atmospheric turbulence, the laser beam proved to be insensitive to the stream velocity's cross component to the path of the beam. The aeroelasticity portion of the study consisted of measurements of the dynamic response of a straight, 18.25 inch span, 4.00 inch chord, NACA 0006 airfoil thickness profile, one sided wing to in-flight aircraft maneuvers, landing gear buffeting, unsteady aerodynamics, atmospheric turbulence, and aircraft vibration in general. These measurements were accomplished through the use of accelerometers, strain gauges and in-flight video cameras. Data collected will be used to compute in-flight root loci for the wing as functions of the aircraft's stream velocity. The data may also be used to calibrate data collected by the gust-monitoring system flown, and help verify the
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.
Overview of the Aeroelastic Prediction Workshop
NASA Technical Reports Server (NTRS)
Heeg, Jennifer; Chwalowski, Pawel; Schuster, David M.; Dalenbring, Mats
2013-01-01
The AIAA Aeroelastic Prediction Workshop (AePW) was held in April, 2012, bringing together communities of aeroelasticians and computational fluid dynamicists. The objective in conducting this workshop on aeroelastic prediction was to assess state-of-the-art computational aeroelasticity methods as practical tools for the prediction of static and dynamic aeroelastic phenomena. No comprehensive aeroelastic benchmarking validation standard currently exists, greatly hindering validation and state-of-the-art assessment objectives. The workshop was a step towards assessing the state of the art in computational aeroelasticity. This was an opportunity to discuss and evaluate the effectiveness of existing computer codes and modeling techniques for unsteady flow, and to identify computational and experimental areas needing additional research and development. Three configurations served as the basis for the workshop, providing different levels of geometric and flow field complexity. All cases considered involved supercritical airfoils at transonic conditions. The flow fields contained oscillating shocks and in some cases, regions of separation. The computational tools principally employed Reynolds-Averaged Navier Stokes solutions. The successes and failures of the computations and the experiments are examined in this paper.
NASA Technical Reports Server (NTRS)
Mason, Homer P.
1953-01-01
Three rocket-propelled buffet-research models have been flight tested to determine the buffeting characteristics of a swept-wing- airplane configuration with the horizontal tail operating near the wing wake. The models consisted of parabolic bodies having 45deg sweptback wings of aspect ratio 3.56, at aspect ratio of 0.3, NACA 64A007 airfoil sections, and tail surfaces of geometry and section identical to the wings. Two tests were conducted with the horizontal tail located in the wing chord plane with fixed incidence angles of -1.5deg on one model and 0deg on the other model. The third test was conducted with no horizontal tail. Results of these tests are presented as incremental accelerations in the body due to buffeting, trim angles of attack, trim normal- and side-force coefficients, wing-tip helix angles, static-directional-stability derivatives , and drag coefficients plotted against Mach number. These data indicate that mild low-lift buffeting was experienced by all models over a range of Mach number from approximately 0.7 to 1.4. It is further indicated that this buffeting was probably induced by wing-body interference and was amplified at transonic speeds by the horizontal tail operating in the wing wake. A longitudinal trim change was encountered by the tail-on models at transonic speeds, but no large changes in side force and no wing dropping were indicated.
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.
Flight testing the fixed-wing configuration of the Rotor Systems Research Aircraft (RSRA)
NASA Technical Reports Server (NTRS)
Hall, G. W.; Morris, P. M.
1985-01-01
The Rotor Systems Research Aircraft (RSRA) is a unique research aircraft designed to flight test advanced helicopter rotor system. Its principal flight test configuration is as a compound helicopter. The fixed wing configuration of the RSRA was primarily considered an energy fly-home mode in the event it became necessary to sever an unstable rotor system in flight. While it had always been planned to flight test the fixed wing configuration, the selection of the RSRA as the flight test bed for the X-wing rotor accelerated this schedule. This paper discusses the build-up to, and the test of, the RSRA fixed wing configuration. It is written primarily from the test pilot's perspective.
NASA Technical Reports Server (NTRS)
Pines, S.
1981-01-01
The methods used to compute the mass, structural stiffness, and aerodynamic forces in the form of influence coefficient matrices as applied to a flutter analysis of the Drones for Aerodynamic and Structural Testing (DAST) Aeroelastic Research Wing. The DAST wing was chosen because wind tunnel flutter test data and zero speed vibration data of the modes and frequencies exist and are available for comparison. A derivation of the equations of motion that can be used to apply the modal method for flutter suppression is included. A comparison of the open loop flutter predictions with both wind tunnel data and other analytical methods is presented.
Flight Dynamics of Flexible Aircraft with Aeroelastic and Inertial Force Interactions
NASA Technical Reports Server (NTRS)
Nguyen, Nhan T.; Tuzcu, Ilhan
2009-01-01
This paper presents an integrated flight dynamic modeling method for flexible aircraft that captures coupled physics effects due to inertial forces, aeroelasticity, and propulsive forces that are normally present in flight. The present approach formulates the coupled flight dynamics using a structural dynamic modeling method that describes the elasticity of a flexible, twisted, swept wing using an equivalent beam-rod model. The structural dynamic model allows for three types of wing elastic motion: flapwise bending, chordwise bending, and torsion. Inertial force coupling with the wing elasticity is formulated to account for aircraft acceleration. The structural deflections create an effective aeroelastic angle of attack that affects the rigid-body motion of flexible aircraft. The aeroelastic effect contributes to aerodynamic damping forces that can influence aerodynamic stability. For wing-mounted engines, wing flexibility can cause the propulsive forces and moments to couple with the wing elastic motion. The integrated flight dynamics for a flexible aircraft are formulated by including generalized coordinate variables associated with the aeroelastic-propulsive forces and moments in the standard state-space form for six degree-of-freedom flight dynamics. A computational structural model for a generic transport aircraft has been created. The eigenvalue analysis is performed to compute aeroelastic frequencies and aerodynamic damping. The results will be used to construct an integrated flight dynamic model of a flexible generic transport aircraft.
NASA Technical Reports Server (NTRS)
Cox, T. H.; Gilyard, G. B.
1986-01-01
The drones for aerodynamic and structural testing (DAST) project was designed to control flutter actively at high subsonic speeds. Accurate knowledge of the structural model was critical for the successful design of the control system. A ground vibration test was conducted on the DAST vehicle to determine the structural model characteristics. This report presents and discusses the vibration and test equipment, the test setup and procedures, and the antisymmetric and symmetric mode shape results. The modal characteristics were subsequently used to update the structural model employed in the control law design process.
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.
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).
Optimal Topology of Aircraft Rib and Spar Structures under Aeroelastic Loads
NASA Technical Reports Server (NTRS)
Stanford, Bret K.; Dunning, Peter D.
2014-01-01
Several topology optimization problems are conducted within the ribs and spars of a wing box. It is desired to locate the best position of lightening holes, truss/cross-bracing, etc. A variety of aeroelastic metrics are isolated for each of these problems: elastic wing compliance under trim loads and taxi loads, stress distribution, and crushing loads. Aileron effectiveness under a constant roll rate is considered, as are dynamic metrics: natural vibration frequency and flutter. This approach helps uncover the relationship between topology and aeroelasticity in subsonic transport wings, and can therefore aid in understanding the complex aircraft design process which must eventually consider all these metrics and load cases simultaneously.
Level-Set Topology Optimization with Aeroelastic Constraints
NASA Technical Reports Server (NTRS)
Dunning, Peter D.; Stanford, Bret K.; Kim, H. Alicia
2015-01-01
Level-set topology optimization is used to design a wing considering skin buckling under static aeroelastic trim loading, as well as dynamic aeroelastic stability (flutter). The level-set function is defined over the entire 3D volume of a transport aircraft wing box. Therefore, the approach is not limited by any predefined structure and can explore novel configurations. The Sequential Linear Programming (SLP) level-set method is used to solve the constrained optimization problems. The proposed method is demonstrated using three problems with mass, linear buckling and flutter objective and/or constraints. A constraint aggregation method is used to handle multiple buckling constraints in the wing skins. A continuous flutter constraint formulation is used to handle difficulties arising from discontinuities in the design space caused by a switching of the critical flutter mode.
An Overview of Recent Developments in Computational Aeroelasticity
NASA Technical Reports Server (NTRS)
Bennett, Robert M.; Edwards, John W.
2004-01-01
The motivation for Computational Aeroelasticity (CA) and the elements of one type of the analysis or simulation process are briefly reviewed. The need for streamlining and improving the overall process to reduce elapsed time and improve overall accuracy is discussed. Further effort is needed to establish the credibility of the methodology, obtain experience, and to incorporate the experience base to simplify the method for future use. Experience with the application of a variety of Computational Aeroelasticity programs is summarized for the transonic flutter of two wings, the AGARD 445.6 wing and a typical business jet wing. There is a compelling need for a broad range of additional flutter test cases for further comparisons. Some existing data sets that may offer CA challenges are presented.
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.
Data Comparisons and Summary of the Second Aeroelastic Prediction Workshop
NASA Technical Reports Server (NTRS)
Heeg, Jennifer; Wieseman, Carol D.; Chwalowski, Pawel
2016-01-01
This paper presents the computational results generated by participating teams of the second Aeroelastic Prediction Workshop and compare them with experimental data. Aeroelastic and rigid configurations of the Benchmark Supercritical Wing (BSCW) wind tunnel model served as the focus for the workshop. The comparison data sets include unforced ("steady") system responses, forced pitch oscillations and coupled fluid-structure responses. Integrated coefficients, frequency response functions, and flutter onset conditions are compared. The flow conditions studied were in the transonic range, including both attached and separated flow conditions. Some of the technical discussions that took place at the workshop are summarized.
Probabilistic Aeroelastic Analysis 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.
NASA Technical Reports Server (NTRS)
Hathaway, Michael D.; DelRasario, Ruben; Madavan, Nateri K.
2013-01-01
This paper presents an overview of the propulsion research and technology portfolio of NASA Fundamental Aeronautics Program Fixed Wing Project. The research is aimed at significantly reducing the thrust specific fuel/energy consumption of notional advanced fixed wing aircraft (by 60 % relative to a baseline Boeing 737-800 aircraft with CFM56-7B engines) in the 2030-2035 time frame. The research investments described herein are aimed at improving propulsive efficiency through higher bypass ratio fans, improving thermal efficiency through compact high overall pressure ratio gas generators, and exploring the potential benefits of boundary layer ingestion propulsion and hybrid gas-electric propulsion concepts.
NASA Technical Reports Server (NTRS)
Hathaway, Michael D.; Rosario, Ruben Del; Madavan, Nateri K.
2013-01-01
This paper presents an overview of the propulsion research and technology portfolio of NASA Fundamental Aeronautics Program Fixed Wing Project. The research is aimed at significantly reducing the thrust specific fuel/energy consumption of notional advanced fixed wing aircraft (by 60 percent relative to a baseline Boeing 737-800 aircraft with CFM56-7B engines) in the 2030 to 2035 time frame. The research investments described herein are aimed at improving propulsive efficiency through higher bypass ratio fans, improving thermal efficiency through compact high overall pressure ratio gas generators, and exploring the potential benefits of boundary layer ingestion propulsion and hybrid gas-electric propulsion concepts.
Accomplishments of the Abrupt Wing Stall (AWS) Program and Future Research Requirements
NASA Technical Reports Server (NTRS)
Hall, Robert M.; Woodson, Shawn H.; Chambers, Joseph R.
2003-01-01
The Abrupt Wing Stall (AWS) Program has addressed the problem of uncommanded lateral motions, such as wing drop and wing rock, at transonic speeds. The genesis of this Program was the experience of the F/A-18E/F Program in the late 199O's, when wing drop was discovered in the heart of the maneuver envelope for the pre-production aircraft. While the F/A-18E/F problem was subsequently corrected by a leading-edge flap scheduling change and the addition of a porous door to the wing fold fairing, the AWS Program was initiated as a national response to the lack of technology readiness available at the time of the F/A-18E/F Development Program. The AWS Program objectives were to define causal factors for the F/A-18E/F experience, to gain insights into the flow physics associated with wing drop, and to develop methods and analytical tools so that future programs could identify this type of problem before going to flight test. The paper reviews, for the major goals of the AWS Program, the status of the technology before the program began, the program objectives, accomplishments, and impacts. Lessons learned are presented for the benefit of future programs that must assess whether a vehicle will have uncommanded lateral motions before going to flight test. Finally, recommended future research needs are presented in light of the AWS Program experience.
Development of a composite UAV wing test-bed for structural health monitoring research
NASA Astrophysics Data System (ADS)
Oliver, J. A.; Kosmatka, J. B.; Farrar, Charles R.; Park, Gyuhae
2007-04-01
In order to facilitate damage detection and structural health monitoring (SHM) research for composite unmanned aerial vehicles (UAV) a specialized test-bed has been developed. This test-bed consists of four 2.61 m all-composite test-pieces emulating composite UAV wings, a series of detailed finite element models of the test-pieces and their components, and a dynamic testing setup including a mount for simulating the cantilevered operation configuration of real wings. Two of the wings will have bondline damage built in; one undamaged and one damaged wing will also be fitted with a range of embedded and attached sensors-piezoelectric patches, fiber-optics, and accelerometers. These sensors will allow collection of realistic data; combined with further modal testing they will allow comparison of the physical impact of the sensors on the structure compared to the damage-induced variation, evaluation of the sensors for implementation in an operational structure, and damage detection algorithm validation. At the present time the pieces for four wings have been fabricated and modally tested and one wing has been fully assembled and re-tested in a cantilever configuration. The component part and assembled wing finite element models, created for MSC.Nastran, have been correlated to their respective structures using the modal information. This paper details the design and manufacturing of the test-pieces, the finite element model construction, and the dynamic testing setup. Measured natural frequencies and mode shapes for the assembled cantilevered wing are reported, along with finite element model undamaged modal response, and response with a small disbond at the root of the top main spar-skin bondline.
Investigation of the aeroelastic stability of the AFW wind-tunnel model using CAP-TSD
NASA Technical Reports Server (NTRS)
Silva, Walter A.; Bennett, Robert M.
1992-01-01
The Computational Aeroelasticity Program - Transonic Small Disturbance (CAP-TSD) code is applied to the Active Flexible Wing (AFW) wind tunnel model for prediction of the model's transonic aeroelastic behavior. A semi-span computational model is used for evaluation of symmetric motions and a full span model is used for evaluation of antisymmetric motions. Static aeroelastic solutions using CAP-TSD are computed. Dynamic flutter analyses are then performed as perturbations about the static aeroelastic deformations and presented as flutter boundaries in terms of Mach number and dynamic pressure. Flutter boundaries that take into account modal refinements, vorticity, and entropy corrections, antisymmetric motions and sensitivity to the modeling of the wing tip ballast stores are also presented and compared with experimental flutter results.
NASA Astrophysics Data System (ADS)
Ardelean, Emil Valentin
Flutter is a rather spectacular phenomenon of aeroelastic instability that affects lifting and control surfaces, yet can also lead to catastrophic consequences for the aircraft. The idea of controlling flutter by using the same energy that causes it, namely airflow energy, through changing the aerodynamics in a controlled manner is not new. In the case of fixed wings, the use of trailing edge control surfaces (flaps) is an extremely effective method to alter the aerodynamics. This research presents the development of an actuation system for trailing edge control surfaces (flaps) used for aeroelastic flutter control of a typical section wing model. In order to be effective for aeroelastic control of flutter, flap deflection of +/-5-6° with adequate bandwidth (up to 25--30 Hz) is required. Classical solutions for flap actuation do not have the capabilities required for this task. Therefore actuation systems using active materials became the focus of this investigation. A new piezoelectric actuator (V-Stack Piezoelectric Actuator) was developed. This actuator meets the requirements for trailing edge flap actuation in both stroke and force over the bandwidth of interest. It is compact, simple, sturdy, and leverages stroke geometrically with minimum force penalties, while displaying linearity over a wide range of stroke. Integration of the actuator inside an existing structure requires minimal modifications of the structure. The shape of the actuator makes it very suitable for trailing edge flap actuation, eliminating the need for a push rod. The actuation solution presented here stands out because of its simplicity, compactness, small mass (compared to that of the actuated structure) and high reliability. Although the actuator was designed for flap actuation, other applications can also benefit from its capabilities. In order to demonstrate the actuation concept, a typical section prototype was constructed and tested experimentally in the wind tunnel at Duke
ASTROP2 users manual: A program for aeroelastic stability analysis of propfans
NASA Technical Reports Server (NTRS)
Narayanan, G. V.; Kaza, K. R. V.
1991-01-01
A user's manual is presented for the aeroelastic stability and response of propulsion systems computer program called ASTROP2. The ASTROP2 code preforms aeroelastic stability analysis of rotating propfan blades. This analysis uses a two-dimensional, unsteady cascade aerodynamics model and a three-dimensional, normal-mode structural model. Analytical stability results from this code are compared with published experimental results of a rotating composite advanced turboprop model and of nonrotating metallic wing model.
NASA rotor systems research aircraft: Fixed-wing configuration flight-test results
NASA Technical Reports Server (NTRS)
Erickson, R. E.; Cross, J. L.; Kufeld, R. M.; Acree, C. W.; Nguyen, D.; Hodge, R. W.
1986-01-01
The fixed-wing, airplane configuration flight-test results of the Rotor System Research Aircraft (RSRA), NASA 740, at Ames/Dryden Flight Research Center are documented. Fourteen taxi and flight tests were performed from December 1983 to October 1984. This was the first time the RSRA was flown with the main rotor removed; the tail rotor was installed. These tests confirmed that the RSRA is operable as a fixed-wing aircraft. Data were obtained for various takeoff and landing distances, control sensitivity, trim and dynamics stability characteristics, performance rotor-hub drag, and acoustics signature. Stability data were obtained with the rotor hub both installed and removed. The speed envelope was developed to 261 knots true airspeed (KTAS), 226 knots calibrated airspeed (KCAS) at 10,000 ft density altitude. The airplane was configured at 5 deg. wing incidence with 5 deg. wing flaps as a normal configuration. Level-flight data were acquired at 167 KCAS for wing incidence from 0 to 10 deg. Step inputs and doublet inputs of various magnitudes were utilized to acquire dynamic stability and control sensitivity data. Sine-wave inputs of constantly increasing frequency were used to generate parameter identification data. The maximum load factor attained was 2.34 g at 206 KCAS.
F-8 oblique wing structural feasibility study
NASA Technical Reports Server (NTRS)
Koltko, E.; Katz, A.; Bell, M. A.; Smith, W. D.; Lauridia, R.; Overstreet, C. T.; Klapprott, C.; Orr, T. F.; Jobe, C. L.; Wyatt, F. G.
1975-01-01
The feasibility of fitting a rotating oblique wing on an F-8 aircraft to produce a full scale manned prototype capable of operating in the transonic and supersonic speed range was investigated. The strength, aeroelasticity, and fatigue life of such a prototype are analyzed. Concepts are developed for a new wing, a pivot, a skewing mechanism, control systems that operate through the pivot, and a wing support assembly that attaches in the F-8 wing cavity. The modification of the two-place NTF-8A aircraft to the oblique wing configuration is discussed.
Modal Response of Trapezoidal Wing Structures Using Second Order Shape Sensitivities
NASA Technical Reports Server (NTRS)
Liu, Youhua; Kapania, Rakesh K.
2000-01-01
The modal response of wing structures is very important for assessing their dynamic response including dynamic aeroelastic instabilities. Moreover, in a recent study an efficient structural optimization approach was developed using structural modes to represent the static aeroelastic wing response (both displacement and stress). In this paper, the modal response of general trapezoidal wing structures is approximated using shape sensitivities up to the 2nd order. Also different approaches of computing the derivatives are investigated.
Wind Tunnel to Atmospheric Mapping for Static Aeroelastic Scaling
NASA Technical Reports Server (NTRS)
Heeg, Jennifer; Spain, Charles V.; Rivera, J. A.
2004-01-01
Wind tunnel to Atmospheric Mapping (WAM) is a methodology for scaling and testing a static aeroelastic wind tunnel model. The WAM procedure employs scaling laws to define a wind tunnel model and wind tunnel test points such that the static aeroelastic flight test data and wind tunnel data will be correlated throughout the test envelopes. This methodology extends the notion that a single test condition - combination of Mach number and dynamic pressure - can be matched by wind tunnel data. The primary requirements for affecting this extension are matching flight Mach numbers, maintaining a constant dynamic pressure scale factor and setting the dynamic pressure scale factor in accordance with the stiffness scale factor. The scaling is enabled by capabilities of the NASA Langley Transonic Dynamics Tunnel (TDT) and by relaxation of scaling requirements present in the dynamic problem that are not critical to the static aeroelastic problem. The methodology is exercised in two example scaling problems: an arbitrarily scaled wing and a practical application to the scaling of the Active Aeroelastic Wing flight vehicle for testing in the TDT.
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.
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.
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.
Some Research on the Lift and Stability of Wing-Body Combinations
NASA Technical Reports Server (NTRS)
Purser, Paul E.; Fields, E. M.
1959-01-01
The present paper summarizes and correlates broadly some of the research results applicable to fin-stabilized ammunition. The discussion and correlation are intended to be comprehensive, rather than detailed, in order to show general trends over the Mach number range up to 7.0. Some discussion of wings, bodies, and wing-body interference is presented, and a list of 179 papers containing further information is included. The present paper is intended to serve more as a bibliography and source of reference material than as a direct source of design information.
Aerostructural Level Set Topology Optimization for a Common Research Model Wing
NASA Technical Reports Server (NTRS)
Dunning, Peter D.; Stanford, Bret K.; Kim, H. Alicia
2014-01-01
The purpose of this work is to use level set topology optimization to improve the design of a representative wing box structure for the NASA common research model. The objective is to minimize the total compliance of the structure under aerodynamic and body force loading, where the aerodynamic loading is coupled to the structural deformation. A taxi bump case was also considered, where only body force loads were applied. The trim condition that aerodynamic lift must balance the total weight of the aircraft is enforced by allowing the root angle of attack to change. The level set optimization method is implemented on an unstructured three-dimensional grid, so that the method can optimize a wing box with arbitrary geometry. Fast matching and upwind schemes are developed for an unstructured grid, which make the level set method robust and efficient. The adjoint method is used to obtain the coupled shape sensitivities required to perform aerostructural optimization of the wing box structure.
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.
NASA Technical Reports Server (NTRS)
Ko, William L.; Gong, Leslie
2001-01-01
Heat transfer, thermal stresses, and thermal buckling analyses were performed on the unconventional wing structures of a Hyper-X hypersonic flight research vehicle (designated as X-43) subjected to nominal Mach 7 aerodynamic heating. A wing midspan cross section was selected for the heat transfer and thermal stress analyses. Thermal buckling analysis was performed on three regions of the wing skin (lower or upper); 1) a fore wing panel, 2) an aft wing panel, and 3) a unit panel at the middle of the aft wing panel. A fourth thermal buckling analysis was performed on a midspan wing segment. The unit panel region is identified as the potential thermal buckling initiation zone. Therefore, thermal buckling analysis of the Hyper-X wing panels could be reduced to the thermal buckling analysis of that unit panel. "Buckling temperature magnification factors" were established. Structural temperature-time histories are presented. The results show that the concerns of shear failure at wing and spar welded sites, and of thermal buckling of Hyper-X wing panels, may not arise under Mach 7 conditions.
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.
Parametric studies for tiltrotor aeroelastic stability in high-speed flight
NASA Technical Reports Server (NTRS)
Nixon, Mark W.
1992-01-01
The influence of several system design parameters on tiltrotor aeroelastic stability is examined for the high-speed (axial) flight mode. Coupling of the rotor flapping modes with the wing elastic modes produces a whirl motion, typical of tiltrotors, that can become unstable at high speeds. The sensitivity of this instability with respect to rotor frequencies, wing stiffness, forward wing sweep, and rotor thrust level is examined. Some important new trends are identified regarding the role of blade lag dynamics and forward wing sweep in tiltrotor aeroelastic stability. The blade lag frequency may be tuned to improve tiltrotor stability, and forward wing sweep is destabilizing because of changes in rotor force components associated with the sweep.
NASA Technical Reports Server (NTRS)
Clark, L. E.; Richie, C. B.
1977-01-01
The hypersonic aerodynamic characteristics of an air-launched, delta-wing research aircraft concept were investigated at Mach 6. The effect of various components such as nose shape, wing camber, wing location, center vertical tail, wing tip fins, forward delta wing, engine nacelle, and speed brakes was also studied. Tests were conducted with a 0.021 scale model at a Reynolds number, based on model length, of 10.5 million and over an angel of attack range from -4 deg to 20 deg. Results show that most configurations with a center vertical tail have static longitudinal stability at trim, static directional stability at angles of attack up to 12 deg, and static lateral stability throughout the angle of attack range. Configurations with wing tip fins generally have static longitudinal stability at trim, have lateral stability at angles of attack above 8 deg, and are directionally unstable over the angle of attack range.
The benchmark aeroelastic models program: Description and highlights of initial results
NASA Technical Reports Server (NTRS)
Bennett, Robert M.; Eckstrom, Clinton V.; Rivera, Jose A., Jr.; Dansberry, Bryan E.; Farmer, Moses G.; Durham, Michael H.
1991-01-01
An experimental effort was implemented in aeroelasticity called the Benchmark Models Program. The primary purpose of this program is to provide the necessary data to evaluate computational fluid dynamic codes for aeroelastic analysis. It also focuses on increasing the understanding of the physics of unsteady flows and providing data for empirical design. An overview is given of this program and some results obtained in the initial tests are highlighted. The tests that were completed include measurement of unsteady pressures during flutter of rigid wing with a NACA 0012 airfoil section and dynamic response measurements of a flexible rectangular wing with a thick circular arc airfoil undergoing shock boundary layer oscillations.
NASA Technical Reports Server (NTRS)
Gern, Frank H.; Naghshineh, Amir H.; Sulaeman, Erwin; Kapania, Rakesh K.; Haftka, Raphael T.
2000-01-01
This paper describes a structural and aeroelastic model for wing sizing and weight calculation of a strut-braced wing. The wing weight is calculated using a newly developed structural weight analysis module considering the special nature of strut-braced wings. A specially developed aeroelastic model enables one to consider wing flexibility and spanload redistribution during in-flight maneuvers. The structural model uses a hexagonal wing-box featuring skin panels, stringers, and spar caps, whereas the aerodynamics part employs a linearized transonic vortex lattice method. Thus, the wing weight may be calculated from the rigid or flexible wing spanload. The calculations reveal the significant influence of the strut on the bending material weight of the wing. The use of a strut enables one to design a wing with thin airfoils without weight penalty. The strut also influences wing spanload and deformations. Weight savings are not only possible by calculation and iterative resizing of the wing structure according to the actual design loads. Moreover, as an advantage over the cantilever wing, employment of the strut twist moment for further load alleviation leads to increased savings in structural weight.
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
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.
Development of wing and tail configurations for low altitude unmanned research aircraft (LAURA)
NASA Technical Reports Server (NTRS)
Mangalam, S. M.; Harvey, W. D.; Siddiqi, S.
1987-01-01
The Low Altitude/Airspeed Unmanned Research Aircraft (LAURA) is being developed by the U.S. Navy for flight test research using low-Reynolds number airfoils. This vehicle consists of a standard modular fuselage designed to accept the installation of several wings/tails having low Reynolds number airfoils, and various planform shapes. Design constraints include shipboard storage, long flight endurance at very low airspeeds and sea-skimming cruise altitude. The stringent design constraints require the development of high-performance low Reynolds number (LRN) airfoils, suitable lifting surface configuration, and advanced airframe-propulsion systems. The present paper describes ongoing efforts to develop wing and tail configurations for LAURA using airfoils designed at NASA Langley Research Center.
Wing design for a civil tiltrotor transport aircraft: A preliminary study
NASA Technical Reports Server (NTRS)
Rais-Rohani, Masoud
1993-01-01
A preliminary study was conducted on the design of the wing-box structure for a civil tiltrotor transport aircraft. The wing structural weight is to be minimized subject to structural and aeroelastic constraints. The composite wing-box structure is composed of skin, stringers, ribs, and spars. The design variables include skin ply thicknesses and orientations and spar cap and stringer cross-sectional areas. With the total task defined, an initial study was conducted to learn more about the intricate dynamic and aeroelastic characteristics of the tiltrotor aircraft and their roles in the wing design. Also, some work was done on the wing finite-element modeling (via PATRAN) which would be used in structural analysis and optimization. Initial studies indicate that in order to limit the wing/rotor aeroelastic and dynamic interactions in the preliminary design, the cruise speed, rotor system, and wing geometric attributes must all be held fixed.
The Concorde and aeronautical research
NASA Technical Reports Server (NTRS)
Salmon, M.
1983-01-01
Theoretical and experimental work carried out in various research centers, and particularly at ONERA, which led to the conception and to the main technical solutions included in the design of Concorde: plane form, twist and camber of the wing, lift augmentation by upper surface vortices, kinetic heating, air intakes and jet exhausts, materials, aeroelasticity. The development of research, and the numerous tests carried out for the benefit of the designers since the beginning of the project, are also outlined.
Improved Aerodynamic Influence Coefficients for Dynamic Aeroelastic Analyses
NASA Astrophysics Data System (ADS)
Gratton, Patrice
2011-12-01
Currently at Bombardier Aerospace, aeroelastic analyses are performed using the Doublet Lattice Method (DLM) incorporated in the NASTRAN solver. This method proves to be very reliable and fast in preliminary design stages where wind tunnel experimental results are often not available. Unfortunately, the geometric simplifications and limitations of the DLM, based on the lifting surfaces theory, reduce the ability of this method to give reliable results for all flow conditions, particularly in transonic flow. Therefore, a new method has been developed involving aerodynamic data from high-fidelity CFD codes which solve the Euler or Navier-Stokes equations. These new aerodynamic loads are transmitted to the NASTRAN aeroelastic module through improved aerodynamic influence coefficients (AIC). A cantilevered wing model is created from the Global Express structural model and a set of natural modes is calculated for a baseline configuration of the structure. The baseline mode shapes are then combined with an interpolation scheme to deform the 3-D CFD mesh necessary for Euler and Navier-Stokes analyses. An uncoupled approach is preferred to allow aerodynamic information from different CFD codes. Following the steady state CFD analyses, pressure differences ( DeltaCp), calculated between the deformed models and the original geometry, lead to aerodynamic loads which are transferred to the DLM model. A modal-based AIC method is applied to the aerodynamic matrices of NASTRAN based on a least-square approximation to evaluate aerodynamic loads of a different wing configuration which displays similar types of mode shapes. The methodology developed in this research creates weighting factors based on steady CFD analyses which have an equivalent reduced frequency of zero. These factors are applied to both the real and imaginary part of the aerodynamic matrices as well as all reduced frequencies used in the PK-Method which solves flutter problems. The modal-based AIC method
NASA Technical Reports Server (NTRS)
Hurst, Janet
2011-01-01
A brief overview is presented of the current materials and structures research geared toward propulsion applications for NASA s Subsonic Fixed Wing Project one of four projects within the Fundamental Aeronautics Program of the NASA Aeronautics Research Mission Directorate. The Subsonic Fixed Wing (SFW) Project has selected challenging goals which anticipate an increasing emphasis on aviation s impact upon the global issue of environmental responsibility. These goals are greatly reduced noise, reduced emissions and reduced fuel consumption and address 25 to 30 years of technology development. Successful implementation of these demanding goals will require development of new materials and structural approaches within gas turbine propulsion technology. The Materials and Structures discipline, within the SFW project, comprise cross-cutting technologies ranging from basic investigations to component validation in laboratory environments. Material advances are teamed with innovative designs in a multidisciplinary approach with the resulting technology advances directed to promote the goals of reduced noise and emissions along with improved performance.
Integrated aerodynamic-structural-control wing design
NASA Technical Reports Server (NTRS)
Rais-Rohani, M.; Haftka, R. T.; Grossman, B.; Unger, E. R.
1992-01-01
The aerodynamic-structural-control design of a forward-swept composite wing for a high subsonic transport aircraft is considered. The structural analysis is based on a finite-element method. The aerodynamic calculations are based on a vortex-lattice method, and the control calculations are based on an output feedback control. The wing is designed for minimum weight subject to structural, performance/aerodynamic and control constraints. Efficient methods are used to calculate the control-deflection and control-effectiveness sensitivities which appear as second-order derivatives in the control constraint equations. To suppress the aeroelastic divergence of the forward-swept wing, and to reduce the gross weight of the design aircraft, two separate cases are studied: (1) combined application of aeroelastic tailoring and active controls; and (2) aeroelastic tailoring alone. The results of this study indicated that, for this particular example, aeroelastic tailoring is sufficient for suppressing the aeroelastic divergence, and the use of active controls was not necessary.
NASA Technical Reports Server (NTRS)
Dehoff, R. L.; Reed, W. B.; Trankle, T. L.
1977-01-01
The development and validation of a spey engine model is described. An analysis of the dynamical interactions involved in the propulsion unit is presented. The model was reduced to contain only significant effects, and was used, in conjunction with flight data obtained from an augmentor wing jet STOL research aircraft, to develop initial estimates of parameters in the system. The theoretical background employed in estimating the parameters is outlined. The software package developed for processing the flight data is described. Results are summarized.
NASA Technical Reports Server (NTRS)
Mercer, C. E.; Carson, G. T., Jr.
1979-01-01
The influence of upper-surface nacelle exhaust flow on the aerodynamic characteristics of a supersonic cruise aircraft research configuration was investigated in a 16 foot transonic tunnel over a range of Mach numbers from 0.60 to 1.20. The arrow-wing transport configuration with engines suspended over the wing was tested at angles of attack from -4 deg to 6 deg and jet total pressure ratios from 1 to approximately 13. Wing-tip leading edge flap deflections of -10 deg to 10 deg were tested with the wing-body configuration. Various nacelle locations (chordwise, spanwise, and vertical) were tested over the ranges of Mach numbers, angles of attack, and jet total-pressure ratios. The results show that reflecting the wing-tip leading edge flap from 0 deg to -10 deg increased the maximum lift-drag ratio by 1.0 at subsonic speeds. Jet exhaust interference effects were negligible.
NASA Technical Reports Server (NTRS)
Roskam, J.
1972-01-01
Expressions are derived for computing the aerodynamic influence coefficient matrix for nonplanar wing-body-tail configurations. An aerodynamic influence coefficient is defined as the load in lbs. induced on a panel as a result of a unit angle of attack on another panel. Fuselage, wing and tail thickness are assumed to be small with the result that the thickness effect on the flow-field is negligible. The method for determining the aerodynamic influence coefficient matrix is based on the lifting solution to the small perturbation, steady potential flow equation.
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.
Three-dimensional time-marching aeroelastic analyses using an unstructured-grid Euler method
NASA Technical Reports Server (NTRS)
Rausch, Russ D.; Batina, John T.; Yang, Henry T. Y.
1992-01-01
Modifications to a three dimensional, implicit, upwind, unstructured-grid Euler code for aeroelastic analysis of complete aircraft configurations are described. The modifications involve the addition of the structural equations of motion for their simultaneous time integration with the governing flow equations. The paper presents a detailed description of the time marching aeroelastic procedure and presents comparisons with experimental data to provide an assessment of the capability. Flutter results are shown for an isolated 45 degree swept-back wing and a supersonic transport configuration with a fuselage, clipped delta wing, and two identical rearward-mounted nacelles. Comparisons between computed and experimental flutter characteristics show good agreement, giving confidence in the accuracy of the aeroelastic capability that was developed.
Evaluation of an aeroelastic model technique for predicting airplane buffet loads
NASA Technical Reports Server (NTRS)
Hanson, P. W.
1973-01-01
A wind-tunnel technique which makes use of a dynamically scaled aeroelastic model to predict full-scale airplane buffet loads during buffet boundary penetration is evaluated. A 1/8-scale flutter model of a fighter airplane with remotely controllable variable-sweep wings and trimming surfaces was used for the evaluation. The model was flown on a cable-mount system which permitted high lift forces comparable to those in maneuvering flight. Bending moments and accelerations due to buffet were measured on the flutter model and compared with those measured on the full-scale airplane in an independent flight buffet research study. It is concluded that the technique can provide valuable information on airplane buffet load characteristics not available from any other source except flight test.
NASA,FAA,ONERA Swept-Wing Icing and Aerodynamics: Summary of Research and Current Status
NASA Technical Reports Server (NTRS)
Broeren, Andy
2015-01-01
NASA, FAA, ONERA, and other partner organizations have embarked on a significant, collaborative research effort to address the technical challenges associated with icing on large scale, three-dimensional swept wings. These are extremely complex phenomena important to the design, certification and safe operation of small and large transport aircraft. There is increasing demand to balance trade-offs in aircraft efficiency, cost and noise that tend to compete directly with allowable performance degradations over an increasing range of icing conditions. Computational fluid dynamics codes have reached a level of maturity that they are being proposed by manufacturers for use in certification of aircraft for flight in icing. However, sufficient high-quality data to evaluate their performance on iced swept wings are not currently available in the public domain and significant knowledge gaps remain.
Experimental Aerodynamic Characteristics of a Joined-wing Research Aircraft Configuration
NASA Technical Reports Server (NTRS)
Smith, Stephen C.; Stonum, Ronald K.
1989-01-01
A wind-tunnel test was conducted at Ames Research Center to measure the aerodynamic characteristics of a joined-wing research aircraft (JWRA). This aircraft was designed to utilize the fuselage and engines of the existing NASA AD-1 aircraft. The JWRA was designed to have removable outer wing panels to represent three different configurations with the interwing joint at different fractions of the wing span. A one-sixth-scale wind-tunnel model of all three configurations of the JWRA was tested in the Ames 12-Foot Pressure Wind Tunnel to measure aerodynamic performance, stability, and control characteristics. The results of these tests are presented. Longitudinal and lateral-directional characteristics were measured over an angle of attack range of -7 to 14 deg and over an angle of sideslip range of -5 to +2.5 deg at a Mach number of 0.35 and a Reynolds number of 2.2x10(6)/ft. Various combinations of deflected control surfaces were tested to measure the effectiveness and impact on stability of several control surface arrangements. In addition, the effects on stall and post-stall aerodynamic characteristics from small leading-edge devices called vortilons were measured. The results of these tests indicate that the JWRA had very good aerodynamic performance and acceptable stability and control throughout its flight envelope. The vortilons produced a profound improvement in the stall and post-stall characteristics with no measurable effects on cruise performance.
Advanced Aeroelastic Technologies for Turbomachinery Application
NASA Technical Reports Server (NTRS)
DeWitt, Kenneth; Srivastava, Rakesh; Reddy, T. S. R.
2004-01-01
A summary of the work performed under the grant NCC-1068 is presented. More details can be found in the cited references. The summary is presented in two parts to represent two areas of research. In the first part, methods to analyze a high temperature ceramic guide vane subjected to cooling jets are presented, and in the second part, the effect of unsteady aerodynamic forces on aeroelastic stability as implemented into the turbo-REDUCE code are presented
NASA Technical Reports Server (NTRS)
Roskam, J.; Hamler, F. R.; Reynolds, D.
1972-01-01
The procedures used to establish the mass matrices characteristics for the fighter type wings studied are given. A description of the procedure used to find the mass associated with a specific aerodynamic panel is presented and some examples of the application of the procedure are included.
NASA Technical Reports Server (NTRS)
Roskam, J.; Lan, C.; Smith, H.; Gibson, G.
1972-01-01
An explanation is presented of the method used to locate the elastic axis and the method to determine the EI and GJ distributions along the elastic axes of wings with a 2-spar (front and rear) construction or a single torque-box construction.
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.
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.
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.
NASA Technical Reports Server (NTRS)
Dawson, Kenneth S.; Fortin, Paul E.
1987-01-01
The results of an integrated study of structures, aerodynamics, and controls using the STARS program on two advanced airplane configurations are presented. Results for the X-29A include finite element modeling, free vibration analyses, unsteady aerodynamic calculations, flutter/divergence analyses, and an aeroservoelastic controls analysis. Good correlation is shown between STARS results and various other verified results. The tasks performed on the Oblique Wing Research Aircraft include finite element modeling and free vibration analyses.
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.
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
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
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).
Computational Aeroelasticity: Success, Progress, Challenge
NASA Technical Reports Server (NTRS)
Schuster, David M.; Liu, Danny D.; Huttsell, Lawrence J.
2003-01-01
The formal term Computational Aeroelasticity (CAE) has only been recently adopted to describe aeroelastic analysis methods coupling high-level computational fluid dynamics codes with structural dynamics techniques. However, the general field of aeroelastic computations has enjoyed a rich history of development and application since the first hand-calculations performed in the mid 1930 s. This paper portrays a much broader definition of Computational Aeroelasticity; one that encompasses all levels of aeroelastic computation from the simplest linear aerodynamic modeling to the highest levels of viscous unsteady aerodynamics, from the most basic linear beam structural models to state-of-the-art Finite Element Model (FEM) structural analysis. This paper is not written as a comprehensive history of CAE, but rather serves to review the development and application of aeroelastic analysis methods. It describes techniques and example applications that are viewed as relatively mature and accepted, the "successes" of CAE. Cases where CAE has been successfully applied to unique or emerging problems, but the resulting techniques have proven to be one-of-a-kind analyses or areas where the techniques have yet to evolve into a routinely applied methodology are covered as "progress" in CAE. Finally the true value of this paper is rooted in the description of problems where CAE falls short in its ability to provide relevant tools for industry, the so-called "challenges" to CAE.
NASA Technical Reports Server (NTRS)
Roskam, J.; Smith, H.; Gibson, G.
1972-01-01
The method used in computing the structural influence coefficient matrix of the computer program of Reference 1 (appendix A of the Summary Report) is reported. This matrix is computed for complete wing-body-tail configurations by assuming that all major airplane components can be structurally represented by a slender beam called the elastic axis. A structural influence coefficient is defined as the rotation about the Y-stability axis at panel j induced by a unit load on panel k. A description of how a structural breakdown is performed in detail is included.
The design, analysis and experimental evaluation of an elastic model wing
NASA Technical Reports Server (NTRS)
Cavin, R. K., III; Thisayakorn, C.
1974-01-01
An elastic orbiter model was developed to evaluate the effectiveness of aeroelasticity computer programs. The elasticity properties were introduced by constructing beam-like straight wings for the wind tunnel model. A standard influence coefficient mathematical model was used to estimate aeroelastic effects analytically. In general good agreement was obtained between the empirical and analytical estimates of the deformed shape. However, in the static aeroelasticity case, it was found that the physical wing exhibited less bending and more twist than was predicted by theory.
Aeroelastic tailoring in wind-turbine blade applications
Veers, P.; Lobitz, D.; Bir, G.
1998-04-01
This paper reviews issues related to the use of aeroelastic tailoring as a cost-effective, passive means to shape the power curve and reduce loads. Wind turbine blades bend and twist during operation, effectively altering the angle of attack, which in turn affects loads and energy production. There are blades now in use that have significant aeroelastic couplings, either on purpose or because of flexible and light-weight designs. Since aeroelastic effects are almost unavoidable in flexible blade designs, it may be desirable to tailor these effects to the authors advantage. Efforts have been directed at adding flexible devices to a blade, or blade tip, to passively regulate power (or speed) in high winds. It is also possible to build a small amount of desirable twisting into the load response of a blade with proper asymmetric fiber lay up in the blade skin. (Such coupling is akin to distributed {delta}{sub 3} without mechanical hinges.) The tailored twisting can create an aeroelastic effect that has payoff in either better power production or in vibration alleviation, or both. Several research efforts have addressed different parts of this issue. Research and development in the use of aeroelastic tailoring on helicopter rotors is reviewed. Potential energy gains as a function of twist coupling are reviewed. The effects of such coupling on rotor stability have been studied and are presented here. The ability to design in twist coupling with either stretching or bending loads is examined also.
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.
Benchmark Composite Wing Design Including Joint Analysis and Optimization
NASA Astrophysics Data System (ADS)
Albers, Robert G.
A composite wing panel software package, named WING Joint OpTimization and Analysis (WINGJOTA) featuring bolted joint analysis, is created and presented in this research. Three areas of focus were the development of an analytic composite bolted joint analysis suitable for fast evaluation; a more realistic wing design than what has been considered in the open literature; and the application of two optimization algorithms for composite wing design. Optimization results from 14 wing load cases applied to a composite wing panel with joints are presented. The composite bolted joint analysis consists of an elasticity solution that provides the stress state at a characteristic distance away from the bolt holes. The stresses at the characteristic distance are compared to a failure criterion on a ply-by-ply basis that not only determines first ply failure but also the failure mode. The loads in the multi-fastener joints used in this study were determined by an iterative scheme that provides the bearing-bypass loads to the elasticity analysis. A preliminary design of a composite subsonic transport wing was developed, based around a mid-size, twin-aisle aircraft. The benchmark design includes the leading and trailing edge structures and the center box inside the fuselage. Wing masses were included as point loads, and fuel loads were incorporated as distributed loads. The side-of-body boundary condition was modeled using high stiffness springs, and the aerodynamic loads were applied using an approximate point load scheme. The entire wing structure was modeled using the finite element code ANSYS to provide the internal loads needed as boundary conditions for the wing panel analyzed by WINGJOTA. The software package WINGJOTA combines the composite bolted joint analysis, a composite plate finite element analysis, a wing aeroelastic cycle, and two optimization algorithms to form the basis of a computer code for analysis and optimization. Both the Improving Hit-and-Run (IHR) and
Development of Unsteady Aerodynamic and Aeroelastic Reduced-Order Models Using the FUN3D Code
NASA Technical Reports Server (NTRS)
Silva, Walter A.; Vatsa, Veer N.; Biedron, Robert T.
2009-01-01
Recent significant improvements to the development of CFD-based unsteady aerodynamic reduced-order models (ROMs) are implemented into the FUN3D unstructured flow solver. These improvements include the simultaneous excitation of the structural modes of the CFD-based unsteady aerodynamic system via a single CFD solution, minimization of the error between the full CFD and the ROM unsteady aero- dynamic solution, and computation of a root locus plot of the aeroelastic ROM. Results are presented for a viscous version of the two-dimensional Benchmark Active Controls Technology (BACT) model and an inviscid version of the AGARD 445.6 aeroelastic wing using the FUN3D code.
NASA Technical Reports Server (NTRS)
Kempel, Robert W.; Mcneill, Walter E.; Gilyard, Glenn B.; Maine, Trindel A.
1988-01-01
The NASA Ames Research Center developed an oblique-wing research plane from NASA's digital fly-by-wire airplane. Oblique-wing airplanes show large cross-coupling in control and dynamic behavior which is not present on conventional symmetric airplanes and must be compensated for to obtain acceptable handling qualities. The large vertical motion simulator at NASA Ames-Moffett was used in the piloted evaluation of a proposed flight control system designed to provide decoupled handling qualities. Five discrete flight conditions were evaluated ranging from low altitude subsonic Mach numbers to moderate altitude supersonic Mach numbers. The flight control system was effective in generally decoupling the airplane. However, all participating pilots objected to the high levels of lateral acceleration encountered in pitch maneuvers. In addition, the pilots were more critical of left turns (in the direction of the trailing wingtip when skewed) than they were of right turns due to the tendency to be rolled into the left turns and out of the right turns. Asymmetric side force as a function of angle of attack was the primary cause of lateral acceleration in pitch. Along with the lateral acceleration in pitch, variation of rolling and yawing moments as functions of angle of attack caused the tendency to roll into left turns and out of right turns.
In-Flight Wing Pressure Distributions for the NASA F/A-18A High Alpha Research Vehicle
NASA Technical Reports Server (NTRS)
Davis, Mark C.; Saltzman, John A.
2000-01-01
Pressure distributions on the wings of the F/A-18A High Alpha Research Vehicle (HARV) were obtained using both flush-mounted pressure orifices and surface-mounted pressure tubing. During quasi-stabilized 1-g flight, data were gathered at ranges for angle of attack from 5 deg to 70 deg, for angle of sideslip from -12 deg to +12 deg, and for Mach from 0.23 to 0.64, at various engine settings, and with and without the leading edge extension fence installed. Angle of attack strongly influenced the wing pressure distribution, as demonstrated by a distinct flow separation pattern that occurred between the range from 15 deg to 30 deg. Influence by the leading edge extension fence was evident on the inboard wing pressure distribution, but little influence was seen on the outboard portion of the wing. Angle-of-sideslip influence on wing pressure distribution was strongest at low angle of attack. Influence of Mach number was observed in the regions of local supersonic flow, diminishing as angle of attack was increased. Engine throttle setting had little influence on the wing pressure distribution.
Aeroelastic airfoil smart spar
NASA Technical Reports Server (NTRS)
Greenhalgh, Skott; Pastore, Christopher M.; Garfinkle, Moishe
1993-01-01
Aircraft wings and rotor-blades are subject to undesirable bending and twisting excursions that arise from unsteady aerodynamic forces during high speed flight, abrupt maneuvers, or hard landings. These bending excursions can range in amplitude from wing-tip flutter to failure. A continuous-filament construction 'smart' laminated composite box-beam spar is described which corrects itself when subject to undesirable bending excursions or flutter. The load-bearing spar is constructed so that any tendency for the wing or rotor-blade to bend from its normal position is met by opposite twisting of the spar to restore the wing to its normal position. Experimental and theoretical characterization of these spars was made to evaluate the torsion-flexure coupling associated with symmetric lay-ups. The materials used were uniweave AS-4 graphite and a matrix comprised of Shell 8132 resin and U-40 hardener. Experimental tests were conducted on five spars to determine spar twist and bend as a function of load for 0, 17, 30, 45 and 60 deg fiber angle lay-ups. Symmetric fiber lay-ups do exhibit torsion-flexure couplings. Predictions of the twist and bend versus load were made for different fiber orientations in laminated spars using a spline function structural analysis. The analytical results were compared with experimental results for validation. Excellent correlation between experimental and analytical values was found.
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
Design of transonic airfoils and wings using a hybrid design algorithm
NASA Technical Reports Server (NTRS)
Campbell, Richard L.; Smith, Leigh A.
1987-01-01
A method has been developed for designing airfoils and wings at transonic speeds. It utilizes a hybrid design algorithm in an iterative predictor/corrector approach, alternating between analysis code and a design module. This method has been successfully applied to a variety of airfoil and wing design problems, including both transport and highly-swept fighter wing configurations. An efficient approach to viscous airfoild design and the effect of including static aeroelastic deflections in the wing design process are also illustrated.
Recent Applications of Higher-Order Spectral Analysis to Nonlinear Aeroelastic Phenomena
NASA Technical Reports Server (NTRS)
Silva, Walter A.; Hajj, Muhammad R.; Dunn, Shane; Strganac, Thomas W.; Powers, Edward J.; Stearman, Ronald
2005-01-01
Recent applications of higher-order spectral (HOS) methods to nonlinear aeroelastic phenomena are presented. Applications include the analysis of data from a simulated nonlinear pitch and plunge apparatus and from F-18 flight flutter tests. A MATLAB model of the Texas A&MUniversity s Nonlinear Aeroelastic Testbed Apparatus (NATA) is used to generate aeroelastic transients at various conditions including limit cycle oscillations (LCO). The Gaussian or non-Gaussian nature of the transients is investigated, related to HOS methods, and used to identify levels of increasing nonlinear aeroelastic response. Royal Australian Air Force (RAAF) F/A-18 flight flutter test data is presented and analyzed. The data includes high-quality measurements of forced responses and LCO phenomena. Standard power spectral density (PSD) techniques and HOS methods are applied to the data and presented. The goal of this research is to develop methods that can identify the onset of nonlinear aeroelastic phenomena, such as LCO, during flutter testing.
NASA Technical Reports Server (NTRS)
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.
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.
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
Flight measurements of surface pressures on a flexible supercritical research wing
NASA Technical Reports Server (NTRS)
Eckstrom, C. V.
1985-01-01
A flexible supercritical research wing, designated as ARW-1, was flight-tested as part of the NASA Drones for Aerodynamic and Structural Testing Program. Aerodynamic loads, in the form of wing surface pressure measurements, were obtained during flights at altitudes of 15,000, 20,000, and 25,000 feet at Mach numbers from 0.70 to 0.91. Surface pressure coefficients determined from pressure measurements at 80 orifice locations are presented individually as nearly continuous functions of angle of attack for constant values of Mach number. The surface pressure coefficients are also presented individually as a function of Mach number for an angle of attack of 2.0 deg. The nearly continuous values of the pressure coefficient clearly show details of the pressure gradient, which occurred in a rather narrow Mach number range. The effects of changes in angle of attack, Mach number, and dynamic pressure are also shown by chordwise pressure distributions for the range of test conditions experienced. Reynolds numbers for the tests ranged from 5.7 to 8.4 x 1,000,000.
Predicted flight characteristics of the augmentor wing jet STOL research aircraft
NASA Technical Reports Server (NTRS)
Spitzer, R. E.
1972-01-01
An existing deHavilland C-8A airplane has been modified into an augmentor wing flight test vehicle. Research objectives are to verify the augmentor flap concept and to produce data for STOL airworthiness criteria. The Modified C-8A provides the means for jet-STOL flight research down to a 60 knot approach speed. The airplane has a high thrust-to-weight ratio, high-lift flap system, vectored thrust, powerful flight controls, and lateral-directional stability augmentation system. Normal performance and handling qualities are expected to be satisfactory. Analysis and piloted simulator results indicate that stability and control characteristics in conventional flight are rated satisfactory. Handling qualities in the STOL regime are also generally satisfactory, although pilot workload is high about the longitudinal axis.
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.
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.
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.
Aeroelastic Deformation Measurements of Flap, Gap, and Overhang on a Semispan Model
NASA Technical Reports Server (NTRS)
Burner, A. W.; Liu, Tianshu; Garg, Sanjay; Ghee, Terence A.; Taylor, Nigel J.
2000-01-01
Single-camera, single-view videogrammetry has been used to determine static aeroelastic deformation of a slotted flap configuration on a semispan model at the National Transonic Facility (NTF). Deformation was determined by comparing wind-off to wind-on spatial data from targets placed on the main element, shroud, and flap of the model. Digitized video images from a camera were recorded and processed to automatically determine target image plane locations that were then corrected for sensor, lens, and frame grabber spatial errors. The videogrammetric technique has been established at NASA facilities as the technique of choice when high-volume static aeroelastic data with minimum impact on data taking is required. The primary measurement at the NTF with this technique in the past has been the measurement of static aeroelastic wing twist on full span models. The first results using the videogrammetric technique for the measurement of component deformation during semispan testing at the NTF are presented.
Structural Dynamics Modeling of HIRENASD in Support of the Aeroelastic Prediction Workshop
NASA Technical Reports Server (NTRS)
Wieseman, Carol; Chwalowski, Pawel; Heeg, Jennifer; Boucke, Alexander; Castro, Jack
2013-01-01
An Aeroelastic Prediction Workshop (AePW) was held in April 2012 using three aeroelasticity case study wind tunnel tests for assessing the capabilities of various codes in making aeroelasticity predictions. One of these case studies was known as the HIRENASD model that was tested in the European Transonic Wind Tunnel (ETW). This paper summarizes the development of a standardized enhanced analytical HIRENASD structural model for use in the AePW effort. The modifications to the HIRENASD finite element model were validated by comparing modal frequencies, evaluating modal assurance criteria, comparing leading edge, trailing edge and twist of the wing with experiment and by performing steady and unsteady CFD analyses for one of the test conditions on the same grid, and identical processing of results.
NASA Technical Reports Server (NTRS)
Silva, Walter A.; Bennett, Robert M.
1992-01-01
The Computational Aeroelasticity Program-Transonic Small Disturbance (CAP-TSD) code, developed at LaRC, is applied to the active flexible wing wind-tunnel model for prediction of transonic aeroelastic behavior. A semi-span computational model is used for evaluation of symmetric motions, and a full-span model is used for evaluation of antisymmetric motions, and a full-span model is used for evaluation of antisymmetric motions. Static aeroelastic solutions using CAP-TSD are computed. Dynamic deformations are presented as flutter boundaries in terms of Mach number and dynamic pressure. Flutter boundaries that take into account modal refinements, vorticity and entropy corrections, antisymmetric motion, and sensitivity to the modeling of the wing tip ballast stores are also presented with experimental flutter results.
Fabrication methods for YF-12 wing panels for the Supersonic Cruise Aircraft Research Program
NASA Technical Reports Server (NTRS)
Hoffman, E. L.; Payne, L.; Carter, A. L.
1975-01-01
Advanced fabrication and joining processes for titanium and composite materials are being investigated by NASA to develop technology for the Supersonic Cruise Aircraft Research (SCAR) Program. With Lockheed-ADP as the prime contractor, full-scale structural panels are being designed and fabricated to replace an existing integrally stiffened shear panel on the upper wing surface of the NASA YF-12 aircraft. The program involves ground testing and Mach 3 flight testing of full-scale structural panels and laboratory testing of representative structural element specimens. Fabrication methods and test results for weldbrazed and Rohrbond titanium panels are discussed. The fabrication methods being developed for boron/aluminum, Borsic/aluminum, and graphite/polyimide panels are also presented.
Comparisons of simulator and flight results on augmentor-wing jet STOL research aircraft
NASA Technical Reports Server (NTRS)
Innis, R. C.; Anderson, S. B.
1972-01-01
The considerations involved in making a piloted simulator an effective research tool in the design and development of new aircraft are discussed. An assessment of the limitations of the simulator in depicting real flight as well as the problem of recognizing erroneous results when the simulator is supplied with incorrect input data is made. Examples of the ways in which the simulator is used to design and develop the augmentor-wing aircraft are presented. Four areas of investigation are: (1) to design the lateral control system for proper feel and response, (2) determine the effect of engine failure during approach, (3) develop the best technique for controlling flight path during approach, and (4) the significance of lift loss in ground effect and how to compensate for such loss.
Static noise tests on augmentor wing jet STOL research aircraft (C8A Buffalo)
NASA Technical Reports Server (NTRS)
Marrs, C. C.; Harkonen, D. L.; Okeefe, J. V.
1974-01-01
Results are presented for full scale ground static acoustic tests of over-area conical nozzles and a lobe nozzle installed on the Augmentor Wing Jet STOL Research Aircraft, a modified C8A Buffalo. The noise levels and spectrums of the test nozzles are compared against those of the standard conical nozzle now in use on the aircraft. Acoustic evaluations at 152 m (500 ft), 304 m (1000 ft), and 1216 m (4000 ft) are made at various engine power settings with the emphasis on approach and takeoff power. Appendix A contains the test log and propulsion calculations. Appendix B gives the original test plan, which was closely adhered to during the test. Appendix C describes the acoustic data recording and reduction systems, with calibration details.
Application of modern control design methodology to oblique wing research aircraft
NASA Technical Reports Server (NTRS)
Vincent, James H.
1991-01-01
A Linear Quadratic Regulator synthesis technique was used to design an explicit model following control system for the Oblique Wing Research Aircraft (OWRA). The forward path model (Maneuver Command Generator) was designed to incorporate the desired flying qualities and response decoupling. The LQR synthesis was based on the use of generalized controls, and it was structured to provide a proportional/integral error regulator with feedforward compensation. An unexpected consequence of this design approach was the ability to decouple the control synthesis into separate longitudinal and lateral directional designs. Longitudinal and lateral directional control laws were generated for each of the nine design flight conditions, and gain scheduling requirements were addressed. A fully coupled 6 degree of freedom open loop model of the OWRA along with the longitudinal and lateral directional control laws was used to assess the closed loop performance of the design. Evaluations were performed for each of the nine design flight conditions.
NASA Technical Reports Server (NTRS)
Harris, C. D.
1972-01-01
Wind tunnel tests have been conducted on a research airplane model with an NASA supercritical wing to define the general character of the flow over the wing and to aid in structural design of the full scale airplane. Pressure measurements were made at Mach numbers from 0.25 to 1.30 for sideslip angles from -2.50 deg to 2.50 deg over a moderate range of angles of attack and dynamic pressures. Except for representative figures, the results are presented in tabular form without detailed analysis.
Overview of the NASA Subsonic Rotary Wing Aeronautics Research Program in Rotorcraft Crashworthiness
NASA Technical Reports Server (NTRS)
Jackson, Karen E.; Kellas, Sotiris; Fuchs, Yvonne T.
2009-01-01
This paper provides an overview of rotorcraft crashworthiness research being conducted at NASA Langley Research Center under sponsorship of the Subsonic Rotary Wing (SRW) Aeronautics Program. The research is focused in two areas: development of an externally deployable energy attenuating concept and improved prediction of rotorcraft crashworthiness. The deployable energy absorber (DEA) is a composite honeycomb structure, with a unique flexible hinge design that allows the honeycomb to be packaged and remain flat until needed for deployment. The capabilities of the DEA have been demonstrated through component crush tests and vertical drop tests of a retrofitted fuselage section onto different surfaces or terrain. The research on improved prediction of rotorcraft crashworthiness is focused in several areas including simulating occupant responses and injury risk assessment, predicting multi-terrain impact, and utilizing probabilistic analysis methods. A final task is to perform a system-integrated simulation of a full-scale helicopter crash test onto a rigid surface. A brief description of each research task is provided along with a summary of recent accomplishments.
Overview of the NASA Subsonic Rotary Wing Aeronautics Research Program in Rotorcraft Crashworthiness
NASA Technical Reports Server (NTRS)
Jackson, Karen E.; Fuchs, Yvonne T.; Kellas, Sotiris
2008-01-01
This paper provides an overview of rotorcraft crashworthiness research being conducted at NASA Langley Research Center under sponsorship of the Subsonic Rotary Wing (SRW) Aeronautics Program. The research is focused in two areas: development of an externally deployable energy attenuating concept and improved prediction of rotorcraft crashworthiness. The deployable energy absorber (DEA) is a composite honeycomb structure, with a unique flexible hinge design that allows the honeycomb to be packaged and remain flat until needed for deployment. The capabilities of the DEA have been demonstrated through component crush tests and vertical drop tests of a retrofitted fuselage section onto different surfaces or terrain. The research on improved prediction of rotorcraft crashworthiness is focused in several areas including simulating occupant responses and injury risk assessment, predicting multi-terrain impact, and utilizing probabilistic analysis methods. A final task is to perform a system-integrated simulation of a full-scale helicopter crash test onto a rigid surface. A brief description of each research task is provided along with a summary of recent accomplishments.
NASA Technical Reports Server (NTRS)
Carlson, Harry W.; Mann, Michael J.
1992-01-01
A survey of research on drag-due-to-lift minimization at supersonic speeds, including a study of the effectiveness of current design and analysis methods was conducted. The results show that a linearized theory analysis with estimated attainable thrust and vortex force effects can predict with reasonable accuracy the lifting efficiency of flat wings. Significantly better wing performance can be achieved through the use of twist and camber. Although linearized theory methods tend to overestimate the amount of twist and camber required for a given application and provide an overly optimistic performance prediction, these deficiencies can be overcome by implementation of recently developed empirical corrections. Numerous examples of the correlation of experiment and theory are presented to demonstrate the applicability and limitations of linearized theory methods with and without empirical corrections. The use of an Euler code for the estimation of aerodynamic characteristics of a twisted and cambered wing and its application to design by iteration are discussed.
NASA Technical Reports Server (NTRS)
Jordon, D. E.; Patterson, W.; Sandlin, D. R.
1985-01-01
The XV-15 Tilt Rotor Research Aircraft download phenomenon was analyzed. This phenomenon is a direct result of the two rotor wakes impinging on the wing upper surface when the aircraft is in the hover configuration. For this study the analysis proceeded along tow lines. First was a method whereby results from actual hover tests of the XV-15 aircraft were combined with drag coefficient results from wind tunnel tests of a wing that was representative of the aircraft wing. Second, an analytical method was used that modeled that airflow caused gy the two rotors. Formulas were developed in such a way that acomputer program could be used to calculate the axial velocities were then used in conjunction with the aforementioned wind tunnel drag coefficinet results to produce download values. An attempt was made to validate the analytical results by modeling a model rotor system for which direct download values were determinrd..
On the optimization of discrete structures with aeroelastic constraints
NASA Technical Reports Server (NTRS)
Mcintosh, S. C., Jr.; Ashley, H.
1978-01-01
The paper deals with the problem of dynamic structural optimization where constraints relating to flutter of a wing (or other dynamic aeroelastic performance) are imposed along with conditions of a more conventional nature such as those relating to stress under load, deflection, minimum dimensions of structural elements, etc. The discussion is limited to a flutter problem for a linear system with a finite number of degrees of freedom and a single constraint involving aeroelastic stability, and the structure motion is assumed to be a simple harmonic time function. Three search schemes are applied to the minimum-weight redesign of a particular wing: the first scheme relies on the method of feasible directions, while the other two are derived from necessary conditions for a local optimum so that they can be referred to as optimality-criteria schemes. The results suggest that a heuristic redesign algorithm involving an optimality criterion may be best suited for treating multiple constraints with large numbers of design variables.
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.
DAST in Flight just after Structural Failure of Right Wing
NASA Technical Reports Server (NTRS)
1980-01-01
Two BQM-34 Firebee II drones were modified with supercritical airfoils, called the Aeroelastic Research Wing (ARW), for the Drones for Aerodynamic and Structural Testing (DAST) program, which ran from 1977 to 1983. This photo, taken 12 June 1980, shows the DAST-1 (Serial #72-1557) immediately after it lost its right wing after suffering severe wing flutter. The vehicle crashed near Cuddeback Dry Lake. The Firebee II was selected for the DAST program because its standard wing could be removed and replaced by a supercritical wing. The project's digital flutter suppression system was intended to allow lighter wing structures, which would translate into better fuel economy for airliners. Because the DAST vehicles were flown intentionally at speeds and altitudes that would cause flutter, the program anticipated that crashes might occur. These are the image contact sheets for each image resolution of the NASA Dryden Drones for Aerodynamic and Structural Testing (DAST) Photo Gallery. From 1977 to 1983, the Dryden Flight Research Center, Edwards, California, (under two different names) conducted the DAST Program as a high-risk flight experiment using a ground-controlled, pilotless aircraft. Described by NASA engineers as a 'wind tunnel in the sky,' the DAST was a specially modified Teledyne-Ryan BQM-34E/F Firebee II supersonic target drone that was flown to validate theoretical predictions under actual flight conditions in a joint project with the Langley Research Center, Hampton, Virginia. The DAST Program merged advances in electronic remote control systems with advances in airplane design. Drones (remotely controlled, missile-like vehicles initially developed to serve as gunnery targets) had been deployed successfully during the Vietnamese conflict as reconnaissance aircraft. After the war, the energy crisis of the 1970s led NASA to seek new ways to cut fuel use and improve airplane efficiency. The DAST Program's drones provided an economical, fuel-conscious method for
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
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.
Trim angle of attack of flexible wings using non-linear aerodynamics
NASA Astrophysics Data System (ADS)
Cohen, David Erik
Multidisciplinary interactions are expected to play a significant role in the design of future high-performance aircraft (Blended-Wing Body, Truss-Braced wing; High Speed Civil transport, High-Altitude Long Endurance aircraft and future military aircraft). Also, the availability of supercomputers has made it now possible to employ high-fidelity models (Computational Fluid Dynamics for fluids and detailed finite element models for structures) at the preliminary design stage. A necessary step at that stage is to calculate the wing angle-of-attack at which the wing will generate the desired lift for the specific flight maneuver. Determination of this angle, a simple affair when the wing is rigid and the flow regime linear, becomes difficult when the wing is flexible and the flow regime non-linear. To solve this inherently nonlinear problem, a Newton's method type algorithm is developed to simultaneously calculate the deflection and the angle of attack. The developed algorithm is tested for a wing, used for in-house aeroelasticity research at Boeing (previously McDonnell Douglas) Long Beach. The trim angle of attack is calculated for a range of desired lift values. In addition to the Newton's method algorithm, a non derivative method (NDM) based on fixed point iteration, typical of fixed angle of attack calculations in aeroelasticity, is employed. The NDM, which has been extended to be able to calculate trim angle of attack, is used for one of the cases. The Newton's method calculation converges in fewer iterations, but requires more CPU time than the NDM method. The NDM, however, results in a slightly different value of the trim angle of attack. It should be noted that NDM will converge in a larger number of iterations as the dynamic pressure increases. For one value of the desired lift, both viscous and inviscid results were generated. The use of the inviscid flow model while not resulting in a markedly different value for the trim angle of attack, does result in a
An integrated approach to the optimum design of actively controlled composite wings
NASA Technical Reports Server (NTRS)
Livne, E.
1989-01-01
The importance of interactions among the various disciplines in airplane wing design has been recognized for quite some time. With the introduction of high gain, high authority control systems and the design of thin, flexible, lightweight composite wings, the integrated treatment of control systems, flight mechanics and dynamic aeroelasticity became a necessity. A research program is underway now aimed at extending structural synthesis concepts and methods to the integrated synthesis of lifting surfaces, spanning the disciplines of structures, aerodynamics and control for both analysis and design. Mathematical modeling techniques are carefully selected to be accurate enough for preliminary design purposes of the complicated, built-up lifting surfaces of real aircraft with their multiple design criteria and tight constraints. The presentation opens with some observations on the multidisciplinary nature of wing design. A brief review of some available state of the art practical wing optimization programs and a brief review of current research effort in the field serve to illuminate the motivation and support the direction taken in our research. The goals of this research effort are presented, followed by a description of the analysis and behavior sensitivity techniques used. The presentation concludes with a status report and some forecast of upcoming progress.
Full potential unsteady computations including aeroelastic effects
NASA Technical Reports Server (NTRS)
Shankar, Vijaya; Ide, Hiroshi
1989-01-01
A unified formulation is presented based on the full potential framework coupled with an appropriate structural model to compute steady and unsteady flows over rigid and flexible configurations across the Mach number range. The unsteady form of the full potential equation in conservation form is solved using an implicit scheme maintaining time accuracy through internal Newton iterations. A flux biasing procedure based on the unsteady sonic reference conditions is implemented to compute hyperbolic regions with moving sonic and shock surfaces. The wake behind a trailing edge is modeled using a mathematical cut across which the pressure is satisfied to be continuous by solving an appropriate vorticity convection equation. An aeroelastic model based on the generalized modal deflection approach interacts with the nonlinear aerodynamics and includes both static as well as dynamic structural analyses capability. Results are presented for rigid and flexible configurations at different Mach numbers ranging from subsonic to supersonic conditions. The dynamic response of a flexible wing below and above its flutter point is demonstrated.
NASA Technical Reports Server (NTRS)
Dicarlo, Daniel J.; Brown, Philip W.; Hallissy, James B.
1992-01-01
Flight tests of an F-106B aircraft equipped with a leading-edge vortex flap, which represented the culmination of a research effort to examine the effectiveness of the flap, were conducted at the NASA Langley Research Center. The purpose of the flight tests was to establish a data base on the use of a wing leading-edge vortex flap as a means to validate the design and analysis methods associated with the development of such a vortical flow-control concept. The overall experiment included: refinements of the design codes for vortex flaps; numerous wind tunnel entries to aid in verifying design codes and determining basic aerodynamic characteristics; design and fabrication of the flaps, structural modifications to the wing tip and leading edges of the test aircraft; development and installation of an aircraft research instrumentation system, including wing and flap surface pressure measurements and selected structural loads measurements; ground-based simulation to assess flying qualities; and finally, flight testing. This paper reviews the operational aspects associated with the flight experiment, which includes a description of modifications to the research airplane, the overall flight test procedures, and problems encountered. Selected research results are also presented to illustrate the accomplishments of the research effort.
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
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.
Thin tailored composite wing for civil tiltrotor
NASA Technical Reports Server (NTRS)
Rais-Rohani, Masoud
1994-01-01
The tiltrotor aircraft is a flight vehicle which combines the efficient low speed (i.e., take-off, landing, and hover) characteristics of a helicopter with the efficient cruise speed of a turboprop airplane. A well-known example of such vehicle is the Bell-Boeing V-22 Osprey. The high cruise speed and range constraints placed on the civil tiltrotor require a relatively thin wing to increase the drag-divergence Mach number which translates into lower compressibility drag. It is required to reduce the wing maximum thickness-to-chord ratio t/c from 23% (i.e., V-22 wing) to 18%. While a reduction in wing thickness results in improved aerodynamic efficiency, it has an adverse effect on the wing structure and it tends to reduce structural stiffness. If ignored, the reduction in wing stiffness leads to susceptibility to aeroelastic and dynamic instabilities which may consequently cause a catastrophic failure. By taking advantage of the directional stiffness characteristics of composite materials the wing structure may be tailored to have the necessary stiffness, at a lower thickness, while keeping the weight low. The goal of this study is to design a wing structure for minimum weight subject to structural, dynamic and aeroelastic constraints. The structural constraints are in terms of strength and buckling allowables. The dynamic constraints are in terms of wing natural frequencies in vertical and horizontal bending and torsion. The aeroelastic constraints are in terms of frequency placement of the wing structure relative to those of the rotor system. The wing-rotor-pylon aeroelastic and dynamic interactions are limited in this design study by holding the cruise speed, rotor-pylon system, and wing geometric attributes fixed. To assure that the wing-rotor stability margins are maintained a more rigorous analysis based on a detailed model of the rotor system will need to ensue following the design study. The skin-stringer-rib type architecture is used for the wing
A Summary of Data and Findings from the First Aeroelastic Prediction Workshop
NASA Technical Reports Server (NTRS)
Schuster, David M.; Chwalowski, Pawel.; Heeg, Jennifer; Wieseman, Carol D.
2012-01-01
This paper summarizes data and findings from the first Aeroelastic Prediction Workshop (AePW) held in April, 2012. The workshop has been 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. For this initial workshop, three subject configurations have been 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. Keywords: Unsteady Aerodynamics, Aeroelasticity, Computational Fluid Dynamics, Transonic Flow, Separated Flow.
Flutter analysis of low aspect ratio wings
NASA Technical Reports Server (NTRS)
Parnell, L. A.
1986-01-01
Several very low aspect ratio flat plate wing configurations are analyzed for their aerodynamic instability (flutter) characteristics. All of the wings investigated are delta planforms with clipped tips, made of aluminum alloy plate and cantilevered from the supporting vehicle body. Results of both subsonic and supersonic NASTRAN aeroelastic analyses as well as those from another version of the program implementing the supersonic linearized aerodynamic theory are presented. Results are selectively compared with the experimental data; however, supersonic predictions of the Mach Box method in NASTRAN are found to be erratic and erroneous, requiring the use of a separate program.
Aeroelastic 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
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
Accomplishments of the Abrupt Wing Stall (AWS) Program and Future Research Requirements
NASA Technical Reports Server (NTRS)
Hall, Robert M.; Woodson, Shawn H.; Chambers, Joseph R.
2003-01-01
The Abrupt Wing Stall (AWS) Program has addressed the problem of uncommanded lateral motions, such as wing drop and wing rock, at transonic speeds. The genesis of this Program was the experience of the F/A-1 8E/F Program in the late 1990's, when wing drop was discovered in the heart of the maneuver envelope for the pre-production aircraft. While the F/A-1 8E/F problem was subsequently corrected by a leading-edge flap scheduling change and the addition of a porous door to the wing fold fairing, the AWS Program was initiated as a national response to the lack of technology readiness available at the time of the F/A-18E/F Development Program. The AWS Program objectives were to define causal factors for the F/A-18E/F experience, to gain insights into the flow physics associated with wing drop, and to develop methods and analytical tools so that future programs could identify this type of problem before going to flight test. The paper reviews, for the major goals of the AWS Program, the status of the technology before the
NASA Technical Reports Server (NTRS)
Woodrow Whitlow, Jr. (Editor); Todd, Emily N. (Editor)
1999-01-01
These proceedings represent a collection of the latest advances in aeroelasticity and structural dynamics from the world community. Research in the areas of unsteady aerodynamics and aeroelasticity, structural modeling and optimization, active control and adaptive structures, landing dynamics, certification and qualification, and validation testing are highlighted in the collection of papers. The wide range of results will lead to advances in the prediction and control of the structural response of aircraft and spacecraft.
Unsteady aerodynamics and flow control for flapping wing flyers
NASA Astrophysics Data System (ADS)
Ho, Steven; Nassef, Hany; Pornsinsirirak, Nick; Tai, Yu-Chong; Ho, Chih-Ming
2003-11-01
The creation of micro air vehicles (MAVs) of the same general sizes and weight as natural fliers has spawned renewed interest in flapping wing flight. With a wingspan of approximately 15 cm and a flight speed of a few meters per second, MAVs experience the same low Reynolds number (10 4-10 5) flight conditions as their biological counterparts. In this flow regime, rigid fixed wings drop dramatically in aerodynamic performance while flexible flapping wings gain efficacy and are the preferred propulsion method for small natural fliers. Researchers have long realized that steady-state aerodynamics does not properly capture the physical phenomena or forces present in flapping flight at this scale. Hence, unsteady flow mechanisms must dominate this regime. Furthermore, due to the low flight speeds, any disturbance such as gusts or wind will dramatically change the aerodynamic conditions around the MAV. In response, a suitable feedback control system and actuation technology must be developed so that the wing can maintain its aerodynamic efficiency in this extremely dynamic situation; one where the unsteady separated flow field and wing structure are tightly coupled and interact nonlinearly. For instance, birds and bats control their flexible wings with muscle tissue to successfully deal with rapid changes in the flow environment. Drawing from their example, perhaps MAVs can use lightweight actuators in conjunction with adaptive feedback control to shape the wing and achieve active flow control. This article first reviews the scaling laws and unsteady flow regime constraining both biological and man-made fliers. Then a summary of vortex dominated unsteady aerodynamics follows. Next, aeroelastic coupling and its effect on lift and thrust are discussed. Afterwards, flow control strategies found in nature and devised by man to deal with separated flows are examined. Recent work is also presented in using microelectromechanical systems (MEMS) actuators and angular speed
Aerodynamic effects of flexibility in flapping wings
Zhao, Liang; Huang, Qingfeng; Deng, Xinyan; Sane, Sanjay P.
2010-01-01
Recent work on the aerodynamics of flapping flight reveals fundamental differences in the mechanisms of aerodynamic force generation between fixed and flapping wings. When fixed wings translate at high angles of attack, they periodically generate and shed leading and trailing edge vortices as reflected in their fluctuating aerodynamic force traces and associated flow visualization. In contrast, wings flapping at high angles of attack generate stable leading edge vorticity, which persists throughout the duration of the stroke and enhances mean aerodynamic forces. Here, we show that aerodynamic forces can be controlled by altering the trailing edge flexibility of a flapping wing. We used a dynamically scaled mechanical model of flapping flight (Re ≈ 2000) to measure the aerodynamic forces on flapping wings of variable flexural stiffness (EI). For low to medium angles of attack, as flexibility of the wing increases, its ability to generate aerodynamic forces decreases monotonically but its lift-to-drag ratios remain approximately constant. The instantaneous force traces reveal no major differences in the underlying modes of force generation for flexible and rigid wings, but the magnitude of force, the angle of net force vector and centre of pressure all vary systematically with wing flexibility. Even a rudimentary framework of wing veins is sufficient to restore the ability of flexible wings to generate forces at near-rigid values. Thus, the magnitude of force generation can be controlled by modulating the trailing edge flexibility and thereby controlling the magnitude of the leading edge vorticity. To characterize this, we have generated a detailed database of aerodynamic forces as a function of several variables including material properties, kinematics, aerodynamic forces and centre of pressure, which can also be used to help validate computational models of aeroelastic flapping wings. These experiments will also be useful for wing design for small robotic
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)
Mcgehee, C. R.
1986-01-01
This is Part 2-Appendices of a study conducted under Drones for Aerodynamic and Structural Testing (DAST) Program to accomplish the final design and hardware fabrication for four active control systems compatible with and ready for installation in the NASA Aeroelastic Research Wing No. 2 (ARW-2) and Firebee II drone flight test vehicle. The wing structure was designed so that Active Control Systems (ACS) are required in the normal flight envelope by integrating control system design with aerodynamics and structure technologies. The DAST ARW-2 configuration uses flutter suppression, relaxed static stability, and gust and maneuver load alleviation ACS systems, and an automatic flight control system. Performance goals and criteria were applied to individual systems and the systems collectively to assure that vehicle stability margins, flutter margins, flying qualities, and load reductions were achieved.
NASA Technical Reports Server (NTRS)
Mcgehee, C. R.
1986-01-01
A study was conducted under Drones for Aerodynamic and Structural Testing (DAST) program to accomplish the final design and hardware fabrication for four active control systems compatible with and ready for installation in the NASA Aeroelastic Research Wing No. 2 (ARW-2) and Firebee II drone flight test vehicle. The wing structure was designed so that Active Control Systems (ACS) are required in the normal flight envelope by integrating control system design with aerodynamics and structure technologies. The DAST ARW-2 configuration uses flutter suppression, relaxed static stability, and gust and maneuver load alleviation ACS systems, and an automatic flight control system. Performance goals and criteria were applied to individual systems and the systems collectively to assure that vehicle stability margins, flutter margins, flying qualities and load reductions are achieved.
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
NASA Technical Reports Server (NTRS)
Balough, D. L.; Sandlin, D. R.
1986-01-01
The purpose of this report is to establish linear, decoupled models of rigid body motion for the fixed wing configuration of the Rotor Systems Research Aircraft (RSRA). Longitudinal and lateral control surface fixed linear models were created from aircraft time histories using current system identification techniques. Models were obtained from computer simulation at 160 KCAS and 200 KCAS, and from flight data at 160 KCAS. Comparisons were performed to examine modeling accuracy, variation of dynamics with airspeed and correlation of simulation and flight data results. The results showed that the longitudinal and lateral linear models accurately predicted RSRA dynamics. The flight data results showed that no significant handling qualities problems were present in the RSRA fixed wing aircraft at the flight speed tested.
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.
A nonlinear computational aeroelasticity model for aircraft wings
NASA Astrophysics Data System (ADS)
Feng, Zhengkun
Cette these presente le developpement d'un code d'aeroelasticite nonlineaire base sur un solveur CFD robuste afin de l'appliquer aux ailes flexibles en ecoulement transsonique. Le modele mathematique complet est base sur les equations du mouvement des structures et les equations d'Euler pour les ecoulements transsoniques non-visqueux. La strategie de traiter tel systeme complexe par un couplage etage presente des avantages pour le developpement d'un code modulaire et facile a faire evoluer. La non-correspondance entre les deux grilles de calcul a l'interface fluide-structure, due aux differences des tailles et des types des elements utilises par la resolution de l'ecoulement et de la structure, est resolue par l'ajout d'un module specifique. Les transferts des informations entre ces deux grilles satisfont la loi de la conservation de l'energie. Le modele nonlineaire de la dynamique du fluide base sur la description Euler-Lagrange est discretise dans le maillage mobile. Le modele pour le calcul des structures est suppose lineaire dans lequel la methode de superposition modale est appliquee pour reduire le temps de calcul et la dimension de la memoire. Un autre modele pour la structure base directement sur la methode des elements finis est aussi developpe. Il est egalement couple dans le code pour prouver son extension future aux applications plus generales. La nonlinearite est une autre source de complexite du systeme bien que celle-ci est prevue uniquement dans le modele aerodynamique. L'algorithme GMRES nonlineaire avec le preconditioneur ILUT est implemente dans le solveur CFD ou un capteur de choc pour les ecoulements transsoniques et la technique de stabilisation numerique SUPG pour des ecoulements domines par la convection sont appliques. Un schema du second ordre est utilise pour la discretisation temporelle. Les composants de ce code sont valides par des tests numeriques. Le modele complet est applique et valide sur l'aile aeroelastique AGARD 445.6 dans le cas du nombre de Mach 0.96 qui est une valeur critique en flottement. Les simulations de flottement donnent des resultats numeriques satisfaisants en comparaison avec ceux experimentaux.
Parametric design study of an aeroelastic flutter energy harvester
NASA Astrophysics Data System (ADS)
Bryant, Matthew; Wolff, Eric; Garcia, Ephrahim
2011-03-01
This paper investigates a novel mechanism for powering wireless sensors or low power electronics by extracting energy from an ambient fluid flow using a piezoelectric energy harvester driven 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. This value sets the lower bound on the operating wind speed and frequency range of the system. A system of coupled equations that describe the structural, aerodynamic, and electromechanical aspects of the system are used to model the system dynamics. The model uses unsteady aerodynamic modeling to predict the aerodynamic forces and moments acting on the structure and to account for the effects of vortices shed by the flapping wing, while a modal summation technique is used to model the flexible piezoelectric structure. This model is applied to examine the effects on the cut-in wind speed of the system when several design parameters are tuned and the size and mass 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. Experimental wind tunnel results are included to validate the model predictions.
NASA Technical Reports Server (NTRS)
Bartels, Robert E.
1999-01-01
This paper presents a modification of the spring analogy scheme which uses axial linear spring stiffness with selective spring stiffening/relaxation. An alternate approach to solving the geometric conservation law is taken which eliminates the need for storage of metric Jacobians at previous time steps. Efficiency and verification are illustrated with several unsteady 2-D airfoil Euler computations. The method is next applied to the computation of the turbulent flow about a 2-D airfoil and wing with two and three- dimensional moving spoiler surfaces, and the results compared with Benchmark Active Controls Technology (BACT) experimental data. The aeroelastic response at low dynamic pressure of an airfoil to a single large scale oscillation of a spoiler surface is computed. This study confirms that it is possible to achieve accurate solutions with a very large time step for aeroelastic problems using the fluid solver and aeroelastic integrator as discussed in this paper.
Comparison of supercritical and conventional wing flutter characteristics
NASA Technical Reports Server (NTRS)
Farmer, M. G.; Hanson, P. W.; Wynne, E. C.
1976-01-01
A wind-tunnel study was undertaken to directly compare the measured flutter boundaries of two dynamically similar aeroelastic models which had the same planform, maximum thickness-to-chord ratio, and as nearly identical stiffness and mass distributions as possible, with one wing having a supercritical airfoil and the other a conventional airfoil. The considerations and problems associated with flutter testing supercritical wing models at or near design lift coefficients are discussed, and the measured transonic boundaries of the two wings are compared with boundaries calculated with a subsonic lifting surface theory.
Suppression of bending-torsion wing flutter using self-straining controllers
NASA Astrophysics Data System (ADS)
Lin, Jensen Cheng-Sheng
Flutter is an instability endemic to aircraft that occurs at high enough air speed. Suppression of flutter is in the interest of safety and economy. In this study, we propose a purely analytical approach to the problem flutter suppression. Counter to the commercially available numerical schemes, mathematical precisions are provided to gain a better understanding of the flutter phenomenon and the controller performance. We model the wing structure and aerodynamics with a pair of time-invariant linear partial differential equations. The control action of the self-straining material is easily incorporated into the structural model as boundary control. This model faithfully captures the flutter phenomenon as well as the control action. A State Space representation is carefully chosen for the aeroelastic model. The problem of flutter analysis is reduced to evaluating the resolvent of the aeroelastic operator. We also present a Laplace-Fourier Transform version of the Possio equation in the theory of Unsteady Subsonic Aerodynamics. This new version enables us to obtain explicit formulas for the lift and moment, which in turn afford us to analyze the flutter problem more readily. Analyses reveal the torsion controllers are effective in extending the flutter boundary while the bending controllers are not. A series of experiments were designed to validate our theoretical models for flutter analysis and to test the performance of self-straining actuators. An aeroelastic wing with self-straining sensors and actuators were designed to flutter within the speed limit of the vehicle as well as the assumptions of our theoretical model. The NASA Ground Research Vehicle, the "Roadrunner" served as the platform for these experiments. The processed data from the field tests showed the theoretical prediction of flutter speed is accurate. Theoretical calculations for both of the frequencies and damping as function of air speed were also found to be within the experimental error. However
NASA Technical Reports Server (NTRS)
Welch, Gerard E.; Hathaway, Michael D.; Skoch, Gary J.; Snyder, Christopher A.
2012-01-01
Technical challenges of compressors for future rotorcraft engines are driven by engine-level and component-level requirements. Cycle analyses are used to highlight the engine-level challenges for 3000, 7500, and 12000 SHP-class engines, which include retention of performance and stability margin at low corrected flows, and matching compressor type, axial-flow or centrifugal, to the low corrected flows and high temperatures in the aft stages. At the component level: power-to-weight and efficiency requirements impel designs with lower inherent aerodynamic stability margin; and, optimum engine overall pressure ratios lead to small blade heights and the associated challenges of scale, particularly increased clearance-to-span ratios. The technical challenges associated with the aerodynamics of low corrected flows and stability management impel the compressor aero research and development efforts reviewed herein. These activities include development of simple models for clearance sensitivities to improve cycle calculations, full-annulus, unsteady Navier-Stokes simulations used to elucidate stall, its inception, and the physics of stall control by discrete tip-injection, development of an actuator-duct-based model for rapid simulation of nonaxisymmetric flow fields (e.g., due inlet circumferential distortion), advanced centrifugal compressor stage development and experimentation, and application of stall control in a T700 engine.
Structural dynamic and aeroelastic considerations for hypersonic vehicles
NASA Technical Reports Server (NTRS)
Cazier, F. W., Jr.; Doggett, Robert V., Jr.; Ricketts, Rodney H.
1991-01-01
The specific geometrical, structural, and operational environment characteristics of hypersonic vehicles are discussed with particular reference to aerospace plane type configurations. A discussion of the structural dynamic and aeroelastic phenomena that must be addressed for this class of vehicles is presented. These phenomena are in the aeroservothermoelasticity technical area. Some illustrative examples of recent experimental and analytical work are given. Some examples of current research are pointed out.
Aeroelastic Deformation Measurements of Flap, Gap, and Overhang on a Semispan Model
NASA Technical Reports Server (NTRS)
Burner, A. W.; Liu, Tian-Shu; Garg, Sanjay; Ghee, Terence A.; Taylor, Nigel J.
2001-01-01
Single-camera, single-view videogrammetry has been used for the first time to determine static aeroelastic deformation of a slotted flap configuration on a semispan model at the National Transonic Facility (NTF). Deformation was determined by comparing wind-off to wind-on spatial data from targets placed on the main element, shroud, and flap of the model. Digitized video images from a camera were recorded and processed to automatically determine target image plane locations that were then corrected for sensor, lens, and frame grabber spatial errors. The videogrammetric technique used for the measurements presented here has been established at NASA facilities as the technique of choice when high-volume static aeroelastic data with minimum impact on data taking is required. However, the primary measurement at the NTF with this technique in the past has been the measurement of the static aeroelastic wing twist of the main wing element on full span models rather than for the measurement of component deformation. Considerations for using the videogrammetric technique for semispan component deformation measurements as well as representative results are presented.
Results of including geometric nonlinearities in an aeroelastic model of an F/A-18
NASA Technical Reports Server (NTRS)
Buttrill, Carey S.
1989-01-01
An integrated, nonlinear simulation model suitable for aeroelastic modeling of fixed-wing aircraft has been developed. While the author realizes that the subject of modeling rotating, elastic structures is not closed, it is believed that the equations of motion developed and applied herein are correct to second order and are suitable for use with typical aircraft structures. The equations are not suitable for large elastic deformation. In addition, the modeling framework generalizes both the methods and terminology of non-linear rigid-body airplane simulation and traditional linear aeroelastic modeling. Concerning the importance of angular/elastic inertial coupling in the dynamic analysis of fixed-wing aircraft, the following may be said. The rigorous inclusion of said coupling is not without peril and must be approached with care. In keeping with the same engineering judgment that guided the development of the traditional aeroelastic equations, the effect of non-linear inertial effects for most airplane applications is expected to be small. A parameter does not tell the whole story, however, and modes flagged by the parameter as significant also need to be checked to see if the coupling is not a one-way path, i.e., the inertially affected modes can influence other modes.
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 Technical Reports Server (NTRS)
Liu, Tianshu; Kuykendoll, K.; Rhew, R.; Jones, S.
2004-01-01
This paper describes the avian wing geometry (Seagull, Merganser, Teal and Owl) extracted from non-contact surface measurements using a three-dimensional laser scanner. The geometric quantities, including the camber line and thickness distribution of airfoil, wing planform, chord distribution, and twist distribution, are given in convenient analytical expressions. Thus, the avian wing surfaces can be generated and the wing kinematics can be simulated. The aerodynamic characteristics of avian airfoils in steady inviscid flows are briefly discussed. The avian wing kinematics is recovered from videos of three level-flying birds (Crane, Seagull and Goose) based on a two-jointed arm model. A flapping seagull wing in the 3D physical space is re-constructed from the extracted wing geometry and kinematics.
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.
Aeroelastic flutter of feathers, flight and the evolution of non-vocal communication in birds.
Clark, Christopher J; Prum, Richard O
2015-11-01
Tonal, non-vocal sounds are widespread in both ordinary bird flight and communication displays. We hypothesized these sounds are attributable to an aerodynamic mechanism intrinsic to flight feathers: aeroelastic flutter. Individual wing and tail feathers from 35 taxa (from 13 families) that produce tonal flight sounds were tested in a wind tunnel. In the wind tunnel, all of these feathers could flutter and generate tonal sound, suggesting that the capacity to flutter is intrinsic to flight feathers. This result implies that the aerodynamic mechanism of aeroelastic flutter is potentially widespread in flight of birds. However, the sounds these feathers produced in the wind tunnel replicated the actual flight sounds of only 15 of the 35 taxa. Of the 20 negative results, we hypothesize that 10 are false negatives, as the acoustic form of the flight sound suggests flutter is a likely acoustic mechanism. For the 10 other taxa, we propose our negative wind tunnel results are correct, and these species do not make sounds via flutter. These sounds appear to constitute one or more mechanism(s) we call 'wing whirring', the physical acoustics of which remain unknown. Our results document that the production of non-vocal communication sounds by aeroelastic flutter of flight feathers is widespread in birds. Across all birds, most evolutionary origins of wing- and tail-generated communication sounds are attributable to three mechanisms: flutter, percussion and wing whirring. Other mechanisms of sound production, such as turbulence-induced whooshes, have evolved into communication sounds only rarely, despite their intrinsic ubiquity in ordinary flight. PMID:26385327
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 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.
Witteveen, Jeroen A.S. Bijl, Hester
2009-10-01
The Unsteady Adaptive Stochastic Finite Elements (UASFE) method resolves the effect of randomness in numerical simulations of single-mode aeroelastic responses with a constant accuracy in time for a constant number of samples. In this paper, the UASFE framework is extended to multi-frequency responses and continuous structures by employing a wavelet decomposition pre-processing step to decompose the sampled multi-frequency signals into single-frequency components. The effect of the randomness on the multi-frequency response is then obtained by summing the results of the UASFE interpolation at constant phase for the different frequency components. Results for multi-frequency responses and continuous structures show a three orders of magnitude reduction of computational costs compared to crude Monte Carlo simulations in a harmonically forced oscillator, a flutter panel problem, and the three-dimensional transonic AGARD 445.6 wing aeroelastic benchmark subject to random fields and random parameters with various probability distributions.
Determining XV-15 aeroelastic modes from flight data with frequency-domain methods
NASA Technical Reports Server (NTRS)
Acree, C. W., Jr.; Tischler, Mark B.
1993-01-01
The XV-15 tilt-rotor wing has six major aeroelastic modes that are close in frequency. To precisely excite individual modes during flight test, dual flaperon exciters with automatic frequency-sweep controls were installed. The resulting structural data were analyzed in the frequency domain (Fourier transformed). All spectral data were computed using chirp z-transforms. Modal frequencies and damping were determined by fitting curves to frequency-response magnitude and phase data. The results given in this report are for the XV-15 with its original metal rotor blades. Also, frequency and damping values are compared with theoretical predictions made using two different programs, CAMRAD and ASAP. The frequency-domain data-analysis method proved to be very reliable and adequate for tracking aeroelastic modes during flight-envelope expansion. This approach required less flight-test time and yielded mode estimations that were more repeatable, compared with the exponential-decay method previously used.
Non-linear aeroelastic prediction for aircraft applications
NASA Astrophysics Data System (ADS)
de C. Henshaw, M. J.; Badcock, K. J.; Vio, G. A.; Allen, C. B.; Chamberlain, J.; Kaynes, I.; Dimitriadis, G.; Cooper, J. E.; Woodgate, M. A.; Rampurawala, A. M.; Jones, D.; Fenwick, C.; Gaitonde, A. L.; Taylor, N. V.; Amor, D. S.; Eccles, T. A.; Denley, C. J.
2007-05-01
Current industrial practice for the prediction and analysis of flutter relies heavily on linear methods and this has led to overly conservative design and envelope restrictions for aircraft. Although the methods have served the industry well, it is clear that for a number of reasons the inclusion of non-linearity in the mathematical and computational aeroelastic prediction tools is highly desirable. The increase in available and affordable computational resources, together with major advances in algorithms, mean that non-linear aeroelastic tools are now viable within the aircraft design and qualification environment. The Partnership for Unsteady Methods in Aerodynamics (PUMA) Defence and Aerospace Research Partnership (DARP) was sponsored in 2002 to conduct research into non-linear aeroelastic prediction methods and an academic, industry, and government consortium collaborated to address the following objectives: To develop useable methodologies to model and predict non-linear aeroelastic behaviour of complete aircraft. To evaluate the methodologies on real aircraft problems. To investigate the effect of non-linearities on aeroelastic behaviour and to determine which have the greatest effect on the flutter qualification process. These aims have been very effectively met during the course of the programme and the research outputs include: New methods available to industry for use in the flutter prediction process, together with the appropriate coaching of industry engineers. Interesting results in both linear and non-linear aeroelastics, with comprehensive comparison of methods and approaches for challenging problems. Additional embryonic techniques that, with further research, will further improve aeroelastics capability. This paper describes the methods that have been developed and how they are deployable within the industrial environment. We present a thorough review of the PUMA aeroelastics programme together with a comprehensive review of the relevant research
Ground vibration test of the XV-15 Tiltrotor Research Aircraft and pretest predictions
NASA Technical Reports Server (NTRS)
Studebaker, Karen; Abrego, Anita
1994-01-01
The first comprehensive ground vibration survey was performed on the XV-15 Tiltrotor Research Aircraft to measure the vibration modes of the airframe and to provide data critical for determining whirl flutter stability margins. The aircraft was suspended by the wings with bungee cords and cables. A NASTRAN finite element model was used in the design of the suspension system to minimize its interference with the wing modes. The primary objective of the test was to measure the dynamic characteristics of the wings and pylons for aeroelastic stability analysis. In addition, over 130 accelerometers were placed on the airframe to characterize the fuselage, wing, and tail vibration. Pretest predictions were made with the NASTRAN model as well as correlations with the test data. The results showed that the suspension system provided the isolation necessary for modal measurements.
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
Aerodynamic Analysis of the Truss-Braced Wing Aircraft Using Vortex-Lattice Superposition Approach
NASA Technical Reports Server (NTRS)
Ting, Eric Bi-Wen; Reynolds, Kevin Wayne; Nguyen, Nhan T.; Totah, Joseph J.
2014-01-01
aeroelasticity and flutter. These studies have sought to develop tools and methods to analyze aeroelastic effects by laying the foundation for more modern high aspect ratio wing aircraft such as the Truss-Braced Wing (TBW).1-3 Originally suggested by Northrop Grumman for the development of a long-range bomber, the idea of using truss structures to alleviate the bending moments of an ultra-high aspect ratio wing has culminated in more than a decade of work focused on understanding the aeroelastic properties and structural weight penalties due to the more aerodynamically efficient wing. The Subsonic Ultra Green Aircraft Research (SUGAR) Truss-Braced Wing (TBW) aircraft concept is a Boeingdeveloped N+3 aircraft configuration funded by NASA ARMD Fixed Wing Project.4, 5 The TBW aircraft concept is designed to be aerodynamically efficient by employing an aspect ratio on the order of 14, which is significantly greater than those of conventional aircraft wings. As a result, intermediate structural supports are required. The main wings The development of the TBW aircraft is supported through a collaboration between the NASA FixedWing Project, Boeing Research and Technology, and a number of other organizations. Multidisciplinary design analysis and optimization (MDAO) studies have been conducted at each stage to improve the wing aerodynamics, structural efficiency, and flight performance using advanced N+4 turbofan engines. These MDAO studies have refined the geometry of the wing and configuration layout and have involved trade studies involving minimizing induced drag with wing span, minimizing profile drag at lower Reynolds numbers, and minimizing wave drag due to the addition of the strut and brace. The chart in Fig. 2 summarizes progression of the past revisions of the TBW aircraft design at various developmental stage This paper presents an initial aerodynamic analysis of the TBW aircraft using a conceptual vortex-lattice aerodynamic tool VORLAX coupled with the aerodynamic
Comparison of supercritical and conventional wing flutter characteristics
NASA Technical Reports Server (NTRS)
Farmer, M. G.; Hanson, P. W.
1976-01-01
A wind-tunnel study is described in which it was attempted to compare the measured flutter boundaries of two dynamically similar aeroelastic models with identical planform, maximum thickness-to-chord ratio, and as nearly identical stiffness and mass distributions as possible, but with one wing having a supercritical airfoil and the other a conventional one. At subsonic Mach numbers, the flutter boundary for the supercritical wing was above that of the conventional wing, as predicted by flutter calculations using subsonic lifting theory. In the transonic region, however, the supercritical wing boundary decreases more rapidly and the minimum flutter point occurs at a dynamic pressure below the conventional wing boundary. Airfoil shape effects may account for some of the difference in the flutter boundaries of the two airfoils.
Physical properties of the benchmark models program supercritical wing
NASA Technical Reports Server (NTRS)
Dansberry, Bryan E.; Durham, Michael H.; Bennett, Robert M.; Turnock, David L.; Silva, Walter A.; Rivera, Jose A., Jr.
1993-01-01
The goal of the Benchmark Models Program is to provide data useful in the development and evaluation of aeroelastic computational fluid dynamics (CFD) codes. To that end, a series of three similar wing models are being flutter tested in the Langley Transonic Dynamics Tunnel. These models are designed to simultaneously acquire model response data and unsteady surface pressure data during wing flutter conditions. The supercritical wing is the second model of this series. It is a rigid semispan model with a rectangular planform and a NASA SC(2)-0414 supercritical airfoil shape. The supercritical wing model was flutter tested on a flexible mount, called the Pitch and Plunge Apparatus, that provides a well-defined, two-degree-of-freedom dynamic system. The supercritical wing model and associated flutter test apparatus is described and experimentally determined wind-off structural dynamic characteristics of the combined rigid model and flexible mount system are included.
NASA Technical Reports Server (NTRS)
Hughes, Christopher E.
2009-01-01
An overview of the NASA Fundamental Aeronautics Program (FAP) mission and goals is presented. One of the subprograms under the FAP, the Subsonic Fixed Wing Project (SFW), is the focus of the presentation. The SFW system environmental metrics are discussed, along with highlights of planned, systematic approach to research to reduce the environmental impact of commercial aircraft in the areas of acoustics, fuel burn and emissions. The presentation then focuses on collaborative research being conducted with U.S. Industry on the Ultra High Bypass (UHB) engine cycle, the propulsion cycle selected by the SFW to meet the system goals. The partnerships with General Electric Aviation to investigate Open Rotor propulsion concepts and with Pratt & Whitney to investigate the Geared Turbofan UHB engine are highlighted, including current and planned future collaborative research activities with NASA and each organization.
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
Investigation of the Flutter Suppression by Fuzzy Logic Control for Hypersonic Wing
NASA Astrophysics Data System (ADS)
Li, Dongxu; Luo, Qing; Xu, Rui
This paper presents a fundamental study of flutter characteristics and control performance of an aeroelastic system based on a two-dimensional double wedge wing in the hypersonic regime. Dynamic equations were established based on the modified third order nonlinear piston theory and some nonlinear structural effects are also included. A set of important parameters are observed. And then aeroelastic control law is designed to suppress the amplitude of the LCOs for the system in the sub/supercritical speed range by applying fuzzy logic control on the input of the deflection of the flap. The overall effects of the parameters on the aeroelastic system were outlined. Nonlinear aeroelastic responses in the open- and closed-loop system are obtained through numerical methods. The simulations show fuzzy logic control methods are effective in suppressing flutter and provide a smart approach for this complicated system.
Wind Tunnel Testing of the NASA-DFRC Flutterometer using a Two DOF Wing Section
NASA Technical Reports Server (NTRS)
Strganac, Thomas W.
2001-01-01
Flutter of an aeroelastic structure is potentially destructive aeroelastic instability. This phenomenon has motivated research within the aeroelastic community to develop methods that can accurately predict aeroelastic instabilities. The Flutterometer method used herein, and as developed by NASA DFRC, is based upon the mu method which has been coupled with wavelet filtering processes in estimating aeroelastic models from flight data. The approach leads to a methodology to predict the occurrence of flutter boundaries, and may prove to reliably predict flutter boundaries during flight tests. An analytical model is used as the first estimate of the aeroelastic structural dynamics, and uncertainty operators are introduced into the system to model variations between the theoretical system and the physical system. The modelling uncertainties are then updated from experimental data. Although the model used did not work well with this particular experiment, a sensitivity analysis was additionally performed and improvements suggested.
Transonic calculations for a flexible supercritical wing and comparison with experiment
NASA Technical Reports Server (NTRS)
Bennett, R. M.; Seidel, D. A.; Sandford, M. C.
1985-01-01
Pressure data measured on the flexible DAST ARW-2 wing are compared with results calculated using the transonic small perturbation code XTRAN3S. A brief description of the analysis is given and a recently-developed grid coordinate transformation is described. Calculations are presented for the rigid and flexible wing for Mach numbers from 0.60 to 0.90 and dynamic pressures from 0 to 1000 psf. Calculated and measured static pressures and wing deflections are compared, and calculated static aeroelastic trends are given. Attempts to calculate the transonic instability boundary of the wing are described.
A study of high-altitude manned research aircraft employing strut-braced wings of high-aspect-ratio
NASA Technical Reports Server (NTRS)
Smith, P. M.; Deyoung, J.; Lovell, W. A.; Price, J. E.; Washburn, G. F.
1981-01-01
The effect of increased wing aspect ratio of subsonic aircraft on configurations with and without strut bracing. Results indicate that an optimum cantilever configuration, with a wing aspect ratio of approximately 26, has a 19% improvement in cruise range when compared to a baseline concept with a wing aspect ratio of approximately 10. An optimum strut braced configuration, with a wing aspect ratio of approximately 28, has a 31% improvment in cruise range when compared to the same baseline concept. This improvement is mainly due to the estimated reduction in wing weight resulting from use of lifting struts. All configurations assume the same mission payload and fuel. The drag characteristics of the wings are enhanced with the use of laminar flow airfoils. A method for determining the extent of attainable natural laminar flow, and methods for preliminary structural design and for aerodynamic analysis of wings lifting struts are presented.
NASA Technical Reports Server (NTRS)
Hood, Manley J; White, James A
1933-01-01
Some preliminary results of full scale wind tunnel testing to determine the best means of reducing the tail buffeting and wing-fuselage interference of a low-wing monoplane are given. Data indicating the effects of an engine cowling, fillets, auxiliary airfoils of short span, reflexes trailing edge, propeller slipstream, and various combinations of these features are included. The best all-round results were obtained by the use of fillets together with the National Advisory Committee for Aeronautics (NACA) cowling. This combination reduced the tail buffeting oscillations to one-fourth of their original amplitudes, increased the maximum lift 11 percent, decreased the minimum drag 9 percent, and increased the maximum ratio of lift to drag 19 percent.
NASA Technical Reports Server (NTRS)
Margoulis, W
1922-01-01
To sum up, Professor Joukowski's theory of supporting wings renders it possible to calculate the coefficient of lift in terms of the angle of attack, and Prandtl's coefficient of induced drag and the correction of the angle of attack in terms of the disposition and aspect ratio of the wings.
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.
NASA Technical Reports Server (NTRS)
Shovlin, M. D.
1977-01-01
Interior and exterior fuselage noise levels were measured on NASA's C-8A Augmentor Wing Jet-STOL Research Aircraft in order to provide design information for the Quiet Short-Haul Research Aircraft (QSRA), which will use a modified C-8A fuselage. The noise field was mapped by 11 microphones located internally and externally in three areas: mid-fuselage, aft fuselage, and on the flight deck. Noise levels were recorded at four power settings varying from takeoff to flight idle and were plotted in one-third octave band spectra. The overall sound pressure levels of the external noise field were compared to previous tests and found to correlate well with engine primary thrust levels. Fuselage values were 145 + or - 3 dB over the aircraft's normal STOL operating range.
Modern wing flutter analysis by computational fluid dynamics methods
NASA Technical Reports Server (NTRS)
Cunningham, Herbert J.; Batina, John T.; Bennett, Robert M.
1988-01-01
The application and assessment of the recently developed CAP-TSD transonic small-disturbance code for flutter prediction is described. The CAP-TSD code has been developed for aeroelastic analysis of complete aircraft configurations and was previously applied to the calculation of steady and unsteady pressures with favorable results. Generalized aerodynamic forces and flutter characteristics are calculated and compared with linear theory results and with experimental data for a 45 deg sweptback wing. These results are in good agreement with the experimental flutter data which is the first step toward validating CAP-TSD for general transonic aeroelastic applications. The paper presents these results and comparisons along with general remarks regarding modern wing flutter analysis by computational fluid dynamics methods.
Novel Control Effectors for Truss Braced Wing
NASA Technical Reports Server (NTRS)
White, Edward V.; Kapania, Rakesh K.; Joshi, Shiv
2015-01-01
At cruise flight conditions very high aspect ratio/low sweep truss braced wings (TBW) may be subject to design requirements that distinguish them from more highly swept cantilevered wings. High aspect ratio, short chord length and relative thinness of the airfoil sections all contribute to relatively low wing torsional stiffness. This may lead to aeroelastic issues such as aileron reversal and low flutter margins. In order to counteract these issues, high aspect ratio/low sweep wings may need to carry additional high speed control effectors to operate when outboard ailerons are in reversal and/or must carry additional structural weight to enhance torsional stiffness. The novel control effector evaluated in this study is a variable sweep raked wing tip with an aileron control surface. Forward sweep of the tip allows the aileron to align closely with the torsional axis of the wing and operate in a conventional fashion. Aft sweep of the tip creates a large moment arm from the aileron to the wing torsional axis greatly enhancing aileron reversal. The novelty comes from using this enhanced and controllable aileron reversal effect to provide roll control authority by acting as a servo tab and providing roll control through intentional twist of the wing. In this case the reduced torsional stiffness of the wing becomes an advantage to be exploited. The study results show that the novel control effector concept does provide roll control as described, but only for a restricted class of TBW aircraft configurations. For the configuration studied (long range, dual aisle, Mach 0.85 cruise) the novel control effector provides significant benefits including up to 12% reduction in fuel burn.
NASA Astrophysics Data System (ADS)
Kuhn, Lisa M.
2011-07-01
In recent years, much research has been motivated by the idea of biologically-inspired flight. It is a conjecture of the United States Air Force that incorporating characteristics of biological flight into air vehicles will significantly improve the maneuverability and performance of modern aircraft. Although there are studies which involve the aerodynamics, structural dynamics, modeling, and control of flexible wing micro aerial vehicles (MAVs), issues of control and vehicular modeling as a whole are largely unexplored. Modeling with such dynamics lends itself to systems of partial differential equations (PDEs) with nonlinearities, and limited control theory is available for such systems. In this work, a multiple component structure consisting of two Euler-Bernoulli beams connected to a rigid mass is used to model the heave dynamics of an aeroelastic wing MAV, which is acted upon by a nonlinear aerodynamic lift force. We seek to employ tools from distributed parameter modeling and linear control theory in an effort to achieve agile flight potential of flexible, morphable wing MAV airframes. Theoretical analysis of the model is conducted, which includes generating solutions to the eigenvalue problem for the system and determining well-posedness and the attainment of a C 0-semigroup for the linearly approximated model. In order to test the model's ability to track to a desired state and to gain insight into optimal morphing trajectories, two control objectives are employed on the model: target state tracking and morphing trajectory over time.
Update on HCDstruct - A Tool for Hybrid Wing Body Conceptual Design and Structural Optimization
NASA Technical Reports Server (NTRS)
Gern, Frank H.
2015-01-01
HCDstruct is a Matlab® based software tool to rapidly build a finite element model for structural optimization of hybrid wing body (HWB) aircraft at the conceptual design level. The tool uses outputs from a Flight Optimization System (FLOPS) performance analysis together with a conceptual outer mold line of the vehicle, e.g. created by Vehicle Sketch Pad (VSP), to generate a set of MSC Nastran® bulk data files. These files can readily be used to perform a structural optimization and weight estimation using Nastran’s® Solution 200 multidisciplinary optimization solver. Initially developed at NASA Langley Research Center to perform increased fidelity conceptual level HWB centerbody structural analyses, HCDstruct has grown into a complete HWB structural sizing and weight estimation tool, including a fully flexible aeroelastic loads analysis. Recent upgrades to the tool include the expansion to a full wing tip-to-wing tip model for asymmetric analyses like engine out conditions and dynamic overswings, as well as a fully actuated trailing edge, featuring up to 15 independently actuated control surfaces and twin tails. Several example applications of the HCDstruct tool are presented.
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.
Computational Aeroelastic Modeling of Airframes and TurboMachinery: Progress and Challenges
NASA Technical Reports Server (NTRS)
Bartels, R. E.; Sayma, A. I.
2006-01-01
Computational analyses such as computational fluid dynamics and computational structural dynamics have made major advances toward maturity as engineering tools. Computational aeroelasticity is the integration of these disciplines. As computational aeroelasticity matures it too finds an increasing role in the design and analysis of aerospace vehicles. This paper presents a survey of the current state of computational aeroelasticity with a discussion of recent research, success and continuing challenges in its progressive integration into multidisciplinary aerospace design. This paper approaches computational aeroelasticity from the perspective of the two main areas of application: airframe and turbomachinery design. An overview will be presented of the different prediction methods used for each field of application. Differing levels of nonlinear modeling will be discussed with insight into accuracy versus complexity and computational requirements. Subjects will include current advanced methods (linear and nonlinear), nonlinear flow models, use of order reduction techniques and future trends in incorporating structural nonlinearity. Examples in which computational aeroelasticity is currently being integrated into the design of airframes and turbomachinery will be presented.
Identification of XV-15 aeroelastic modes using frequency-domain methods
NASA Technical Reports Server (NTRS)
Acree, Cecil W., Jr.; Tischler, Mark B.
1989-01-01
The XV-15 Tilt-Rotor wing has six major aeroelastic modes that are close in frequency. To precisely excite individual modes during flight test, dual flaperon exciters with automatic frequency-sweep controls were installed. The resulting structural data were analyzed in the frequency domain (Fourier transformed) with cross spectral and transfer function methods. Modal frequencies and damping were determined by performing curve fits to transfer function magnitude and phase data and to cross spectral magnitude data. Results are given for the XV-15 with its original metal rotor blades. Frequency and damping values are also compared with earlier predictions.
Using frequency-domain methods to identify XV-15 aeroelastic modes
NASA Technical Reports Server (NTRS)
Acree, C. W., Jr.; Tischler, Mark B.
1987-01-01
The XV-15 Tilt-Rotor wing has six major aeroelastic modes that are close in frequency. To precisely excite individual modes during flight test, dual flaperon exciters with automatic frequency-sweep controls were installed. The resulting structural data were analyzed in the frequency domain (Fourier transformed) with cross spectral and transfer function methods. Modal frequencies and damping were determined by performing curve fits to transfer function magnitude and phase data and to cross spectral magnitude data. Results are given for the XV-15 with its original metal rotor blades. Frequency and damping values are also compared with earlier predictions.
NASA Technical Reports Server (NTRS)
Doggett, R. V., Jr.; Abel, I.; Ruhlin, C. L.
1976-01-01
A status report and review of wind tunnel model experimental techniques that have been developed to study and validate the use of active control technology for the minimization of aeroelastic response are presented. Modeling techniques, test procedures, and data analysis methods used in three model studies are described. The studies include flutter mode suppression on a delta-wing model, flutter mode suppression and ride quality control on a 1/30-size model of the B-52 CCV airplane, and an active lift distribution control system on a 1/22 size C-5A model.
NASA Technical Reports Server (NTRS)
Friedmann, P. P.; Venkatesan, C.; Yuan, K.
1992-01-01
This paper describes the development of a new structural optimization capability aimed at the aeroelastic tailoring of composite rotor blades with straight and swept tips. The primary objective is to reduce vibration levels in forward flight without diminishing the aeroelastic stability margins of the blade. In the course of this research activity a number of complicated tasks have been addressed: (1) development of a new, aeroelastic stability and response analysis; (2) formulation of a new comprehensive sensitive analysis, which facilitates the generation of the appropriate approximations for the objective and the constraints; (3) physical understanding of the new model and, in particular, determination of its potential for aeroelastic tailoring, and (4) combination of the newly developed analysis capability, the sensitivity derivatives and the optimizer into a comprehensive optimization capability. The first three tasks have been completed and the fourth task is in progress.
NASA Technical Reports Server (NTRS)
Howard, Anna K. T.
1999-01-01
The tiltrotor offers the best mix of hovering and cruise flight of any of the current V/STOL configurations. One possible improvement on the tiltrotors of today designs would be using a soft-inplane hingeless hub. The advantages to a soft-inplane hingeless hub range from reduced weight and maintenance to reduced vibration and loads. However, soft-inplane rotor systems are inherently in danger of the aeromechanical instabilities of ground and air resonance. Furthermore tiltrotors can be subject to whirl flutter. At least in part because of the potential for air and ground resonance in a soft-inplane rotor, the Bell XV-15, the Bell-Boeing V-22 Osprey, and the new Bell Augusta 609 have stiff-inplane, gimballed rotors which do not experience these instabilities. In order to design soft-inplane V/STOL aircraft that do not experience ground or air resonance, it is important to be able to predict these instabilities accurately. Much of the research studying the stability of tiltrotors has been focused on the understanding and prediction of whirl flutter. As this instability is increasingly well understood, air and ground resonance for a tiltrotor need to be investigated. Once we understand the problems of air and ground resonance in a tiltrotor, we must look for solutions to these instabilities. Other researchers have found composite or kinematic couplings in the blades of a helicopter helpful for ground and air resonance stability. Tiltrotor research has shown composite couplings in the wing to be helpful for whirl flutter. Therefore, this project will undertake to model ground and air resonance of a soft-inplane hingeless tiltrotor to understand the mechanisms involved and to evaluate whether aeroelastic couplings in the wing or kinematic couplings in the blades would aid in stabilizing these instabilities in a tiltrotor.
Design, realization and structural testing of a compliant adaptable wing
NASA Astrophysics Data System (ADS)
Molinari, G.; Quack, M.; Arrieta, A. F.; Morari, M.; Ermanni, P.
2015-10-01
This paper presents the design, optimization, realization and testing of a novel wing morphing concept, based on distributed compliance structures, and actuated by piezoelectric elements. The adaptive wing features ribs with a selectively compliant inner structure, numerically optimized to achieve aerodynamically efficient shape changes while simultaneously withstanding aeroelastic loads. The static and dynamic aeroelastic behavior of the wing, and the effect of activating the actuators, is assessed by means of coupled 3D aerodynamic and structural simulations. To demonstrate the capabilities of the proposed morphing concept and optimization procedure, the wings of a model airplane are designed and manufactured according to the presented approach. The goal is to replace conventional ailerons, thus to achieve controllability in roll purely by morphing. The mechanical properties of the manufactured components are characterized experimentally, and used to create a refined and correlated finite element model. The overall stiffness, strength, and actuation capabilities are experimentally tested and successfully compared with the numerical prediction. To counteract the nonlinear hysteretic behavior of the piezoelectric actuators, a closed-loop controller is implemented, and its capability of accurately achieving the desired shape adaptation is evaluated experimentally. Using the correlated finite element model, the aeroelastic behavior of the manufactured wing is simulated, showing that the morphing concept can provide sufficient roll authority to allow controllability of the flight. The additional degrees of freedom offered by morphing can be also used to vary the plane lift coefficient, similarly to conventional flaps. The efficiency improvements offered by this technique are evaluated numerically, and compared to the performance of a rigid wing.
NASA Technical Reports Server (NTRS)
Yeager, William T., Jr.; Kvaternik, Raymond G.
2001-01-01
A historical account of the contributions of the Aeroelasticity Branch (AB) and the Langley Transonic Dynamics Tunnel (TDT) to rotorcraft technology and development since the tunnel's inception in 1960 is presented. The paper begins with a summary of the major characteristics of the TDT and a description of the unique capability offered by the TDT for testing aeroelastic models by virtue of its heavy gas test medium. This is followed by some remarks on the role played by scale models in the design and development of rotorcraft vehicles and a review of the basic scaling relationships important for designing and building dynamic aeroelastic models of rotorcraft vehicles for testing in the TDT. Chronological accounts of helicopter and tiltrotor research conducted in AB/TDT are then described in separate sections. Both experimental and analytical studies are reported and include a description of the various physical and mathematical models employed, the specific objectives of the investigations, and illustrative experimental and analytical results.
Integrated aerodynamic/structural design of a sailplane wing
NASA Technical Reports Server (NTRS)
Grossman, B.; Gurdal, Z.; Haftka, R. T.; Strauch, G. J.; Eppard, W. M.
1986-01-01
Using lifting-line theory and beam analysis, the geometry (planiform and twist) and composite material structural sizes (skin thickness, spar cap, and web thickness) were designed for a sailplane wing, subject to both structural and aerodynamic constraints. For all elements, the integrated design (simultaneously designing the aerodynamics and the structure) was superior in terms of performance and weight to the sequential design (where the aerodynamic geometry is designed to maximize the performance, following which a structural/aeroelastic design minimizes the weight). Integrated designs produced less rigid, higher aspect ratio wings with favorable aerodynamic/structural interactions.
NASA Technical Reports Server (NTRS)
Stanford, Bret K.; Jutte, Christine V.
2016-01-01
A series of aeroelastic optimization problems are solved on a high aspect ratio wingbox of the Common Research Model, in an effort to minimize structural mass under coupled stress, buckling, and flutter constraints. Two technologies are of particular interest: tow steered composite laminate skins and curvilinear stiffeners. Both methods are found to afford feasible reductions in mass over their non-curvilinear structural counterparts, through both distinct and shared mechanisms for passively controlling aeroelastic performance. Some degree of diminishing returns are seen when curvilinear stiffeners and curvilinear fiber tow paths are used simultaneously.
Hybrid state vector methods for structural dynamic and aeroelastic boundary value problems
NASA Technical Reports Server (NTRS)
Lehman, L. L.
1982-01-01
A computational technique is developed that is suitable for performing preliminary design aeroelastic and structural dynamic analyses of large aspect ratio lifting surfaces. The method proves to be quite general and can be adapted to solving various two point boundary value problems. The solution method, which is applicable to both fixed and rotating wing configurations, is based upon a formulation of the structural equilibrium equations in terms of a hybrid state vector containing generalized force and displacement variables. A mixed variational formulation is presented that conveniently yields a useful form for these state vector differential equations. Solutions to these equations are obtained by employing an integrating matrix method. The application of an integrating matrix provides a discretization of the differential equations that only requires solutions of standard linear matrix systems. It is demonstrated that matrix partitioning can be used to reduce the order of the required solutions. Results are presented for several example problems in structural dynamics and aeroelasticity to verify the technique and to demonstrate its use. These problems examine various types of loading and boundary conditions and include aeroelastic analyses of lifting surfaces constructed from anisotropic composite materials.
NASA Technical Reports Server (NTRS)
Ferris, J. C.
1973-01-01
The Langley 8-foot transonic pressure tunnel to determine the wing chordwise pressure distribution for a 0.09-scale model of a research airplane incorporating a 17-percent-thick supercritical wing. Airfoil profile drag was determined from wake pressure measurements at the 42-percent-semispan wing station. The investigation was conducted at Mach numbers from 0.30 to 0.80 over an angle-of-attack range sufficient to include buffet onset. The Reynolds number based on the mean geometric chord varied from 2 x 10 to the 6th power at Mach number 0.30 to 3.33 x 10 to the 6th power at Mach number 0.65 and was maintained at a constant value of 3.86 x 10 to the 6th power at Mach numbers from 0.70 to 0.80. Pressure coefficients for four wing semispan stations and wing-section normal-force and pitching-moment coefficients for two semispan stations are presented in tabular form over the Mach number range from 0.30 to 0.80. Plotted chordwise pressure distributions and wake profiles are given for a selected range of section normal-force coefficients over the same Mach number range.
Gooding, Benjamin W. T.; Geoghegan, John M.; Wallace, W. Angus; Manning, Paul A.
2013-01-01
This review explores the causes of scapula winging, with overview of the relevant anatomy, proposed aetiology and treatment. Particular focus is given to lesions of the long thoracic nerve, which is reported to be the most common aetiological factor.
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 Technical Reports Server (NTRS)
Fox, C. H., Jr.
1978-01-01
A general research fighter model was tested in the Langley 7 by 10 foot high speed tunnel at a Mach number of 0.3. Strakes with exposed semi-spans of 10 percent, 20 percent, and 30 percent of the wing reference semi-span were tested in combination with wings having leading edge sweep angles of 30, 44, and 60 degrees. The angle of attack range was from -4 degrees to approximately 48 degrees at sideslip angles of 0, -5, and 5 degrees. The data are presented without analysis in order to expedite publication.
NASA Technical Reports Server (NTRS)
Boyden, R. P.
1974-01-01
The aerodynamic damping in pitch, yaw, and roll and the oscillatory stability in pitch and yaw of a supercritical-wing research airplane model were determined for Mach numbers of 0.25 to 1.20 by using the small-amplitude forced-oscillation technique. The angle-of-attack range was from -2 deg to 20 deg. The effects of the underwing leading-edge vortex generators and the contributions of the wing, vertical tail, and horizontal tail to the appropriate damping and stability were measured.
Mertz, Leslie
2012-01-01
When the Defense Advanced Research Projects Agency (DARPA) asks research questions, it goes big. This is, after all, the same agency that put together teams of scientists and engineers to find a way to connect the worlds computers and, in doing so, developed the precursor to the Internet. DARPA, the experimental research wing of the U.S. Department of Defense, funds the types of research queries that scientists and engineers dream of tackling. Unlike a traditional granting agency that conservatively metes out its funding and only to projects with a good chance of success, DARPA puts its money on massive, multi-institutional projects that have no guarantees, but have enormous potential. In the 1990s, DARPA began its biological and medical science research to improve the safety, health, and well being of military personnel, according to DARPA program manager and Army Colonel Geoffrey Ling, Ph.D., M.D. More recently, DARPA has entered the realm of neuroscience and neurotechnology. Its focus with these projects is on its prime customer, the U.S. Department of Defense, but Ling acknowledged that technologies developed in its programs "certainly have potential to cascade into civilian uses." PMID:22344946
Effects of leading-edge tubercles on wing flutter speeds.
Ng, B F; New, T H; Palacios, R
2016-06-01
The dynamic aeroelastic effects on wings modified with bio-inspired leading-edge (LE) tubercles are examined in this study. We adopt a state-space aeroelastic model via the coupling of unsteady vortex-lattice method and a composite beam to evaluate stability margins as a result of LE tubercles on a generic wing. The unsteady aerodynamics and spanwise mass variations due to LE tubercles have counteracting effects on stability margins with the former having dominant influence. When coupled, flutter speed is observed to be 5% higher, and this is accompanied by close to 6% decrease in reduced frequencies as an indication of lower structural stiffness requirements for wings with LE tubercles. Both tubercle amplitude and wavelength have similar influences over the change in flutter speeds, and such modifications to the LE would have minimal effect on stability margins when concentrated inboard of the wing. Lastly, when used in sweptback wings, LE tubercles are observed to have smaller impacts on stability margins as the sweep angle is increased. PMID:27070824
A Numerical Model of Unsteady, Subsonic Aeroelastic Behavior. Ph.D Thesis
NASA Technical Reports Server (NTRS)
Strganac, Thomas W.
1987-01-01
A method for predicting unsteady, subsonic aeroelastic responses was developed. The technique accounts for aerodynamic nonlinearities associated with angles of attack, vortex-dominated flow, static deformations, and unsteady behavior. The fluid and the wing together are treated as a single dynamical system, and the equations of motion for the structure and flow field are integrated simultaneously and interactively in the time domain. The method employs an iterative scheme based on a predictor-corrector technique. The aerodynamic loads are computed by the general unsteady vortex-lattice method and are determined simultaneously with the motion of the wing. Because the unsteady vortex-lattice method predicts the wake as part of the solution, the history of the motion is taken into account; hysteresis is predicted. Two models are used to demonstrate the technique: a rigid wing on an elastic support experiencing plunge and pitch about the elastic axis, and an elastic wing rigidly supported at the root chord experiencing spanwise bending and twisting. The method can be readily extended to account for structural nonlinearities and/or substitute aerodynamic load models. The time domain solution coupled with the unsteady vortex-lattice method provides the capability of graphically depicting wing and wake motion.
Role of HPC in Advancing Computational Aeroelasticity
NASA Technical Reports Server (NTRS)
Guruswamy, Guru P.
2004-01-01
On behalf of the High Performance Computing and Modernization Program (HPCMP) and NASA Advanced Supercomputing Division (NAS) a study is conducted to assess the role of supercomputers on computational aeroelasticity of aerospace vehicles. The study is mostly based on the responses to a web based questionnaire that was designed to capture the nuances of high performance computational aeroelasticity, particularly on parallel computers. A procedure is presented to assign a fidelity-complexity index to each application. Case studies based on major applications using HPCMP resources are presented.
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.
Method of performing computational aeroelastic analyses
NASA Technical Reports Server (NTRS)
Silva, Walter A. (Inventor)
2011-01-01
Computational aeroelastic analyses typically use a mathematical model for the structural modes of a flexible structure and a nonlinear aerodynamic model that can generate a plurality of unsteady aerodynamic responses based on the structural modes for conditions defining an aerodynamic condition of the flexible structure. In the present invention, a linear state-space model is generated using a single execution of the nonlinear aerodynamic model for all of the structural modes where a family of orthogonal functions is used as the inputs. Then, static and dynamic aeroelastic solutions are generated using computational interaction between the mathematical model and the linear state-space model for a plurality of periodic points in time.
NASA Technical Reports Server (NTRS)
Skillen, Michael D.; Crossley, William A.
2008-01-01
This report documents a series of investigations to develop an approach for structural sizing of various morphing wing concepts. For the purposes of this report, a morphing wing is one whose planform can make significant shape changes in flight - increasing wing area by 50% or more from the lowest possible area, changing sweep 30 or more, and / or increasing aspect ratio by as much as 200% from the lowest possible value. These significant changes in geometry mean that the underlying load-bearing structure changes geometry. While most finite element analysis packages provide some sort of structural optimization capability, these codes are not amenable to making significant changes in the stiffness matrix to reflect the large morphing wing planform changes. The investigations presented here use a finite element code capable of aeroelastic analysis in three different optimization approaches -a "simultaneous analysis" approach, a "sequential" approach, and an "aggregate" approach.
NASA Technical Reports Server (NTRS)
Florance, Jennifer P.; Burner, Alpheus W.; Fleming, Gary A.; Martin, Christopher A.
2003-01-01
An overview of the contributions of the NASA Langley Research Center (LaRC) to the DARPA/AFRL/NASA/ Northrop Grumman Corporation (NGC) Smart Wing program is presented. The overall objective of the Smart Wing program was to develop smart** technologies and demonstrate near-flight-scale actuation systems to improve the aerodynamic performance of military aircraft. NASA LaRC s roles were to provide technical guidance, wind-tunnel testing time and support, and Computational Fluid Dynamics (CFD) analyses. The program was divided into two phases, with each phase having two wind-tunnel entries in the Langley Transonic Dynamics Tunnel (TDT). This paper focuses on the fourth and final wind-tunnel test: Phase 2, Test 2. During this test, a model based on the NGC Unmanned Combat Air Vehicle (UCAV) concept was tested at Mach numbers up to 0.8 and dynamic pressures up to 150 psf to determine the aerodynamic performance benefits that could be achieved using hingeless, smoothly-contoured control surfaces actuated with smart materials technologies. The UCAV-based model was a 30% geometric scale, full-span, sting-mounted model with the smart control surfaces on the starboard wing and conventional, hinged control surfaces on the port wing. Two LaRC-developed instrumentation systems were used during the test to externally measure the shapes of the smart control surface and quantify the effects of aerodynamic loading on the deflections: Videogrammetric Model Deformation (VMD) and Projection Moire Interferometry (PMI). VMD is an optical technique that uses single-camera photogrammetric tracking of discrete targets to determine deflections at specific points. PMI provides spatially continuous measurements of model deformation by computationally analyzing images of a grid projected onto the model surface. Both the VMD and PMI measurements served well to validate the use of on-board (internal) rotary potentiometers to measure the smart control surface deflection angles. Prior to the final
Anisotropy and non-homogeneity of an Allomyrina Dichotoma beetle hind wing membrane.
Ha, N S; Jin, T L; Goo, N S; Park, H C
2011-12-01
Biomimetics is one of the most important paradigms as researchers seek to invent better engineering designs over human history. However, the observation of insect flight is a relatively recent work. Several researchers have tried to address the aerodynamic performance of flapping creatures and other natural properties of insects, although there are still many unsolved questions. In this study, we try to answer the questions related to the mechanical properties of a beetle's hind wing, which consists of a stiff vein structure and a flexible membrane. The membrane of a beetle's hind wing is small and flexible to the point that conventional methods cannot adequately quantify the material properties. The digital image correlation method, a non-contact displacement measurement method, is used along with a specially designed mini-tensile testing system. To reduce the end effects, we developed an experimental method that can deal with specimens with as high an aspect ratio as possible. Young's modulus varies over the area in the wing and ranges from 2.97 to 4.5 GPa in the chordwise direction and from 1.63 to 2.24 GPa in the spanwise direction. Furthermore, Poisson's ratio in the chordwise direction is 0.63-0.73 and approximately twice as large as that in the spanwise direction (0.33-0.39). From these results, we can conclude that the membrane of a beetle's hind wing is an anisotropic and non-homogeneous material. Our results will provide a better understanding of the flapping mechanism through the formulation of a fluid-structure interaction analysis or aero-elasticity analysis and meritorious data for biomaterial properties database as well as a creative design concept for a micro aerial flapper that mimics an insect. PMID:21992989
Aeroelastic Stability of a Four-Bladed Semi-Articulated Soft-Inplane Tiltrotor Model
NASA Technical Reports Server (NTRS)
Nixon, Mark W.; Langston, Chester W.; Singleton, Jeffrey D.; Piatak, David J.; Kvaternik, Raymond G.; Corso, Lawrence M.; Brown, Ross K.
2003-01-01
A new four-bladed, semi-articulated, soft-inplane rotor system, designed as a candidate for future heavy-lift rotorcraft, was tested at model scale on the Wing and Rotor Aeroelastic Testing System (WRATS), a 1/5-size aeroelastic wind-tunnel model based on the V-22. The experimental investigation included a hover test with the model in helicopter mode subject to ground resonance conditions, and a forward flight test with the model in airplane mode subject to whirl-flutter conditions. An active control system designed to augment system damping was also tested as part of this investigation. Results of this study indicate that the new four-bladed, soft-inplane rotor system in hover has adequate damping characteristics and is stable throughout its rotor-speed envelope. However, in airplane mode it produces very low damping in the key wing beam-bending mode, and has a low whirl-flutter stability boundary with respect to airspeed. The active control system was successful in augmenting the damping of the fundamental system modes, and was found to be robust with respect to changes in rotor speed and airspeed. Finally, conversion-mode dynamic loads were measured on the rotor and these were found to be signi.cantly lower for the new soft-inplane hub than for the previous baseline stiff - inplane hub.
Fuzzy Model-based Pitch Stabilization and Wing Vibration Suppression of Flexible Wing Aircraft.
NASA Technical Reports Server (NTRS)
Ayoubi, Mohammad A.; Swei, Sean Shan-Min; Nguyen, Nhan T.
2014-01-01
This paper presents a fuzzy nonlinear controller to regulate the longitudinal dynamics of an aircraft and suppress the bending and torsional vibrations of its flexible wings. The fuzzy controller utilizes full-state feedback with input constraint. First, the Takagi-Sugeno fuzzy linear model is developed which approximates the coupled aeroelastic aircraft model. Then, based on the fuzzy linear model, a fuzzy controller is developed to utilize a full-state feedback and stabilize the system while it satisfies the control input constraint. Linear matrix inequality (LMI) techniques are employed to solve the fuzzy control problem. Finally, the performance of the proposed controller is demonstrated on the NASA Generic Transport Model (GTM).
Simulation of tail buffet using delta wing-vertical tail configuration
NASA Technical Reports Server (NTRS)
Kandil, Osama A.; Kandil, Hamdy A.; Massey, Steven J.
1993-01-01
Computational simulation of the vertical tail buffet problem is accomplished using a delta wing-vertical tail configuration. Flow conditions are selected such that the wing primary-vortex cores experience vortex breakdown and the resulting flow interacts with the vertical tail. This multidisciplinary problem is solved successively using three sets of equations for the fluid flow, aeroelastic deflections and grid displacements. For the fluid dynamics part, the unsteady, compressible, full Navier-Stokes equations are solved accurately in time using an implicit, upwind, flux-difference splitting, finite-volume scheme. For the aeroelastic part, the aeroelastic equation for bending vibrations is solved accurately in time using the Galerkin method and the four-stage Runge-Kutta scheme. The grid for the fluid dynamics computations is updated every few time steps using a third set of interpolation equations. The computational application includes a delta wing of aspect ratio 1 and a rectangular vertical tail of aspect ratio 2, which is placed at 0.5 root-chord length downstream of the wing trailing edge. The wing angle of attack is 35 deg and the flow Mach number and Reynolds number are 0.4 and 10,000, respectively.
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.
Aeroelastic Stability and Response of Rotating Structures
NASA Technical Reports Server (NTRS)
Keith, Theo G., Jr.; Reddy, T. S. R.
1998-01-01
A summary of the work performed from 1996 to 1997 is presented. More details can be found in the cited references. This grant led to the development of aeroelastic analyses methods for predicting flutter and forced response in fans, compressors, and turbines using computational
Aeroelastic Stability & Response of Rotating Structures
NASA Technical Reports Server (NTRS)
Keith, Theo G., Jr.; Reddy, T. S. R.
2001-01-01
A summary of the work performed under NASA grant NCC3-605 is presented. More details can be found in the cited references. This grant led to the development of relatively faster aeroelastic analyses methods for predicting flutter and forced response in fans, compressors, and turbines using computational fluid dynamic (CFD) methods.
Wind-tunnel roll-damping measurements of a winged space shuttle configuration in launch attitude
NASA Technical Reports Server (NTRS)
Hess, R. W.; Davenport, E. E.
1973-01-01
Ground-wind load studies were conducted on three model configurations to assess the importance of aeroelastic instabilities of erected space shuttle vehicles. Roll damping was measured on a fuselage-alone model, which had a D cross section, and a fuselage and tail surfaces in combination with either a clipped-delta wing or a low-sweep tapered wing as the primary lifting surface. The largest negative roll-damping coefficients were measured with the fuselage-alone configuration and were a function of wind azimuth. At the wind azimuths at which the wing-fuselage configuration was unstable, the negative roll-damping coefficients were a function of reduced frequency.
Design, testing, and damage tolerance study of bonded stiffened composite wing cover panels
NASA Technical Reports Server (NTRS)
Madan, Ram C.; Sutton, Jason O.
1988-01-01
Results are presented from the application of damage tolerance criteria for composite panels to multistringer composite wing cover panels developed under NASA's Composite Transport Wing Technology Development contract. This conceptual wing design integrated aeroelastic stiffness constraints with an enhanced damage tolerance material system, in order to yield optimized producibility and structural performance. Damage tolerance was demonstrated in a test program using full-sized cover panel subcomponents; panel skins were impacted at midbay between stiffeners, directly over a stiffener, and over the stiffener flange edge. None of the impacts produced visible damage. NASTRAN analyses were performed to simulate NDI-detected invisible damage.
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.
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.
Self-induced wing rock of slender delta wings
NASA Technical Reports Server (NTRS)
Nguyen, L. T.; Yip, L. P.; Chambers, J. R.
1981-01-01
As part of a research program aimed at exploring basic mechanisms that cause wing rock in combat aircraft, an investigation was conducted to study the aerodynamic factors which cause the low-speed wing rock exhibited by slender delta wings. A flat-plate delta wing with 80 deg leading-edge sweep was subjected to conventional static-force tests and dynamic wind-tunnel experiments which included forced-oscillation, rotary, and free-to-roll tests. In addition, visualization of the flow phenomena involved was obtained by observing tuft patterns and using a helium-bubble technique. This paper summarizes the results of this study. Fundamental information is presented on the aerodynamic mechanisms that cause the wing rock and the problem of mathematically modeling the aerodynamics and motions is discussed.
NASA Technical Reports Server (NTRS)
Sevart, F. D.; Patel, S. M.; Wattman, W. J.
1972-01-01
Testing and evaluation of stability augmentation systems for aircraft flight control were conducted. The flutter suppression system analysis of a scale supersonic transport wing model is described. Mechanization of the flutter suppression system is reported. The ride control synthesis for the B-52 aeroelastic model is discussed. Model analyses were conducted using equations of motion generated from generalized mass and stiffness data.
NASA Technical Reports Server (NTRS)
Mendoza, Jeff M.; Brooks, Thomas F.; Humphreys, William M.
2002-01-01
Aeroacoustic evaluations of high-lift devices have been carried out in the Quiet Flow Facility of the NASA Langley Research Center. The present paper deals with detailed flow and acoustic measurements that have been made to understand, and to possibly predict and reduce, the noise from a wing leading edge slat configuration. The acoustic database is obtained by a moveable Small Aperture Directional Array (SADA) of microphones designed to electronically steer to different portions of models under study. The slat is shown to be a uniform distributed noise source. The data was processed such that spectra and directivity were determined with respect to a one-foot span of slat. The spectra are normalized in various fashions to demonstrate slat noise character. In order to equate portions of the spectra to different slat noise components, trailing edge noise predictions using measured slat boundary layer parameters as inputs are compared to the measured slat noise spectra.
NASA Technical Reports Server (NTRS)
Eppel, J. C.; Shovlin, M. D.; Jaynes, D. N.; Englar, R. J.; Nichols, J. H., Jr.
1982-01-01
Full scale static investigations were conducted on the Quiet Short Haul Research Aircraft (QSRA) to determine the thrust deflecting capabilities of the circulation control wing/upper surface blowing (CCW/USB) concept. This scheme, which combines favorable characteristics of both the A-6/CCW and QSRA, employs the flow entrainment properties of CCW to pneumatically deflect engine thrust in lieu of the mechanical USB flap system. Results show that the no moving parts blown system produced static thrust deflections in the range of 40 deg to 97 deg (depending on thrust level) with a CCW pressure of 208,900 Pa (30.3 psig). In addition, the ability to vary horizontal forces from thrust to drag while maintaining a constant vertical (or lift) value was demonstrated by varying the blowing pressure. The versatility of the CCW/USB system, if applied to a STOL aircraft, was confirmed, where rapid conversion from a high drag approach mode to a thrust recovering waveoff or takeoff configuration could be achieved by nearly instantaneous blowing pressure variation.
Applications of a transonic wing design method
NASA Technical Reports Server (NTRS)
Campbell, Richard L.; Smith, Leigh A.
1989-01-01
A method for designing wings and airfoils at transonic speeds using a predictor/corrector approach was developed. The procedure iterates between an aerodynamic code, which predicts the flow about a given geometry, and the design module, which compares the calculated and target pressure distributions and modifies the geometry using an algorithm that relates differences in pressure to a change in surface curvature. The modular nature of the design method makes it relatively simple to couple it to any analysis method. The iterative approach allows the design process and aerodynamic analysis to converge in parallel, significantly reducing the time required to reach a final design. Viscous and static aeroelastic effects can also be accounted for during the design or as a post-design correction. Results from several pilot design codes indicated that the method accurately reproduced pressure distributions as well as the coordinates of a given airfoil or wing by modifying an initial contour. The codes were applied to supercritical as well as conventional airfoils, forward- and aft-swept transport wings, and moderate-to-highly swept fighter wings. The design method was found to be robust and efficient, even for cases having fairly strong shocks.
Fiber-optically sensorized composite wing
NASA Astrophysics Data System (ADS)
Costa, Joannes M.; Black, Richard J.; Moslehi, Behzad; Oblea, Levy; Patel, Rona; Sotoudeh, Vahid; Abouzeida, Essam; Quinones, Vladimir; Gowayed, Yasser; Soobramaney, Paul; Flowers, George
2014-04-01
Electromagnetic interference (EMI) immune and light-weight, fiber-optic sensor based Structural Health Monitoring (SHM) will find increasing application in aerospace structures ranging from aircraft wings to jet engine vanes. Intelligent Fiber Optic Systems Corporation (IFOS) has been developing multi-functional fiber Bragg grating (FBG) sensor systems including parallel processing FBG interrogators combined with advanced signal processing for SHM, structural state sensing and load monitoring applications. This paper reports work with Auburn University on embedding and testing FBG sensor arrays in a quarter scale model of a T38 composite wing. The wing was designed and manufactured using fabric reinforced polymer matrix composites. FBG sensors were embedded under the top layer of the composite. Their positions were chosen based on strain maps determined by finite element analysis. Static and dynamic testing confirmed expected response from the FBGs. The demonstrated technology has the potential to be further developed into an autonomous onboard system to perform load monitoring, SHM and Non-Destructive Evaluation (NDE) of composite aerospace structures (wings and rotorcraft blades). This platform technology could also be applied to flight testing of morphing and aero-elastic control surfaces.
Soap film flow visualization investigations of oscillating wing energy harvesters
NASA Astrophysics Data System (ADS)
Kirschmeier, Benjamin; Bryant, Matthew
2015-03-01
With increasing population and proliferation of wireless electronics, significant research attention has turned to harvesting energy from ambient sources such as wind and water flows at scales ranging from micro-watt to mega-watt levels. One technique that has recently attracted attention is the application of bio-inspired flapping wings for energy harvesting. This type of system uses a heaving and pitching airfoil to extract flow energy and generate electricity. Such a device can be realized using passive devices excited by aeroelastic flutter phenomena, kinematic mechanisms driven by mechanical linkages, or semi-active devices that are actively controlled in one degree of freedom and passively driven in another. For these types of systems, numerical simulations have showed strong dependence on efficiency and vortex interaction. In this paper we propose a new apparatus for reproducing arbitrary pitch-heave waveforms to perform flow visualization experiments in a soap film tunnel. The vertically falling, gravity driven soap film tunnel is used to replicate flows with a chord Reynolds number on the order of 4x104. The soap film tunnel is used to investigate leading edge vortex (LEV) and trailing edge vortex (TEV) interactions for sinusoidal and non-sinusoidal waveforms. From a qualitative analysis of the fluid structure interaction, we have been able to demonstrate that the LEVs for non-sinusoidal motion convect faster over the airfoil compared with sinusoidal motion. Signifying that optimal flapping frequency is dependent on the motion profile.
NASA Astrophysics Data System (ADS)
Goltsch, Mandy
Design denotes the transformation of an identified need to its physical embodiment in a traditionally iterative approach of trial and error. Conceptual design plays a prominent role but an almost infinite number of possible solutions at the outset of design necessitates fast evaluations. The corresponding practice of empirical equations and low fidelity analyses becomes obsolete in the light of novel concepts. Ever increasing system complexity and resource scarcity mandate new approaches to adequately capture system characteristics. Contemporary concerns in atmospheric science and homeland security created an operational need for unconventional configurations. Unmanned long endurance flight at high altitudes offers a unique showcase for the exploration of new design spaces and the incidental deficit of conceptual modeling and simulation capabilities. Structural and aerodynamic performance requirements necessitate light weight materials and high aspect ratio wings resulting in distinct structural and aeroelastic response characteristics that stand in close correlation with natural vibration modes. The present research effort evolves around the development of an efficient and accurate optimization algorithm for high aspect ratio wings subject to natural frequency constraints. Foundational corner stones are beam dimensional reduction and modal perturbation redesign. Local and global analyses inherent to the former suggest corresponding levels of local and global optimization. The present approach departs from this suggestion. It introduces local level surrogate models to capacitate a methodology that consists of multi level analyses feeding into a single level optimization. The innovative heart of the new algorithm originates in small perturbation theory. A sequence of small perturbation solutions allows the optimizer to make incremental movements within the design space. It enables a directed search that is free of costly gradients. System matrices are decomposed
NASA Astrophysics Data System (ADS)
Kim, Dae-Kwan; Lee, Jun-Seong; Han, Jae-Hung
2009-07-01
The sweep-back effect of a flexible flapping wing is investigated through fluid-structure interaction analysis. The aeroelastic analysis is carried out by using an efficient fluid-structure interaction analysis tool, which is based on the modified strip theory and the flexible multibody dynamics. To investigate the sweep-back effect, the aeroelastic analysis is performed on various sweep-back wing models defined by sweep-chord ratio and sweep-span ratio, and then the sweep-back effect on the aerodynamic performance is discussed. The aeroelastic results of the sweep-back wing analysis clearly confirm that the sweep-back angle can help a flexible flapping wing to generate greater twisting motion, resulting in the aerodynamic improvement of thrust and input power for all flapping-axis angle regimes. The propulsive efficiency can also be increased by the sweep-back effect. The sweep angle of a flapping wing should be considered as an important design feature for artificial flexible flapping wings.
X-29A forward-swept-wing flight research program status
NASA Technical Reports Server (NTRS)
Trippensee, Gary A.; Lux, David P.
1988-01-01
The X-29A aircraft is a fascinating combination of integrated technologies incorporated into a unique research aircraft. The X-29A program is a multiple agency program with management and other responsibilities divided among NASA, DARPA, the U.S. Air Force, and the Grumman Corporation. An overview of the recently completed X-29A flight research program, objectives achieved, and a discussion of its future is presented. Also discussed are the flight test approach expanding the envelope, typical flight maneuvers performed, X-29A program accomplishments, lessons learned for the Number One aircraft, and future plans with the Number Two aircraft. A schedule for both aircraft is presented. A description of the unique technologies incorporated into the X-29A aircraft is given, along with descriptions of the onboard instrumentation system. The X-29A aircraft research program has proven highly successful. Using high fly rates from a very reliable experimental aircraft, the program has consistently met or exceeded its design and research goals.
X-29A forward-swept-wing flight research program status
NASA Technical Reports Server (NTRS)
Trippensee, Gary A.; Lux, David P.
1987-01-01
The X-29A aircraft is a fascinating combination of integrated technologies incorporated into a unique research aircraft. The X-29A program is multiple agency program with management and other responsibilities divided among NASA, DARPA, the U.S. Air Force, and the Grumman Corporation. An overview of the recently completed X-29A flight research program, objectives achieved, and a discussion of its future is presented. Also discussed are the flight test approach expanding the envelope, typical flight maneuvers performed, X-29A program accomplishments, lessons learned for the Number One aircraft, and future plans with the Number Two aircraft. A schedule for both aircraft is presented. A description of the unique technologies incorporated into the X-29A aircraft is given, along with descriptions of the onboard instrumentation system. The X-29A aircraft research program has proven highly successful. Using high fly rates from a very reliable experimental aircraft, the program has consistently met or exceeded its design and research goals.
Control of a swept wing tailless aircraft through wing morphing
NASA Astrophysics Data System (ADS)
Guiler, Richard W.
Inspired by flight in nature, work done by Lippisch, the Hortens, and Northrop offered insight to achieving the efficiency of bird flight with swept-wing tailless aircraft. Tailless designs must incorporate aerodynamic compromises for control, which have inhibited potential advantages. A morphing mechanism, capable of changing the twist of wing and that can also provide pitch, roll and yaw control for a tailless swept wing aircraft is the first step to a series of morphing techniques, which will lead to more fluid, bird-like flight. This research focuses on investigating the design of a morphing wing to improve the flight characteristics of swept wing Horten type tailless aircraft. Free flight demonstrators, wind tunnel flow visualization, wind-tunnel force and moment data along with CFD studies have been used to evaluate the stability, control and efficiency of a morphing swept wing tailless aircraft. A wing morphing mechanism for the control of a swept wing tailless aircraft has been developed. This new control technique was experimentally and numerically compared to an existing elevon equipped tailless aircraft and has shown the potential for significant improvement in efficiency. The feasibility of this mechanism was also validated through flight testing of a flight weight version. In the process of comparing the Horten type elevon equipped aircraft and the morphing model, formal wind tunnel verification of wingtip induced thrust, found in Horten (Bell Shaped Lift distribution) type swept wing tailless aircraft was documented. A more complete physical understanding of the highly complex flow generated in the control region of the morphing tailless aircraft has been developed. CFD models indicate the possibility of the presence of a Leading Edge Vortex (LEV) on the control section morphing wing when the tip is twisted between +3.5 degrees and +7 degrees. The presence of this LEV causes a reduction of drag while lift is increased. Similar LEVs have been
Experimental unsteady pressures at flutter on the Supercritical Wing Benchmark Model
NASA Technical Reports Server (NTRS)
Dansberry, Bryan E.; Durham, Michael H.; Bennett, Robert M.; Rivera, Jose A.; Silva, Walter A.; Wieseman, Carol D.; Turnock, David L.
1993-01-01
This paper describes selected results from the flutter testing of the Supercritical Wing (SW) model. This model is a rigid semispan wing having a rectangular planform and a supercritical airfoil shape. The model was flutter tested in the Langley Transonic Dynamics Tunnel (TDT) as part of the Benchmark Models Program, a multi-year wind tunnel activity currently being conducted by the Structural Dynamics Division of NASA Langley Research Center. The primary objective of this program is to assist in the development and evaluation of aeroelastic computational fluid dynamics codes. The SW is the second of a series of three similar models which are designed to be flutter tested in the TDT on a flexible mount known as the Pitch and Plunge Apparatus. Data sets acquired with these models, including simultaneous unsteady surface pressures and model response data, are meant to be used for correlation with analytical codes. Presented in this report are experimental flutter boundaries and corresponding steady and unsteady pressure distribution data acquired over two model chords located at the 60 and 95 percent span stations.
A progress report on the development of an augmentor wing jet STOL research aircraft.
NASA Technical Reports Server (NTRS)
Quigley, H. C.; Sinclair, S. R. M.; Nark, T. C., Jr.; O'Keefe, J. V.
1971-01-01
The development of the aircraft has progressed to the point where the design of the modifications to the de Havilland C-8A Buffalo is complete and the engines are being tested. The predicted performance shows that the aircraft will be able to take off and land in less than 1500 ft. Simulation studies indicate that the handling qualities of the aircraft, with stability augmentation, will be acceptable for STOL research missions. Special techniques were required, however, for flight path control and transition from cruise to landing configuration .
Priddy, Tommy G.
1988-01-01
An inflatable wing is formed from a pair of tapered, conical inflatable tubes in bonded tangential contact with each other. The tubes are further connected together by means of top and bottom reinforcement boards having corresponding longitudinal edges lying in the same central diametral plane passing through the associated tube. The reinforcement boards are made of a stiff reinforcement material, such as Kevlar, collapsible in a direction parallel to the spanwise wing axis upon deflation of the tubes. The stiff reinforcement material cooperates with the inflated tubes to impart structural I-beam characteristics to the composite structure for transferring inflation pressure-induced tensile stress from the tubes to the reinforcement boards. A plurality of rigid hoops shaped to provide airfoil definition are spaced from each other along the spanwise axis and are connected to the top and bottom reinforcement boards. Tension lines are employed for stabilizing the hoops along the trailing and leading edges thereof.
Further Investigation of the Support System Effects and Wing Twist on the NASA Common Research Model
NASA Technical Reports Server (NTRS)
Rivers, Melissa B.; Hunter, Craig A.; Campbell, Richard L.
2012-01-01
An experimental investigation of the NASA Common Research Model was conducted in the NASA Langley National Transonic Facility and NASA Ames 11-foot Transonic Wind Tunnel Facility for use in the Drag Prediction Workshop. As data from the experimental investigations was collected, a large difference in moment values was seen between the experiment and computational data from the 4th Drag Prediction Workshop. This difference led to a computational assessment to investigate model support system interference effects on the Common Research Model. The results from this investigation showed that the addition of the support system to the computational cases did increase the pitching moment so that it more closely matched the experimental results, but there was still a large discrepancy in pitching moment. This large discrepancy led to an investigation into the shape of the as-built model, which in turn led to a change in the computational grids and re-running of all the previous support system cases. The results of these cases are the focus of this paper.
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 instability stoppers for wind tunnel models
NASA Technical Reports Server (NTRS)
Doggett, R. V., Jr.; Ricketts, R. H. (Inventor)
1981-01-01
A mechanism for constraining models or sections thereof, was wind tunnel tested, deployed at the onset of aeroelastic instability, to forestall destructive vibrations in the model is described. The mechanism includes a pair of arms pivoted to the tunnel wall and straddling the model. Rollers on the ends of the arms contact the model, and are pulled together against the model by a spring stretched between the arms. An actuator mechanism swings the arms into place and back as desired.
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.
Unsteady transonic flow simulation on a full-span-wing-body configuration
NASA Technical Reports Server (NTRS)
Guruswamy, Guru P.; Goorjian, Peter M.
1987-01-01
The presence of a body influences both the aerodynamic and aeroelastic performance of wings. Such effects are more pronounced in the transonic regime. To accurately account for the effect of the body, particularly when the wings are experiencing asymmetric modal motions, it is necessary to model the full configuration in the nonlinear transonic regime. In this study, full-span-wing-body configurations are simulated for the first time by a theoretical method that uses the unsteady potential equations based on the small-disturbance theory. The body geometry is modeled exactly as the physical shape, instead of as a rectangular box, which has been done in the past. Steady pressure computations for wing-body configurations compare well with the available experimental data. Unsteady pressure computations when the wings are oscillating in asymmetric modes show significant influence of the body.
Aeroelastic modeling for the FIT (Functional Integration Technology) team F/A-18 simulation
NASA Technical Reports Server (NTRS)
Zeiler, Thomas A.; Wieseman, Carol D.
1989-01-01
As part of Langley Research Center's commitment to developing multidisciplinary integration methods to improve aerospace systems, the Functional Integration Technology (FIT) team was established to perform dynamics integration research using an existing aircraft configuration, the F/A-18. An essential part of this effort has been the development of a comprehensive simulation modeling capability that includes structural, control, and propulsion dynamics as well as steady and unsteady aerodynamics. The structural and unsteady aerodynamics contributions come from an aeroelastic mode. Some details of the aeroelastic modeling done for the Functional Integration Technology (FIT) team research are presented. Particular attention is given to work done in the area of correction factors to unsteady aerodynamics data.
Evaluation of installed performance of a wing-tip-mounted pusher turboprop on a semispan wing
NASA Technical Reports Server (NTRS)
Patterson, James C., Jr.; Bartlett, Glynn R.
1987-01-01
An exploratory investigation has been conducted at the Langley Research Center to determine the effect of a wing-tip-mounted pusher turboprop on the aerodynamic characteristics of a semispan wing. Tests were conducted on a semispan model with an upswept, untapered wing and an airdriven motor that powered an SR-2 high-speed propeller located on the tip of the wing as a pusher propeller. All tests were conducted at a Mach number of 0.70 over an angle-of-attack range from approximately -2 to 4 deg at a Reynolds number of 3.82 x 10 to the 6th based on the wing reference chord of 13 in. The data indicate that, as a result of locating the propeller behind the wing trailing edge at the wing tip in the crossflow of the wing-tip vortex, it is possible to improve propeller performance and simultaneously reduce the lift-induced drag.
The development of an augmentor wing jet STOL research airplane (modified C-8A). Volume 1: Summary
NASA Technical Reports Server (NTRS)
Ashleman, R. H.; Kavdahl, H.
1972-01-01
A project to develop an experimental aircraft for use as an inflight demonstrator of the augmentor wing, short takeoff concept is discussed. The required modifications were made on a de Havilland C-8A aircraft. The modifications to the aircraft are explained and the performance of the modified aircraft is reported.
14 CFR 25.629 - Aeroelastic stability requirements.
Code of Federal Regulations, 2013 CFR
2013-01-01
... 14 Aeronautics and Space 1 2013-01-01 2013-01-01 false Aeroelastic stability requirements. 25.629 Section 25.629 Aeronautics and Space FEDERAL AVIATION ADMINISTRATION, DEPARTMENT OF TRANSPORTATION AIRCRAFT AIRWORTHINESS STANDARDS: TRANSPORT CATEGORY AIRPLANES Design and Construction General § 25.629 Aeroelastic stability requirements. (a)...
NASA Technical Reports Server (NTRS)
Straub, F. K.; Johnston, R. A.
1987-01-01
A 27% dynamically scaled model of the YAH-64 Advanced Attack Helicopter main rotor and hub has been designed and fabricated. The model will be tested in the NASA Langley Research Center V/STOL wind tunnel using the General Rotor Model System (GRMS). This report documents the studies performed to ensure dynamic similarity of the model with its full scale parent. It also contains a preliminary aeroelastic and aeromechanical substantiation for the rotor installation in the wind tunnel. From the limited studies performed no aeroelastic stability or load problems are projected. To alleviate a projected ground resonance problem, a modification of the roll characteristics of the GRMS is recommended.
Flutter of wings involving a locally distributed flexible control surface
NASA Astrophysics Data System (ADS)
Mozaffari-Jovin, S.; Firouz-Abadi, R. D.; Roshanian, J.
2015-11-01
This paper undertakes to facilitate appraisal of aeroelastic interaction of a locally distributed, flap-type control surface with aircraft wings operating in a subsonic potential flow field. The extended Hamilton's principle serves as a framework to ascertain the Euler-Lagrange equations for coupled bending-torsional-flap vibration. An analytical solution to this boundary-value problem is then accomplished by assumed modes and the extended Galerkin's method. The developed aeroelastic model considers both the inherent flexibility of the control surface displaced on the wing and the inertial coupling between these two flexible bodies. The structural deformations also obey the Euler-Bernoulli beam theory, along with the Kelvin-Voigt viscoelastic constitutive law. Meanwhile, the unsteady thin-airfoil and strip theories are the tools of producing the three-dimensional airloads. The origin of aerodynamic instability undergoes analysis in light of the oscillatory loads as well as the loads owing to arbitrary motions. After successful verification of the model, a systematic flutter survey was conducted on the theoretical effects of various control surface parameters. The results obtained demonstrate that the flapping modes and parameters of the control surface can significantly impact the flutter characteristics of the wings, which leads to a series of pertinent conclusions.
The DAST-1 remotely piloted research vehicle development and initial flight testing
NASA Technical Reports Server (NTRS)
Kotsabasis, A.
1981-01-01
The development and initial flight testing of the DAST (drones for aerodynamic and structural testing) remotely piloted research vehicle, fitted with the first aeroelastic research wing ARW-I are presented. The ARW-I is a swept supercritical wing, designed to exhibit flutter within the vehicle's flight envelope. An active flutter suppression system (FSS) designed to increase the ARW-I flutter boundary speed by 20 percent is described. The development of the FSS was based on prediction techniques of structural and unsteady aerodynamic characteristics. A description of the supporting ground facilities and aircraft systems involved in the remotely piloted research vehicle (RPRV) flight test technique is given. The design, specification, and testing of the remotely augmented vehicle system are presented. A summary of the preflight and flight test procedures associated with the RPRV operation is given. An evaluation of the blue streak test flight and the first and second ARW-I test flights is presented.
Study of the single body yawed-wing aircraft concept
NASA Technical Reports Server (NTRS)
Kulfan, R. M.; Nisbet, J. W.; Neuman, F. D.; Hamilton, E. J.; Murakami, J. K.; Mcbarron, J. P.; Kumasaka, K.
1974-01-01
Areas relating to the development and improvement of the single-fuselage, yawed-wing transonic transport concept were investigated. These included: (1) developing an alternate configuration with a simplified engine installation;(2) determining a structural design speed placard that would allow the engine-airframe match for optimum airplane performance; and (3) conducting an aeroelastic stability and control analysis of the yawed-wing configuration with a flexible wing. A two-engine, single-fuselage, yawed-wing configuration was developed that achieved the Mach 1.2 design mission at 5560 km (3000 nmi) and payload of 18,140 kg (40,000 lb) with a gross weight of 217,700 kg (480,000 lb). This airplane was slightly heavier than the aft-integrated four-engine configuration that had been developed in a previous study. A modified structural design speed placard, which was determined, resulted in a 6% to 8% reduction in the gross weight of the yawed-wing configurations. The dynamic stability characteristics of the single-fuselage yawed-wing configuration were found to be very dependent on the magnitude of the pitch/roll coupling, the static longitudinal stability, and the dihedral effect.
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.
NASA Technical Reports Server (NTRS)
Guruswamy, P.; Goorjian, P. M.
1982-01-01
Comparisons were made of computed and experimental data in three-dimensional unsteady transonic aerodynamics, including aeroelastic applications. The computer code LTRAN3, which is based on small-disturbance aerodynamic theory, was used to obtain the aerodynamic data. A procedure based on the U-g method was developed to compute flutter boundaries by using the unsteady aerodynamic coefficients obtained from LTRAN3. The experimental data were obtained from available NASA publications. All the studies were conducted for thin, unswept, rectangular wings with circular-arc cross sections. Numerical and experimental steady and unsteady aerodynamic data were compared for a wing with an aspect ratio of 3 and a thickness ratio of 5% at Mach numbers of 0.7 and 0.9. Flutter data were compared for a wing with an aspect ratio of 5. Two thickness ratios, 6% at Mach numbers of 0.715, 0.851, and 0.913, and 4% at Mach number of 0.904, were considered. Based on the unsteady aerodynamic data obtained from LTRAN3, flutter boundaries were computed; they were compared with those obtained from experiments and the code NASTRAN, which uses linear aerodynamics.
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.
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.
Wind-tunnel experiments on divergence of forward-swept wings
NASA Technical Reports Server (NTRS)
Ricketts, R. H.; Doggett, R. V., Jr.
1980-01-01
An experimental study to investigate the aeroelastic behavior of forward-swept wings was conducted in the Langley Transonic Dynamics Tunnel. Seven flat-plate models with varying aspect ratios and wing sweep angles were tested at low speeds in air. Three models having the same planform but different airfoil sections (i.e., flat-plate, conventional, and supercritical) were tested at transonic speeds in Freon 12. Linear analyses were performed to provide predictions to compare with the measured aeroelastic instabilities which include both static divergence and flutter. Six subcritical response testing techniques were formulated and evaluated at transonic speeds for accuracy in predicting static divergence. Two "divergence stoppers" were developed and evaluated for use in protecting the model from structural damage during tests.
Wing flutter calculations with the CAP-TSD unsteady transonic small disturbance program
NASA Technical Reports Server (NTRS)
Bennett, Robert M.; Batina, John T.; Cunningham, Herbert J.
1988-01-01
The application and assessment is described of CAP-TSD (Computational Aeroelasticity Program - Transonic Small Disturbance) code for flutter prediction. The CAP-TSD program was developed for aeroelastic analysis of complete aircraft configurations and was previously applied to the calculation of steady and unsteady pressures. Flutter calculations are presented for two thin, swept-and-tapered wing planforms with well defined modal properties. The calculations are for Mach numbers from low subsonic to low supersonic values, including the transonic range, and are compared with subsonic linear theory and experimental flutter data. The CAP-TSD flutter results are generally in good agreement with the experimental values and are in good agreement with subsonic linear theory when wing thickness is neglected.
NASA Astrophysics Data System (ADS)
Newman, James Charles, III
1997-10-01
practical interest, the sensitivity analysis and shape optimization has been performed for several two- and three-dimensional cases. In twodimensions, an initially symmetric NACA-0012 airfoil and a high-lift multielement airfoil were examined. For the three-dimensional configurations, an initially rectangular wing with uniform NACA-0012 cross-sections was optimized; in addition, a complete Boeing 747-200 aircraft was studied. Furthermore, the current study also examines the effect of inconsistency in the order of spatial accuracy between the nonlinear fluid and linear shape sensitivity equations. The second step was to develop a computationally efficient, high-fidelity, integrated static aeroelastic analysis procedure. To accomplish this, a structural analysis code was coupled with the aforementioned unstructured grid aerodynamic analysis solver. The use of an unstructured grid scheme for the aerodynamic analysis enhances the interaction compatibility with the wing structure. The structural analysis utilizes finite elements to model the wing so that accurate structural deflections may be obtained. In the current work, parameters have been introduced to control the interaction of the computational fluid dynamics and structural analyses; these control parameters permit extremely efficient static aeroelastic computations. To demonstrate and evaluate this procedure, static aeroelastic analysis results for a flexible wing in low subsonic, high subsonic (subcritical), transonic (supercritical), and supersonic flow conditions are presented.
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.
Wing Deployment Sequence #2: The deployable, inflatable wing technology demonstrator experiment airc
NASA Technical Reports Server (NTRS)
2001-01-01
Wing Deployment Sequence #2: The deployable, inflatable wing technology demonstrator experiment aircraft's wings continue deploying following separation from its carrier aircraft during a flight conducted by the NASA Dryden Flight Research Center, Edwards, California. The inflatable wing project represented a basic flight research effort by Dryden personnel. Three successful flights of the I2000 inflatable wing aircraft occurred. During the flights, the team air-launched the radio-controlled (R/C) I2000 from an R/C utility airplane at an altitude of 800-1000 feet. As the I2000 separated from the carrier aircraft, its inflatable wings 'popped-out,' deploying rapidly via an on-board nitrogen bottle. The aircraft remained stable as it transitioned from wingless to winged flight. The unpowered I2000 glided down to a smooth landing under complete control.
Wing Deployment Sequence #1: The deployable, inflatable wing technology demonstrator experiment airc
NASA Technical Reports Server (NTRS)
2001-01-01
Wing Deployment Sequence #1: The deployable, inflatable wing technology demonstrator experiment aircraft's wings begin deploying following separation from its carrier aircraft during a flight conducted by the NASA Dryden Flight Research Center, Edwards, California. The inflatable wing project represented a basic flight research effort by Dryden personnel. Three successful flights of the I2000 inflatable wing aircraft occurred. During the flights, the team air-launched the radio-controlled (R/C) I2000 from an R/C utility airplane at an altitude of 800-1000 feet. As the I2000 separated from the carrier aircraft, its inflatable wings 'popped-out,' deploying rapidly via an on-board nitrogen bottle. The aircraft remained stable as it transitioned from wingless to winged flight. The unpowered I2000 glided down to a smooth landing under complete control.
Wing Deployment Sequence #3: The deployable, inflatable wing technology demonstrator experiment airc
NASA Technical Reports Server (NTRS)
2001-01-01
Wing Deployment Sequence #3: The deployable, inflatable wing technology demonstrator experiment aircraft's wings fully deployed during flight following separation from its carrier aircraft during a flight conducted by the NASA Dryden Flight Research Center, Edwards, Californiaornia. The inflatable wing project represented a basic flight research effort by Dryden personnel. Three successful flights of the I2000 inflatable wing aircraft occurred. During the flights, the team air-launched the radio-controlled (R/C) I2000 from an R/C utility airplane at an altitude of 800-1000 feet. As the I2000 separated from the carrier aircraft, its inflatable wings 'popped-out,' deploying rapidly via an on-board nitrogen bottle. The aircraft remained stable as it transitioned from wingless to winged flight. The unpowered I2000 glided down to a smooth landing under complete control.
NASA Technical Reports Server (NTRS)
Weick, Fred E; Wenzinger, Carl J
1933-01-01
This report is the first in a series in which it is intended to compare the relative merits of all ordinary and some special forms of ailerons and other lateral control devices in regard to their effect on lateral controllability, lateral stability, and airplane performance. The comparisons are based on wind-tunnel test data, all the control devices being fitted to model wings having the same span, area, and airfoil section, and being subjected to the same series of force and rotation tests. The results are given for five different aileron movements: one with equal up-and-down deflection, one with average and one with extreme differential motion, one with upward deflection only, and one with the ailerons arranged to float with respect to the wing.
An overview of the active flexible wing program
NASA Technical Reports Server (NTRS)
Cole, Stanley R.; Perry, Boyd, III; Miller, Gerald D.
1991-01-01
An outline of the Active Flexible Wing (AFW) project that was meant to serve as an introduction to an entire session of the Computational Control Workshop is presented. Following background information on the project is a description of the AFW wind tunnel model and results from the initial wind tunnel test of the AFW model under the current project. Emphasis is on major project accomplishments. The AFW project is an effort to demonstrate aeroelastic control through the application of digital controls technology. Active flutter suppression and active control of maneuver loads during high speed rolling maneuvers are examined.
Delta wing flutter based on doublet lattice method in NASTRAN
NASA Technical Reports Server (NTRS)
Jew, H.
1975-01-01
The subsonic doublet-lattice method (DLM) aeroelastic analysis in NASTRAN was successfully applied to produce subsonic flutter boundary data in parameter space for a large delta wing configuration. Computed flow velocity and flutter frequency values as functions of air density ratio, flow Mach number, and reduced frequency are tabulated. The relevance and the meaning of the calculated results are discussed. Several input-deck problems encountered and overcome are cited with the hope that they may be helpful to NASTRAN Rigid Format 45 users.
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.
Calculations in bridge aeroelasticity via CFD
Brar, P.S.; Raul, R.; Scanlan, R.H.
1996-12-31
The central focus of the present study is the numerical calculation of flutter derivatives. These aeroelastic coefficients play an important role in determining the stability or instability of long, flexible structures under ambient wind loading. A class of Civil Engineering structures most susceptible to such an instability are long-span bridges of the cable-stayed or suspended-span variety. The disastrous collapse of the Tacoma Narrows suspension bridge in the recent past, due to a flutter instability, has been a big impetus in motivating studies in flutter of bridge decks.
Applications of the unsteady vortex-lattice method in aircraft aeroelasticity and flight dynamics
NASA Astrophysics Data System (ADS)
Murua, Joseba; Palacios, Rafael; Graham, J. Michael R.
2012-11-01
The unsteady vortex-lattice method provides a medium-fidelity tool for the prediction of non-stationary aerodynamic loads in low-speed, but high-Reynolds-number, attached flow conditions. Despite a proven track record in applications where free-wake modelling is critical, other less-computationally expensive potential-flow models, such as the doublet-lattice method and strip theory, have long been favoured in fixed-wing aircraft aeroelasticity and flight dynamics. This paper presents how the unsteady vortex-lattice method can be implemented as an enhanced alternative to those techniques for diverse situations that arise in flexible-aircraft dynamics. A historical review of the methodology is included, with latest developments and practical applications. Different formulations of the aerodynamic equations are outlined, and they are integrated with a nonlinear beam model for the full description of the dynamics of a free-flying flexible vehicle. Nonlinear time-marching solutions capture large wing excursions and wake roll-up, and the linearisation of the equations lends itself to a seamless, monolithic state-space assembly, particularly convenient for stability analysis and flight control system design. The numerical studies emphasise scenarios where the unsteady vortex-lattice method can provide an advantage over other state-of-the-art approaches. Examples of this include unsteady aerodynamics in vehicles with coupled aeroelasticity and flight dynamics, and in lifting surfaces undergoing complex kinematics, large deformations, or in-plane motions. Geometric nonlinearities are shown to play an instrumental, and often counter-intuitive, role in the aircraft dynamics. The unsteady vortex-lattice method is unveiled as a remarkable tool that can successfully incorporate all those effects in the unsteady aerodynamics modelling.
Aeroelastic Response of the Adaptive Compliant Trailing Edge Transtition Section
NASA Technical Reports Server (NTRS)
Herrera, Claudia Y.; Spivey, Natalie D.; Lung, Shun-fat
2016-01-01
The Adaptive Compliant Trailing Edge demonstrator was a joint task under the Environmentally Responsible Aviation Project in partnership with the Air Force Research Laboratory and FlexSys, Inc. (Ann Arbor, Michigan), chartered by the National Aeronautics and Space Administration to develop advanced technologies that enable environmentally friendly aircraft, such as continuous mold-line technologies. The Adaptive Compliant Trailing Edge demonstrator encompassed replacing the Fowler flaps on the SubsoniC Aircraft Testbed, a Gulfstream III (Gulfstream Aerospace, Savannah, Georgia) aircraft, with control surfaces developed by FlexSys, Inc., a pair of uniquely-designed, unconventional flaps to be used as lifting surfaces during flight-testing to substantiate their structural effectiveness. The unconventional flaps consisted of a main flap section and two transition sections, inboard and outboard, which demonstrated the continuous mold-line technology. Unique characteristics of the transition sections provided a challenge to the airworthiness assessment for this part of the structure. A series of build-up tests and analyses were conducted to ensure the data required to support the airworthiness assessment were acquired and applied accurately. The transition sections were analyzed both as individual components and as part of the flight-test article assembly. Instrumentation was installed in the transition sections based on the analysis to best capture the in-flight aeroelastic response. Flight-testing was conducted and flight data were acquired to validate the analyses. This paper documents the details of the aeroelastic assessment and in-flight response of the transition sections of the unconventional Adaptive Compliant Trailing Edge flaps.
Wing rotation and lift in SUEX flapping wing mechanisms
NASA Astrophysics Data System (ADS)
Mateti, Kiron; Byrne-Dugan, Rory A.; Tadigadapa, Srinivas A.; Rahn, Christopher D.
2013-01-01
This research presents detailed modeling and experimental testing of wing rotation and lift in the LionFly, a low cost and mass producible flapping wing mechanism fabricated monolithically from SUEX dry film and powered by piezoelectric bimorph actuators. A flexure hinge along the span of the wing allows the wing to rotate in addition to flapping. A dynamic model including aerodynamics is developed and validated using experimental testing with a laser vibrometer in air and vacuum, stroboscopic photography and high definition image processing, and lift measurement. The 112 mg LionFly produces 46° flap and 44° rotation peak to peak with 12° phase lag, which generates a maximum average lift of 71 μN in response to an applied sinusoidal voltage of 75 V AC and 75 V DC at 37 Hz. Simulated wing trajectories accurately predict measured wing trajectories at small voltage amplitudes, but slightly underpredict amplitude and lift at high voltage amplitudes. By reducing the length of the actuator, reducing the mechanism amplification and tuning the rotational hinge stiffness, a redesigned device is simulated to produce a lift to weight ratio of 1.5.
Development of an aeroelastic methodology for surface morphing rotors
NASA Astrophysics Data System (ADS)
Cook, James R.
Helicopter performance capabilities are limited by maximum lift characteristics and vibratory loading. In high speed forward flight, dynamic stall and transonic flow greatly increase the amplitude of vibratory loads. Experiments and computational simulations alike have indicated that a variety of active rotor control devices are capable of reducing vibratory loads. For example, periodic blade twist and flap excitation have been optimized to reduce vibratory loads in various rotors. Airfoil geometry can also be modified in order to increase lift coefficient, delay stall, or weaken transonic effects. To explore the potential benefits of active controls, computational methods are being developed for aeroelastic rotor evaluation, including coupling between computational fluid dynamics (CFD) and computational structural dynamics (CSD) solvers. In many contemporary CFD/CSD coupling methods it is assumed that the airfoil is rigid to reduce the interface by single dimension. Some methods retain the conventional one-dimensional beam model while prescribing an airfoil shape to simulate active chord deformation. However, to simulate the actual response of a compliant airfoil it is necessary to include deformations that originate not only from control devices (such as piezoelectric actuators), but also inertial forces, elastic stresses, and aerodynamic pressures. An accurate representation of the physics requires an interaction with a more complete representation of loads and geometry. A CFD/CSD coupling methodology capable of communicating three-dimensional structural deformations and a distribution of aerodynamic forces over the wetted blade surface has not yet been developed. In this research an interface is created within the Fully Unstructured Navier-Stokes (FUN3D) solver that communicates aerodynamic forces on the blade surface to University of Michigan's Nonlinear Active Beam Solver (UM/NLABS -- referred to as NLABS in this thesis). Interface routines are developed for
NASA Technical Reports Server (NTRS)
Kvaternik, Raymond G.; Piatak, David J.; Nixon, Mark W.; Langston, Chester W.; Singleton, Jeffrey D.; Bennett, Richard L.; Brown, Ross K.
2001-01-01
The results of a joint NASA/Army/Bell Helicopter Textron wind-tunnel test to assess the potential of Generalized Predictive Control (GPC) for actively controlling the swashplate of tiltrotor aircraft to enhance aeroelastic stability in the airplane mode of flight are presented. GPC is an adaptive time-domain predictive control method that uses a linear difference equation to describe the input-output relationship of the system and to design the controller. The test was conducted in the Langley Transonic Dynamics Tunnel using an unpowered 1/5-scale semispan aeroelastic model of the V-22 that was modified to incorporate a GPC-based multi-input multi-output control algorithm to individually control each of the three swashplate actuators. Wing responses were used for feedback. The GPC-based control system was highly effective in increasing the stability of the critical wing mode for all of the conditions tested, without measurable degradation of the damping in the other modes. The algorithm was also robust with respect to its performance in adjusting to rapid changes in both the rotor speed and the tunnel airspeed.
NASA Technical Reports Server (NTRS)
Whitlow, Jr., Woodrow (Editor); Todd, Emily N. (Editor)
1999-01-01
The proceedings of a workshop sponsored by the Confederation of European Aerospace Societies (CEAS), the American Institute of Aeronautics and Astronautics (AIAA), the National Aeronautics and Space Administration (NASA), Washington, D.C., and the Institute for Computer Applications in Science and Engineering (ICASE), Hampton, Virginia, and held in Williamsburg, Virginia June 22-25, 1999 represent a collection of the latest advances in aeroelasticity and structural dynamics from the world community. Research in the areas of unsteady aerodynamics and aeroelasticity, structural modeling and optimization, active control and adaptive structures, landing dynamics, certification and qualification, and validation testing are highlighted in the collection of papers. The wide range of results will lead to advances in the prediction and control of the structural response of aircraft and spacecraft.
Innovative scaling laws for aeroelastic and aeroservoelastic problems in compressible flow
NASA Astrophysics Data System (ADS)
Presente, Eyal
Active flutter suppression of a two dimensional wing section in subsonic flow is studied. The equations of motion of a typical cross section are presented in nondimensional form. A two degree of freedom problem, with pitch and plunge dynamics, combined with a trailing-edge control surface is considered. Aerodynamic loads are expressed in the time-domain using Roger's approximation. Augmented aerodynamic states are reconstructed using a Kalman filter, and linear optimal control is used to design a full-state feedback regulator for flutter suppression. Recent advances in the area of adaptive materials, smart structures, have led to the use of such materials as actuators for aeroservoelastic applications. The attractiveness of such materials consists of their potential to introduce continuous structural deformations of the lifting surface that can be exploited to manipulate the unsteady aerodynamic loads and prevent undesirable aeroelastic effects such as flutter. A general formulation of the aerodynamic loads, based on thin airfoil theory, and the deformation of a flat plate wing section are used to calculate the amount of power required to twist a wing along its span with piezoelectric patches. Composite materials enhance bend/twist coupling, which is used to modify the aerodynamic loads for the purpose of flutter suppression. Scaling laws of aeroservoelastic systems are addressed. Scaling parameters required for maintaining similarity between a full-scale system and a model are studied. An innovative two-pronged approach is used to obtain "similarity solutions" of the aeroservoelastic problem. Changes of structural and aerodynamic variables between a full scale configuration and its scaled models facilitate similarity between the systems. Two cases of scaled models are examined, a geometrically scaled model and an aeroelastically scaled one. Flutter suppression of a typical cross section employing a trailing edge control surface is compared with that of a typical
Scaling law and enhancement of lift generation of an insect-size hovering flexible wing.
Kang, Chang-kwon; Shyy, Wei
2013-08-01
We report a comprehensive scaling law and novel lift generation mechanisms relevant to the aerodynamic functions of structural flexibility in insect flight. Using a Navier-Stokes equation solver, fully coupled to a structural dynamics solver, we consider the hovering motion of a wing of insect size, in which the dynamics of fluid-structure interaction leads to passive wing rotation. Lift generated on the flexible wing scales with the relative shape deformation parameter, whereas the optimal lift is obtained when the wing deformation synchronizes with the imposed translation, consistent with previously reported observations for fruit flies and honeybees. Systematic comparisons with rigid wings illustrate that the nonlinear response in wing motion results in a greater peak angle compared with a simple harmonic motion, yielding higher lift. Moreover, the compliant wing streamlines its shape via camber deformation to mitigate the nonlinear lift-degrading wing-wake interaction to further enhance lift. These bioinspired aeroelastic mechanisms can be used in the development of flapping wing micro-robots. PMID:23760300
NASA Technical Reports Server (NTRS)
Mullen, J., Jr.
1978-01-01
The implementation of the changes to the program for Wing Aeroelastic Design and the development of a program to estimate aircraft fuselage weights are described. The equations to implement the modified planform description, the stiffened panel skin representation, the trim loads calculation, and the flutter constraint approximation are presented. A comparison of the wing model with the actual F-5A weight material distributions and loads is given. The equations and program techniques used for the estimation of aircraft fuselage weights are described. These equations were incorporated as a computer code. The weight predictions of this program are compared with data from the C-141.
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,
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.
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.
Optimization of nonlinear aeroelastic tailoring criteria
NASA Technical Reports Server (NTRS)
Abdi, F.; Ide, H.; Shankar, V. J.; Sobieszczanski-Sobieski, J.
1988-01-01
A static flexible fighter aircraft wing configuration is presently addressed by a multilevel optimization technique, based on both a full-potential concept and a rapid structural optimization program, which can be applied to such aircraft-design problems as maneuver load control, aileron reversal, and lift effectiveness. It is found that nonlinearities are important in the design of an aircraft whose flight envelope encompasses the transonic regime, and that the present structural suboptimization produces a significantly lighter wing by reducing ply thicknesses.
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.
Computational Aeroelastic Analyses of a Low-Boom Supersonic Configuration
NASA Technical Reports Server (NTRS)
Silva, Walter A.; Sanetrik, Mark D.; Chwalowski, Pawel; Connolly, Joseph
2015-01-01
An overview of NASA's Commercial Supersonic Technology (CST) Aeroservoelasticity (ASE) element is provided with a focus on recent computational aeroelastic analyses of a low-boom supersonic configuration developed by Lockheed-Martin and referred to as the N+2 configuration. The overview includes details of the computational models developed to date including a linear finite element model (FEM), linear unsteady aerodynamic models, unstructured CFD grids, and CFD-based aeroelastic analyses. In addition, a summary of the work involving the development of aeroelastic reduced-order models (ROMs) and the development of an aero-propulso-servo-elastic (APSE) model is provided.
Nonlinear Aerodynamics and the Design of Wing Tips
NASA Technical Reports Server (NTRS)
Kroo, Ilan
1991-01-01
The analysis and design of wing tips for fixed wing and rotary wing aircraft still remains part art, part science. Although the design of airfoil sections and basic planform geometry is well developed, the tip regions require more detailed consideration. This is important because of the strong impact of wing tip flow on wing drag; although the tip region constitutes a small portion of the wing, its effect on the drag can be significant. The induced drag of a wing is, for a given lift and speed, inversely proportional to the square of the wing span. Concepts are proposed as a means of reducing drag. Modern computational methods provide a tool for studying these issues in greater detail. The purpose of the current research program is to improve the understanding of the fundamental issues involved in the design of wing tips and to develop the range of computational and experimental tools needed for further study of these ideas.
Aerodynamic Analysis of a Hale Aircraft Joined-Wing Configuration
NASA Astrophysics Data System (ADS)
Sivaji, Rangarajan; Ghia, Urmila; Ghia, Karman; Thornburg, Hugh
2003-11-01
Aerodynamic analysis of a high-aspect ratio, joined wing of a High-Altitude Long Endurance (HALE) aircraft is performed. The requirement of high lift over extended flight periods for the HALE aircraft leads to high-aspect ratio wings experiencing significant deflections necessitating consideration of aeroelastic effects. The finite-volume solver COBALT, with Reynolds-averaged Navier-Stokes (RANS) and Detached Eddy Simulation (DES) capabilities, is used for the flow simulations. Calculations are performed at á = 0° and 12° for M = 0.6, at an altitude of 30,000 feet, at a Re per unit length of 5.6x106. The wing cross sections are NACA 4421 airfoils. Because of the high lift-to-drag ratio wings, an inviscid flow analysis is also performed. The inviscid surface pressure coefficient (Cp) is compared with the corresponding viscous Cp to examine the feasibility of the use of the inviscid pressure loads as an estimate of the total fluid loads on the structure. The viscous and inviscid Cp results compare reasonably only at á = 0°. The viscous flow is examined in detail via surface and field velocity vectors, vorticity, density and pressure contours. For á = 12°, the unsteady DES solutions show a weak shock at the aft-wing trailing edge. Also, the flow near the joint exhibits a region of mild separation.
NASA Technical Reports Server (NTRS)
Byrdsong, T. A.; Brooks, C. W., Jr.
1983-01-01
Wind-tunnel measurements were made of the wing-surface static-pressure distributions on a 0.237 scale model of a remotely piloted research vehicle equipped with a thick, high-aspect-ratio supercritical wing. Data are presented for two model configurations (with and without a ventral pod) at Mach numbers from 0.70 to 0.92 at angles of attack from -4 deg to 8 deg. Large variations of wing-surface local pressure distributions were developed; however, the characteristic supercritical-wing pressure distribution occurred near the design condition of 0.80 Mach number and 2 deg angle of attack. The significant variations of the local pressure distributions indicated pronounced shock-wave movements that were highly sensitive to angle of attack and Mach number. The effect of the vertical pod varied with test conditions; however at the higher Mach numbers, the effects on wing flow characteristics were significant at semispan stations as far outboard as 0.815. There were large variations of the wing loading in the range of test conditions, both model configurations exhibited a well-defined peak value of normal-force coefficient at the cruise angle of attack (2 deg) and Mach number (0.80).
Freight Wing Trailer Aerodynamics
Graham, Sean; Bigatel, Patrick
2004-10-17
Freight Wing Incorporated utilized the opportunity presented by this DOE category one Inventions and Innovations grant to successfully research, develop, test, patent, market, and sell innovative fuel and emissions saving aerodynamic attachments for the trucking industry. A great deal of past scientific research has demonstrated that streamlining box shaped semi-trailers can significantly reduce a truck's fuel consumption. However, significant design challenges have prevented past concepts from meeting industry needs. Market research early in this project revealed the demands of truck fleet operators regarding aerodynamic attachments. Products must not only save fuel, but cannot interfere with the operation of the truck, require significant maintenance, add significant weight, and must be extremely durable. Furthermore, SAE/TMC J1321 tests performed by a respected independent laboratory are necessary for large fleets to even consider purchase. Freight Wing used this information to create a system of three practical aerodynamic attachments for the front, rear and undercarriage of standard semi trailers. SAE/TMC J1321 Type II tests preformed by the Transportation Research Center (TRC) demonstrated a 7% improvement to fuel economy with all three products. If Freight Wing is successful in its continued efforts to gain market penetration, the energy and environmental savings would be considerable. Each truck outfitted saves approximately 1,100 gallons of fuel every 100,000 miles, which prevents over 12 tons of CO2 from entering the atmosphere. If all applicable trailers used the technology, the country could save approximately 1.8 billion gallons of diesel fuel, 18 million tons of emissions and 3.6 billion dollars annually.
Real-time simulation of aeroelastic rotor loads for horizontal axis wind turbines
NASA Astrophysics Data System (ADS)
Marnett, M.; Wellenberg, S.; Schröder, W.
2014-06-01
Wind turbine drivetrain research and test facilities with hardware-in-the-loop capabilities require a robust and accurate aeroelastic real-time rotor simulation environment. Recent simulation environments do not guarantee a computational response at real-time. Which is why a novel simulation tool has been developed. It resolves the physical time domain of the turbulent wind spectra and the operational response of the turbine at real-time conditions. Therefore, there is a trade-off between accuracy of the physical models and the computational costs. However, the study shows the possibility to preserve the necessary computational accuracy while simultaneously granting dynamic interaction with the aeroelastic rotor simulation environment. The achieved computational costs allow a complete aeroelastic rotor simulation at a resolution frequency of 100 Hz on standard computer platforms. Results obtained for the 5-MW reference wind turbine by the National Renewable Energy Laboratory (NREL) are discussed and compared to NREL's fatigue, aerodynamics, structures, and turbulence (FAST)- Code. The rotor loads show a convincing match. The novel simulation tool is applied to the wind turbine drivetrain test facility at the Center for Wind Power Drives (CWD), RWTH Aachen University to show the real-time hardware-in-the-loop capabilities.
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.
View east, showing Northwest Wing (Wing 5) and rear elevations ...
View east, showing Northwest Wing (Wing 5) and rear elevations of facade and tis flaking wings (Wings 1 and 2) - Hospital for Sick Children, 1731 Bunker Hill Road, Northeast, Washington, District of Columbia, DC
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)
Calise, A. J.; Kadushin, I.; Kramer, F.
1981-01-01
The current status of research on the application of variable structure system (VSS) theory to design aircraft flight control systems is summarized. Two aircraft types are currently being investigated: the Augmentor Wing Jet STOL Research Aircraft (AWJSRA), and AV-8A Harrier. The AWJSRA design considers automatic control of longitudinal dynamics during the landing phase. The main task for the AWJSRA is to design an automatic landing system that captures and tracks a localizer beam. The control task for the AV-8A is to track velocity commands in a hovering flight configuration. Much effort was devoted to developing computer programs that are needed to carry out VSS design in a multivariable frame work, and in becoming familiar with the dynamics and control problems associated with the aircraft types under investigation. Numerous VSS design schemes were explored, particularly for the AWJSRA. The approaches that appear best suited for these aircraft types are presented. Examples are given of the numerical results currently being generated.
Subtractive Structural Modification of Morpho Butterfly Wings.
Shen, Qingchen; He, Jiaqing; Ni, Mengtian; Song, Chengyi; Zhou, Lingye; Hu, Hang; Zhang, Ruoxi; Luo, Zhen; Wang, Ge; Tao, Peng; Deng, Tao; Shang, Wen
2015-11-11
Different from studies of butterfly wings through additive modification, this work for the first time studies the property change of butterfly wings through subtractive modification using oxygen plasma etching. The controlled modification of butterfly wings through such subtractive process results in gradual change of the optical properties, and helps the further understanding of structural optimization through natural evolution. The brilliant color of Morpho butterfly wings is originated from the hierarchical nanostructure on the wing scales. Such nanoarchitecture has attracted a lot of research effort, including the study of its optical properties, its potential use in sensing and infrared imaging, and also the use of such structure as template for the fabrication of high-performance photocatalytic materials. The controlled subtractive processes provide a new path to modify such nanoarchitecture and its optical property. Distinct from previous studies on the optical property of the Morpho wing structure, this study provides additional experimental evidence for the origination of the optical property of the natural butterfly wing scales. The study also offers a facile approach to generate new 3D nanostructures using butterfly wings as the templates and may lead to simpler structure models for large-scale man-made structures than those offered by original butterfly wings. PMID:26397977
Unsteady aerodynamic modeling and active aeroelastic control
NASA Technical Reports Server (NTRS)
Edwards, J. W.
1977-01-01
Unsteady aerodynamic modeling techniques are developed and applied to the study of active control of elastic vehicles. The problem of active control of a supercritical flutter mode poses a definite design goal stability, and is treated in detail. The transfer functions relating the arbitrary airfoil motions to the airloads are derived from the Laplace transforms of the linearized airload expressions for incompressible two dimensional flow. The transfer function relating the motions to the circulatory part of these loads is recognized as the Theodorsen function extended to complex values of reduced frequency, and is termed the generalized Theodorsen function. Inversion of the Laplace transforms yields exact transient airloads and airfoil motions. Exact root loci of aeroelastic modes are calculated, providing quantitative information regarding subcritical and supercritical flutter conditions.
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.
Wing flutter boundary prediction using unsteady Euler aerodynamic method
NASA Technical Reports Server (NTRS)
Lee-Rausch, Elizabeth M.; Batina, John T.
1993-01-01
Modifications to an existing 3D implicit upwind Euler/Navier-Stokes code for the aeroelastic analysis of wings are described. These modifications include the incorporation of a deforming mesh algorithm and the addition of the structural equations of motion for their simultaneous time-integration with the governing flow equations. The paper gives a brief description of these modifications and presents unsteady calculations which check the modifications to the code. Euler flutter results for an isolated 45 deg swept-back wing are compared with experimental data for seven freestream Mach numbers which define the flutter boundary over a range of Mach number from 0.499 to 1.14. These comparisons show good agreement in flutter characteristics for freestream Mach numbers below unity. For freestream Mach numbers above unity, the computed aeroelastic results predict a premature rise in the flutter boundary as compared with the experimental boundary. Steady and unsteady contours of surface Mach number and pressure are included to illustrate the basic flow characteristics of the time-marching flutter calculations and to aid in identifying possible causes for the premature rise in the computational flutter boundary.
Wing flutter boundary prediction using an unsteady Euler aerodynamic method
NASA Technical Reports Server (NTRS)
Lee-Rausch, Elizabeth M.; Batina, John T.
1993-01-01
Modifications to an existing three-dimensional, implicit, upwind Euler/Navier-Stokes code (CFL3D Version 2.1) for the aeroelastic analysis of wings are described. These modifications, which were previously added to CFL3D Version 1.0, include the incorporation of a deforming mesh algorithm and the addition of the structural equations of motion for their simultaneous time-integration with the government flow equations. The paper gives a brief description of these modifications and presents unsteady calculations which check the modifications to the code. Euler flutter results for an isolated 45 degree swept-back wing are compared with experimental data for seven freestream Mach numbers which define the flutter boundary over a range of Mach number from 0.499 to 1.14. These comparisons show good agreement in flutter characteristics for freestream Mach numbers below unity. For freestream Mach numbers above unity, the computed aeroelastic results predict a premature rise in the flutter boundary as compared with the experimental boundary. Steady and unsteady contours of surface Mach number and pressure are included to illustrate the basic flow characteristics of the time-marching flutter calculations and to aid in identifying possible causes for the premature rise in the computational flutter boundary.
View east, showing Northwest Wing (Wing 5), west wall of ...
View east, showing Northwest Wing (Wing 5), west wall of the North Wing (Wing 2) and rear elevations of the facade and its flanking wings (Wings 1 and 2) - Hospital for Sick Children, 1731 Bunker Hill Road, Northeast, Washington, District of Columbia, DC
NASA Dryden Flight Research Center
NASA Technical Reports Server (NTRS)
Navarro, Robert
2009-01-01
This DVD has several short videos showing some of the work that Dryden is involved in with experimental aircraft. These are: shots showing the Active AeroElastic Wing (AAW) loads calibration tests, AAW roll maneuvers, AAW flight control surface inputs, Helios flight, and takeoff, and Pathfinder takeoff, flight and landing.
X-29 Research Pilot Rogers Smith
NASA Technical Reports Server (NTRS)
1988-01-01
prevented stalling (loss of lift) at high angles of attack. Introduction of composite materials in the 1970s opened a new field of aircraft construction. It also made possible the construction of the X-29's thin supercritical wing. State-of-the-art composites allowed aeroelastic tailoring which, in turn, allowed the wing some bending but limited twisting and eliminated structural divergence within the flight envelope (i.e. deformation of the wing or the wing breaking off in flight). Additionally, composite materials allowed the wing to be sufficiently rigid for safe flight without adding an unacceptable weight penalty. The X-29 project consisted of two phases plus the follow-on vortex-control phase. Phase 1 demonstrated that the forward sweep of the X-29 wings kept the wing tips unstalled at the moderate angles of attack flown in that phase (a maximum of 21 degrees). Phase I also demonstrated that the aeroelastic tailored wing prevented structural divergence of the wing within the flight envelope, and that the control laws and control-surface effectiveness were adequate to provide artificial stability for an otherwise unstable aircraft. Phase 1 further demonstrated that the X-29 configuration could fly safely and reliably, even in tight turns. During Phase 2 of the project, the X-29, flying at an angle of attack of up to 67 degrees, demonstrated much better control and maneuvering qualities than computational methods and simulation models had predicted . During 120 research flights in this phase, NASA, Air Force, and Grumman project pilots reported the X-29 aircraft had excellent control response to an angle of attack of 45 degrees and still had limited controllability at a 67-degree angle of attack. This controllability at high angles of attack can be attributed to the aircraft's unique forward-swept wing- canard design. The NASA/Air Force-designed high-gain flight control laws also contributed to the good flying qualities. During the Air Force-initiated vortex
Gust Buffeting and Aeroelastic Behaviour of Poles and Monotubular Towers
NASA Astrophysics Data System (ADS)
Solari, G.; Pagnini, L. C.
1999-10-01
The evolution in the constructional field and the realization of ever more slender and light structures have emphasized the increasing difficulty of properly evaluating the actions and effects of wind on poles and monotubular towers. Faced with this situation the Italian constructors, united in a consortium coordinated by ACS ACAI Servizi, entrusted the Department of Structural and Geotechnical Engineering of Genova University with the task of formulating an ad hoc calculation procedure for this type of structure. This gave rise to a wide-ranging research project in which theoretical models, experimental evaluations and engineering methods were developed in parallel through an effective and quite a unique co-operation between researchers, designers and builders. This paper illustrates the physical aspects, the general principles and the basic formulation of the method proposed, with special emphasis on gust buffeting and aeroelastic phenomena. Preliminary results of full-scale measurements of the structural damping are also presented. The conclusions highlight the scientific and technical perspectives of this research.
NASA Technical Reports Server (NTRS)
Bobbitt, P. J.; Manro, M. E.; Kulfan, R. M.
1980-01-01
Wind tunnel tests of an arrow wing body configuration consisting of flat, twisted, and cambered twisted wings were conducted at Mach numbers from 0.40 to 2.50 to provide an experimental data base for comparison with theoretical methods. A variety of leading and trailing edge control surface deflections were included in these tests, and in addition, the cambered twisted wing was tested with an outboard vertical fin to determine its effect on wing and control surface loads. Theory experiment comparisons show that current state of the art linear and nonlinear attached flow methods were adequate at small angles of attack typical of cruise conditions. The incremental effects of outboard fin, wing twist, and wing camber are most accurately predicted by the advanced panel method PANAIR. Results of the advanced panel separated flow method, obtained with an early version of the program, show promise that accurate detailed pressure predictions may soon be possible for an aeroelasticity deformed wing at high angles of attack.
NASA Technical Reports Server (NTRS)
Weick, Fred E; Harris, Thomas A
1934-01-01
This report covers the sixth of a series of investigations in which various lateral control devices are compared with particular reference to their effectiveness at high angles of attack. The present report deals with flap-type ailerons hinged about axes having an angle with respect to the leading and trailing edges of the wing. Tests were made on four different skewed ailerons, including two different angles of skew and two sizes of ailerons. At the high angles of attack, all the skewed ailerons tested were slightly inferior with respect to rolling and yawing moments to straight ailerons having the same span and average chord. Computations indicate that the skewed ailerons are also inferior with respect to hinge moments.
Numerical computation of aeroelastically corrected transonic loads
NASA Technical Reports Server (NTRS)
Chipman, R.; Waters, C.; Mackenzie, D.
1979-01-01
A numerical scheme is presented for the computation of transonic aerodynamic loads on flexible wings. The method consists of iteratively applying the loads computed by a 3D transonic aerodynamics code to a structural model to obtain elastic twist, and then recomputing the loads. Because this iteration is performed concurrently with the iterations performed in computing the aerodynamics, flexible loads are obtained in roughly the same amount of computing time as required to obtain rigid loads. Applications of this method to a flexible supercritical transonic transport wing are presented and compared with model test data.
Aeroelastic instability stoppers for wind tunnel models
NASA Technical Reports Server (NTRS)
Doggett, R. V., Jr.; Ricketts, R. H. (Inventor)
1981-01-01
A mechanism for diverting the flow in a wind tunnel from the wing of a tested model is described. The wing is mounted on the wall of a tunnel. A diverter plate is pivotally mounted on the tunnel wall ahead of the model. An actuator fixed to the tunnel is pivotably connected to the diverter plate, by plunger. When the model is about to become unstable during the test the actuator moves the diverter plate from the tunnel wall to divert maintaining stable model conditions. The diverter plate is then retracted to enable normal flow.