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At least 1,063 records · Page 59

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NASA NTRS2007

Twist Model Development and Results from the Active Aeroelastic Wing F/A-18 Aircraft

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.

Lizotte, Andrew M.↗

NASA NTRS1993

Lift augmentation for highly swept wing aircraft

A pair of spaced slots, disposed on each side of an aircraft centerline and spaced well inboard of the wing leading edges, are provided in the wing upper surfaces and directed tangentially spanwise toward thin sharp leading wing edges of a highly swept, delta wing aircraft. The slots are individually connected through separate plenum chambers to separate compressed air tanks and serve, collectively, as a system for providing aircraft lift augmentation. A compressed air supply is tapped from the aircraft turbojet power plant. Suitable valves, under the control of the aircraft pilot, serve to selective provide jet blowing from the individual slots to provide spanwise sheets of jet air closely adjacent to the upper surfaces and across the aircraft wing span to thereby create artificial vortices whose suction generate additional lift on the aircraft. When desired, or found necessary, unequal or one-side wing blowing is employed to generate rolling moments for augmented lateral control. Trailing flaps are provided that may be deflected differentially, individually, or in unison, as needed for assistance in take-off or landing of the aircraft.

Rao, Dhanvada M.↗

NASA NTRS2008

Fiber Optic Wing Shape Sensing on NASA's Ikhana UAV

This document discusses the development of fiber optic wing shape sensing on NASA's Ikhana vehicle. The Dryden Flight Research Center's Aerostructures Branch initiated fiber-optic instrumentation development efforts in the mid-1990s. Motivated by a failure to control wing dihedral resulting in a mishap with the Helios aircraft, new wing displacement techniques were developed. Research objectives for Ikhana included validating fiber optic sensor measurements and real-time wing shape sensing predictions; the validation of fiber optic mathematical models and design tools; assessing technical viability and, if applicable, developing methodology and approaches to incorporate wing shape measurements within the vehicle flight control system; and, developing and flight validating approaches to perform active wing shape control using conventional control surfaces and active material concepts.

Richards, Lance↗

NASA NTRS2009

Hybrid Wing Body Configuration System Studies

The objective of this study was to develop a hybrid wing body (HWB) sizing and analysis capability, apply that capability to estimate the fuel burn potential for an HWB concept, and identify associated technology requirements. An advanced tube with wings concept was also developed for comparison purposes. NASA s Flight Optimization System (FLOPS) conceptual aircraft sizing and synthesis software was modified to enable the sizing and analysis of HWB concepts. The noncircular pressurized centerbody of the HWB concept was modeled, and several options were created for defining the outboard wing sections. Weight and drag estimation routines were modified to accommodate the unique aspects of an HWB configuration. The resulting capability was then utilized to model a proprietary Boeing blended wing body (BWB) concept for comparison purposes. FLOPS predicted approximately a 15 percent greater drag, mainly caused by differences in compressibility drag estimation, and approximately a 5 percent greater takeoff gross weight, mainly caused by the additional fuel required, as compared with the Boeing data. Next, a 777-like reference vehicle was modeled in FLOPS and calibrated to published Boeing performance data; the same mission definition was used to size an HWB in FLOPS. Advanced airframe and propulsion technology assumptions were applied to the HWB to develop an estimate for potential fuel burn savings from such a concept. The same technology assumptions, where applicable, were then applied to an advanced tube-with-wings concept. The HWB concept had a 39 percent lower block fuel burn than the reference vehicle and a 12 percent lower block fuel burn than the advanced tube-with-wings configuration. However, this fuel burn advantage is partially derived from assuming the high-risk technology of embedded engines with boundary-layer-ingesting inlets. The HWB concept does have the potential for significantly reduced noise as a result of the shielding advantages that are inherent with an over-body engine installation.

Nickol, Craig L.↗

NASA NTRS1934

Wing-Nacelle-Propeller Tests - Comparative Tests of Liquid-Cooled and Air-Cooled Engine Nacelles

This report gives the results of measurements of the lift, drag, and propeller characteristics of several wing and nacelle combinations with a tractor propeller. The nacelles were so located that the propeller was about 31% of the wing chord directly ahead of the leading edge of the wing, a position which earlier tests (NASA Report No. 415) had shown to be efficient. The nacelles were scale models of an NACA cowled nacelle for a radial air-cooled engine, a circular nacelle with the V-type engine located inside and the radiator for the cooling liquid located inside and the radiator for the type, and a nacelle shape simulating the housing which would be used for an extension shaft if the engine were located entirely within the wing. The propeller used in all cases was a 4-foot model of Navy No. 4412 adjustable metal propeller. The results of the tests indicate that, at the angles of attack corresponding to high speeds of flight, there is no marked advantage of one type of nacelle over the others as far as low drag is concerned, since the drag added by any of the nacelles in the particular location ahead of the wing is very small. The completely cowled nacelle for a radial air-cooled engine appears to have the highest drag, the liquid-cooled engine appears to have the highest drag, the liquid-cooled engine nacelle with external radiator slightly less drag. The liquid-cooled engine nacelle with radiator in the cowling hood has about half the drag of the cowled radial air-cooled engine nacelle. The extension-shaft housing shows practically no increase in drag over that of the wing alone. A large part of the drag of the liquid-cooled engine nacelle appears to be due to the external radiator. The maximum propulsive efficiency for a given propeller pitch setting is about 2% higher for the liquid-cooled engine nacelle with the radiator in the cowling hood than that for the other cowling arrangements.

Wood, Donald H.↗

NASA NTRS2010

Slot Nozzle Effects for Reduced Sonic Boom on a Generic Supersonic Wing Section

NASA has conducted research programs to reduce or eliminate the operational restrictions of supersonic aircraft over populated areas. Restrictions are due to the disturbance from the sonic boom, caused by the coalescence of shock waves formed off the aircraft. Results from two-dimensional computational fluid dynamic (CFD) analyses (performed on a baseline Mach 2.0 nozzle in a simulated Mach 2.2 flow) indicate that over-expanded and under-expanded operation of the nozzle has an effect on the N-wave boom signature. Analyses demonstrate the feasibility of reducing the magnitude of the sonic boom N-wave by controlling the nozzle plume interaction with the nozzle boat tail shock structure. This work was extended to study the impact of integrating a high aspect ratio exhaust nozzle or long slot nozzle on the trailing edge of a supersonic wing. The nozzle is operated in a highly under-expanded condition, creating a large exhaust plume and a shock at the trailing edge of the wing. This shock interacts with and suppresses the expansion wave caused by the wing, a major contributor to the sonic boom signature. The goal was to reduce the near field pressures caused by the expansion using a slot nozzle located at the wing trailing edge. Results from CFD analysis on a simulated wing cross-section and a slot nozzle indicate potential reductions in sonic boom signature compared to a baseline wing with no propulsion or trailing edge exhaust. Future studies could investigate if this effect could be useful on a supersonic aircraft for main propulsion, auxiliary propulsion, or flow control.

Castner, Raymond S.↗

NASA NTRS2010

Receptivity and Transition of Supersonic Boundary Layers Over Swept Wings

The receptivity, stability, and transition of three-dimensional supersonic boundary layers over (1) a swept cylinder, (2) a swept wing with a sharp leading edge, and (3) a swept wing with a blunt leading edge are numerically investigated for a free-stream Mach number of 3. These computations are compared to an earlier experimental and computational study performed by Archambaud et al.1 The steady flow fields with and without roughness elements are obtained by solving the full Navier-Stokes equations. The N-factors computed in this study at the transition onset locations reported in Ref. 1 for flow over the swept cylinder are approximately 16.5 for traveling crossflow disturbances and 9 for stationary disturbances. The N-factors for the traveling crossflow are high based on our past experiences. However, they are comparatively smaller than those reported by Archambaud et al., who found N-factor values in the range of 20 to 25 for traveling disturbances and 13 to 20 for stationary disturbances. Similarly, the N-factors computed in this study for the traveling and stationary disturbances for the flow over the sharp wing are approximately 7 and 2.5, respectively, and for the flow over the blunt wing are 6.5 and 4.8, respectively. Using the envelope method, Archambaud et al. obtained values of approximately 8.0 and 4.0 for the sharp wing case and 16.0 and 12.0 for the blunt wing case.

Balakumar, Ponnampalam↗

NASA NTRS2010

Slot Nozzle Effects for Reduced Sonic Boom on a Generic Supersonic Wing Section

Caster, Raymond S.↗

NASA NTRS2010

Methods for In-Flight Wing Shape Predictions of Highly Flexible Unmanned Aerial Vehicles: Formulation of Ko Displacement Theory

The Ko displacement theory is formulated for a cantilever tubular wing spar under bending, torsion, and combined bending and torsion loading. The Ko displacement equations are expressed in terms of strains measured at multiple sensing stations equally spaced on the surface of the wing spar. The bending and distortion strain data can then be input to the displacement equations to calculate slopes, deflections, and cross-sectional twist angles of the wing spar at the strain-sensing stations for generating the deformed shapes of flexible aircraft wing spars. The displacement equations have been successfully validated for accuracy by finite-element analysis. The Ko displacement theory that has been formulated could also be applied to calculate the deformed shape of simple and tapered beams, plates, and tapered cantilever wing boxes. The Ko displacement theory and associated strain-sensing system (such as fiber optic sensors) form a powerful tool for in-flight deformation monitoring of flexible wings and tails, such as those often employed on unmanned aerial vehicles. Ultimately, the calculated displacement data can be visually displayed in real time to the ground-based pilot for monitoring the deformed shape of unmanned aerial vehicles during flight.

Ko, William L.↗

NASA NTRS2011

Analysis of Asymmetric Aircraft Aerodynamics Due to an Experimental Wing Glove

Aerodynamic analysis on a business jet with a wing glove attached to one wing is presented and discussed. If a wing glove is placed over a portion of one wing, there will be asymmetries in the aircraft as well as overall changes in the forces and moments acting on the aircraft. These changes, referred to as deltas, need to be determined and quantified to make sure the wing glove does not have a drastic effect on the aircraft flight characteristics. TRANAIR, a non-linear full potential solver was used to analyze a full aircraft, with and without a glove, at a variety of flight conditions and angles of attack and sideslip. Changes in the aircraft lift, drag and side force, along with roll, pitch and yawing moment are presented. Span lift and moment distributions are also presented for a more detailed look at the effects of the glove on the aircraft. Aerodynamic flow phenomena due to the addition of the glove and its fairing are discussed. Results show that the glove used here does not present a drastic change in forces and moments on the aircraft, but an added torsional moment around the quarter-chord of the wing may be a cause for some structural concerns.

Hartshorn, Fletcher↗

NASA NTRS2013

A CFD Database for Airfoils and Wings at Post-Stall Angles of Attack

This paper presents selected results from an ongoing effort to develop an aerodynamic database from Reynolds-Averaged Navier-Stokes (RANS) computational analysis of airfoils and wings at stall and post-stall angles of attack. The data obtained from this effort will be used for validation and refinement of a low-order post-stall prediction method developed at NCSU, and to fill existing gaps in high angle of attack data in the literature. Such data could have potential applications in post-stall flight dynamics, helicopter aerodynamics and wind turbine aerodynamics. An overview of the NASA TetrUSS CFD package used for the RANS computational approach is presented. Detailed results for three airfoils are presented to compare their stall and post-stall behavior. The results for finite wings at stall and post-stall conditions focus on the effects of taper-ratio and sweep angle, with particular attention to whether the sectional flows can be approximated using two-dimensional flow over a stalled airfoil. While this approximation seems reasonable for unswept wings even at post-stall conditions, significant spanwise flow on stalled swept wings preclude the use of two-dimensional data to model sectional flows on swept wings. Thus, further effort is needed in low-order aerodynamic modeling of swept wings at stalled conditions.

Petrilli, Justin↗

NASA NTRS2014

Static Aeroelastic Scaling and Analysis of a Sub-Scale Flexible Wing Wind Tunnel Model

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.

Aeroelasticity↗

NASA NTRS1960

A Method of Reducing Aeroelastic Effects of Highly Swept Wings

The greatest efficiency for a lifting surface at supersonic speeds, according to the theoretical considerations of reference 1, can be attained if the leading edge is swept well behind the Mach cone and the highest aspect ratio which is structurally possible is employed. Such a wing, designed for a Mach number of 3.0, would have 80 deg. of sweepback. Aeroelastic effects have < been shown 3 to be considerable for a wing with 60deg of sweepback and designed for a Mach number of 2.0. The wing shown was found theoretically to have considerable loss in maximum lift-drag ratio attributable to aeroelasticity. This wing has 12-per cent-thick Clark-Y airfoils normal to the wing leading edge. If it were of solid aluminum and flying at a dynamic pressure of 2,400 lbs./sq.ft. (flexibility parameter qb(exp. 4) /El(0) = 7.8), analysis indicates that the wing would deflect so as to reduce the maximum lift-drag ratio about 30 per cent.

Hopkins, Edward J.↗

NASA NTRS2015

Development of Variable Camber Continuous Trailing Edge Flap for Performance Adaptive Aeroelastic Wing

This paper summarizes the recent development of an adaptive aeroelastic wing shaping control technology called variable camber continuous trailing edge flap (VCCTEF). As wing flexibility increases, aeroelastic interactions with aerodynamic forces and moments become an increasingly important consideration in aircraft design and aerodynamic performance. Furthermore, aeroelastic interactions with flight dynamics can result in issues with vehicle stability and control. The initial VCCTEF concept was developed in 2010 by NASA under a NASA Innovation Fund study entitled "Elastically Shaped Future Air Vehicle Concept," which showed that highly flexible wing aerodynamic surfaces can be elastically shaped in-flight by active control of wing twist and bending deflection in order to optimize the spanwise lift distribution for drag reduction. A collaboration between NASA and Boeing Research & Technology was subsequently funded by NASA from 2012 to 2014 to further develop the VCCTEF concept. This paper summarizes some of the key research areas conducted by NASA during the collaboration with Boeing Research and Technology. These research areas include VCCTEF design concepts, aerodynamic analysis of VCCTEF camber shapes, aerodynamic optimization of lift distribution for drag minimization, wind tunnel test results for cruise and high-lift configurations, flutter analysis and suppression control of flexible wing aircraft, and multi-objective flight control for adaptive aeroelastic wing shaping control.

Aircraft↗

NASA NTRS2015

Three-Dimensional Piecewise-Continuous Class-Shape Transformation of Wings

Class-Shape Transformation (CST) is a popular method for creating analytical representations of the surface coordinates of various components of aerospace vehicles. A wide variety of two- and three-dimensional shapes can be represented analytically using only a modest number of parameters, and the surface representation is smooth and continuous to as fine a degree as desired. This paper expands upon the original two-dimensional representation of airfoils to develop a generalized three-dimensional CST parametrization scheme that is suitable for a wider range of aircraft wings than previous formulations, including wings with significant non-planar shapes such as blended winglets and box wings. The method uses individual functions for the spanwise variation of airfoil shape, chord, thickness, twist, and reference axis coordinates to build up the complete wing shape. An alternative formulation parameterizes the slopes of the reference axis coordinates in order to relate the spanwise variation to the tangents of the sweep and dihedral angles. Also discussed are methods for fitting existing wing surface coordinates, including the use of piecewise equations to handle discontinuities, and mathematical formulations of geometric continuity constraints. A subsonic transport wing model is used as an example problem to illustrate the application of the methodology and to quantify the effects of piecewise representation and curvature constraints.

Olson, Erik D.↗

NASA NTRS2017

Multi-Objective Flight Control for Drag Minimization and Load Alleviation of High-Aspect Ratio Flexible Wing Aircraft

As aircraft wings become much more flexible due to the use of light-weight composites material, adverse aerodynamics at off-design performance can result from changes in wing shapes due to aeroelastic deflections. Increased drag, hence increased fuel burn, is a potential consequence. Without means for aeroelastic compensation, the benefit of weight reduction from the use of light-weight material could be offset by less optimal aerodynamic performance at off-design flight conditions. Performance Adaptive Aeroelastic Wing (PAAW) technology can potentially address these technical challenges for future flexible wing transports. PAAW technology leverages multi-disciplinary solutions to maximize the aerodynamic performance payoff of future adaptive wing design, while addressing simultaneously operational constraints that can prevent the optimal aerodynamic performance from being realized. These operational constraints include reduced flutter margins, increased airframe responses to gust and maneuver loads, pilot handling qualities, and ride qualities. All of these constraints while seeking the optimal aerodynamic performance present themselves as a multi-objective flight control problem. The paper presents a multi-objective flight control approach based on a drag-cognizant optimal control method. A concept of virtual control, which was previously introduced, is implemented to address the pair-wise flap motion constraints imposed by the elastomer material. This method is shown to be able to satisfy the constraints. Real-time drag minimization control is considered to be an important consideration for PAAW technology. Drag minimization control has many technical challenges such as sensing and control. An initial outline of a real-time drag minimization control has already been developed and will be further investigated in the future. A simulation study of a multi-objective flight control for a flight path angle command with aeroelastic mode suppression and drag minimization demonstrates the effectiveness of the proposed solution. In-flight structural loads are also an important consideration. As wing flexibility increases, maneuver load and gust load responses can be significant and therefore can pose safety and flight control concerns. In this paper, we will extend the multi-objective flight control framework to include load alleviation control. The study will focus initially on maneuver load minimization control, and then subsequently will address gust load alleviation control in future work.

Aerolasticity↗

NASA NTRS2017

Computational Analysis of Powered Lift Augmentation for the LEAPTech Distributed Electric Propulsion Wing

A computational study of a distributed electric propulsion wing with a 40deg flap deflection has been completed using FUN3D. Two lift-augmentation power conditions were compared with the power-off configuration on the high-lift wing (40deg flap) at a 73 mph freestream flow and for a range of angles of attack from -5 degrees to 14 degrees. The computational study also included investigating the benefit of corotating versus counter-rotating propeller spin direction to powered-lift performance. The results indicate a large benefit in lift coefficient, over the entire range of angle of attack studied, by using corotating propellers that all spin counter to the wingtip vortex. For the landing condition, 73 mph, the unpowered 40deg flap configuration achieved a maximum lift coefficient of 2.3. With high-lift blowing the maximum lift coefficient increased to 5.61. Therefore, the lift augmentation is a factor of 2.4. Taking advantage of the fullspan lift augmentation at similar performance means that a wing powered with the distributed electric propulsion system requires only 42 percent of the wing area of the unpowered wing. This technology will allow wings to be 'cruise optimized', meaning that they will be able to fly closer to maximum lift over drag conditions at the design cruise speed of the aircraft.

Deere, Karen A.↗

NASA NTRS2017

Comparison of High-Fidelity Computational Tools for Wing Design of a Distributed Electric Propulsion Aircraft

A variety of tools, from fundamental to high order, have been used to better understand applications of distributed electric propulsion to aid the wing and propulsion system design of the Leading Edge Asynchronous Propulsion Technology (LEAPTech) project and the X-57 Maxwell airplane. Three high-fidelity, Navier-Stokes computational fluid dynamics codes used during the project with results presented here are FUN3D, STAR-CCM+, and OVERFLOW. These codes employ various turbulence models to predict fully turbulent and transitional flow. Results from these codes are compared for two distributed electric propulsion configurations: the wing tested at NASA Armstrong on the Hybrid-Electric Integrated Systems Testbed truck, and the wing designed for the X-57 Maxwell airplane. Results from these computational tools for the high-lift wing tested on the Hybrid-Electric Integrated Systems Testbed truck and the X-57 high-lift wing presented compare reasonably well. The goal of the X-57 wing and distributed electric propulsion system design achieving or exceeding the required 𝐶 (sub L) = 3.95 for stall speed was confirmed with all of the computational codes.

Deere, Karen A.↗