Browsing by Author "Dr. Ashok Gopalarathnam, Committee Chair"

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Advancements in Aerodynamic Technologies for Airfoils and Wings
(2006-12-08) Jepson, Jeffrey Keith; Dr. Jeffrey A. Joines, Committee Member; Dr. Charles E. Hall, Committee Member; Dr. Hassan A. Hassan, Committee Member; Dr. Ashok Gopalarathnam, Committee Chair
Although aircraft operate over a wide range of ﬂight conditions, current fixed geometry aircraft are optimized for only a few of these conditions. By altering the shape of the aircraft, adaptive aerodynamics can be used to increase the safety and performance of an aircraft by tailoring the aircraft for multiple light conditions. Of the various shape adaptation concepts currently being studied, the use of multiple trailing-edge ﬂaps along the span of a wing offers a relatively high possibility of being incorporated on aircraft in the near future. Multiple trailing-edge ﬂaps allow for effective spanwise camber adaptation with resulting drag beneﬁts over a large speed range and load alleviation at high-g conditions. The research presented in this dissertation focuses on the development of this concept of using trailing-edge ﬂaps to tailor an aircraft for multiple liight conditions. One of the major tasks involved in implementing trailing-edge ﬂaps is in designing the airfoil to incorporate the ﬂap. The ﬁrst part of this dissertation presents a design formulation that incorporates aircraft performance considerations in the inverse design of low-speed laminar-ﬂow adaptive airfoils with trailing-edge cruise ﬂaps. The beneﬁt of using adaptive airfoils is that the size of the low-drag region of the drag polar can be effectively increased without increasing the maximum thickness of the airfoil. Two aircraft performance parameters are considered: level-ﬂight maximum speed and maximum range. It is shown that the lift coefficients for the lower and upper corners of the airfoil low-drag range can be appropriately adjusted to tailor the airfoil for these two aircraft performance parameters. The design problem is posed as a part of a multidimensional Newton iteration in an existing conformal-mapping based inverse design code, PROFOIL. This formulation automatically adjusts the lift coefficients for the corners of the low-drag range for a given ﬂap deﬂection as required for the airfoil-aircraft matching. Examples are presented to illustrate the ﬂapped-airfoil design approach for a general aviation aircraft and the results are validated by comparison with results from post-design aircraft performance computations. Once the airfoil is designed to incorporate a TE ﬂap, it is important to determine the most suitable ﬂap angles along the wing for different flight conditions. The second part of this dissertation presents a method for determining the optimum flap angles to minimize drag based on pressures measured at select locations on the wing. Computational ﬂow simulations using a panel method are used "in the loop" for demonstrating closed-loop control of the ﬂaps. Examples in the paper show that the control algorithm is successful in correctly adapting the wing to achieve the target lift distributions for minimizing induced drag while adjusting the wing angle of attack for operation of the wing in the drag bucket. It is shown that the "sense-and-adapt" approach developed is capable of handling varying and unpredictable inﬂow conditions. Such a capability could be useful in adapting long-span ﬂexible wings that may experience signiﬁcant and unknown atmospheric inﬂow variations along the span. To further develop the "sense-and-adapt" approach, the method was tested experimentally in the third part of the research. The goal of the testing was to see if the same results found computationally can be obtained experimentally. The North Carolina State University subsonic wind tunnel was used for the wind tunnel tests. Results from the testing showed that the "sense-and-adapt" approach has the same performance experimentally as it did computationally. The research presented in this dissertation is a stepping stone towards further development of the concept, which includes modeling the system in the Simulink environment and flight experiments using uninhabited aerial vehicles.
Analytical and Computational Investigations of Airfoils Undergoing High-Frequency Sinusoidal Pitch and Plunge Motions at Low Reynolds Numbers
(2008-11-10) McGowan, Gregory Zar; Dr. Harvey Charlton, Committee Member; Dr. Jack Edwards, Committee Member; Dr. Hassan Hassan, Committee Member; Dr. Ashok Gopalarathnam, Committee Chair
Current interests in Micro Air Vehicle (MAV)technologies call for the development of aerodynamic-design tools that will aid in the design of more efficient platforms that will also have adequate stability and control for flight in gusty environments. Influenced largely by nature MAVs tend to be very small, have low flight speeds, and utilize flapping motions for propulsion. For these reasons the focus is, specifically, on high-frequency motions at low Reynolds numbers. Toward the goal of developing design tools, it is of interest to explore the use of elementary flow solutions for simple motions such as pitch and plunge oscillations to predict aerodynamic performance for more complex motions. In the early part of this research, a validation effort was undertaken. Computations from the current effort were compared with experiments conducted in a parallel, collaborative effort at the Air Force Research Laboratory (AFRL). A set of pure-pitch and pure-plunge sinusoidal oscillations of the SD7003 airfoil were examined. Phase-averaged measurements using particle image velocimetry in a water tunnel were compared with computations using two flow solvers: i) an incompressible Navier-Stokes Immersed Boundary Method and ii) an unsteady compressible Reynolds-Averaged Navier-Stokes (RANS) solver. The motions were at a reduced frequency of $k = 3.93$, and pitch-angle amplitudes were chosen such that a kinematic equivalence in amplitudes of effective angle of attack (from plunge) was obtained. Plunge cases showed good qualitative agreement between computation and experiment, but in the pitch cases, the wake vorticity in the experiment was substantially different from that predicted by both computations. Further, equivalence between the pure-pitch and pure-plunge motions was not attained through matching effective angle of attack. With the failure of pitch/plunge equivalence using equivalent amplitudes of effective angle of attack, the effort shifted to include pitch-rate and wake-effect terms through the use of analytical methods including quasi-steady thin-airfoil theory (QSTAT) and Theodorsen's theory. These theories were used to develop three analytical approaches for determining pitch motions equivalent to plunge motions. A study of variation in plunge height was then examined and followed by a study of the effect of rotation point using the RANS solver. For the range of plunge heights studied, it was observed that kinematic matching between plunge and pitch using QSTAT gave outstanding similarities in flow field, while the matching performed using Theodorsen's theory gave the best equivalence in lift coefficients for all cases. The variation of rotation point revealed that, for the given plunge height, with rotation point in front of the mid-chord location, all three schemes matched flow-field vorticity well, and with rotation point aft of the mid-chord no scheme matched vorticity fields. However, for all rotation points (except for the mid-chord location), CFD prediction of lift coefficients from the Theodorsen matching scheme matched the lift time histories closely to CFD predictions for pure-pitch. Combined pitch and plunge motions were then examined using kinematic parameters obtained from the three schemes. The results showed that QSTAT nearly cancels the vortices emanating from the trailing edge. Theodorsen's matching approach was successful in generating a lift that was close to constant over the entire cycle. Additionally this approach showed the presence of the reverse Karman vortex sheet through the wake. Combined pitch/plunge motions were then analyzed, computationally and experimentally, with a non-zero mean angle of attack. All computational results compared excellently with experiments, capturing vorticity production on the airfoil's surface and through the wake. Lift coefficient through a cycle was shown to tend toward a constant using Theodorsen's parameters, with the constant being dependent on the initial angle of attack. This result points to the possibility of designing an unsteady motion to match a given flight-condition requirement.
Analytical and Experimental Approaches to Airfoil-Aircraft Design Integration
(2002-06-07) McAvoy, Christopher William; Dr. Kailash C. Misra, Committee Member; Dr. Ndaona Chokani, Committee Member; Dr. Ashok Gopalarathnam, Committee Chair
The aerodynamic characteristics of the wing airfoil are critical to achieving desired aircraft performance. However, even with all of the advances in airfoil and aircraft design, there remains little guidance on how to tailor an airfoil to suit a particular aircraft. Typically a trial-and-error approach is used to select the most-suitable airfoil. An airfoil thus selected is optimized for only a narrow range of flight conditions. Some form of geometry change is needed to adapt the airfoil for other flight conditions and it is desirable to automate this geometry change to avoid an increase in pilot workload. To make progress in these important aeronautical needs, the research described in this thesis is the result of seeking answers to two questions: (1) how does one efficiently tailor an airfoil to suit an aircraft? and (2) how can an airfoil be adapted for a wide range of flight conditions without increased pilot workload? The first part of the thesis presents a two-pronged approach to tailoring an airfoil for an aircraft: (1) an approach in which aircraft performance simulations are used to study the effects of airfoil changes and to guide the airfoil design and (2) an analytical approach to determine expressions that provide guidance in sizing and locating the airfoil low-drag range. The analytical study shows that there is an ideal value for the lift coefficient for the lower corner of the airfoil low-drag range when the airfoil is tailored for aircraft level-flight maximum speed. Likewise there is an ideal value for the lift coefficient for the upper corner of the low-drag range when the airfoil is tailored for maximizing the aircraft range. These ideal locations are functions of the amount of laminar flow on the upper and lower surfaces of the airfoil and also depend on the geometry, drag, and power characteristics of the aircraft. Comparison of the results from the two approaches for a hypothetical general aviation aircraft are presented to validate the expressions derived in the analytical approach. The second part of the thesis examines the use of a small trailing-edge flap, often referred to as a 'cruise flap,' that can be used to extend the low-drag range of a natural-laminar-flow airfoil. Automation of such a cruise flap is likely to result in improved aircraft performance over a large speed range without an increase in the pilot work load. An approach for the automation is presented here using two pressure-based schemes for determining the optimum flap angle for any given airfoil lift coefficient. The schemes use the pressure difference between two pressure sensors on the airfoil surface close to the leading edge. In each of the schemes, for a given lift coefficient this nondimensionalized pressure difference is brought to a predetermined target value by deflecting the flap. It is shown that the drag polar is then shifted to bracket the given lift coefficient. This non-dimensional pressure difference can, therefore, be used to determine and set the optimum flap angle for a specified lift coefficient. The two schemes differ in the method used for the nondimensionalization. The effectiveness of the two schemes are verified using computational and wind-tunnel results for two NASA laminar flow airfoils. To further validate the effectiveness of the two schemes in an automatic flap system, a closed-loop control system is developed and demonstrated for an airfoil in a wind tunnel. The control system uses a continuously-running Newton iteration to adjust the airfoil angle of attack and flap deflection. Finally, the aircraft performance-simulation approach developed in the first part of the thesis is used to analyze the potential aircraft performance benefits of an automatic cruise flap system while addressing trim drag considerations.
Computational Analysis of Circulation Control Airfoils
(2004-10-26) McGowan, Gregory Zar; Dr. Harvey Charlton, Committee Member; Dr. Ashok Gopalarathnam, Committee Chair; Dr. William Roberts, Committee Member
Current projections for future aircraft concepts call for stringent requirements on high-lift and low cruise-drag. The purpose of this study is to examine the use of circulation control, through trailing edge blowing, to meet both requirements. This study was conducted in two stages: (i) validation of computational fluid dynamic procedures on a general aviation circulation control airfoil and (ii) a study of an adaptive circulation control airfoil for controlling lift coefficients in the low-drag range. In an effort to validate computational fluid dynamics procedures for calculating flows around circulation control airfoils, the commercial flow solver FLUENT was utilized to study the flow around a general aviation circulation control airfoil. The results were compared to experimental and computational fluid dynamics results conducted at the NASA Langley Research Center. This effort was conducted in three stages: (i) a comparison of the results for free-air conditions to those from previously conducted experiments, (ii) a study of wind-tunnel wall effects, and (iii) a study of the stagnation-point behavior. In general, the trends in the results from the current work agreed well with those from experiments, some differences in magnitude were present between computations and experiments. For the cases examined, FLUENT computations showed no noticeable effect on the results due to the presence of wind-tunnel walls. The study also showed that the leading-edge stagnation point moves in a systematic manner with changes to the jet blowing coefficient and angle of attack, indicating that this location can be sensed for use in closed-loop control of such airfoil flows. The focus of the second part of the study was to examine the use of adaptive circulation control on a natural laminar flow airfoil for controlling the lift coefficient of the low-drag range. In this effort, adaptive circulation control was achieved through blowing over a small mechanical flap that can be deflected up or down. Such a blown trailing-edge flap allows for control of the jet direction to be independent of the jet momentum coefficient. This study was performed in two stages. In the first study, a two-dimensional thin-airfoil thin-jet theory and accompanying computer program was developed. With this method, changes to the airfoil ideal lift coefficient were studied for various jet blowing rates and angles showing that the ideal lift coefficient could be adjusted by varying either the blowing rate or the flap angle. In the second stage, a hybrid computational study was conducted. This hybrid method involved the use of the CFL3D Reynolds-averaged Navier-Stokes code in conjunction with an integral boundary layer method. The surface pressure distributions for the airfoil were determined using CFL3D. Using these pressure distributions, the boundary layer transition locations were calculated using the integral boundary layer method. The transition-location data was then used to determine the lift-coefficient range in which extended laminar flow could be achieved for cases with and without blowing. The results of this study confirmed that, in addition to flap angle, blowing across the trailing edge flap can be used to adjust the range of lift coefficients over which extensive laminar flow can be achieved. The blown trailing-edge flap was shown to be more effective at altering the location of the low-drag range than a cruise flap with no blowing. In addition, the blown flap eliminates separation off the flap at high flap angles.
Ideal Lift Distributions and Flap Settings for Adaptive Tailless Aircraft
(2006-01-11) Cusher, Aaron Anthony; Dr. Ashok Gopalarathnam, Committee Chair; Dr. Robert T. Nagel, Committee Member; Dr. Larry M. Silverberg, Committee Member
With ever increasing maturity in the field of subsonic aircraft design, there exists the desire to tailor the performance of an aircraft to suit specific flight conditions. This has led to several adaptive-wing approaches which seek to improve aircraft performance by changing the wing shape in flight, resulting in drag reduction. One such adaptive-wing approach that has gained considerable popularity is the use of multiple spanwise trailing-edge flaps which are used to optimally distribute the lift of the wing such that drag is minimized. Recent research has been conducted utilizing such a technique applied to an aircraft with a wing-tail configuration and discussed the need to extend these methods to tailless, or all-wing, aircraft, thereby improving design possibilities to include unconventional configurations. The current work explores tailless aircraft configurations which utilize multiple trailing-edge flaps for the purpose of wing adaptation and drag reduction. As with all tailless aircraft design, the trailing-edge flap settings, and thus wing lift distribution, must be solved while satisfying a longitudinal-pitching-moment constraint in order to ensure longitudinal stability and trim. This is due to the lack of a secondary horizontal surface, such as a tail or canard, which is typically used for stability and trim purposes. The current work implements a numerical approach which was developed to solve for the optimal flap scheduling of a wing with multiple trailing-edge flaps for various flight conditions. Theory presented by R.T. Jones was used as a starting point to solve for the target lift distribution resulting in minimized induced drag with a pitching moment constraint. Also utilized were the ideas of basic and additional lift, as well as thin airfoil theory relations in order to reduce both induced and profile drag by the redistribution of wing lift along its span. The cases were solved with longitudinal trim and lift constraints. The results were presented for planar, tapered wings with multiple quarter-chord sweep angles as well as multiple airfoil sections in order to verify the theory and gain insight into design capabilities and trends. It has been shown by these results that such adaptive wing methods are applicable and beneficial to tailless aircraft configurations, as reductions in both induced and profile drag have been achieved. In addition, the method is successful for achieving longitudinal trim, and was explored successfully for multiple static margins in order to test the consequence of different longitudinal stability considerations.
Modeling and Analysis of Active Turbulators on Low Reynolds Number Unmanned Aerial Vehicles
(2008-11-07) Short, Seth Ryan; Dr. Larry Silverberg, Committee Member; Dr, Pierre A. Gremaud, Committee Member; Dr. Ashok Gopalarathnam, Committee Chair
ABSTRACT SETH RYAN SHORT. Modeling and Analysis of Active Turbulators on Low Reynolds Number Unmanned Aerial Vehicles. (Under the direction of Dr. Ashok Gopalarathnam.) The current research explores an approach for obtaining performance gains on small unmanned aerial vehicles (UAVs) operating in a low Reynolds number flight regime. Performance gains are sought through the use of an adaptive â€œtrippingâ€ or turbulator system for reduction of the drag resulting from laminar separation bubbles. Because laminar bubbles change in strength and chordwise location with changes in aircraft operating conditions, an adaptive bubble-control system is necessary to eliminate or reduce the adverse effects of the bubble over a large speed range. In the current effort, the active system is modeled using the concept of an active ideal turbulator. This active ideal turbulator model, developed in the current effort, is implemented using the XFOIL code. The results from this model are compared with experiments for fixed and active turbulators and it is shown that the model is sufficiently good for use in assessing the impact of the technology. The effect of the drag reduction on the aircraft performance is studied in this effort using three notional unmanned aerial vehicles of different sizes. For this pur- pose, models for determining the aircraft drag and power-system characteristics have been developed. The improvements in aircraft endurance, range and rate of climb are studied. The results show that the active system can be used to achieve significant performance improvements when the bubble drag on the baseline air- foil is large. Results are also presented for the weight and power-consumption penalties of the active system at which the drag-reduction benefits are negated.
On the Use of Wing Adaptation and Formation for Improved Aerodynamic Efficiency
(2005-05-11) King, Rachel Marie; Dr. Ashok Gopalarathnam, Committee Chair; Dr. Hassan Hassan, Committee Member; Dr. Jack Edwards, Committee Member; Dr. Xiao-Biao Lin, Committee Member
There is a continuous effort to improve the performance and efficiency of today's aircraft, and the reduction of aircraft drag has been the primary focus of many aerodynamicists. In the current research, two different and innovative approaches for aircraft drag reduction are examined. These approaches are: (1) multiple spanwise trailing-edge flaps, and (2) formation and ground-effect flight. The main goal of this dissertation was to assess the drag benefits of the two approaches, in an effort to explore their potential for use on future aircraft. The first approach of using multiple trailing-edge flaps has the potential for application on aircraft in the near future. By using multiple trailing-edge flaps along the wing span, it is possible to redistribute the spanwise lift distribution to suit the flight condition. For this research, a numerical approach was developed for determining optimum lift distributions on a wing with multiple trailing-edge flaps for various flight conditions. The objective of the approach was to determine the flap angles that will reduce the drag at 1-g flight conditions, and constrain the wing root-bending moment at high-g conditions to not exceed a specified value. The approach uses the concept of additional and basic lift distributions, and the proper use of a trailing-edge flap for redistributing the aerodynamic loads to bring about a minimum in profile and induced drag. The results for the flap-angle distributions are presented for a planar and a nonplanar wing, along with post-design analysis and aircraft performance simulations used to validate the optimum flap-angle distributions determined using the numerical approach. It is shown that the approach is effective in determining optimum flap angles for reducing both profile and induced drag over a wide range of flight conditions. Performance benefits due to using the optimum flap angles are shown when compared to the zero-flap case. In addition, the trailing-edge flaps were found to be successful in relieving the wing root-bending moment at high-g flight conditions, which can be used to reduce wing weight. When examining formation and ground-effect flight as another approach for aircraft drag reduction, an optimum-downwash approach using a vortex-lattice implementation was used to study formations of wings loaded optimally for minimum induced drag with roll trim. An exact approach was also developed to examine the drag of elliptically-loaded wings in formation. The exact approach allows for decomposition of the benefits by considering the mutual-interference contributions from different pairs of wings in a formation. The results show that elliptically-loaded wing formations have nearly the same drag as optimally-loaded wing formations. For a formation of planar wings, in or out of ground effect, the optimum lateral separation corresponds to a 9%-span overlap of wing tips. At this optimum lateral separation, a formation of 25 elliptically-loaded wings flying out of ground effect experiences an 81% drag reduction compared to 25 wings flying in isolation. For large formations, in or out of ground effect, multiple local optima are seen for the lateral separation. Large formations experience small additional benefits due to ground effect even at relatively large ground clearances of four wing spans. The shape of vee-formations, for equipartition of drag benefits, is found to be nearly independent of flight in or out of ground effect. Overall, both approaches for aircraft drag reduction show potential for significant drag savings. It is believed that the presented research will further increase interest in such flight techniques, and thus advance their progression toward becoming viable solutions for drag reduction on future aircraft.
Optimum Flap Angles for Roll Control on Wings with Multiple Trailing-Edge Flaps
(2007-12-04) Segawa, Hidehiro; Dr. Agnes Szanto, Committee Member; Dr. Ashok Gopalarathnam, Committee Chair; Dr. Charles Hall, Committee Member
This research effort explores the use of multiple trailing-edge flaps for efficiently generating rolling moment on aircraft. Using the concept of basic and additional lift distributions, the induced drag of the wing is expressed in terms of the flap angles. The theory of relative extrema is then used to determine the optimum flaps angles for minimum induced drag with a constraint on the rolling moment. By setting the mean of the flap angle for operation of the wing within the low-drag range, profile drag is also minimized. The general methodology can also be used on tailless aircraft and to study the effect of failure modes such as a stuck flap. The results show that multiple flaps can be used to generate rolling moments with lower drag than when ailerons are used. They also provide redundancy that helps efficiently handle control failures such as stuck flaps. The current research serves as a starting point for further investigation into the use of multiple flaps for efficient aircraft control.
The Stability and Control of an Aircraft with an Adaptive Wing
(2005-02-02) Vosburg, Victor Jay; Dr. Arkady Kheyfets, Committee Member; Dr. Ashok Gopalarathnam, Committee Chair; Dr. Charles E. Hall, Jr., Committee Member
With increasing interest in the use of adaptive lifting surfaces for improved aircraft performance, it is necessary to study the stability and control characteristics of an aircraft with an adaptive wing. This research builds on recent development of an automated cruise flap for adapting a wing shape to achieve low drag over a large lift range. Such an automated cruise flap system was shown to have unusual lift and pitching moment curves, which prompted the need for studying the effect on aircraft stability and control. In this thesis, the static stability considerations are used to show that when an automated cruise flap is used on an airfoil to continuously adjust the flap to the optimum angle, there is a need for an accompanying controller that achieves the desired lift coefficient by adjusting the airfoil angle of attack. Likewise, when an automated flap is used on an aircraft wing, there is a need for an accompanying airspeed controller to adjust the elevator. The thesis presents Simulink models for analyzing the dynamic behavior of an aircraft with an automated flap. Two schemes were studied for the flap and elevator controllers. The most desirable results were obtained when the flap and elevator controllers were coupled so that trim changes due to the flap are immediately compensated by adjustment of the elevator angle. The results of the simulation show that the aircraft does not exhibit any undesirable behavior with the automated cruise flap. The study, therefore, provides the confidence needed to implement the system on an uninhabited aerial vehicle.
Static Aeroelasticity Considerations in Aerodynamic Adaptation of Wings for Low Drag
(2006-05-10) Shipley, Edward Nicholas Jr.; Dr. Charles E. Hall, Jr., Committee Member; Dr. Pierre A. Gremaud, Committee Member; Dr. Ashok Gopalarathnam, Committee Chair
This thesis presents a methodology to calculate the flap angles necessary to reproduce a desired lift distribution for an elastic wing with multiple trailing-edge flaps. The methodology presented builds upon the adaptive wing methodology for rigid wings developed by King and Gopalarathnam. Relevant equations from thin airfoil theory and beam theory are presented, and are then used to develop the solution procedure for the determination of flap angles for an elastic wing. Examples using an elliptic lift distribution are presented, and the effect of wing elasticity on the calculated flap angles and lift distributions is shown. Post-design analysis shows that the flap angles from the current method are successful in achieving an elliptical loading while accounting for using torsional deformations. The methodology is verified by comparing the twist predicted by the methodology to the twist predicted by a finite element analysis. Finally, an example demonstrating how the methodology handles flap effectiveness and flap reversal is provided.