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.
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.
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.
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
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.
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.
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.
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.
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.
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.
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.
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
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.
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.
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.
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.
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.
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.
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.
Airplane wing vibrations due to atmospheric turbulence
NASA Technical Reports Server (NTRS)
Pastel, R. L.; Caruthers, J. E.; Frost, W.
1981-01-01
The magnitude of error introduced due to wing vibration when measuring atmospheric turbulence with a wind probe mounted at the wing tip was studied. It was also determined whether accelerometers mounted on the wing tip are needed to correct this error. A spectrum analysis approach is used to determine the error. Estimates of the B-57 wing characteristics are used to simulate the airplane wing, and von Karman's cross spectrum function is used to simulate atmospheric turbulence. It was found that wing vibration introduces large error in measured spectra of turbulence in the frequency's range close to the natural frequencies of the wing.
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 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.
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.
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 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.
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
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.
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)
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.
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)
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.
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.
Buffet characteristics of the F-8 supercritical wing airplane
NASA Technical Reports Server (NTRS)
Deangelis, V. M.; Monaghan, R. C.
1977-01-01
The buffet characteristics of the F-8 supercritical wing airplane were investigated. Wing structural response was used to determine the buffet characteristics of the wing and these characteristics are compared with wind tunnel model data and the wing flow characteristics at transonic speeds. The wingtip accelerometer was used to determine the buffet onset boundary and to measure the buffet intensity characteristics of the airplane. The effects of moderate trailing edge flap deflections on the buffet onset boundary are presented. The supercritical wing flow characteristics were determined from wind tunnel and flight static pressure measurements and from a dynamic pressure sensor mounted on the flight test airplane in the vicinity of the shock wave that formed on the upper surface of the wing at transonic speeds. The comparison of the airplane's structural response data to the supercritical flow characteristics includes the effects of a leading edge vortex generator.
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.
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 16â€…mâ€…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)
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)
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.
Development of tailless and all-wing gliders and airplanes
NASA Technical Reports Server (NTRS)
Lademann, Robert W E
1932-01-01
Tailless airplanes are characterized by having all their control surfaces, especially the elevator, incorporated in the wings. This paper provides a discussion of the history of their development and current state of development.
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.
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.
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.
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.
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.
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.
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.
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.
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.
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.
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.
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.
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.
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.
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.
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).
NASA Technical Reports Server (NTRS)
Higdon, Donald T.
1959-01-01
An examination of oscillatory stability for a straight-winged airplane with large concentrated wing-tip masses was made using wing-bending and airplane-pitching degrees of freedom and considering only quasi-steady aerodynamic forces. It was found that instability caused by coupling of airplane pitching and wing bending occurred for large ratios of effective wing-tip mass to total airplane mass and for coupled wing-bending frequencies near or below the uncoupled pitching frequency. Boundaries for this instability are given in terms of two quantities: (1) the ratio of effective tip mass to airplane mass, which can be estimated, and (2) the ratio of the coupled bending frequency to the uncoupled pitch frequency, which can be measured in flight. These boundaries are presented for various values of several airplane parameters.
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.
A possible method for preventing the autorotation of airplane wings
NASA Technical Reports Server (NTRS)
Schrenk, Oskar
1930-01-01
At the suggestion of Professor Betz, the following device was tested with the object of reducing the autorotational speed of airplane wings. The model of a normal wing with the Gottigen profile 420, 1 meter span and 0.2 meter chord was provided with a pair of symmetrical slots on the suction side, connected with each other inside the wing. The arrangement of the testing equipment and models are given and the effect of the slots can be seen in the experimental curves that are included.
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.
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.
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.
NASA Technical Reports Server (NTRS)
Whitlow, W., Jr.; Bennett, R. M.
1982-01-01
Since the aerodynamic theory is nonlinear, the method requires the coupling of two iterative processes - an aerodynamic analysis and a structural analysis. A full potential analysis code, FLO22, is combined with a linear structural analysis to yield aerodynamic load distributions on and deflections of elastic wings. This method was used to analyze an aeroelastically-scaled wind tunnel model of a proposed executive-jet transport wing and an aeroelastic research wing. The results are compared with the corresponding rigid-wing analyses, and some effects of elasticity on the aerodynamic loading are noted.
NASA Technical Reports Server (NTRS)
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.
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.
Stresses Produced in Airplane Wings by Gusts
NASA Technical Reports Server (NTRS)
Kussner, Hans Georg
1932-01-01
Accurate prediction of gust stress being out of the question because of the multiplicity of the free air movements, the exploration of gust stress is restricted to static method which must be based upon: 1) stress measurements in free flight; 2) check of design specifications of approved type airplanes. With these empirical data the stress must be compared which can be computed for a gust of known intensity and structure. This "maximum gust" then must be so defined as to cover the whole ambit of empiricism and thus serve as prediction for new airplane designs.
The Design of Airplane Wing Ribs
NASA Technical Reports Server (NTRS)
Newlin, J A; Trayer, George W
1931-01-01
The purpose of this investigation was to obtain information for use in the design of truss and plywood forms, particularly with reference to wing ribs. Tests were made on many designs of wing ribs, comparing different types in various sizes. Many tests were also made on parallel-chord specimens of truss and plywood forms in place of the actual ribs and on parts of wing ribs, such as truss diagonals and sections of cap strips. It was found that for ribs of any size or proportions, when they were designed to obtain a well-balanced construction and were carefully manufactured, distinct types are of various efficiencies; the efficiency is based on the strength per unit of weight. In all types of ribs the heavier are the stronger per unit of weight. Reductions in the weight of wing ribs are accompanied even in efficient designs by a much greater proportional reduction in strength.
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.
NASA Technical Reports Server (NTRS)
Rhyne, Richard H
1958-01-01
A flight investigation was made on a large sweptback-wing bomber airplane and the results are compared with data previously obtained at low altitude (5,000 feet). The effects of wing flexibility on the wing strains were, on the average, about 20 percent larger at the higher altitude.
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.
The right wing of the LEFT airplane
NASA Technical Reports Server (NTRS)
Powell, Arthur G.
1987-01-01
The NASA Leading-Edge Flight Test (LEFT) program addressed the environmental issues which were potential problems in the application of Laminar Flow Control (LFC) to transport aircraft. These included contamination of the LFC surface due to dirt, rain, insect remains, snow, and ice, in the critical leading-edge region. Douglas Aircraft Company designed and built a test article which was mounted on the right wing of the C-140 JetStar aircraft. The test article featured a retractable leading-edge high-lift shield for contamination protection and suction through perforations on the upper surface for LFC. Following a period of developmental flight testing, the aircraft entered simulated airline service, which included exposure to airborne insects, heavy rain, snow, and icing conditions both in the air and on the ground. During the roughly 3 years of flight testing, the test article has consistently demonstrated laminar flow in cruising flight. The experience with the LEFT experiment was summarized with emphasis on significant test findings. The following items were discussed: test article design and features; suction distribution; instrumentation and transition point reckoning; problems and fixes; system performance and maintenance requirements.
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.
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.
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)
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.
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.
NASA Technical Reports Server (NTRS)
Crowley, J W , Jr; Green, M W
1929-01-01
This investigation was conducted by the National Advisory Committee for Aeronautics at Langley Field, Va., at the request of the Army Air Corps, for the purpose of comparing the full scale lift and drag characteristics of an airplane equipped with several sets of wings of commonly used airfoil sections. A Sperry Messenger Airplane with wings of R.A.F.-15, U.S.A.-5, U.S.A.-27, and Gottingen 387 airfoil sections was flown and the lift and drag characteristics of the airplane with each set of wings were determined by means of glide tests. The results are presented in tabular and curve form. (author)
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.
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.
Effects of airplane flexibility on wing bending strains in rough air
NASA Technical Reports Server (NTRS)
Coleman, Thomas L; Press, Harry; Shufflebarger, C C
1957-01-01
Some results on the effects of wing flexibility on wing bending strains as determined from flight tests of a Boeing B-29 and a Boeing B-47A airplane in rough air are presented. Results from an analytical study of the flexibility effects on the B-29 wing strains are compared with the experimental results. Both the experimental and calculated results are presented as frequency-response functions of the bending strains at various spanwise wing stations to gust disturbances. In addition, some indirect evidence of the effect of spanwise variations in turbulence on the response of the B-47A airplane is presented.
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.
Determination of the Profile Drag of an Airplane Wing in Flight at High Reynolds Numbers
NASA Technical Reports Server (NTRS)
Bicknell, Joseph
1939-01-01
Flight tests were made to determine the profile-drag coefficients of a portion of the original wing surface of an all-metal airplane and of a portion of the wing made aerodynamically smooth and more nearly fair than the original section. The wing section was approximately the NACA 2414.5. The tests were carried out over a range of airplane speeds giving a maximum Reynolds number of 15,000,000. Tests were also carried out to locate the point of transition from laminar to turbulent boundary layer and to determine the velocity distribution along the upper surface of the wing. The profile-drag coefficients of the original and of the smooth wing portions at a Reynolds number of 15,000,000 were 0.0102 and 0.0068, respectively; i.e., the surface irregularities on the original wing increased the profile-drag coefficient 50 percent above that of the smooth wing.
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.
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.
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.
NASA Technical Reports Server (NTRS)
Friend, E. L.; Sakamoto, G. M.
1978-01-01
A flight research program was conducted to investigate the improvements in maneuverability of an F-111A airplane equipped with a supercritical wing. In this configuration the aircraft is known as the F-111 TACT (transonic aircraft technology) airplane. The variable-wing-sweep feature permitted an evaluation of the supercritical wing in many configurations. The primary emphasis was placed on the transonic Mach number region, which is considered to be the principal air combat arena for fighter aircraft. An agility study was undertaken to assess the maneuverability of the F-111A aircraft with a supercritical wing at both design and off-design conditions. The evaluation included an assessment of aerodynamic and maneuver performance in conjunction with an evaluation of precision controllability during tailchase gunsight tracking tasks.
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.
Application of a performance modeling technique to an airplane with variable sweep wings
NASA Technical Reports Server (NTRS)
Redin, P. C.
1981-01-01
A performance modeling concept previously applied to an F-104F G and a YF-12C airplane was applied to an F-111A airplane. This application extended the concept to an airplane with variable sweep wings. The performance model adequately matched flight test data for maneuvers flown at different wing sweep angles at maximum afterburning and intermediate power settings. For maneuvers flown at less than intermediate power, including dynamic maneuvers, the performance model was not validated because the method used to correlate model and in-flight power setting was not adequate. Individual dynamic maneuvers were matched sucessfully by using adjustments unique to each maneuver.
Effect of the Surface Condition of a Wing on the Aerodynamic Characteristics of an Airplane
NASA Technical Reports Server (NTRS)
Defrance, S J
1934-01-01
In order to determine the effect of the surface conditions of a wing on the aerodynamic characteristics of an airplane, tests were conducted in the N.A.C.A. full-scale wind tunnel on the Fairchild F-22 airplane first with normal commercial finish of wing surface and later with the same wing polished. Comparison of the characteristics of the airplane with the two surface conditions shows that the polish caused a negligible change in the lift curve, but reduced the minimum drag coefficient by 0.001. This reduction in drag if applied to an airplane with a given speed of 200 miles per hour and a minimum drag coefficient of 0.025 would increase the speed only 2.9 miles per hour, but if the speed remained the same, the power would be reduced 4 percent.
Analysis of Nonplanar Wing-tip-mounted Lifting Surfaces on Low-speed Airplanes
NASA Technical Reports Server (NTRS)
Vandam, C. P.; Roskam, J.
1983-01-01
Nonplanar wing tip mounted lifting surfaces reduce lift induced drag substantially. Winglets, which are small, nearly vertical, winglike surfaces, are an example of these devices. To achieve reduction in lift induced drag, winglets produce significant side forces. Consequently, these surfaces can seriously affect airplane lateral directional aerodynamic characteristics. Therefore, the effects of nonplanar wing tip mounted surfaces on the lateral directional stability and control of low speed general aviation airplanes were studied. The study consists of a theoretical and an experimental, in flight investigation. The experimental investigation involves flight tests of winglets on an agricultural airplane. Results of these tests demonstrate the significant influence of winglets on airplane lateral directional aerodynamic characteristics. It is shown that good correlations exist between experimental data and theoretically predicted results. In addition, a lifting surface method was used to perform a parametric study of the effects of various winglet parameters on lateral directional stability derivatives of general aviation type wings.
Approximations for column effect in airplane wing spars
NASA Technical Reports Server (NTRS)
Warner, Edward P; Short, Mac
1927-01-01
The significance attaching to "column effect" in airplane wing spars has been increasingly realized with the passage of time, but exact computations of the corrections to bending moment curves resulting from the existence of end loads are frequently omitted because of the additional labor involved in an analysis by rigorously correct methods. The present report represents an attempt to provide for approximate column effect corrections that can be graphically or otherwise expressed so as to be applied with a minimum of labor. Curves are plotted giving approximate values of the correction factors for single and two bay trusses of varying proportions and with various relationships between axial and lateral loads. It is further shown from an analysis of those curves that rough but useful approximations can be obtained from Perry's formula for corrected bending moment, with the assumed distance between points of inflection arbitrarily modified in accordance with rules given in the report. The discussion of general rules of variation of bending stress with axial load is accompanied by a study of the best distribution of the points of support along a spar for various conditions of loading.
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.
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.
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)
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)
Ashley, H.
1984-01-01
Graduate research activity in the following areas is reported: the divergence of laminated composite lifting surfaces, subsonic propeller theory and aeroelastic analysis, and cross sectional resonances in wind tunnels.
NASA Technical Reports Server (NTRS)
Lovell, Powell M , Jr; Parlett, Lysle P
1956-01-01
An investigation of the stability and control of a low-wing four-engine transport vertical-take-off airplane during the transition from hovering to normal forward flight has been conducted with a remotely controlled free-flight model. The model had four propellers distributed along the wing with the thrust axes in the wing-chord plane. The wing could be rotated to 90 degrees incidence so that the propeller thrust axes were vertical for hovering flight.
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.
An analysis of life expectancy of airplane wings in normal cruising flight
NASA Technical Reports Server (NTRS)
Putnam, Abbott A
1945-01-01
In order to provide a basis for judging the relative importance of wing failure by fatigue and by single intense gusts, an analysis of wing life for normal cruising flight was made based on data on the frequency of atmospheric gusts. The independent variables considered in the analysis included stress-concentration factor, stress-load relation, wing loading, design and cruising speeds, design gust velocity, and airplane size. Several methods for estimating fatigue life from gust frequencies are discussed. The procedure selected for the analysis is believed to be simple and reasonably accurate, though slightly conservative.
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.
Analysis of Effect of Rolling Pull-Outs on Wing and Aileron Loads of a Fighter Airplane
NASA Technical Reports Server (NTRS)
Pearson, Henry A.; Aiken, William S.
1946-01-01
An analysis was made to determine the effect of rolling pull-out maneuvers on the wing and aileron loads of a typical fighter airplane, the P-47B. The results obtained indicate that higher loads are imposed upon wings and ailerons because of the rolling pull-out maneuver, than would be obtained by application of the loading requirements to which the airplane was designed. An increase of 102 lb or 15 percent of wing weight would be required if the wing were designed for rolling pull-out maneuver. It was also determined that the requirements by which the aileron was originally designed were inadequate.
Measurement and Analysis of Wing and Tail Buffeting Loads on a Fighter Airplane
NASA Technical Reports Server (NTRS)
Huston, Wilber B; Skopinski, T H
1955-01-01
The buffeting loads measured on the wing and tail of a fighter airplane during 194 maneuvers are given in tabular form, along with the associated flight conditions. Measurements were made at altitudes of 30,000 to 10,000 feet and at speeds up to a Mach number of 0.8. Least-squares methods have been used for a preliminary analysis of the data. The agreement between the results of this analysis and the loads measured in stalls is sufficiently good to suggest the examination of the buffeting of other airplanes on the same basis.
NASA Technical Reports Server (NTRS)
Silva, Walter A.; Perry, Boyd III; Chwalowski, Pawel
2014-01-01
Reduced-order modeling (ROM) methods are applied to the CFD-based aeroelastic analysis of the AGARD 445.6 wing in order to gain insight regarding well-known discrepancies between the aeroelastic analyses and the experimental results. The results presented include aeroelastic solutions using the inviscid CAP-TSD code and the FUN3D code (Euler and Navier-Stokes). Full CFD aeroelastic solutions and ROM aeroelastic solutions, computed at several Mach numbers, are presented in the form of root locus plots in order to better reveal the aeroelastic root migrations with increasing dynamic pressure. Important conclusions are drawn from these results including the ability of the linear CAP-TSD code to accurately predict the entire experimental flutter boundary (repeat of analyses performed in the 1980's), that the Euler solutions at supersonic conditions indicate that the third mode is always unstable, and that the FUN3D Navier-Stokes solutions stabilize the unstable third mode seen in the Euler solutions.
NASA Technical Reports Server (NTRS)
Seidman, Oscar; Neihouse, A I
1940-01-01
The reported tests are a continuation of an NACA investigation being made in the free-spinning wind tunnel to determine the effects of independent variations in load distribution, wing and tail arrangement, and control disposition on the spin characteristics of airplanes. The standard series of tests was repeated to determine the effect of airplane relative density. Tests were made at values of the relative-density parameter of 6.8, 8.4 (basic), and 12.0; and the results were analyzed. The tested variations in the relative-density parameter may be considered either as variations in the wing loading of an airplane spun at a given altitude, with the radii of gyration kept constant, or as a variation of the altitude at which the spin takes place for a given airplane. The lower values of the relative-density parameter correspond to the lower wing loadings or to the lower altitudes of the spin.
Crash tests of four identical high-wing single-engine airplanes
NASA Technical Reports Server (NTRS)
Vaughan, V. L., Jr.; Hayduk, R. J.
1980-01-01
Four identical four place, high wing, single engine airplane specimens with nominal masses of 1043 kg were crash tested at the Langley Impact Dynamics Research Facility under controlled free flight conditions. These tests were conducted with nominal velocities of 25 m/sec along the flight path angles, ground contact pitch angles, and roll angles. Three of the airplane specimens were crashed on a concrete surface; one was crashed on soil. Crash tests revealed that on a hard landing, the main landing gear absorbed about twice the energy for which the gear was designed but sprang back, tending to tip the airplane up to its nose. On concrete surfaces, the airplane impacted and remained in the impact attitude. On soil, the airplane flipped over on its back. The crash impact on the nose of the airplane, whether on soil or concrete, caused massive structural crushing of the forward fuselage. The liveable volume was maintained in both the hard landing and the nose down specimens but was not maintained in the roll impact and nose down on soil specimens.
Influence of elliptical distribution of lift on strength of airplane wings
NASA Technical Reports Server (NTRS)
DORAND
1922-01-01
Hitherto it has been generally assumed, in calculating the fall of an airplane, that the forces withstood by the latter were distributed uniformly throughout the whole length of the wing. In reality this is not the case and German engineers in particular are now assuming an elliptical distribution of the forces. The latter hypothesis has made it possible to carry out a certain number of calculations which have been verified by experiment.
High-Speed Tests of a Model Twin-Engine Low-Wing Transport Airplane
NASA Technical Reports Server (NTRS)
Becker, John V; LEONARD LLOYD H
1942-01-01
Report presents the results of force tests made of a 1/8-scale model of a twin-engine low-wing transport airplane in the NACA 8-foot high-speed tunnel to investigate compressibility and interference effects of speeds up to 450 miles per hour. In addition to tests of the standard arrangement of the model, tests were made with several modifications designed to reduce the drag and to increase the critical speed.
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.
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.
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.
Dynamic Ground Effect for a Cranked Arrow Wing Airplane
NASA Technical Reports Server (NTRS)
Curry, Robert E.
1997-01-01
Flight-determined ground effect characteristics for an F-16XL airplane are presented and correlated with wind tunnel predictions and similar flight results from other aircraft. Maneuvers were conducted at a variety of flightpath angles. Conventional ground effect flight test methods were used, with the exception that space positioning data were obtained using the differential global positioning system (DGPS). Accuracy of the DGPS was similar to that of optical tracking methods, but it was operationally more attractive. The dynamic flight determined lift and drag coefficient increments were measurably lower than steady-state wind-tunnel predictions. This relationship is consistent with the results of other aircraft for which similar data are available. Trends in the flight measured lift increments caused by ground effect as a function of flightpath angle were evident but weakly correlated. An engineering model of dynamic ground effect was developed based on linear aerodynamic theory and super-positioning of flows. This model was applied to the F-16XL data set and to previously published data for an F-15 airplane. In both cases, the model provided an engineering estimate of the ratio between the steady-state and dynamic data sets.
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.
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.
Development of a Composite Tailoring Procedure for Airplane Wings
NASA Technical Reports Server (NTRS)
Chattopadhyay, Aditi
2000-01-01
The quest for finding optimum solutions to engineering problems has existed for a long time. In modern times, the development of optimization as a branch of applied mathematics is regarded to have originated in the works of Newton, Bernoulli and Euler. Venkayya has presented a historical perspective on optimization in [1]. The term 'optimization' is defined by Ashley [2] as a procedure "...which attempts to choose the variables in a design process so as formally to achieve the best value of some performance index while not violating any of the associated conditions or constraints". Ashley presented an extensive review of practical applications of optimization in the aeronautical field till about 1980 [2]. It was noted that there existed an enormous amount of published literature in the field of optimization, but its practical applications in industry were very limited. Over the past 15 years, though, optimization has been widely applied to address practical problems in aerospace design [3-5]. The design of high performance aerospace systems is a complex task. It involves the integration of several disciplines such as aerodynamics, structural analysis, dynamics, and aeroelasticity. The problem involves multiple objectives and constraints pertaining to the design criteria associated with each of these disciplines. Many important trade-offs exist between the parameters involved which are used to define the different disciplines. Therefore, the development of multidisciplinary design optimization (MDO) techniques, in which different disciplines and design parameters are coupled into a closed loop numerical procedure, seems appropriate to address such a complex problem. The importance of MDO in successful design of aerospace systems has been long recognized. Recent developments in this field have been surveyed by Sobieszczanski-Sobieski and Haftka [6].
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)
Friend, E. L.; Sefic, W. J.
1972-01-01
A flight program was conducted on the F-104 airplane to investigate the effects of moderate deflections of wing leading- and trailing-edge flaps on the buffet characteristics at subsonic and transonic Mach numbers. Selected deflections of the wing leading and trailing-edge flaps, individually and in combination, were used to assess buffet onset, intensity, and frequency; lift curves; and wing-rock characteristics for each configuration. Increased deflection of the trailing-edge flap delayed the buffet onset and buffet intensity rise to a significantly higher airplane normal-force coefficient. Deflection of the leading-edge flap produced some delay in buffet onset and the resulting intensity rise at low subsonic speeds. Increased deflection of the trailing-edge flap provided appreciable lift increments in the angle-of-attack range covered, whereas the leading-edge flap provided lift increments only at high angles-of-attack. The pilots appreciated the increased maneuvering envelope provided by the flaps because of the improved turn capability.
An Airplane Design having a Wing with Fuselage Attached to Each Tip
NASA Technical Reports Server (NTRS)
Spearman, Leroy M.
2001-01-01
This paper describes the conceptual design of an airplane having a low aspect ratio wing with fuselages that are attached to each wing tip. The concept is proposed for a high-capacity transport as an alternate to progressively increasing the size of a conventional transport design having a single fuselage with cantilevered wing panels attached to the sides and tail surfaces attached at the rear. Progressively increasing the size of conventional single body designs may lead to problems in some area's such as manufacturing, ground-handling and aerodynamic behavior. A limited review will be presented of some past work related to means of relieving some size constraints through the use of multiple bodies. Recent low-speed wind-tunnel tests have been made of models representative of the inboard-wing concept. These models have a low aspect ratio wing with a fuselage attached to each tip. Results from these tests, which included force measurements, surface pressure measurements, and wake surveys, will be presented herein.
Crash tests of three identical low-wing single-engine airplane
NASA Technical Reports Server (NTRS)
Castle, C. B.; Alfaro-Bou, E.
1983-01-01
Three identical four place, low wing single engine airplane specimens with nominal masses of 1043 kg were crash tested under controlled free flight conditions. The tests were conducted at the same nominal velocity of 25 m/sec along the flight path. Two airplanes were crashed on a concrete surface (at 10 and 30 deg pitch angles), and one was crashed on soil (at a -30 deg pitch angle). The three tests revealed that the specimen in the -30 deg test on soil sustained massive structural damage in the engine compartment and fire wall. Also, the highest longitudinal cabin floor accelerations occurred in this test. Severe damage, but of lesser magnitude, occurred in the -30 deg test on concrete. The highest normal cabin floor accelerations occurred in this test. The least structural damage and lowest accelerations occurred in the 10 deg test on concrete.
The Pressure Distribution over the Wings and Tail Surfaces of a PW-9 Pursuit Airplane in Flight
NASA Technical Reports Server (NTRS)
Rhode, Richard
1931-01-01
This report presents the results of an investigation to determine (1) the magnitude and distribution of aerodynamic loads over the wings and tail surfaces of a pursuit-type airplane in the maneuvers likely to impose critical loads on the various subassemblies of the airplane structure. (2) To study the phenomenon of center of pressure movement and normal force coefficient variation in accelerated flight, and (3) to measure the normal accelerations at the center of gravity, wing-tip, and tail, in order to determine the nature of the inertia forces acting simultaneously with the critical aerodynamic loads. The results obtained throw light on a number of important questions involving structural design. Some of the more interesting results are discussed in some detail, but in general the report is for the purpose of making this collection of airplane-load data obtained in flight available to those interested in airplane structures.
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.
Summary of Results Obtained in Full-Scale Tunnel Investigation of the Ryan Flex-Wing Airplane
NASA Technical Reports Server (NTRS)
Johnson, Joseph L., Jr.; Hassell, James L., Jr.
1962-01-01
The performance and static stability and control characteristics of the Ryan Flex-Wing airplane were determined in an investigation conducted in the Langley full-scale tunnel through an angle-of-attack range of the keel from about 14 to 44 deg. for power-on and -off conditions. Comparisons of the wind-tunnel data with flight-test data obtained with the same airplane by the Ryan Aeronautical Company were made in a number of cases.
A simple method for increasing the lift of airplane wings by means of flaps
NASA Technical Reports Server (NTRS)
Gruschwitz, Eugen; Schrenk, Oskar
1933-01-01
Aerodynamic considerations led us, not long ago, to investigate a device which seemed to promise a contribution to the problem of reducing the landing speed of an airplane. We have subsequently learned that similar devices had already been proposed and investigated by others, but it seems advisable, nevertheless, to report our results. The problem is to create, in landing, a region of turbulence on the lower side of the wing near the trailing edge by some obstacle to the air flow. The devices tested by us consisted of flaps of varying chord and position, the chord s being equal to the distance of the pivot from the trailing edge.
The Strength of One-Piece Solid, Build-Up and Laminated Wood Airplane Wing Beams
NASA Technical Reports Server (NTRS)
Nelson, John H
1920-01-01
The purpose of this report is to summarize the results of all wood airplane wing beams tested to date in the Bureau of Standards Laboratory in order that the various kinds of wood and methods of construction may be compared. All beams tested were of an I section and the majority were somewhat similar in size and cross section to the front wing beam of the Curtiss JN-4 machine. Construction methods may be classed as (1) solid beams cut from solid stock; (2) three-piece beams, built up of three pieces, web and flanges glued together by a tongue-and-groove joint and (3) laminated beams built up of thin laminations of wood glued together.
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.
Pressure distribution for the wing of the YAV-8B airplane; with and without pylons
NASA Technical Reports Server (NTRS)
Saltzman, Edwin J.; Delfrate, John H.; Sabsay, Catherine M.; Yarger, Jill M.
1992-01-01
Pressure distribution data have been obtained in flight at four span stations on the wing panel of the YAV-8B airplane. Data obtained for the supercritical profiled wing, with and without pylons installed, ranged from Mach 0.46 to 0.88. The altitude ranged from approximately 20,000 to 40,000 ft and the resultant Reynolds numbers varied from approximately 7.2 million to 28.7 million based on the mean aerodynamic chord. Pressure distribution data and flow visualization results show that the full-scale flight wing performance is compromised because the lower surface cusp region experiences flow separation for some important transonic flight conditions. This condition is aggravated when local shocks occur on the lower surface of the wing (mostly between 20 and 35 percent chord) when the pylons are installed for Mach 0.8 and above. There is evidence that convex fairings, which cover the pylon attachment flanges, cause these local shocks. Pressure coefficients significantly more negative than those for sonic flow also occur farther aft on the lower surface (near 60 percent chord) whether or not the pylons are installed for Mach numbers greater than or equal to 0.8. These negative pressure coefficient peaks and associated local shocks would be expected to cause increasing wave and separation drag at transonic Mach number increases.
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.
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.
Preliminary study of a large span-distributed-load flying-wing cargo airplane concept
NASA Technical Reports Server (NTRS)
Jernell, L. S.
1978-01-01
An aircraft capable of transporting containerized cargo over intercontinental distances is analyzed. The specifications for payload weight, density, and dimensions in essence configure the wing and establish unusually low values of wing loading and aspect ratio. The structural weight comprises only about 18 percent of the design maximum gross weight. Although the geometric aspect ratio is 4.53, the winglet effect of the wing-tip-mounted vertical tails, increase the effective aspect ratio to approximately 7.9. Sufficient control power to handle the large rolling moment of inertia dictates a relatively high minimum approach velocity of 315 km/hr (170 knots). The airplane has acceptable spiral, Dutch roll, and roll-damping modes. A hardened stability augmentation system is required. The most significant noise source is that of the airframe. However, for both take-off and approach, the levels are below the FAR-36 limit of 108 db. The design mission fuel efficiency is approximately 50 percent greater than that of the most advanced, currently operational, large freighter aircraft. The direct operating cost is significantly lower than that of current freighters, the advantage increasing as fuel price increases.
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.
NASA Technical Reports Server (NTRS)
Snider, H. L.; Reeder, F. L.; Dirkin, W. J.
1972-01-01
Fourteen C-130 airplane center wings, each containing service-imposed fatigue damage resulting from 4000 to 13,000 accumulated flight hours, were tested to determine their fatigue crack propagation and static residual strength characteristics. Eight wings were subjected to a two-step constant amplitude fatigue test prior to static testing. Cracks up to 30 inches long were generated in these tests. Residual static strengths of these wings ranged from 56 to 87 percent of limit load. The remaining six wings containing cracks up to 4 inches long were statically tested as received from field service. Residual static strengths of these wings ranged from 98 to 117 percent of limit load. Damage-tolerant structural design features such as fastener holes, stringers, doublers around door cutouts, and spanwise panel splices proved to be effective in retarding crack propagation.
NASA Technical Reports Server (NTRS)
Muse, Thomas C.
1941-01-01
An investigation was made in the NACA 19-foot pressure wind tunnel to determine the effect of various win-gun installation on the aerodynamic characteristics of a model with an NACA low-drag wing. Measurements were made of lift and drag over an angle-of-attack range and for several values of dynamic pressure on a four-tenths scale model of a high-speed airplane equipped with the low-drag wing and with various wing-gun installations. Two installations were tested: one in which the blast tube and part of the gun barrel protrude ahead of the wing and another in which the guns is mounted wholly within the wing. Two types of openings for the latter installation were tested. For each installation three simulated guns were mounted in each wing. The results are given in the form of nondimensional coefficients. The installations tested appear to have little effect on the maximum-lift coefficient of the model. However, the drag coefficient shows a definite change. The least adverse effect was obtained with the completely internal mounting and small nose entrance. The results indicate that a properly designed wing-gun installation will have very little adverse effect on the aerodynamic characteristics of the low-drag wing.
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.
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)
Clousing, Lawrence A; Turner, William N; Rolls, L Stewart
1946-01-01
Pressure-distribution measurements were made on the right wing of a pursuit-type airplane at values of Mach number up to 0.80. The results showed that a considerable portion of the lift was carried by components of the airplane other than the wings, and that the proportion of lift carried by the wings may vary considerably with Mach number, thus changing the bending moment at the wing root whether or not there is a shift in the lateral position of the center of pressure. It was also shown that the center of pressure does not necessarily move outward at high Mach numbers, even though the wing-thickness ratio decreases toward the wing tip. The wing pitching-moment coefficient increased sharply in a negative direction at a Mach lift-curve slope increased with Mach number up to values of above the critical value. Pressures inside the wing were small and negative.
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
Tests of Aluminum-alloy Stiffened-sheet Specimens Cut from an Airplane Wing
NASA Technical Reports Server (NTRS)
Holt, Marshall
1943-01-01
The specimens used in the present tests were cut from an actual airplane wing of the stressed-skin type. The specimens thus obtained were not representative of the usual type of laboratory specimens because the stiffeners were not exactly parallel nor evenly spaced and, in one case, the skin consisted of pieces of sheet of different thicknesses. The test data obtained indicate that the buckling strain of stiffened curved sheet can be computed with reasonable accuracy by the equation given by Wenzek. The ultimate loads of the specimens when tested as flat sheet were within +/-11 percent of the product of the compressive yield strength and the cross-sectional area of the stiffeners. A rivet spacing equal to 98 times the sheet thickness was a source of weakness, and rivet spacings up to 36 times the sheet thickness appeared satisfactory.
An in-flight investigation of ground effect on a forward-swept wing airplane
NASA Technical Reports Server (NTRS)
Curry, Robert E.; Moulton, Bryan J.; Kresse, John
1989-01-01
A limited flight experiment was conducted to document the ground-effect characteristics of the X-29A research airplane. This vehicle has an aerodynamic platform which includes a forward-swept wing and close-coupled, variable incidence canard. The flight-test program obtained results for errors in the airdata measurement and for incremental normal force and pitching moment caused by ground effect. Correlations with wind-tunnel and computational analyses were made. The results are discussed with respect to the dynamic nature of the flight measurements, similar data from other configurations, and pilot comments. The ground-effect results are necessary to obtain an accurate interpretation of the vehicle's landing characteristics. The flight data can also be used in the development of many modern aircraft systems such as autoland and piloted simulations.
Upper-surface blowing nacelle design study for a swept wing airplane at cruise conditions
NASA Technical Reports Server (NTRS)
Gillette, W. B.; Mohn, L. W.; Ridley, H. G.; Nark, T. C.
1974-01-01
A study was made to design two types of overwing nacelles for an existing wing-body at a design condition of Mach = 0.8 and C sub L = 0.2. Internal and external surface contours were developed for nacelles having either a D-shaped nozzle or a high-aspect-ratio nozzle for upper-surface blowing in the powered-lift mode of operation. The goal of the design was the development of external nacelle lines that would minimize high-speed aerodynamic interference effects. Each nacelle type was designed for both two- and four-engine airplanes using an iterative process of aerodynamic potential flow analysis. Incremental nacelle drag estimates were made for flow-through wind tunnel models of each configuration.
NASA Technical Reports Server (NTRS)
Stough, H. Paul, III; Dicarlo, Daniel J.; Patton, James M., Jr.
1987-01-01
Flight tests were performed to investigate the change in stall/spin characteristics due to the addition of an outboard wing-leading-edge modification to a four-place, low-wing, single-engine, T-tail, general aviation research airplane. Stalls and attempted spins were performed for various weights, center of gravity positions, power settings, flap deflections, and landing-gear positions. Both stall behavior and wind resistance were improved compared with the baseline airplane. The latter would readily spin for all combinations of power settings, flap deflections, and aileron inputs, but the modified airplane did not spin at idle power or with flaps extended. With maximum power and flaps retracted, the modified airplane did enter spins with abused loadings or for certain combinations of maneuver and control input. The modified airplane tended to spin at a higher angle of attack than the baseline airplane.
Whirling Arm Tests on the Effect of Ground Proximity to an Airplane Wing
NASA Technical Reports Server (NTRS)
Long, M. E.
1944-01-01
This report gives the results of tests on a rectangular wing model with a 20% full spun split flap, conducted on the whirling arm at the Daniel Guggenheim Airship Institute in Akron, Ohio. The effect of a ground board on the lift and pitching moment was measured. The ground board consisted of an inclined ramp rising up in the test channel to a level floor extending for some distance parallel to the model path. The path of the wing model with respect to the ground board accordingly represented with comparative exactness an airplane coming in for a landing. The ground clearances over the level portion of the board varied from 0 6 to 1,6 chord lengths. Results are given in the standard dimensionless coefficients plotted versus angle of attack for a particular ground clearance. The effect of the ground board is to increase the lift coefficient for a given angle of attack all the way up the stall. The magnitude of the increase varies both with the ground clearance and the angle of attack. The effect on the pitching moment coefficient is not so readily apparent due to experimental difficulties but, in general, the diving moment increases over the ground board. This effect is apparent principally at the high angles of attack. An exception to this effect occurs with flaps deflected at the lowest ground clearance (0.6 chords). Here the diving moment decreases over the ground board.
NASA Technical Reports Server (NTRS)
Abeyounis, W. K.; Patterson, J. C., Jr.
1985-01-01
As part of a propulsion/airframe integration program, tests were conducted in the Langley 16-Foot Transonic Tunnel to determine the longitudinal aerodynamic effects of installing flow through engine nacelles in the aft underwing position of a high wing transonic transfer airplane. Mixed flow nacelles with circular and D-shaped inlets were tested at free stream Mach numbers from 0.70 to 0.85 and angles of attack from -2.5 deg to 4.0 deg. The aerodynamic effects of installing antishock bodies on the wing and nacelle upper surfaces as a means of attaching and supporting nacelles in an extreme aft position were investigated.
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.
Pressure distribution on wing ribs of the VE-7 and TS airplanes in flight Part II : pull-ups
NASA Technical Reports Server (NTRS)
Rhode, R V
1928-01-01
This paper is the second of a series of notes, each of which presents the complete results of pressure distribution tests made at Langley Field by the National Advisory Committee for Aeronautics, on wing and tail ribs of the VE-7 and TS airplanes for a particular maneuver of flight. The results for pull-ups are presented in the form of curves which show the variation of pressure distribution, total loads, normal acceleration and center of pressure with respect to time.
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.
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.
Geometric Model for a Parametric Study of the Blended-Wing-Body Airplane
NASA Technical Reports Server (NTRS)
Mastin, C. Wayne; Smith, Robert E.; Sadrehaghighi, Ideen; Wiese, Micharl R.
1996-01-01
A parametric model is presented for the blended-wing-body airplane, one concept being proposed for the next generation of large subsonic transports. The model is defined in terms of a small set of parameters which facilitates analysis and optimization during the conceptual design process. The model is generated from a preliminary CAD geometry. From this geometry, airfoil cross sections are cut at selected locations and fitted with analytic curves. The airfoils are then used as boundaries for surfaces defined as the solution of partial differential equations. Both the airfoil curves and the surfaces are generated with free parameters selected to give a good representation of the original geometry. The original surface is compared with the parametric model, and solutions of the Euler equations for compressible flow are computed for both geometries. The parametric model is a good approximation of the CAD model and the computed solutions are qualitatively similar. An optimal NURBS approximation is constructed and can be used by a CAD model for further refinement or modification of the original geometry.
NASA Technical Reports Server (NTRS)
Dicarlo, D. J.; Stough, H. P., III; Patton, J. M., Jr.
1980-01-01
Wind tunnel and flight tests were conducted to determine the effects of several discontinuous drooped wing leading-edge configurations on the spinning characteristics of a light, single-engine, low-wing research airplane. Particular emphasis was placed on the identification of modifications which would improve the spinning characteristics. The spanwise length of a discontinuous outboard droop was varied and several additional inboard segments were added to determine the influence of such leading-edge configurations on the spin behavior. Results of the study indicated that the use of only the discontinuous outboard droop, over a specific spanwise area, was most effective towards improving spin and spin recovery characteristics, whereas the segmented configurations having both inboard and outboard droop exhibited a tendency to enter a flat spin.
NASA Technical Reports Server (NTRS)
Murphy, A. C.
1981-01-01
Experimental data and correlative analytical results on the flutter and gust response characteristics of a torsion-free-wing (TFW) fighter airplane model are presented. TFW consists of a combined wing/boom/canard surface and was tested with the TFW free to pivot in pitch and with the TFW locked to the fuselage. Flutter and gust response characteristics were measured in the Langley Transonic Dynamics Tunnel with the complete airplane model mounted on a cable mount system that provided a near free flying condition. Although the lowest flutter dynamic pressure was measured for the wing free configuration, it was only about 20 deg less than that for the wing locked configuration. However, no appreciable alleviation of the gust response was measured by freeing the wing.
Distribution of Pressure Over Model of the Upper Wing and Aileron of a Fokker D-VII Airplane
NASA Technical Reports Server (NTRS)
Fairbanks, A J
1927-01-01
This report describes tests made for the purpose of determining the distribution of pressure over a model of the tapered portion of the upper wing and the aileron of a Fokker D-VII Airplane. Normal pressures were measured simultaneously at 74 points distributed over the wing and aileron. Tests were made throughout the useful range of angles of attack with aileron setting ranging from -20 degrees to +20 degrees. The results are presented graphically. It was found that the pressure distribution along the chord is in general similar to that of thick tapered airfoils previously tested. The maximum resultant pressure recorded was five times the dynamic pressure. The distribution of the air load along the span may be assumed to be uniform for design purposes. Aileron displacements affect the pressures forward to the leading edge of the wing and may increase the air load on the outer portion of the wing by a considerable amount. With the wing at large angles of attack, the overhanging portion of the aileron creates usually a burble flow and therefore a large drag. The balance reduces the control stick forces at small angles of attack for all aileron displacements. At large angles of attack it does this for small displacements only. With the airplane at its maximum speed, an angle of attack of 18 degrees, and a down aileron displacement of 20 degrees, the bending moment tending to break off the overhanging portion of the aileron will be greater than that caused by a uniform static load of 35 pounds per square foot.
Pressure Distribution over a Wing and Tail Rib of a VE-7 and of a TS Airplane in Flight
NASA Technical Reports Server (NTRS)
Crowley, J W , Jr
1928-01-01
This investigation was made to determine the pressure distribution over a rib of the wing and over a rib of the horizontal tail surface of an airplane in flight and to obtain information as to the time correlation of the loads occurring on these ribs. Two airplanes, VE-7 and TS, were selected in order to obtain the information for a thin and a thick wing section. In each case the pressure distribution was recorded for the full range of angle of attack in level flight and throughout violent maneuvers. The results show: (a) that the present rib load specifications in use by the Army Air Corps and the Bureau of Aeronautics, Navy Department, are in fair agreement with the loads actually occurring in flight, but could be slightly improved; (b) that there appears to be no definite sequence in which wing and tail surface ribs reach their respective maximum loads in different maneuvers; (c) that in accelerated flight, at air speeds less than or equal to 60 per cent of the maximum speed, the accelerations measured agree very closely with the theoretically possible maximum accelerations. In maneuvers at higher air speeds the observed accelerations were smaller than those theoretically possible. (author)
NASA Technical Reports Server (NTRS)
Ayers, T. G.
1973-01-01
An investigation was conducted in the Langley 8 foot transonic pressure tunnel and the Langley Unitary Plan wind tunnel to evaluate the effectiveness of three variations of the NASA supercritical airfoil as applied to a model of a variable wing sweep fighter airplane. Wing panels incorporating conventional NACA 64A series airfoil with 0.20 and 0.40 camber were used as bases of reference for this evaluation. Static force and moment measurements were obtained for wing leading edge sweep angles of 26, 33, 39, and 72.5 degrees. Fluctuating wing root bending moment data were obtained at subsonic speeds to determine buffet characteristics. Subsonic data were also obtained for determining the effects of wing transition location and spoiler deflection. Limited lateral directional data are included for the conventional 0.20 cambered wing and the supercritical wing.
NASA Technical Reports Server (NTRS)
Lovell, Powell M , Jr; Parlett, Lysle P
1957-01-01
An investigation of the stability and control of a high-wing transport vertical-take-off airplane with four engines during constant-altitude transitions from hovering to normal forward flight was conducted with a remotely controlled free-flight model. The model had four propellers distributed along the wing with the thrust axes in the wing chord plane. The wing could be rotated to 90 degrees incidence so that the propeller thrust axes were vertical for hovering flight. An air jet at the rear of the fuselage provided pitch and yaw control for hovering and low-speed flight.
NASA Technical Reports Server (NTRS)
Abeyounis, William K.; Patterson, James C., Jr.
1990-01-01
As part of a propulsion/airframe integration program at Langley Research Center, tests were conducted in the Langley 16-Foot Transonic Tunnel to determine the effects of locating flow-through mixed flow engine nacelles in several aft underwing positions on the longitudinal aerodynamics of a high wing transport airplane. D-shaped inlet nacelles were used in the test. Some configurations with antishock bodies and with nacelle toe-in were also tested. Data were obtained for a free stream Mach number range of 0.70 to 0.85 and a model angle-of-attack range from -2.5 to 4.0 degrees.
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.
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)
Pyle, J. S.; Steers, L. L.
1975-01-01
Flight measurements obtained with a TF-8A airplane modified with a supercritical wing are presented for altitudes from 7.6 kilometers (25,000 feet) to 13.7 kilometers (45,000 feet), Mach numbers from 0.6 to 1.2, and Reynolds numbers from 0.8 x 10 to the 7th power to 2.3 x 10 to the 7th power. Flight results for the airplane with and without area-rule fuselage fairings are compared. The techniques used to determine the lift and drag characteristics of the airplane are discussed. Flight data are compared with wind-tunnel model results, where applicable.
NASA Technical Reports Server (NTRS)
Murri, Daniel G.; Jordan, Frank L., Jr.
1987-01-01
An investigation was conducted in the Langley 30- by 60-Foot Wind Tunnel to evaluate the performance, stability, and control characteristics of a full-scale general aviation airplane equipped with an advanced laminar flow wing. The study focused on the effects of natural laminar flow and advanced boundary layer transition on performance, stability, and control, and also on the effects of several wing leading edge modifications on the stall/departure resistance of the configuration. Data were measured over an angle-of-attack range from -6 to 40 deg and an angle-of-sideslip range from -6 to 20 deg. The Reynolds number was varied from 1.4 to 2.4 x 10 to the 6th power based on the mean aerodynamic chord. Additional measurements were made using hot-film and sublimating chemical techniques to determine the condition of the wing boundary layer, and wool tufts were used to study the wing stall characteristics. The investigation showed that large regions of natural laminar flow existed on the wing which would significantly enhance cruise performance. Also, because of the characteristics of the airfoil section, artificially tripping the wing boundary layer to a turbulent condition did not significantly effect the lift, stability, and control characteristics. The addition of a leading-edge droop arrangement was found to increase the stall angle of attack at the wingtips and, therefore, was considered to be effective in improving the stall/departure resistance of the configuration. Also the addition of the droop arrangement resulted in only minor increases in drag.
NASA Technical Reports Server (NTRS)
Mitcham, Grady L; Stevens, Joseph E; Norris, Harry P
1956-01-01
A flight investigation of rocket-powered models of a tailless triangular-wing airplane configuration was made through the transonic and low supersonic speed range at the Langley Pilotless Aircraft Research Station at Wallops Island, Va. An analysis of the aerodynamic coefficients, stability derivatives, and flying qualities based on the results obtained from the successful flight tests of three models is presented.
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)
Riley, D. R.
1985-01-01
A six-degree-of-freedom nonlinear simulation was developed for a two-place, single-engine, low-wing general aviation airplane for the stall and initial departure regions of flight. Two configurations, one with and one without an outboard wing-leading-edge modification, were modeled. The math models developed are presented simulation predictions and flight-test data for validation purposes and simulation results for the two configurations for various maneuvers and power settings are compared to show the beneficial influence of adding the wing-leading-edge modification.
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.
An analytical comparison of two wing structures for Mach 5 cruise airplanes
NASA Technical Reports Server (NTRS)
Taylor, A. H.; Jackson, L. R.; Cerro, J. A.; Scotti, S. J.
1983-01-01
Mach 5 cruise research conducted by NASA is related to aerodynamics, propulsion, and structures. The study of structures includes the propulsion system, fuselage, and wings. Various studies have shown that the achievement of adequate range is largely dependent on a low structural mass fraction. The present investigation is concerned with a study of two wing structure configurations for Mach 5 aircraft. An uprated version (Ti-6242 replacing B-120 titanium) of the YF-12/SR-71 wing structure is considered. The B-120 titanium structure represents the current art of high speed aircraft wing structures. The YF-12 wing structure was designed about 20 years ago when the analytical methods for calculating thermal stresses were limited. The second wing structural configuration studied in the present investigation also used Ti-6242 materials but replaced the corrugated-beaded panels with diffusion bonded honeycomb-core sandwich panels, and replaced the z-stiffened shear webs with sine-wave stiffened shear webs.
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.
Design of a Large Span-Distributed Load Flying-Wing Cargo Airplane
NASA Technical Reports Server (NTRS)
Jernell, L. S.; Quartero, C. B.
1977-01-01
The design and operation of very large, long-range, subsonic cargo aircraft are considered. A design concept which distributes the payload along the wingspan to counterbalance the aerodynamic loads, with a resultant decrease in the in-flight wing bending moments and shear forces, is described. The decreased loading of the wing structure, coupled with the very thick wing housing the cargo, results in a relatively low overall structural weight in comparison to that of conventional aircraft.
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.
NASA Astrophysics Data System (ADS)
Staelens, Yann Daniel
During the first century of flight few major changes have been made to the configuration of subsonic airplanes. A distinct fuselage with wings, a tail, engines and a landing gear persists as the dominant arrangement. During WWII some companies developed tailless all-wing airplanes. However the concept failed to advance till the late 80's when the B-2, the only flying wing to enter production to date, illustrated its benefits at least for a stealth platform. The advent of the Blended-Wing-Body (BWB) addresses the historical shortcomings of all-wing designs, specifically poor volume utility and excess wetted area as a result. The BWB is now poised to become the new standard for large subsonic airplanes. Major aerospace companies are studying the concept for next generation of passenger airplanes. But there are still challenges. One is the BWB's short control lever-arm pitch. This affects rotation and go-around performances. This study presents a possible solution by using a novel type of control surface, a belly-flap, on the under side of the wing to enhance its lift and pitching moment coefficient during landing, go-around and takeoff. Increases of up to 30% in lift-off CL and 8% in positive pitching moment have been achieved during wind tunnel tests on a generic BWB-model with a belly-flap. These aerodynamic improvements when used in a mathematical simulation of landing, go-around and takeoff procedure were showing reduction in landing-field-length by up to 22%, in takeoff-field-length by up to 8% and in loss in altitude between initiation of rotation and actual rotation during go-around by up to 21.5%.
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
Crash tests of four low-wing twin-engine airplanes with truss-reinforced fuselage structure
NASA Technical Reports Server (NTRS)
Williams, M. S.; Fasanella, E. L.
1982-01-01
Four six-place, low-wing, twin-engine, general aviation airplane test specimens were crash tested under controlled free flight conditions. All airplanes were impacted on a concrete test surface at a nomial flight path velocity of 27 m/sec. Two tests were conducted at a -15 deg flight path angle (0 deg pitch angle and 15 deg pitch angle), and two were conducted at a -30 deg flight path angle (-30 deg pitch angle). The average acceleration time histories (crash pulses) in the cabin area for each principal direction were calculated for each crash test. In addition, the peak floor accelerations were calculated for each test as a function of aircraft fuselage longitudinal station number. Anthropomorphic dummy accelerations were analyzed using the dynamic response index and severity index (SI) models. Parameters affecting the dummy restraint system were studied; these parameters included the effect of no upper torso restraint, measurement of the amount of inertia-reel strap pullout before locking, measurement of dummy chest forward motion, and loads in the restraints. With the SI model, the dummies with no shoulder harness received head impacts above the concussive threshold.
General problem of the airplane
NASA Technical Reports Server (NTRS)
Richard, Maurice; Richard, Paul
1922-01-01
A series of equations relating to airplanes are given and examples listed. Some of the equations listed include: the speed, altitude and carrying capacity of various airplanes; weight of an airplane; weight of various parts of an airplane; the polars of the wings; speeds of airplanes; radius of action.
NASA Technical Reports Server (NTRS)
Brewer, Gerald W.
1946-01-01
An investigation of a 1/7-scale powered model of the Kaiser Fleetwing all-wing airplane was made in the Langley full-scale tunnel to provide data for an estimation of the flying qualities of the airplane. The analysis of the stability and control characteristics of the airplane has been made as closely as possible in accordance with the requirements of the Bureau of Aeronautics, Navy Department's specifications, and a summary of the more significant conclusions is presented as follows. With the normal center of gravity located at 20 percent of the mean aerodynamic chord, the airplane will have adequate static longitudinal stability, elevator fixed, for all flight conditions except for low-power operation at low speeds where the stability will be about neutral. There will not be sufficient down-elevator deflection available for trim above speeds of about 130 miles per hour. It is probable that the reduction in the up-elevator deflections required for trim will be accompanied by reduced elevator hinge moments for low-power operation at low flight speeds. The static directional stability for this airplane will be low for all rudder-fixed or rudder-free flight conditions. The maximum rudder deflection of 30 deg will trim only about 15 deg yaw for most flight conditions and only 10 deg yaw for the condition with low power at low speeds. Also, at low powers and low speeds, it is estimated that the rudders will not trim the total adverse yaw resulting from an abrupt aileron roll using maximum aileron deflection. The airplane will meet the requirements for stability and control for asymmetric power operation with one outboard engine inoperative. The airplane would have no tendency for directional divergence but would probably be spirally unstable, with rudders fixed. The static lateral stability of the airplane will probably be about neutral for the high-speed flight conditions and will be only slightly increased for the low-power operation in low-speed flight. The
NASA Astrophysics Data System (ADS)
Rabenstein, L.; Lobach, J.; Haas, C.
2007-12-01
Regular observation of Arctic and Antarctic sea ice thickness is of high importance for a better understanding of processes of climate change in polar regions. For regular and accurate observations of polar sea ice thickness a long range airborne device is necessary. Airborne electromagnetic induction (AEM) sounding was found to be an ideal method for accurate and wide area sea ice thickness measurements. As a consequence of five years of successful helicopter electromagnetic (HEM) sea ice thickness measurements and to overcome helicopter range restrictions, the Alfred Wegener Institute (AWI) constructed a new airplane based fixed wing EM system. The first test flights were carried out in 2006 over the North Sea and in April 2007 in Svalbard, where the system's performance was proven under arctic conditions. The system operates in frequency domain with 1990 Hz and a vertical coplanar coil configuration. Thus the system produces a horizontal dipole. The coils are mounted beneath the wings with a separation of 11.6 meters. The airplane, a Dornier 228, is also equipped with a laser altimeter to determine the altitude of the instrument with an accuracy of 2cm. The compensation of the transmitter signal at the receiver coil is done electronically. Flights over open sea are used for the calibration of the system, because the ocean functions as a homogeneous half space with well known conductivity. A data acquisition computer records four voltages with a sample rate of 10 Hz. These are the reference voltage of the transmitter, the compensated and raw receiver voltages and the compensation signal. The laser height is recorded with a sample rate of 100 Hz to account for surface roughness. EM instruments for sea ice thickness sounding should have a vertical resolution of 10cm but due to the electrical noise caused by the airplane engines this was not easy to achieve. To account for the noise a time average filter is used. Alternatively, in order to keep the original
Flight tests of a radio-controlled airplane mode with a free-wing, free-canard configuration
NASA Technical Reports Server (NTRS)
Gee, S. W.
1978-01-01
Flight characteristics, controllability, and potential operating problems were investigated in a radio-controlled airplane model in which the wing is so attached to the fuselage that it is free to pivot about a spanwise axis forward of its aerodynamic center and is subject only to aerodynamic pitching moments imposed by lift and drag forces and a control surface. A simple technique of flying the test vehicle in formation with a pickup truck was used to obtain trim data. The test vehicle was flown through a series of maneuvers designed to permit evaluation of certain characteristics by observation. The free-wing free-canard concept was determined to be workable. Stall/spin characteristics were considered to be excellent, and no effect on longitudinal stability was observed when center of gravity changes were made. Several problems were encountered during the early stages of flight testing, such as aerodynamic lockup of the free canard and excessive control sensitivity. Lack of onboard instrumentation precluded any conclusions about gust alleviation or ride qualities.
NASA Technical Reports Server (NTRS)
Hunton, Lynn W.; Dew, Joseph K.
1948-01-01
Wind-tunnel tests of a full-scale model of the Republic XP-91 airplane were conducted to determine the longitudinal and lateral characteristics of the wing alone and the wing-fuselage combination, the characteristics of the aileron, and the damping in roll af the wing alone. Various high-lift devices were investigated including trailing-edge split flaps and partial- and full-span leading-edge slats and Krueger-type nose flaps. Results of this investigation showed that a very significant gain in maximum lift could be achieved through use of the proper leading-edge device, The maximum lift coefficient of the model with split flaps and the original partial-span straight slats was only 1.2; whereas a value of approximately 1.8 was obtained by drooping the slat and extending it full span, Improvement in maximum lift of approximately the same amount resulted when a full-span nose flap was substituted for the original partial-span slat.
NASA Technical Reports Server (NTRS)
Hallissy, J. B.; Harris, C. D.
1974-01-01
Wind-tunnel tests have been conducted at Mach numbers of 0.85, 0.88, and 0.90 to determine the aerodynamic load distribution for the 39 deg swept-wing configuration of a variable-wing-sweep fighter airplane with a NASA supercritical airfoil. Chordwise pressure distributions were measured at two wing stations. Also measured were the overall longitudinal aerodynamic force and moment characteristics and the buffet characteristics. The analysis indicates that localized regions of shock-induced flow separation may exist on the rearward portions of the supercritical wing at high subsonic speeds, and caution must be exercised in the prediction of buffet onset when using variations in trailing-edge pressure coefficients at isolated locations.
NASA Technical Reports Server (NTRS)
Kramer, Max
1932-01-01
Wind-tunnel tests are described, in which the angle of attack of a wing model was suddenly increased (producing the effect of a vertical gust) and the resulting forces were measured. It was found that the maximum lift coefficient increases in proportion to the rate of increase in the angle of attack. This fact is important for the determination of the gust stresses of airplanes with low wing loading. The results of the calculation of the corrective factor are given for a high-performance glider and a light sport plane of conventional type.
NASA Technical Reports Server (NTRS)
1979-01-01
Configurations with full-span and segmented leading-edge flaps and full-span and segmented leading-edge droop were tested. Studies were conducted with wind-tunnel models, with an outdoor radio-controlled model, and with a full-scale airplane. Results show that wing-leading-edge modifications can produce large effects on stall/spin characteristics, particularly on spin resistance. One outboard wing-leading-edge modification tested significantly improved lateral stability at stall, spin resistance, and developed spin characteristics.
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.
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 Â§ 25.629 Aeroelastic stability requirements. (a)...
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.
NASA Technical Reports Server (NTRS)
Ruhlin, C. L.; Murphy, A. C.
1981-01-01
An exploratory study of a 1/5.5 size, complete airplane version of a torsion free wing (TFW) fighter aircraft was conducted. The TFW consisted of a wing/boom/canard assembly on each fuselage side that was interconnected by a common pivot shaft so that the TFW could rotate freely in pitch. The effect of the TFW was evaluated by comparing data obtained with the TFW free and the TFW locked to the fuselage. With the model mounted on cables to simulate an airplane free flying condition, flutter boundaries were measured at Mach number (M) from 0.85 to 1.0 and gust responses at M = 0.65 and 0.90. The critical flutter mode for the TFW free configuration was found experimentally to occur at M = 0.95 and had the rigid TFW pitch mode as its apparent aerodynamic driver.
NASA Technical Reports Server (NTRS)
Kelly, Mark W; Anderson, Seth B; Innis, Robert C
1958-01-01
A wind-tunnel investigation was made to determine the effects on the aerodynamic characteristics of a 35 degree swept-wing airplane of applying blowing-type boundary-layer control to the trailing-edge flaps. Flight tests of a similar airplane were then conducted to determine the effects of boundary-layer control on the handling qualities and operation of the airplane, particularly during landing and take-off. The wind-tunnel and flight tests indicated that blowing over the flaps produced large increases in flap lift increment, and significant increases in maximum lift. The use of blowing permitted reductions in the landing approach speeds of as much as 12 knots.
NASA Technical Reports Server (NTRS)
Brandon, Jay M.; Brown, Philip W.; Wunschel, Alfred J.
1987-01-01
A piloted-simulation study was conducted to investigate the effects of vortex flaps on low-speed handling qualities of a delta-wing airplane. The simulation math model was developed from wind tunnel tests of a 0.15 scale model of the F-106B airplane. Pilot evaluations were conducted using a six-degree-of-freedom motion base simulator. The results of the investigation showed that the reduced static longitudinal stability caused by the vortex flaps significantly degraded handling qualities in the approach-to-landing task. Acceptable handling qualities could be achieved by limiting the aft center-of-gravity location, consequently reducing the operational envelope of the airplane. Further improvement were possible by modifying the flight control force-feel system to reduce pitch-control sensitivity.
NASA Technical Reports Server (NTRS)
Stough, H. Paul, III; Patton, James M., Jr.; Sliwa, Steven M.
1987-01-01
Flight tests were performed to investigate the stall, spin, and recovery characteristics of a low-wing, single-engine, light airplane with four interchangeable tail configurations. The four tail configurations were evaluated for effects of varying mass distribution, center-of-gravity position, and control inputs. The airplane tended to roll-off at the stall. Variations in tail configuration produced spins ranging from 40 deg to 60 deg angle of attack and turn rates of about 145 to 208 deg/sec. Some unrecoverable flat spins were encountered which required use of the airplane spin chute for recovery. For recoverable spins, antispin rudder followed by forward wheel with ailerons centered provided the quickest spin recovery. The moderate spin modes agreed very well with those predicted from spin-tunnel model tests, however, the flat spin was at a lower angle of attack and a slower rotation rate than indicated by the model tests.
NASA Technical Reports Server (NTRS)
Cook, Woodrow L; Anderson, Seth B; Cooper, George E
1958-01-01
A wind-tunnel investigation was made to determine the effects on the aerodynamic characteristics of a 35 degree swept-wing airplane of applying area-suction boundary-layer control to the trailing-edge flaps. Flight tests of a similar airplane were then conducted to determine the effect of boundary-layer control in the handling qualities and operation of the airplane, particularly during landing. The wind-tunnel and flight tests indicated that area suction applied to the trailing-edge flaps produced significant increases in flap lift increment. Although the flap boundary-layer control reduced the stall speed only slightly, a reduction in minimum comfortable approach speed of about 12 knots was obtained.
NASA Technical Reports Server (NTRS)
Cole, Henry A , Jr; Brown, Stuart C; Holleman, Euclid C
1957-01-01
Measured and predicted dynamic response characteristics of a large flexible swept-wing airplane to control surface inputs are presented for flight conditions of 0.6 to 0.85 Mach number at an altitude of 35,000 feet. The report is divided into two parts. The first part deals with the response of the airplane to elevator control inputs with principal responses contained in a band of frequencies including the longitudinal short-period mode and several symmetrical structural modes. The second part deals with the response of the airplane to aileron and rudder control inputs with principal responses contained in a band of frequencies including the dutch roll mode, the rolling mode, and three antisymmetrical structural modes.
NASA Technical Reports Server (NTRS)
Barnes, Robert H.
1947-01-01
The tests reported herein were made for the purpose of determining the high-speed load distribution on the wing of a 3/16 scale model of a scout-bomber airplane. Comparisons are made between the root bending-moment and section torsional-moment coefficients as obtained experimentally and derived analytically. The results show good correlation for the bending-moment coefficients but considerable disagreement for the torsional-moment coefficients.
NASA Technical Reports Server (NTRS)
Sadoff, Melvin; Matteson, Frederick H.; Van Dyke, Rudolph D., Jr.
1954-01-01
An investigation was conducted on a 35 deg swept-wing fighter airplane to determine the effects of several blunt-trailing-edge modifications to the wing and tail on the high-speed stability and control characteristics and tracking performance. The results indicated significant improvement in the pitch-up characteristics for the blunt-aileron configuration at Mach numbers around 0.90. As a result of increased effectiveness of the blunt-trailing-edge aileron, the roll-off, customarily experienced with the unmodified airplane in wings-level flight between Mach numbers of about 0.9 and 1.0 was eliminated, The results also indicated that the increased effectiveness of the blunt aileron more than offset the large associated aileron hinge moment, resulting in significant improvement in the rolling performance at Mach numbers between 0.85 and 1.0. It appeared from these results that the tracking performance with the blunt-aileron configuration in the pitch-up and buffeting flight region at high Mach numbers was considerably improved over that of the unmodified airplane; however, the tracking errors of 8 to 15 mils were definitely unsatisfactory. A drag increment of about O.OOl5 due to the blunt ailerons was noted at Mach numbers to about 0.85. The drag increment was 0 at Mach numbers above 0.90.
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.
Free-spinning-tunnel Investigation of a 1/30 Scale Model of a Twin-jet-swept-wing Fighter Airplane
NASA Technical Reports Server (NTRS)
Bowman, James S., Jr.; Healy, Frederick M.
1960-01-01
An investigation has been made in the Langley 20-foot free-spinning tunnel to determine the erect and inverted spin and recovery characteristics of a 1/30-scale dynamic model of a twin-jet swept-wing fighter airplane. The model results indicate that the optimum erect spin recovery technique determined (simultaneous rudder reversal to full against the spin and aileron deflection to full with the spin) will provide satisfactory recovery from steep-type spins obtained on the airplane. It is considered that the air-plane will not readily enter flat-type spins, also indicated as possible by the model tests, but developed-spin conditions should be avoided in as much as the optimum recovery procedure may not provide satisfactory recovery if the airplane encounters a flat-type developed spin. Satisfactory recovery from inverted spins will be obtained on the airplane by neutralization of all controls. A 30-foot- diameter (laid-out-flat) stable tail parachute having a drag coefficient of 0.67 and a towline length of 27.5 feet will be satisfactory for emergency spin recovery.
NASA Technical Reports Server (NTRS)
Goodman, Harold R.
1949-01-01
Measurements were made, in dives to transonic speeds, of the static-pressure position error at a distance of one chord ahead of the McDonnell XF-88 airplane. The airplane incorporates a wing which is swept back 35 deg along the 0.22 chord line and utilizes a 65-series airfoil with a 9-percent-thick section perpendicular to the 0.25-chord line. The section in the stream direction is approximately 8-percent thick. Data up to a Mach number of about 0.97 were obtained within an airplane normal-force-coefficient range from about 0.05 to about 0.68. Data at Mach numbers above about 0.97 were obtained within an airplane normal-force-coefficient range from about 0.05 to about 0.68. Results of the measurements indicate that the static-pressure error, within the accuracy of measurement, is negligible from a Mach number of 0.65 to a Mach number of about 0.97. With a further increase in Mach number, the static-pressure error increases rapidly; at the highest Mach number attained in these tests (about M = 1.038), the error increases to about 8 percent of the impact pressure. Above a Mach number of about 0.975, the recorded Mach number remains substantially constant with increasing true Mach number; the installation is of no value between a Mach number of about 0.975 and at least 1.038, as the true Mach number cannot be obtained from the recorded Mach number in this range. Previously published data have shown that at 0.96 chord ahead of the wing tip of the straight-wing X-l airplanes, a rapid rise of position error started at a Mach number of about 0.8. In the case of the XF-88 airplane, this rise of position error was delayed, presumably by the sweep of the wing, to a Mach number of about 0.97.
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.
NASA Technical Reports Server (NTRS)
Rainey, A Gerald; Igoe, William B
1958-01-01
The buffeting loads acting on the wing and horizontal tail of a 1/4-scale model of the X-1E airplane have been measured in the Langley 16-foot transonic tunnel in the Mach number range from 0.40 to 0.90. When the buffeting loads were reduced to a nondimensional aerodynamic coefficient of buffeting intensity, it was found that the maximum buffeting intensity of the horizontal tail was about twice as large as that of the wing. Comparison of power spectra of buffeting loads acting on the horizontal tail of the airplaneand of the model indicated that the model horizontal tail, which was of conventional force-test-model design, responded in an entirely different mode than did the airplane.This result implied that if quantitative extrapolation of model data to flight conditions were desired a dynamically scaled model of the rearward portion of the fuselage and empennage would be required. A study of the sources of horizontal-tail buffeting of the model indicated that the wing wake contributed a large part of the total buffeting load. At one condition it was found that removal of the wing wake would reduce the buffeting loads on the horizontal tail to about one-third of the original value.
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.
NASA Technical Reports Server (NTRS)
Fisher, Lewis R.; Williams, James L.
1958-01-01
A research model of an airplane with a configuration suitable for supersonic flight was tested at transonic speeds in order to establish the effects on longitudinal and lateral stability of certain changes in both wing sweep and height of the horizontal tail. Two wings of aspect ratio 3 and taper ratio 0.15, one having the quarter-chord line swept back 30 deg and the other 45 deg, were each tested with the horizontal tail of the model in a low and in a high position. One configuration was also tested with fuselage strakes. The tests were made at Mach numbers from 0.60 to 1.17 and Reynolds numbers from 1.9 x 10(exp 6) to 2.6 x 10(exp 6). The results indicated that a low horizontal-tail position (below the wing-chord plane) gave positive longitudinal stability for the model for all angles of attack used (angles of attack up to 24 deg); whereas, a higher tail position (above the wing-chord plane) resulted in a large reduction in stability at moderate angles of attack. With the higher horizontal tail, the 30 deg-swept-wing model had somewhat more stability than the 45 deg-swept-wing model at subsonic Mach numbers. With the lower tail, the 45 deg-swept-wing model had slightly more stability at all Mach numbers. The model with the 30 deg swept wing had greater directional stability with the tail in the higher rather than the lower position, but the opposite was true for the 45 deg-swept-wing model. The directional stability decreased sharply at high angles of attack; this characteristic was alleviated by the use of fuselage strakes which, however, proved to be detrimental to the longitudinal stability of the model tested.
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.
Wave formation on a liquid layer for de-icing airplane wings
NASA Technical Reports Server (NTRS)
Yih, Chia-Shun
1990-01-01
Wave formation on a thin liquid layer used for deicing aircraft wings is investigated by studying the stability of airflow over a liquid-coated flat plate at zero angle of incidence. The ratio of the viscosity of the liquid to that of air is very high (over 500,000), and the Reynolds number based on liquid depth and air viscosity is of the order of a few thousand. Under these circumstances the analysis gives two formulas, in closed form, for the growth rate and phase velocity of the waves in terms of the wavenumber and other relevant parameters. The expected wavenumber is that for which the growth rate is the maximum. The instability is one in which the viscosity difference between the two fluids (air and liquid) plays the dominant role, and is of the kind found by Yih (1967).
NASA Technical Reports Server (NTRS)
Thompson, Jim Rogers; Bray, Richard S; COOPER GEORGE E
1950-01-01
The calibrations of four airspeed systems installed in a North American F-86A airplane have been determined in flight at Mach numbers up to 1.04 by the NACA radar-phototheodolite method. The variation of the static-pressure error per unit indicated impact pressure is presented for three systems typical of those currently in use in flight research, a nose boom and two different wing-tip booms, and for the standard service system installed in the airplane. A limited amount of information on the effect of airplane normal-force coefficient on the static-pressure error is included. The results are compared with available theory and with results from wind-tunnel tests of the airspeed heads alone. Of the systems investigated, a nose-boom installation was found to be most suitable for research use at transonic and low supersonic speeds because it provided the greatest sensitivity of the indicated Mach number to a unit change in true Mach number at very high subsonic speeds, and because it was least sensitive to changes in airplane normal-force coefficient. The static-pressure error of the nose-boom system was small and constant above a Mach number of 1.03 after passage of the fuselage bow shock wave over the airspeed head.
NASA Technical Reports Server (NTRS)
Ziff, Howard L; Rathert, George A; Gadeberg, Burnett L
1953-01-01
Standard air-to-air-gunnery tracking runs were conducted with F-51H, F8F-1, F-86A, and F-86E airplanes equipped with fixed gunsights. The tracking performances were documented over the normal operating range of altitude, Mach number, and normal acceleration factor for each airplane. The sources of error were studied by statistical analyses of the aim wander.
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.
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.
NASA Technical Reports Server (NTRS)
Johnson, J. L., Jr.; Newsom, W. A.; Satran, D. R.
1980-01-01
The paper presents the results of a recent investigation to determine the effects of wing leading-edge modifications on the high angle-of-attack aerodynamic characteristics of a low-wing general aviation airplane in the Langley Full-Scale Wind Tunnel. The investigation was conducted to provide aerodynamic information for correlation and analysis of flight-test results obtained for the configuration. The wind-tunnel investigation consisted of force and moment measurements, wing pressure measurements, flow surveys, and flow visualization studies utilizing a tuft grid, smoke and nonintrusive mini-tufts which were illuminated by ultra-violet light. In addition to the tunnel scale system which measured overall forces and moments, the model was equipped with an auxiliary strain-gage balance within the left wing panel to measure lift and drag forces on the outer wing panel independent of the tunnel scale system. The leading-edge modifications studied included partial- and full-span leading-edge droop arrangements as well as leading-edge slats.
NASA Technical Reports Server (NTRS)
Kuhn, Richard E.
1959-01-01
The analysis presented uses the momentum theory as a starting point in developing semiempirical expressions for calculating the effect of propeller thrust and slipstream on the lift and drag characteristics of wing-flap configurations that would be suitable for vertical-take-off-and-landing (VTOL) and short-take-off-and-landing (STOL) airplanes. The method uses power-off forward-speed information and measured slipstream deflection data at zero forward speed to provide a basis for estimating the lift and drag at combined forward speed and power-on conditions. A correlation of slipstream deflection data is also included. The procedure is applicable only in the unstalled flight regime; nevertheless, it should be useful in preliminary design estimates of the performance that may be expected of VTOL and STOL airplanes.
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.
Study of the Mutual Interaction Between a Wing Wake and an Encountering Airplane
NASA Technical Reports Server (NTRS)
Walden, A. B.; vanDam, C. P.
1996-01-01
In an effort to increase airport productivity, several wind-tunnel and flight-test programs are currently underway to determine safe reductions in separation standards between aircraft. These programs are designed to study numerous concepts from the characteristics and detection of wake vortices to the wake-vortex encounter phenomenon. As part of this latter effort, computational tools are being developed and utilized as a means of modeling and verifying wake-vortex hazard encounters. The objective of this study is to assess the ability of PMARC, a low-order potential-flow panel method, to predict the forces and moments imposed on a following business-jet configuration by a vortex interaction. Other issues addressed include the investigation of several wake models and their ability to predict wake shape and trajectory, the validity of the velocity field imposed on the following configuration, modeling techniques and the effect of the high-lift system and the empennage. Comparisons with wind-tunnel data reveal that PMARC predicts the characteristics for the clean wing-body following configuration fairly well. Non-linear effects produced by the addition of the high-lift system and empennage, however, are not so well predicted.
NASA Technical Reports Server (NTRS)
Bird, John D; Jaquet, Byron M
1951-01-01
An investigation was made to determine the accuracy with which the lateral flight motions of a swept-wing airplane could be predicted from experimental stability derivatives, determined in the 6-foot-diameter rolling-flow test section and 6 by 6-foot curved-flow test section of the Langley stability tunnel. In addition, determination of the significance of including the nonlinear aerodynamic effects of sideslip in the calculations of the motions was desired. All experimental aerodynamic data necessary for prediction of the lateral flight motions are presented along with a number of comparisons between flight and calculated motions caused by rudder and aileron disturbances.
NASA Technical Reports Server (NTRS)
Hallissy, James B.; Phillips, Pamela S.
1989-01-01
A wind tunnel test was conducted to evaluate the aerodynamic characteristics and wing pressure distributions of a variable wing sweep aircraft having wing panels that are modified to promote laminar flow. The modified wing section shapes were incorporated over most of the exposed outer wing panel span and were obtained by extending the leading edge and adding thickness to the existing wing upper surface forward of 60 percent chord. Two different wing configurations, one each for Mach numbers 0.7 and 0.8, were tested on the model simultaneously, with one wing configuration on the left side and the other on the right. The tests were conducted at Mach numbers 0.20 to 0.90 for wing sweep angles of 20, 25, 30, and 35 degrees. Longitudinal, lateral and directional aerodynamic characteristics of the modified and baseline configurations, and selected pressure distributions for the modified configurations, are presented in graphical form without analysis. A tabulation of the pressure data for the modified configuration is available as microfiche.
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.
Flight-determined stability and control coefficients of the F-111A airplane
NASA Technical Reports Server (NTRS)
Iliff, K. W.; Maine, R. E.; Steers, S. T.
1978-01-01
A complete set of linear stability and control derivatives of the F-111A airplane was determined with a modified maximum likelihood estimator. The derivatives were determined at wing sweep angles of 26 deg, 35 deg, and 58 deg. The flight conditions included a Mach number range of 0.63 to 1.43 and an angle of attack range of 2 deg to 15 deg. Maneuvers were performed at normal accelerations from 0.9g to 3.8g during steady turns to assess the aeroelastic effects on the stability and control characteristics. The derivatives generally showed consistent trends and reasonable agreement with the wind tunnel estimates. Significant Mach effects were observed for Mach numbers as low as 0.82. No large effects attributable to aeroelasticity were noted.
NASA Technical Reports Server (NTRS)
Berman, Theodore; Wilson, Jack H.
1948-01-01
An investigation of the spin and recovery characteristics of a 1/24-scale model of the Grumman XF9F-2 airplane with wing-tip tanks installed has been conducted-in the Langley 20-foot free-spinning tunnel. The effects of control settings and movements on the erect spin and recovery characteristics of the model for a range of possible loadings of the tip tanks were determined. Spin and recovery characteristics without tanks were determined in a previous investigation. The model results indicated that the airplane spins will generally be oscillatory and that recoveries will be satisfactory for all loadings by normal recovery technique (full rudder reversal followed approximately one-half turn later by moving the elevator down). The rudder force necessary for recovery should be within the physical capability of the pilot but the elevator force may be excessive so that some type of balance or booster might be necessary, or it might be necessary to jettison the wing-tip tanks.
NASA Technical Reports Server (NTRS)
Weick, Fred E; Wood, Donald H
1929-01-01
An investigation was conducted to determine the effect of the wings on propulsive efficiency. The wings are shown to cause a reduction of 1 percent to 3 percent in propulsive efficiency, which is about the same for monoplane as well as biplane wings.
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)
Holzhauser, Curt A; Bray, Richard S
1956-01-01
An investigation was undertaken to determine the increase in maximum lift coefficient that could be obtained by applying area suction near the leading edge of a wing. This investigation was performed first with a 35 degree swept-wing model in the wind tunnel, and then with an operational 35 degree swept-wing airplane which was modified in accord with the wind-tunnel results. The wind-tunnel and flight tests indicated that the maximum lift coefficient was increased more than 50 percent by the use of area suction. Good agreement was obtained in the comparison of the wind-tunnel results with those measured in flight.
NASA Technical Reports Server (NTRS)
Tambor, Ronald
1960-01-01
The lift and drag characteristics of a Boeing KC-135 airplane were determined during maneuvering flight over the Mach number range from 0.70 to 0.85 for the airplane in the clean configuration at an altitude of 26,000 feet. Data were also obtained over the speed range of 130 knots to 160 knots at 9,000 feet for various flap deflections with gear down.
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.