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NASA NTRS2005

Flight Test of the F/A-18 Active Aeroelastic Wing Airplane

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.

Clarke, Robert↗

NASA NTRS2005

Flight Test of the F/A-18 Active Aeroelastic Wing Airplane

Clarke, Robert↗

NASA NTRS2005

Loads Model Development and Analysis for the F/A-18 Active Aeroelastic Wing Airplane

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.

Allen, Michael J.↗

NASA NTRS2005

Development and Testing of Control Laws for the Active Aeroelastic Wing Program

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.

Dibley, Ryan P.↗

NASA NTRS1950

Investigation of Downwash, Sidewash, and Mach Number Distribution behind a Rectangular Wing at a Mach Number of 2.41

An investigation of the nature of the flow field behind a rectangular circular-arc wing has been conducted in the Langley 9-inch supersonic tunnel. Pitot- and static-pressure surveys covering a region of flow behind the wing have been made together with detailed pitot surveys throughout the region of the wake. In addition, the flow direction has been measured using a weathercocking vane measurements. Theoretical calculations of the variation of both downwash and sidewash with angle of attack using Lagerstrom's superposition method have been made. In addition the effect of the wing thickness on the sidewash with the wing at 0 angle of attack has been evaluated. Near an angle of attack of 0, agreement between theory and experiment is good, particularly for the downwash results, except in the plane of the wing, inboard of the tip. In this region the proximity of the shed vortex sheet and the departure of the spanwise distribution of vorticity from theory would account for the disagreement. At higher angles of attack prediction of downwash depends on a knowledge of the location of the trailing vortex sheet, in order that the downwash may be corrected for its displacement and distortion. The theoretical location of the trailing vortex sheet, based on the theoretical downwash values integrated downstream from the wing trailing edge, is shown to differ widely from the experimental case. The rolling-up of the trailing vortex sheet behind the wing tip is evidenced by both the wake surveys and the flow-angle measurements.

Adamson, D.↗

NASA NTRS1949

Wind-Tunnel Tests of a 0.16-Scale Model of the Douglas MX-656 Airplane at High Subsonic Speeds. II - Wing and Fuselage Pressure Distribution

Measurements of wing and fuselage pressure distributions were made at low and high subsonic Much numbers on a 0.16-scale model of the projected MX-656 research airplane. The MX-656 is a supersonic design utilizing a low-aspect-ratio wing and tail. Pressure-distribution measurements indicated that, although the critical Mach number of the wing was approximately 0.81 at 0 degree angle of attack, compressibility effects were of little significance below a Mach number of at least 0.90. The principal effect of compressibility was an increase in the pressure gradient over the after 30 percent of the wing chord, causing a tendency for the flow to separate. At 0.40 Mach number, the wing stalled abruptly at approximately 12 deg, angle of attack. The wing-pressure distribution showed this stall was a result of complete separation of the flow from the upper surface of the wing, Deflecting the leading-edge flaps delayed the stall to a higher angle of attack with some increase in the maximum section normal force,

Cleary, Joseph W.↗

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 NTRS2012

Computational Optimization of a Natural Laminar Flow Experimental Wing Glove

Computational optimization of a natural laminar flow experimental wing glove that is mounted on a business jet is presented and discussed. The process of designing a laminar flow wing glove starts with creating a two-dimensional optimized airfoil and then lofting it into a three-dimensional wing glove section. The airfoil design process does not consider the three dimensional flow effects such as cross flow due wing sweep as well as engine and body interference. Therefore, once an initial glove geometry is created from the airfoil, the three dimensional wing glove has to be optimized to ensure that the desired extent of laminar flow is maintained over the entire glove. TRANAIR, a non-linear full potential solver with a coupled boundary layer code was used as the main tool in the design and optimization process of the three-dimensional glove shape. The optimization process uses the Class-Shape-Transformation method to perturb the geometry with geometric constraints that allow for a 2-in clearance from the main wing. The three-dimensional glove shape was optimized with the objective of having a spanwise uniform pressure distribution that matches the optimized two-dimensional pressure distribution as closely as possible. Results show that with the appropriate inputs, the optimizer is able to match the two dimensional pressure distributions practically across the entire span of the wing glove. This allows for the experiment to have a much higher probability of having a large extent of natural laminar flow in flight.

Hartshom, Fletcher↗

NASA NTRS2013

Flight Testing of Novel Compliant Spines for Passive Wing Morphing on Ornithopters

Unmanned Aerial Vehicles (UAVs) are proliferating in both the civil and military markets. Flapping wing UAVs, or ornithopters, have the potential to combine the agility and maneuverability of rotary wing aircraft with excellent performance in low Reynolds number flight regimes. The purpose of this paper is to present new free flight experimental results for an ornithopter equipped with one degree of freedom (1DOF) compliant spines that were designed and optimized in terms of mass, maximum von-Mises stress, and desired wing bending deflections. The spines were inserted in an experimental ornithopter wing spar in order to achieve a set of desired kinematics during the up and down strokes of a flapping cycle. The ornithopter was flown at Wright Patterson Air Force Base in the Air Force Research Laboratory Small Unmanned Air Systems (SUAS) indoor flight facility. Vicon motion tracking cameras were used to track the motion of the vehicle for five different wing configurations. The effect of the presence of the compliant spine on wing kinematics and leading edge spar deflection during flight is presented. Results show that the ornithopter with the compliant spine inserted in its wing reduced the body acceleration during the upstroke which translates into overall lift gains.

Wissa, Aimy↗

NASA NTRS2013

Swept-Wing Ice Accretion Characterization and Aerodynamics

NASA, FAA, ONERA, the University of Illinois and Boeing have embarked on a significant, collaborative research effort to address the technical challenges associated with icing on large-scale, three-dimensional swept wings. The overall goal is to improve the fidelity of experimental and computational simulation methods for swept-wing ice accretion formation and resulting aerodynamic effect. A seven-phase research effort has been designed that incorporates ice-accretion and aerodynamic experiments and computational simulations. As the baseline, full-scale, swept-wing-reference geometry, this research will utilize the 65% scale Common Research Model configuration. Ice-accretion testing will be conducted in the NASA Icing Research Tunnel for three hybrid swept-wing models representing the 20%, 64% and 83% semispan stations of the baseline-reference wing. Three-dimensional measurement techniques are being developed and validated to document the experimental ice-accretion geometries. Artificial ice shapes of varying geometric fidelity will be developed for aerodynamic testing over a large Reynolds number range in the ONERA F1 pressurized wind tunnel and in a smaller-scale atmospheric wind tunnel. Concurrent research will be conducted to explore and further develop the use of computational simulation tools for ice accretion and aerodynamics on swept wings. The combined results of this research effort will result in an improved understanding of the ice formation and aerodynamic effects on swept wings. The purpose of this paper is to describe this research effort in more detail and report on the current results and status to date. 1

Broeren, Andy P.↗

NASA NTRS2013

Swept-Wing Ice Accretion Characterization and Aerodynamics

NASA, FAA, ONERA, the University of Illinois and Boeing have embarked on a significant, collaborative research effort to address the technical challenges associated with icing on large-scale, three-dimensional swept wings. The overall goal is to improve the fidelity of experimental and computational simulation methods for swept-wing ice accretion formation and resulting aerodynamic effect. A seven-phase research effort has been designed that incorporates ice-accretion and aerodynamic experiments and computational simulations. As the baseline, full-scale, swept-wing-reference geometry, this research will utilize the 65 percent scale Common Research Model configuration. Ice-accretion testing will be conducted in the NASA Icing Research Tunnel for three hybrid swept-wing models representing the 20, 64 and 83 percent semispan stations of the baseline-reference wing. Threedimensional measurement techniques are being developed and validated to document the experimental ice-accretion geometries. Artificial ice shapes of varying geometric fidelity will be developed for aerodynamic testing over a large Reynolds number range in the ONERA F1 pressurized wind tunnel and in a smaller-scale atmospheric wind tunnel. Concurrent research will be conducted to explore and further develop the use of computational simulation tools for ice accretion and aerodynamics on swept wings. The combined results of this research effort will result in an improved understanding of the ice formation and aerodynamic effects on swept wings. The purpose of this paper is to describe this research effort in more detail and report on the current results and status to date.

Broeren, Andy P.↗

NASA NTRS2015

Preliminary Weight Savings Estimate for a Commercial Transport Wing Using Rod-Stiffened Stitched Composite Technology

A structural concept called pultruded rod stitched efficient unitized structure (PRSEUS) was developed by the Boeing Company to address the complex structural design aspects associated with a pressurized hybrid wing body (HWB) aircraft configuration. While PRSEUS was an enabling technology for the pressurized HWB structure, limited investigation of PRSEUS for other aircraft structures, such as circular fuselages and wings, has been done. Therefore, a study was undertaken to investigate the potential weight savings afforded by using the PRSEUS concept for a commercial transport wing. The study applied PRSEUS to the Advanced Subsonic Technology (AST) Program composite semi-span test article, which was sized using three load cases. The initial PRSEUS design was developed by matching cross-sectional stiffnesses for each stringer/skin combination within the wing covers, then the design was modified to ensure that the PRSEUS design satisfied the design criteria. It was found that the PRSEUS wing design exhibited weight savings over the blade-stiffened composite AST Program wing of nearly 9%, and a weight savings of 49% and 29% for the lower and upper covers, respectively, compared to an equivalent metallic wing.

Lovejoy, Andrew E.↗

NASA NTRS2015

Active Dihedral Control System for a Torisionally Flexible Wing

A span-loaded, highly flexible flying wing, having horizontal control surfaces mounted aft of the wing on extended beams to form local pitch-control devices. Each of five spanwise wing segments of the wing has one or more motors and photovoltaic arrays, and produces its own lift independent of the other wing segments, to minimize inter-segment loads. Wing dihedral is controlled by separately controlling the local pitch-control devices consisting of a control surface on a boom, such that inboard and outboard wing segment pitch changes relative to each other, and thus relative inboard and outboard lift is varied.

Kendall, Greg T.↗

NASA NTRS2017

Active Dihedral Control System for a Torsionally Flexible Wing

Kendall, Greg T.↗

NASA NTRSDate not supplied

Aeroservoelastic Control Law Development for the Integrated Adaptive Wing Technology Maturation Wind-Tunnel Test

The NASA Advanced Air Transport Technology (AATT) Project seeks to “explore and develop technologies and concepts for improved energy efficiency and environmental compatibility for fixed wing subsonic transports” [1]. The IAWTM project is a subproject within AATT and is a joint Boeing/NASA effort to, in part, demonstrate the active control systems necessary to employ high-aspect ratio wings for improved aerodynamic efficiency. The aircraft configuration under study for the IAWTM project is a generic jet transport based on the NASA Common Research Model (CRM) whose wing aspect ratio has been increased to 13.5 (as opposed to nominal CRM aspect ratio of 9) and whose development is given in Ref. [2]. A semi-span model of the high-aspect-ratio CRM with 10 trailing edge control surfaces is currently being fabricated and will be delivered to the NASA Langley Transonic Dynamics Tunnel (TDT). There, several tests will be conducted to demonstrate maneuver load alleviation (MLA), gust load alleviation (GLA), active flutter suppression (AFS), and drag reduction, all at transonic conditions. A layout of the wind tunnel model is shown in Fig. 1. The wind-tunnel model is a spar/pod wing design with a relatively rigid fuselage attached to a mounting structure (seen in Fig. 1), which in turn is attached to a load balance behind the tunnel wall. Shims exist that can be attached to force the load to go through the fuselage rather than through the balance so as not to damage the balance in the event of flutter. The spar/pod design is similar to that used in the Boeing truss-braced wing aeroelastic model tested in the TDT in 2013 [3]. There is a central spar running the span of the wing to which the pods are attached that form the aerodynamic shape of the wing. The pods are constructed of ribs for attachment to the spar and skins to create the outer mold line. There are ten active control surfaces that will be used for the test objectives discussed above. Three of these (two outboard ailerons and one inboard aileron) are high-speed electrohydraulic actuators, and the remaining seven are low-speed mini-plane flaps, which have slower servo motors that will be used primarily for drag and maneuver load reductions. Control laws must be developed for each test (MLA, GLA, AFS, and drag reduction) and must be able to work in tandem when two or more control strategies are being tested simultaneously (e.g., when control laws to mitigate maneuver loads and gust loads operate simultaneously). The development of these control laws requires accurate mathematical representations of the aerodynamics, structural dynamics, and control surface dynamics of the system. Taken together, this mathematical representation—which is often in state-space form— is the ASE model. Three low-order ASE models of differing fidelities were developed for the IAWTM test and will be presented here. The differences in fidelity appeared in the aerodynamic modeling—one model assumed purely linear aerodynamics, while the other two models used different techniques to linearize the aerodynamics about a nonlinear mean. Furthermore, simple observer-based full-state feedback controllers for each test (MLA, GLA, AFS, and drag reduction) were designed from these low-order ASE models and implemented in a higher-order, nonlinear computational ASE simulation to assess the effectiveness of each controller. Observers were designed from each of the three lower-order ASE models, and each (linear) observer was then used as an estimate of the nonlinear dynamics of the system. Using the estimated states and the full-state feedback controller, control surface commands were fed back to the higher-order, nonlinear ASE system. The insight gained during this development will be applied to control law development for the physical IAWTM wind-tunnel test.

Josiah M Waite↗

NASA NTRSDate not supplied

Aerodynamic and Acoustic Interactions Associated with Inboard Propeller-Wing Configurations

A series of aerodynamic performance and acoustic measurements have been made on a range of inboard propeller-wing interaction configurations in the NASA Langley Low Speed Aeroacoustic Wind Tunnel (LSAWT). The results presented in this paper are part of a more expansive testing campaign encompassing both single propeller-wing and multipropeller-wing interactions, the former of which is discussed in the present work. The primary testing parameters of interest to this study are the axial and vertical positioning of the wing relative to the propeller slipstream under a constant propeller advance ratio. A multi-faceted computational effort was also employed in an effort to identify reflection and scattering effects imposed by both the wing geometry as well as the primary components of the facility test setup. This effort consisted of aerodynamic predictions using high-fidelity computational fluid dynamics (CFD), acoustic predictions using an impermeable Ffowcs Williams and Hawkings (FW-H) solver, and acoustic scattering predictions. Acoustic measurements reveal variations in the acoustic directivity behavior of the propeller blade passage frequency for even modest variations in wing position. CFD-based acoustic predictions reveal discrepancies relative to the experimental data, which is believed to be due to complex acoustic scattering behavior within the test section. Initial attempts at modeling the scattered acoustic field showed functional dependency of the acoustic amplitude variations on the wing position relative to the propeller disk, however discrepancies with experimental data remain.

Nikolas S. Zawodny↗

NASA NTRSDate not supplied

Transition Analysis for the Pazy Wing

The Pazy wing, designed as a benchmark for highly flexible wings that experience large deformations, presents intriguing nonlinear phenomena due to its unique geometry with spanwise varying static deformation and low Reynolds number flow conditions. These phenomena include limit cycle oscillation related to laminar separation bubbles, boundary-layer transition, and dynamic stall. The primary focus of this study is to investigate transition behavior on the sagged Pazy wing under static loading by using the shear-stress-transport (SST)-based Langtry-Menter 𝜸 − 𝑹𝒆𝜽𝒕 transition model in the NASA OVERFLOW and FUN3D solvers. Based on the steady-state results obtained via different variants of the Langtry-Menter model, we investigate the likely transition behavior over a range of incidence angles. The predicted flowfield on the sagged Pazy wing is strongly three dimensional and the application of turbulence index as a viable indicator of the transition front in such flows is also explored. The study demonstrates the limitations of the Langtry-Menter model when applied to low Reynolds number flows. We introduce a spanwise periodic configuration to simulate a narrow-span section of the Pazy wing, specifically centered on one of the 14 nearly equidistant ribs spanning the wing. The computational results demonstrate that the spanwise periodic configuration provides a straightforward environment for studying transition characteristics as well as demonstrating systematic grid convergence and successful code-to-code comparisons. The research offers useful insights into the transition characteristics of flexible wings under various loading conditions, contributing to the field of aeroelastic simulations.

CFD Modeling↗

NASA NTRSDate not supplied

Progress Towards the Design of a Natural Laminar Flow Wing for a Low-Boom Concept using CDISC

The X-59 is being developed to evaluate the ability to reduce sonic boom loudness to acceptable levels, with the ultimate goal of enabling supersonic flight overland. If successful, one of the future research challenges will be focused on reducing fuel burn. In this paper, Natural Laminar Flow (NLF) is proposed as a technology to decrease fuel burn for a low-boom configuration. This will require an evaluation of the compatibility of NLF wings for a low-boom configuration. As a first step towards this goal, the CATNLF method was employed using the CDISC design module to perform NLF design on an isolated wing from a low-boom configuration. The results show that the NLF wing features laminar flow over 46% of the wing upper surface, providing a potential 8.3 drag count reduction relative to the design with no laminar flow. Off-design analyses were additionally performed to evaluate the performance of the NLF wing for changes in angle of attack and Mach number. The results showed that the NLF wing offers a performance benefit over the entire range of off-design conditions considered, with most of the laminar flow maintained for near cruise conditions. Future work will consider the full configuration, which will enable an assessment of the impact of the NLF wing design on sonic boom loudness. Additionally, attachment line transition will be addressed, which was predicted to result in a loss of laminar flow on the inboard 50% span for this work. Finally, future research will focus on developing a method for effectively choosing the frequency-beta pairs used for the stability analysis to improve transition prediction and provide a more accurate estimate of the performance benefit.

Supersonic↗