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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.

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At least 181 records · Page 10
All sourcesNASA NTRSDOE OSTI
NASA NTRS2009

Static Aeroelastic Effects of Formation Flight for Slender Unswept Wings

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.

Hanson, Curtis E.↗
NASA NTRS2020

Jig Twist Optimization of Mach 0.745 Transonic Truss-Braced Wing Aircraft and High-Fidelity CFD Validation

This paper presents a jig twist optimization study of Mach 0.745 Transonic Truss-Braced Wing (TTBW) aircraft using an in-house developed aero-structural analysis solver VSPAERO coupled to BEAM3D. A vortex-lattice model of the TTBW model is developed, and a transonic and viscous flow correction method is implemented in the VSPAERO model to account for transonic and viscous flow effects. A correction method for the wing-strut interference aerodynamics is developed and applied to the VSPAERO solver. Also, a structural dynamic finite-element model of the TTBW aircraft is developed. This finite-element model includes the geometric nonlinear effect due to the tension in the struts which causes a deflection-dependent nonlinear stiffness. The VSPAERO model is coupled to the corresponding finite-element model to provide a rapid aero-structural analysis. A flight condition corresponding to Mach 0.745 at 42000 ft is selected for the TTBW aircraft jig twist optimization to reduce the drag coefficient. After the design is implemented, the drag coefficient of the twist optimized TTBW aircraft is reduced about 8 counts. At the end, a high-fidelity CFD solver FUN3D is used to validate the design.

Xiong, Juntao↗
NASA NTRS2020

Jig Twist Optimization of Mach 0.745 Transonic Truss-Braced Wing Aircraft and High-Fidelity CFD Validation

This paper presents a jig twist optimization study of Mach 0.745 Transonic Truss-Braced Wing (TTBW) aircraft using an in-house developed aero-structural analysis solver VSPAERO coupled to BEAM3D. A vortex-lattice model of the TTBW model is developed, and a transonic and viscous flow correction method is implemented in the VSPAERO model to account for transonic and viscous flow effects. A correction method for the wing-strut interference aerodynamics is developed and applied to the VSPAERO solver. Also, a structural dynamic finite-element model of the TTBW aircraft is developed. This finite-element model includes the geometric nonlinear effect due to the tension in the struts which causes a deflection-dependent nonlinear stiffness. The VSPAERO model is coupled to the corresponding finite-element model to provide a rapid aero-structural analysis. A design flight condition corresponding to Mach 0.745 at 42000 ft is selected for the TTBW aircraft jig twist optimization to reduce the drag coefficient. After the design is implemented, the drag coefficient of the twist optimized TTBW aircraft is reduced about 8 counts. At the end, a high-fidelity CFD solver FUN3D is used to validate the design.

BEAM3D↗
NASA NTRS1994

Effect of leading- and trailing-edge flaps on clipped delta wings with and without wing camber at supersonic speeds

An experimental investigation of the aerodynamic characteristics of thin, moderately swept fighter wings has been conducted to evaluate the effect of camber and twist on the effectiveness of leading- and trailing-edge flaps at supersonic speeds in the Langley Unitary Plan Wind Tunnel. The study geometry consisted of a generic fuselage with camber typical of advanced fighter designs without inlets, canopy, or vertical tail. The model was tested with two wing configurations an uncambered (flat) wing and a cambered and twisted wing. Each wing had an identical clipped delta planform with an inboard leading edge swept back 65 deg and an outboard leading edge swept back 50 deg. The trailing edge was swept forward 25 deg. The leading-edge flaps were deflected 4 deg to 15 deg, and the trailing-edge flaps were deflected from -30 deg to 10 deg. Longitudinal force and moment data were obtained at Mach numbers of 1.60, 1.80, 2.00, and 2.16 for an angle-of-attack range 4 deg to 20 deg at a Reynolds number of 2.16 x 10(exp 6) per foot and for an angle-of-attack range 4 deg to 20 deg at a Reynolds number of 2.0 x 10(exp 6) per foot. Vapor screen, tuft, and oil flow visualization data are also included.

Hernandez, Gloria↗
NASA NTRS1983

An aerodynamic analysis computer program and design notes for low speed wing flap systems

The expanded capabilities for analysis and design of low speed flap systems afforded by recent modifications of an existing computer program is described. The program provides for the simultaneous analysis of up to 25 pairs of leading-edge and trailing-edge flap deflection schedules. Among other new features of the program are a revised attainable thrust estimation method to provide more accurate predictions for low Mach numbers, and a choice of three options for estimation of leading-edge separation vortex flow effects. Comparison of program results with low speed experimental data for an arrow wing supersonic cruise configuration with leading-edge and trailing-edge flaps showed good agreement over most of the range of flap deflections. Other force data comparisons and an independent study of airfoil and wing pressure distributions indicated that wind-tunnel measurements of the aerodynamic performance of twisted and cambered wings and wings with leading-edge flaps can be very sensitive to Reynolds number effects.

Carlson, H. W.↗
NASA NTRS2015

Model Deformation Measurements of Sonic Boom Models in the NASA Ames 9- by 7-Ft Supersonic Wind Tunnel

The deformations of two sonic-boom models were measured by stereo photogrammetry during tests in the 9- by 7-Ft Supersonic Wind Tunnel at NASA Ames Research Center. The models were geometrically similar but one was 2.75 times as large as the other. Deformation measurements were made by simultaneously imaging the upper surfaces of the models from two directions by calibrated cameras that were mounted behind windows of the test section. Bending and twist were measured at discrete points using conventional circular targets that had been marked along the leading and trailing edges of the wings and tails. In addition, continuous distributions of bending and twist were measured from ink speckles that had been applied to the upper surfaces of the model. Measurements were made at wind-on (M = 1.6) and wind-off conditions over a range of angles of attack between 2.5 deg. and 5.0 deg. At each condition, model deformation was determined by comparing the wind-off and wind-on coordinates of each measurement point after transforming the coordinates to reference coordinates tied to the model. The necessary transformations were determined by measuring the positions of a set of targets on the rigid center-body of the models whose model-axes coordinates were known. Smoothly varying bending and twist measurements were obtained at all conditions. Bending displacements increased in proportion to the square of the distance to the centerline. Maximum deflection of the wingtip of the larger model was about 5 mm (2% of the semispan) and that of the smaller model was 0.9 mm (1% of the semispan). The change in wing twist due to bending increased in direct proportion to distance from the centerline and reached a (absolute) maximum of about -1 at the highest angle of attack for both models. The measurements easily resolved bending displacements as small as 0.05 mm and bending-induced changes in twist as small as 0.05 deg.

model deformation↗
NASA NTRS1961

A Method of Determining Aerodynamic-Influence Coefficients from Wind-Tunnel Data for Wings at Supersonic Speeds

A method is described for determining aerodynamic-influence coefficients from wind-tunnel data for calculating the steady-state load distribution on a wing with arbitrary angle-of-attack distribution at supersonic speeds. The method combines linearized theory with empirical adjustments in order to give accurate results over a wide range of angles of attack. The experimented data required are pressure distributions measured on a flat wing of the desired planform at the desired Mach number and over the desired range of angles of attack. The method has been tested by applying it to wind-tunnel data measured at Mach numbers of 1.61 and 2.01 on wings of the same planform but of different surface shapes. Influence coefficients adjusted to fit the flat wing gave good predictions of the spanwise and chord-wise distributions of loadings measured on twisted and cambered wings.

Gainer, Patrick A.↗
NASA NTRSDate not supplied

Multi-point Jig Twist Optimization of Mach 0.745 Transonic Truss-Braced Wing Aircraft and High-Fidelity CFD Validation

This paper presents a multi-point jig twist optimization study of the Mach 0.745 Transonic Truss-Braced Wing (TTBW) aircraft using an in-house developed aero-structural analysis solver VSPAERO coupled to BEAM3D. A vortex-lattice model of the TTBW model is developed, and a transonic small disturbance /integral boundary layer correction method is implemented in the VSPAERO model to account for transonic and viscous flow effects. A correction method for the wing-strut interference aerodynamics is developed and applied to the VSPAERO solver. Also, a structural dynamic finite-element model of the TTBW aircraft is developed. This finite-element model includes the geometric nonlinear effect due to the tension in the struts which causes a deflection-dependent nonlinear stiffness. The VSPAERO model is coupled to the corresponding finite-element model to provide a rapid aero-structural analysis. The VSPAERO model has been validated against wind tunnel test data and FUN3D CFD simulation data. Three different flight conditions corresponding to Mach 0.745 at 40,000 ft are selected for the TTBW aircraft jig twist optimization to reduce the drag coefficients at the design and off-design cruise lift coefficients. The preliminary optimization results show that the TTBW aircraft with the optimized jig twist achieves a drag reduction of about 9 counts, 7 counts, and8 counts corresponding to the lift coefficients 0.68, 0.73, and 0.78, respectively. A high-fidelity CFD solver FUN3D is used to validate the jig twist optimization.

Multi-point↗
NASA NTRS1961

Effects at Mach Numbers of 1.61 and 2.01 of Camber and Twist on the Aerodynamic Characteristics of Three Swept Wings Having the Same Planform

An investigation has been made at Mach numbers of 1.61 and 2.01 to determine the aerodynamic characteristics of three wings having a sweepback of 50 deg at the quarter-chord line, a taper ratio of 0.20, an NACA 65A005 thickness distribution, and an aspect ratio of 3.5. One wing was flat, one had at each spanwise station an a = 0 mean line modified to have a maximum height of 4-percent chord, and one had a linear variation of twist with 6 deg of washout at the tip. Tests were made with natural and fixed transition at Reynolds numbers ranging from 1.2 x 10(exp 6) to 3.6 x 10(exp 6) through an angle-of-attack range of -20 deg to 20 deg. When compared with the flat wing, the effect of the linear variation of twist with 6 deg of washout at the tip was to increase the lift-drag ratio when the leading edge was subsonic; but little increase in lift-drag ratio was obtained when the leading edge was supersonic. Pitching moment was increased and gave a positive trim point without greatly affecting the rate of change of pitching moment with lift coefficient. For the cambered wing the high minimum drag resulted in comparatively low lift-drag ratios. In addition, the pitching moments were decreased so that a negative trim point was obtained.

Landrum, Emma Jean↗
NASA NTRS1993

An inverse method for computation of structural stiffness distributions of aeroelastically optimized wings

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.

Schuster, David M.↗
NASA NTRS1984

An analytical design procedure for the determination of effective leading edge extensions on thick delta wings

An analytical design procedure for leading edge extensions (LEE) was developed for thick delta wings. This LEE device is designed to be mounted to a wing along the pseudo-stagnation stream surface associated with the attached flow design lift coefficient of greater than zero. The intended purpose of this device is to improve the aerodynamic performance of high subsonic and low supersonic aircraft at incidences above that of attached flow design lift coefficient, by using a vortex system emanating along the leading edges of the device. The low pressure associated with these vortices would act on the LEE upper surface and the forward facing area at the wing leading edges, providing an additional lift and effective leading edge thrust recovery. The first application of this technique was to a thick, round edged, twisted and cambered wing of approximately triangular planform having a sweep of 58 deg and aspect ratio of 2.30. The panel aerodynamics and vortex lattice method with suction analogy computer codes were employed to determine the pseudo-stagnation stream surface and an optimized LEE planform shape.

Farhad Ghaffari↗
NASA NTRS1999

Supersonic Aerodynamic Design Improvements of an Arrow-Wing HSCT Configuration Using Nonlinear Point Design Methods

This paper is a discussion of the supersonic nonlinear point design optimization efforts at McDonnell Douglas Aerospace under the High-Speed Research (HSR) program. The baseline for these optimization efforts has been the M2.4-7A configuration which represents an arrow-wing technology for the High-Speed Civil Transport (HSCT). Optimization work on this configuration began in early 1994 and continued into 1996. Initial work focused on optimization of the wing camber and twist on a wing/body configuration and reductions of 3.5 drag counts (Euler) were realized. The next phase of the optimization effort included fuselage camber along with the wing and a drag reduction of 5.0 counts was achieved. Including the effects of the nacelles and diverters into the optimization problem became the next focus where a reduction of 6.6 counts (Euler W/B/N/D) was eventually realized. The final two phases of the effort included a large set of constraints designed to make the final optimized configuration more realistic and they were successful albeit with a loss of performance.

Unger, Eric R.↗