Engineering Papers⌕ Search

SEARCH · Engineering Papers

Results for “aeroelastic tailoring”

Search indexed NASA NTRS and DOE OSTI research on propulsion, heat transfer, battery materials and energy systems. Follow report and document links to the original sources.

Quote a phrase for an exact phrase match. Source license links do not imply unrestricted reuse.

At least 37 records · Page 2
All sourcesNASA NTRSDOE OSTI
NASA NTRS2016

Comparison of Curvilinear Stiffeners and Tow Steered Composites for Aeroelastic Tailoring of Transports

A series of aeroelastic optimization problems are solved on a high aspect ratio wingbox of the Common Research Model, in an effort to minimize structural mass under coupled stress, buckling, and flutter constraints. Two technologies are of particular interest: tow steered composite laminate skins and curvilinear stiffeners. Both methods are found to afford feasible reductions in mass over their non-curvilinear structural counterparts, through both distinct and shared mechanisms for passively controlling aeroelastic performance. Some degree of diminishing returns are seen when curvilinear stiffeners and curvilinear fiber tow paths are used simultaneously.

Stanford, Bret K.↗
NASA NTRS2014

Material and Thickness Grading for Aeroelastic Tailoring of the Common Research Model Wing Box

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.

Stanford, Bret K.↗
NASA NTRS2015

Aeroelastic Tailoring of Transport Wings Including Transonic Flutter Constraints

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.

Stanford, Bret K.↗
NASA NTRS2015

Aeroelastic Tailoring Study of N+2 Low-Boom Supersonic Commercial Transport Aircraft

The Lockheed Martins N+2 Low-boom Supersonic Commercial Transport (LSCT) aircraft is optimized in this study through the use of a multidisciplinary design optimization tool developed at the NASA Armstrong Flight Research Center. A total of 111 design variables are used in the first optimization run. Total structural weight is the objective function in this optimization run. Design requirements for strength, buckling, and flutter are selected as constraint functions during the first optimization run. The MSC Nastran code is used to obtain the modal, strength, and buckling characteristics. Flutter and trim analyses are based on ZAERO code and landing and ground control loads are computed using an in-house code.

aeroelastic tailoring↗
NASA NTRS2015

Aeroelastic Tailoring Study of N+2 Low Boom Supersonic Commerical Transport Aircraft

The Lockheed Martin N+2 Low - boom Supersonic Commercial Transport (LSCT) aircraft was optimized in this study through the use of a multidisciplinary design optimization tool developed at the National Aeronautics and S pace Administration Armstrong Flight Research Center. A total of 111 design variables we re used in the first optimization run. Total structural weight was the objective function in this optimization run. Design requirements for strength, buckling, and flutter we re selected as constraint functions during the first optimization run. The MSC Nastran code was used to obtain the modal, strength, and buckling characteristics. Flutter and trim analyses we re based on ZAERO code, and landing and ground control loads were computed using an in - house code. The w eight penalty to satisfy all the design requirement s during the first optimization run was 31,367 lb, a 9.4% increase from the baseline configuration. The second optimization run was prepared and based on the big-bang big-crunch algorithm. Six composite ply angles for the second and fourth composite layers were selected as discrete design variables for the second optimization run. Composite ply angle changes can't improve the weight configuration of the N+2 LSCT aircraft. However, this second optimization run can create more tolerance for the active and near active strength constraint values for future weight optimization runs.

low boom supersonic civil transport↗
NASA NTRS1986

Aeroelastic tailoring of composite wings with external stores

The use of an integrating matrix solution procedure allows the aeroelastic stability of a large number of wing/store combinations to be studied with relative speed and accuracy. Results are presented which show the effects of tailoring the orientation of the plies of laminated composite wings when external stores are present. The stores utilized take the form of concentrated masses that are varied in terms of size and location, and combinations of stores are also studied. It is seen that there are cases for which small changes in the laminate can produce large changes in the aeroelastic stability boundary, and for such situations care should be used in tailoring a wing. For a given wing, the choice fiber angle to optimize the aeroelastic performance is quite sensitive to both the size and location of the store.

Green, J. A.↗
NASA NTRS1991

Experimental investigation of helicopter vibration reduction using rotor blade aeroelastic tailoring

A wind tunnel investigation has been conducted to parametrically investigate the effect of blade nonstructural mass on helicopter fixed- and rotating-system vibratory loads. The data were obtained using Mach- and aeroelastically-scaled model rotor blades which allowed for the addition of concentrated nonstructural masses at multiple points along the blade radius. Testing was conducted for advance ratios ranging from 0.10 to 0.35 for ten blade mass configurations. Three thrust levels were obtained at representative full-scale shaft angles for each configuration. Results indicate that proper placement of blade nonstructural mass can provide reductions in fixed-system vibratory loads, but that correct mass placement and the loads reduction realized are dependent upon flight condition. The data base obtained with this investigation provides a comprehensive set of fixed-system shears and moments, blade moment, and blade flap and lag response. The data set is well suited for use in the correlation and development of advanced rotorcraft analyses.

Wilbur, Matthew L.↗
NASA NTRS1983

Aeroelastic tailoring of rotor blades for vibration reduction in forward flight

Modern structural optimization techniques are applied to vibration reduction of helicopter rotor blades in forward flight. The objective function minimized consists of the oscillatory vertical hub shears or the hub rolling moments at one particular advance ratio. The behavior constraints are the frequency placements of the blade and the requirement that aeroelastic stability margins, in hover, remain unaffected by the optimization process. The aeroelastic stability and response analysis is based on a fully coupled flap-lag-torsional analysis of the blade. The vertical hub shears and rolling moments used as the objective function are obtained by appropriate integration of the loads acting along the span of the blade combined with a transformation to a hub fixed coordinate system, and a summation over the total number of blades. Numerical results for both a stiff-in-plane and a soft-in-plane configuration are presented, indicating that structural optimization yields the highest benefits when applied to soft-in-plane blade configuration. The results indicate substantial (15-40 percent) reduction in vibration levels, as well as a blade which is 20 percent lighter than the initial design.

Friedmann, P. P.↗
NASA NTRS1979

Characterization of graphite/epoxy laminates for aeroelastic tailoring

A study of interaction between wing bending and twist by graphite/epoxy anisotropic laminates used in aircraft wing skins is presented. The laminates were used as covers for subscale box beams supported as a cantilever and tested in tip shear and tip torque, measuring beam response with a reflected light technique. The results indicated that the in-plane stiffness properties of anisotropic laminates can be predicted if the layer properties of the composite materials are known; thus, the coupled bending/twist response of wing type structures made from these laminates can be determined provided the limits of the laminate linear behavior are not significantly exceeded.

Shyprykevich, P.↗
NASA NTRS1988

Optimization of nonlinear aeroelastic tailoring criteria

A static flexible fighter aircraft wing configuration is presently addressed by a multilevel optimization technique, based on both a full-potential concept and a rapid structural optimization program, which can be applied to such aircraft-design problems as maneuver load control, aileron reversal, and lift effectiveness. It is found that nonlinearities are important in the design of an aircraft whose flight envelope encompasses the transonic regime, and that the present structural suboptimization produces a significantly lighter wing by reducing ply thicknesses.

Abdi, F.↗
NASA NTRS1992

Modeling and analysis methodology for aeroelastically tailored chordwise deformable wings

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.

Rehfield, Lawrence W.↗
NASA NTRS2016

Static and Dynamic Aeroelastic Tailoring With Variable Camber Control

This paper examines the use of a Variable Camber Continuous Trailing Edge Flap (VCCTEF) system for aeroservoelastic optimization of a transport wingbox. The quasisteady and unsteady motions of the flap system are utilized as design variables, along with patch-level structural variables, towards minimizing wingbox weight via maneuver load alleviation and active flutter suppression. The resulting system is, in general, very successful at removing structural weight in a feasible manner. Limitations to this success are imposed by including load cases where the VCCTEF system is not active (open-loop) in the optimization process, and also by including actuator operating cost constraints.

Stanford, Bret K.↗
NASA NTRS2020

Passive Aeroelastic Tailored Wing Modal Test Using the Fixed Base Correction Method

In modal testing and finite element model correlation, analysts desire modal results using free-free or rigid boundary conditions to ease comparisons of test versus analytical data. It is often expensive both in cost and schedule to build and test with boundary conditions that replicate the free-free or rigid boundaries. Static test fixtures for load testing are often large, heavy, and unyielding, but do not provide adequate boundaries for modal tests because they are dynamically too flexible and often contain natural frequencies within the frequency range of interest of the test article. Dynamic coupling between the test article and test fixture complicates the model updating process because significant effort is required to model the test fixture and boundary conditions in addition to the test article. If there were a way to correct the modal results for fixture coupling, then setups used for other structural testing could be adequate for modal testing. In the case described in this paper, a partial static loads testing setup was used, which allowed significant schedule and cost savings by eliminating a unique setup for a modal test. A fixed base correction technique was investigated during modal testing of a flexible wing cantilevered from part of a static test fixture. The technique was successfully used to measure the wing modes de-coupled from the dynamically active test fixture. The technique is promising for future aircraft applications, but more research is needed.

structural dynamics↗