The flow over a straight and a cranked lambda wing planform having a flat upper surface, and a sharp leading edge was investigated experimentally, using flow visualization and particle image velocimetry, while measuring the forces and moments on the model. A single steady jet located just inboard of the crank provided the wing with nose up or down pitch control authority and enabled the wing to be trimmed at high angles of incidence. To achieve this result the jet was rotated at various angles relative to the oncoming free stream therefore disrupting the feedback mechanism existing between the primary leading-edge vortex and the separated shear layer that feeds it with vorticity. In the absence of a crank, this feedback is provided by the vortical layer (secondary vortex) that is forced upward by the primary vortex and is entrained into the separated shear layer. In the presence of the crank, a new vortex created downstream of a crank serves a similar purpose as the shear layer. The experiments were carried out in a low-speed wind tunnel at 25 meters/second therefore providing a Reynold number of approximately 1.2 x 10(6) and a Mach number smaller than 0.1.

Tests were carried out in a university-type atmospheric wind tunnel on a NATO-designed Swept Wing Flow Test (SWiFT) model at incompressible Mach numbers and Reynolds numbers ([Formula: see text]) ranging from [Formula: see text] to [Formula: see text]. The initial purpose was to compare the current results with those obtained in the National Transonic Facility (NTF) to assess the significance of [Formula: see text] on the physics of the flow. Following the excellent agreement in force balance data, the investigation expanded to study the leading-edge vortex (LEV) that impacts the pitch dependence on incidence and the flow unsteadiness after its liftoff from the surface, potentially leading to buffet. Oil flow visualization identified the liftoff location and its effect on pitch, while smoke visualization revealed the unsteadiness of the process. Quantitative analysis using particle image velocimetry (PIV) and hot wire anemometry exposed large periodic oscillations contributing to a turbulence level that is much larger than is known in a typical turbulent boundary layer on the verge of separation. Proper orthogonal decomposition (POD) confirmed the unsteady behavior at incidence angles exceeding vortex liftoff that raised the significance of a convective length scale and Strouhal number on this type of blended-wing–body configuration.

Tests were carried out in a university-scale atmospheric wind tunnel on a NATO designed Swept Wing Flow Test (SWiFT) model at incompressible Mach numbers and Reynolds numbers (Re) ranging from Re=0.2 x 10(6) to Re= 2 x 10(6). The initial purpose was to compare results with those from the National Transonic Facility (NTF) to assess the significance of Re. After comparing the force balance data of the two tests and seeing that an increase in Re did not result in novel phenomena, the evolution of the Leading-Edge Vortex (LEV) and its lift-off from the surface due to the LE crank was investigated. Based on this information, a single small jet was chosen to alter the forces and moments acting on this model and provide information about the interaction between the small jet and the LEV. Preliminary results indicate that the location and orientation of a jet actuator are as important as the jet momentum, therefore requiring optimization for different parts of the flight envelope. To ascertain the significance of the crank angle and its location on the wing relative to other design parameters, a simplified flat-top lambda wing planform with a sharp leading edge was tested at Re= 1.2 x 10(6). Oil flow visualization exposed the differences between these two models that were exposed by probing the LEV by Particle Image Velocimetry (PIV). The use of PIV revealed extremely large periodic oscillations in the flow that were associated with localized separation and vortex lift-off, providing a reason for flutter. These oscillations were observed on both models, and they were attributed mostly to the crank. Consequently, a substantial reduction of these oscillations became an added task for the useful application of Active Flow Control (AFC).

The broad practical purpose of this study was to explore the effectiveness of Active Flow Control (AFC) in controlling a tailless aircraft model called the SWIFT that was designed in the UK under the auspices of NATO's Advanced Vehicle Technology program. Wind tunnel tests focused on the maintenance of trim at high lift coefficients using a single supersonic jet emanating from a small nozzle at an appropriate location and orientation relative to the free stream. Although none of the AFC parameters was properly optimized the trimmed lift coefficient was tripled. To understand the effect of AFC on the flow over such a wing, flow visualization and Particle Image Velocimetry (PIV) were used. Proper Orthogonal Decomposition applied to this data revealed unsteady behavior in the vortex lift-off and its advection over the outer wing that was effectively subdued by AFC thus extending the range of incidence angles that could be flown. Preliminary tests were carried out on yawing and rolling moments exploring the possibility of replacing conventional control surfaces on an airplane by AFC. Wind tunnel data were acquired at Reynolds number of 1.6x10(6) based on the root chord of the semi-span model used.

The pressure distribution on a surface, over which a wall-jet is blowing, is altered by the wall jet's entrainment. It renders the boundary layer approximation – that justifies the use of an inviscid flow solution to determine the pressure over the surface – invalid. Thus, in order for Active Flow Control (AFC) by blowing to become a viable technology, some of the preconceptions associated with Boundary Layer Control (BLC) for many decades must be discarded. In particular, the momentum coefficient used to characterize BLC should be replaced by another variable that represents a conserved quantity that is independent of specific installations. Injected momentum is a vector quantity whose effect on a surface like a wing depends on its specific design, location, and orientation. Therefore, a new approach is proposed based on the AFC system's power consumption and its mass flowrate. Moreover, all flow installations suffer from unavoidable losses, which must be determined in an unambiguous manner, allowing for an impartial comparison of AFC systems. The present article provides examples from tests carried out at various universities and at NASA, exposing some popular misconceptions. It does not provide a design tool due to the complexity of the needed approach, but a method to assess the efficacy and efficiency of an evolving platform that includes AFC is suggested.

The Swept Wing Flow Test (SWIFT) is a tailless unmanned combat aerial vehicle (UCAV) model to be tested at high Reynolds numbers in NASA’s National Transonic Facility. The model is designed around a [Formula: see text]-shaped wing with a single, large crank at its leading edge (LE). It suffers from an unstable nose-up pitch departure resulting from flow separation augmented by the LE crank. A small-scale, modular wind tunnel model ([Formula: see text]) was built that allowed for changes in the crank angle by increasing the outboard wing sweep. Eliminating the crank entirely increased the [Formula: see text] and changed the sign of pitch departure, thus exposing the significance of the LE crank. The model was equipped with sweeping jet actuators that could be individually enabled by valves located at the actuator inlets, allowing one to explore the role of active flow control (AFC) in expanding the model’s longitudinal stability margins and controlling its yaw while being cognizant of the coupling between changes in the model’s planform and their effect on AFC. Test results indicated that selective actuation depending on the model’s attitude modified the flow and dramatically increased the trimmed [Formula: see text], while further suggesting that the actuation should dynamically change with incidence to improve AFC efficacy.

Purpose This paper aims to address shortcomings of current tiltrotor designs, such as the small aspect ratio of the wings, large download and the close proximity of the rotor tips. It also aims to avoid the complex transition of tiltrotors to normal airplane mode. Design/methodology/approach This design combines tiltrotor and tiltwing aircraft designs into a hybrid that is augmented by active flow control, using a gimbaled channel wing for attitude control in hover. Findings The proposed hybrid design is based on experimental results of components that were tested individually for potential use in hover and steep ascend from a stationary position. Originality/value This research was inspired by the extremely short take-off of the V-22, when its rotors were tilted forward. It combines several design approaches in a unique way to achieve extremely short take-off capabilities combined with high-speed and reduced maintenance costs.

Boundary-layer control (BLC) of turbulent flows aims to compensate for mean frictional losses on a frozen configuration, while active flow control (AFC) seeks to alter the mean flow by exploiting instabilities. Since AFC performance is judged by the energy required to achieve a prescribed goal, it must be introduced into the aircraft design process at the early conceptual stages. For blowing applications, deciding whether to draw from engine bleed or install smaller compressors adjacent to the actuators is an important system-level consideration that requires the separation of the loss in the air supply system from the input that triggers and amplifies the exterior flow instabilities. The current paper exposes the inadequacy of the momentum coefficient that has primarily been used in the past to assess AFC efficacy and recommends two easily measurable fluid power coefficients to replace it. These enable direct comparisons of actuator efficiencies and system efficiencies, creating a link between the physics of flow control and system integration. This link is demonstrated first on a two-dimensional configuration that uses steady blowing, which is later extended to the use of sweeping jet actuators on a modular swept wing of finite aspect ratio.

View Video Presentation: https://doi.org/10.2514/6.2022-2426.vid Highly three-dimensional vortical structures and flow separation phenomena present on swept back wings result in complex lift, drag, and pitching moment characteristics. While this may create challenges in the traditional aircraft design process, it provides a promising platform for integrating active flow control (AFC). Subsonic wind tunnel investigations are performed on a highly configurable swept wing model equipped with a spanwise array of sweeping jet actuators at 80% chord. Relating force balance measurements and surface tuft visualizations shed light on how oscillatory blowing interacts with various flow structures over a wide range of geometric and actuator configurations. We show that the lift-enhancing capabilities of AFC is closely related to the strength of vortex lift. We also demonstrate that the unstable pitch behavior of highly swept wings can be mitigated through AFC, significantly extending the flight regime over which the aircraft is trimmed. The optimal actuator distribution depends on the control objective and type of flow conditions present. Detailed understanding of these flow interaction mechanisms will be crucial for effective utilization of AFC technology on future swept wing aircraft.

There is a conceptual difference between Boundary Layer Control (BLC) and Active Flow Control (AFC). The former aims to compensate for frictional losses in the boundary layer, while the latter seeks to exploit instabilities in the flow to alter its character. The momentum coefficient Cμ has been the parameter of choice for energizing the boundary layer and enhancing the circulation by blowing. While it has proven useful in cataloging some data, it is not a universally valid single parameter describing the complex flow over wings. Furthermore, the jet velocity and density required for its calculation are difficult to measure in situ, necessitating unrealistic assumptions or extensive calibrations. The best quantity to be used is one that defines the energy requirements of the entire AFC system while realizing its lack of uniqueness. Experimental investigations on a thick flapped airfoil and a simple swept wing demonstrate that such a quantity is more suitable for design purposes because it shifts the perspective from detailed jet flow to system needs. Although this figure-of-merit approach black boxes flow interaction details between the jet and the freestream, system-based coefficients can still provide various physical insights in their capacity as cataloguing parameters.

The extremely short takeoffs of the V-22 from the USS Iwo-Jima when its rotors were tilted forward by 19º relative to the traditional 88o used in hover, inspired a series of wind tunnel tests simulating the phenomenon that enabled the V-22 to takeoff within 30ft of run distance (only ½ of its body length. It transpired that the download force that normally consumes 12-13% of the rotor thrust was substantially reduced by the tilt forward of the rotor, while the Suck-Back (SB, lift force in wind tunnel coordinates) was increased. The introduction of Active Flow Control (AFC) reduced the download even more while increasing the SB. This suggested that a hybrid AFC augmented airplane configuration consisting of fixed wing attached to a fuselage and a tilt wing attached to the rotor or nacelle are most effective in hover. Namely, it can carry more weight per installed thrust. When some of these ideas were tested on a powered model in hover, its download was reduced by approximately 60% relative to an identical tilt rotor model in the absence of AFC.

Full-scale wind tunnel tests were carried out on a Boeing 757 vertical tail equipped with 37 sweeping jet actuators on the starboard side along the rudder hinge line. The tests were performed at the National Full-Scale Aerodynamics Complex (NFAC) 40- by 80-Foot Wind Tunnel (40 x 80) at NASA Ames Research Center. The model was tested at a nominal airspeed of 100 knots across rudder deflections and sideslip angles that covered the airplane's emergency single engine climb. Active flow control (AFC) was optimized at the maximum rudder deflection of 30 degrees and sideslip angles of 0 degrees and -7.5 degrees. A threshold success criterion of 20% increase in the maximum side force was exceeded at largest rudder deflection using 12 sweeping jet actuators in the absence of sideslip but barely reached it with 18 actuators at the maximum sideslip at a momentum coefficient input of approximately 0.5%; consequently, 31 actuators were selected for most tests. AFC caused significant increases in suction pressure and associated side force on the actuated side. The successful demonstration of this application cleared the way for a subsequent flight demonstration on the Boeing 757 ecoDemonstrator in 2015.

Free AccessIntroductionIntroduction to the Flow Control Virtual CollectionDavid Greenblatt, Edward A. Whalen and Israel J. WygnanskiDavid GreenblattProfessor, Faculty of Mechanical Engineering, Technion–Israel Institute of Technology, 3200003 Haifa, Israel; . Associate Fellow AIAA., Edward A. WhalenFlow Control Actuators Manager, Boeing Research & Technology, Boeing Company, Hazelwood, Missouri 63042. Associate Fellow AIAA. and Israel J. WygnanskiProfessor, Aerospace and Mechanical Engineering Department, The University of Arizona, Tucson, Arizona 85721. Fellow AIAA.Published Online:2 Aug 2019https://doi.org/10.2514/1.J058507SectionsRead Now ToolsAdd to favoritesDownload citationTrack citations ShareShare onFacebookTwitterLinked InRedditEmail AboutFlow Control Virtual Collection: https://arc.aiaa.org/vc/flowcontrolI. Brief Historical Development of Fundamental ConceptsThe study and application of active flow control (AFC) has its roots in the discovery and formulation of the boundary layer by Prandtl in 1904 [1] and, in particular, the control of boundary-layer separation. Traditional boundary layer control (BLC) involved the addition of momentum flux or the removal of the low-momentum layer in order to restore the inviscid pressure distribution associated with attached boundary layers [2]. When added momentum flux produces lift forces that exceeded the inviscid limit, this is termed circulation control or super-circulation [3]. BLC reached its apogee in the mid-20th century [3,4], leading to the "blown flaps," which were installed on some production aircraft (e.g., the Lockheed F-104 and some models of the Mikoyan-Gurevich MiG-21). However, with the dawn of the space race, R&D funding was dramatically cut back and BLC development came to a virtual standstill [5].The 1973 oil crisis led to focused efforts for increasing efficiency and, in particular, boundary-layer drag reduction was researched extensively. Most active methods focused on the large disparities between laminar and turbulent boundary-layer drag, with the objective of "delaying" transition, that is, shifting it further downstream. Apart from obvious geometric (pressure gradient) approaches, active surface suction though a porous skin/slits or surface temperature variations were evaluated. This often involved the identification and elimination of flow instabilities that led to transition. To date, suction though a porous skin has shown the greatest promise, with extensive flight testing [6], but commercial aircraft application has been limited [7]. Many modern approaches aim to delay transition via the introduction of controlled perturbations [8,9], especially on swept back wings. Active methods—such as wall suction or blowing, oscillating walls, and sublayer blowing—have also been applied directly to turbulent boundary-layer drag reduction. These methods are designed to manipulate near-wall low-speed streaks, and net drag reductions of up to 65% have been reported in recent wind tunnel tests [10].Until the 1970s, the meaningful control or manipulation of turbulent flows by means of small perturbations was considered to be impossible, with conventional wisdom asserting that "a turbulent flow forgets its origin." However, dedicated experiments on turbulent shear layers revealed that instability-driven "coherent structures" dominate the flowfield [11] and, as a consequence, "small" or O(ε) perturbations can produce "large" O(1) overall changes [12]. These findings subsequently spawned a large body of research, covering shear layers, jets, separation control on airfoils and flaps, cavity flows, and combustion processes. Although instability-driven flow control has been demonstrated in flight tests [13], it has not yet reached the application stage.For reasons of safety, simplicity, reliability, and power consumption, AFC becomes viable only when a decisive aerodynamic advantage can be demonstrated over passive flow control (PFC) techniques—for example, slotted-flaps and vortex generators (VGs) for separation control [14,15]; chevrons for jet noise reduction [16,17]; and riblets for near-wall drag reduction [18,19]. Moreover, apart from aerodynamic benefits, actuators or actuation techniques must be safe, small, light, and robust with associated low energy consumption [20]. To realize the full potential of AFC, it must be integrated into the vehicle design and not introduced as a postdesign add-on. This requires a full three-dimensional flowfield analysis and hence a breakaway from traditional two-dimensional paradigms. From a control theory standpoint [21], the vast majority of AFC investigations are considered to be the "open loop control." Implicit in this approach is the understanding that any successful and viable technique will ultimately be incorporated into automatic flight control systems.This Virtual Collection of articles was conceived of with the intention of producing a cross section of the diversity and depth of modern flow control research. Contributing authors were encouraged to produce articles that could be used as reference material to both researchers and students and to present the philosophy underpinning their research. All articles in this Virtual Collection have been published online as regular contributions, but here they are presented as a cohesive unit. The Collection is broadly divided into five main sections: fundamental studies (Sec. II), airfoil and wing applications (Sec. III), actuators and actuation methods (Sec. IV), closed-loop control (Sec. V), and flight applications and testing (Sec. VI). Conclusions and outlook are presented in Sec. VII. Although this volume does not cover all aspects of flow control, it does summarize a cross section of activities that make up the broader field.II. Fundamental StudiesFundamental studies relating to drag reduction, shear layer control, separation control, and transition control continue to be the bedrock of flow control research. In the area of drag reduction, Corke and Thomas [10] present a comprehensive review of flow-control-based drag reduction strategies encompassing 50 years of research. These strategies aim to modify either the large-scale outer motions or the near-wall small-scale streaks using riblets, wall mass-flux, and transverse wall oscillations. A new and particularly effective approach is presented that introduces spanwise blowing pulses, achieved with newly developed pulsed-dc plasma actuators, to prevent the lift-up of near-wall streaks. The reduction in wall-normal vorticity that inhibits streak growth results in an unprecedented net drag reduction of up to 65%. McKeon et al. [22] describe the use of the "dynamic roughness" to manipulate turbulent boundary structures. Linear and nonlinear responses to perturbations are studied with a view to modifying the overall turbulent structure. Practical applications will need to leverage advances in microelectromechanical system (MEMS) devices or metamaterial actuators.Fundamental studies on the control of turbulent free shear layers, which are substantially altered by periodic excitation, are examined by Little [23], who studied thermal perturbations of free and reattaching incompressible turbulent shear layers using nanosecond-pulse-driven dielectric barrier discharge (NS-DBD) plasma actuators, where the perturbation mechanism is associated with Joule heating. The research focuses on differences in the flow receptivity to these perturbations when compared with traditional momentum-based actuation. Samimy et al. [24] produced effective control of turbulent supersonic jets using circumferentially mounted arc filament plasma actuators that produce localized short pulse-width (low power), high-bandwidth temperature spikes, and hence compression waves. They demonstrated similarities between perturbations of low-speed shear layers and supersonic jets, where the effects of excitation depend on Mach number and compressibility level. In cavities, where flow oscillations can generate large and potentially damaging pressure fluctuations (e.g., in weapons and landing-gear bays), Zhang et al. [25] showed that jet arrays deployed at the cavity leading edge resulted in significant surface pressure reductions, both experimentally and computationally. In the field of turbomachinery, Bons et al. [26] studied low-pressure turbine blade separation and secondary flow control strategies, including slot blowing, suction, steady, and pulsed vortex-generator jets (VGJs), where shear-layer excitations are exploited in the latter case. They showed effective flow control focusing on the fluid dynamics mechanisms. They point out, however, that flow unsteadiness, high turbulence levels, and secondary flows rendered these problems more challenging.The control of transition remains an important research area on swept and unswept configurations. Both theoretical and experimental approaches have been adopted to attenuate or cancel the amplitude of Tollmien–Schlichting (TS) waves. Using direct numerical simulations (DNS), Dadfar et al. [27] employed a linear quadratic Gaussian (LQG) controller to attenuate a TS wave packet (see Sec. V). The controller design was based on rows of downstream sensors, and plasma actuation was modeled by means of a volumetric force. Simon et al. [28] studied the active cancellation of TS waves by generating "counterwaves" using plasma actuators, with the objective of reducing friction drag by extending the laminar flow regime. Counterwaves cancel the incoming TS waves at approximately one wavelength downstream, with the velocity wave shape playing a secondary role. Saric et al. [29] employed discrete roughness elements (DREs) to study crossflow instabilities leading to boundary-layer transition on a swept wing model mounted on an aircraft. Linear stability analyses were used to determine the test matrix, and infrared thermography was used to determine the transition locations. Surprisingly, the transition location associated with a painted leading edge was unexpectedly farther aft than that on a highly polished one. Further research is required to determine if DREs can be employed as a viable laminar flow control technique.III. Airfoils, Wing, and Bluff BodiesThe original objectives of BLC related primarily to airfoil and wing separation and circulation control as a precursor to flight tests. Recent research adopts conventional methods in unique implementations, studies novel actuator systems, and introduces substantial three-dimensional effects. Warsop and Crowther [30] use the well-known Coandă effect to develop aircraft systems with sufficient control authority so as to eliminate conventional control surfaces. Their extensive body of research encompasses fundamental studies together with system integration and small-scale flying technology demonstrators (also see Sec. VI). Dolgopyat and Seifert [31] also studied flight control where suction and oscillatory blowing (SaOB) actuators are used to achieve this purpose by an unsteady mechanism. The system can increase lift or drag with concomitant minor effects on other aerodynamic parameters. Amitay and Gildersleeve [32] investigated synthetic jet actuators, low-aspect-ratio static and dynamic pin-type actuators, and a novel hybrid (failsafe) actuator that combines the two as they interact with a laminar boundary layer. Synthetic jet, pin, and hybrid actuator parameters were studied in detail and an array of passive cylindrical pins was shown to be effective for separation control. Sweeping jets were also shown to be effective over highly deflected flaps by Pack Melton et al. [33] (also see Sec. IV). Gursul and Wang [34] present a review of tip and edge vortex control by considering lift enhancement, drag reduction, flight control, noise reduction, buffet and wing rock, and the excitation of wake instabilities. A wide variety of techniques have been evaluated based on bleed, exploitation of multiple co-rotating vortices, as well as steady and unsteady blowing configurations. On afterbodies (cylindrical bodies with a slanted base), both passive (Gurney flap) and active (blowing) methods reduced drag.Jentzsch et al. [35] used a combination of sweeping jets and traditional control surfaces to trim and control a tailless aircraft model, showing that significantly smaller flap deflections are required with targeted local actuation. Visbal and Benton [36] demonstrated dynamic stall control on a pitching airfoil using high-fidelity wall-resolved large-eddy simulations and linear stability theory, where high-frequency perturbations are introduced near the leading edge. High-frequency perturbations excite convective instabilities in the laminar separation bubble, thereby affecting the bubble bursting and transition mechanisms. The approach shows potential for significantly ameliorating dynamic loads and moments. Keisar et al. [37] demonstrated the use of well-known pulsed dielectric barrier discharge (DBD) plasma actuators on a full-scale unmanned air vehicle tail panel, where the objective was to increase allowable cross winds during takeoff. The emphasis was on developing a lightweight and low-power system appropriate for flight vehicles. A resulting 65% increase in allowable crosswind takeoff speeds demonstrated that pulsed DBD plasma actuators are a viable and practical solution. Lambert et al. [38] used azimuthally mounted zero-mass-flux (synthetic) jets on the tail end of an axisymmetric bluff-body model with curved (Coandă) surfaces to demonstrate either suppression or augmentation of the motion-induced aerodynamic loads. The model was mounted and controlled by means of mounting wires that connected to servo motors with in-line load cells. Closed-loop control aspects of this research are discussed in Sec. V.IV. Actuators or Actuation MethodsThe development of effective and efficient actuators and actuation methods is critical for practical implementation of flow control [20]. Recently, sweeping jets (SJs), also referred to as fluidic oscillators, have received considerable attention. Hirsch and Gharib [39] employed a novel Schlieren system combined with a high-speed camera to study the internal dynamics of SJs. These data were used to validate a theory based on nondimensional frequency and jet velocity. They also studied two adjacent actuators with several separation distances and pressure ratios that produced underexpanded and transonic jets. Woszidlo et al. [40] measured SJ phase-resolved properties and determined that flow from the feedback channels enters the recirculation bubbles, and this is what drives the jet oscillation. They also developed scaling parameters for actuator sizing and different fluid densities. Pack Melton et al. [33] studied the effect of sweeping and discrete jets over a simply hinged flap of a semispan wing up to deflection angles of 60 deg. Sweeping jets increased lift, but were not able to completely eliminate the separation over the highly deflected flaps. Kara et al. [41] numerically studied the effect of SJ actuation on flow-separation over a wall-mounted hump model using 2-D unsteady RANS (URANS).Crittenden et al. [42] studied both fundamental and application aspects of combustion-powered actuators, namely, the ignition of a mixture of gaseous fuel and air to produce high-impulse jets. Performance was strongly dependent on chemical species composition and the actuator geometry. They then demonstrated the utility of the actuators for selected static and dynamic flow control applications.V. Modeling, Feedforward, and Feedback ControlControl theory, which generally involves some form of closed-loop or feedback, deals with the behavior of dynamical systems and is concerned with automating flow control. Williams and King [43] provide a detailed comprehensive overview of closed-loop flow control, particularly for the alleviation of unsteady aerodynamic loads associated with aircraft and ground vehicles. They describe the various models for unsteady aerodynamic loads (disturbance models), models for the dynamic response to control input, and surrogate measurements of forces and moments. Different closed-loop architectures are described for different feed-forward and feedback control applications. This serves as a framework for determining the practical bandwidth for aircraft gust alleviation using flow control. In a related study, Williams et al. [44] used pitching airfoil lift history with blowing-type flow control to obtain low-dimensional disturbance models for control based on a modified Goman-Khrabrov model, and a plant model for the lift response to blowing that took the form of a time-delay system. A feed-forward controller-based linearized version of the dynamic actuation model was capable of attenuating the lift oscillations by a factor of 10 over a wide range of frequencies.For transition control described in Sec. II, Dadfar et al. [27] used solutions of the linearized Navier–Stokes equations to develop a reduced-order model based on the Eigen-system Realization Algorithm (ERA), and the LQG controller was designed based on this model. Plasma actuation that was modeled by means of volumetric force and feedback was provided via downstream sensors. Seidel et al. [45] describe a heuristic approach to feedback flow control for two different problems, namely, a tangent ogive at large angles of attack and a shear layer over a backward-facing step. On the tangent ogive, using only four pressure sensors, a prescribed side force signal could be tracked. For the shear layer, a reduction of 40% in the optical path difference was achieved. Lambert et al. [38] used the synthetic jets described in Sec. III to control the 6-DOF bluff body motion using a trajectory tracking controller. An inner control loop used a proportional integral differential (PID) controller based on load cell sensors, with the controller driving the motors. An outer loop PID-based controller measured the body's position and facilitated accurate trajectory tracking with disturbance rejection, enabling real-time load and moment measurements.VI. Flight Applications and TestingA major objective of many flow control studies is flight testing and this is performed on both small-scale and large-scale demonstrators. Warsop and Crowther [30] demonstrated their Coandă effect and thrust vectoring systems (see Sec. III) on a series of small-scale flying technology demonstrators. These ranged in size and complexity (from less than 7 kg to 90 kg) and were vital for understanding system integration and estimating full-scale flight performance. Shmilovich and Vasta [46] review computational methods for practical simulations of flow control for flight vehicles, based on both Reynolds-averaged Navier–Stokes and lattice Boltzmann methods. Francis [47] reviews vectored thrust concepts and emphasizes the enhanced control authority at low airspeeds, including poststall maneuverability, tailless flight, and extremely short takeoff and landing. Integrated vectored thrust, used in conjunction with gas-turbine engines, including both mechanical and fluidic approaches, is reviewed.Extensive research was conducted on the Boeing 757 ecoDemonstrator aircraft, including computational simulations of individual SJ actuators (Shmilovich et al. [48]); small and full-scale vertical tail wind tunnel campaigns employing multiple actuator types and configurations; and subsequently a full-scale flight implementation on the ecoDemonstrator. This research was motivated by projected fuel savings attained by reducing the size of the vertical tail on commercial aircraft, which is sized to ensure controllability in the event of engine failure, but is larger than required for control and stability under normal flight conditions. Boeing and NASA investigated the effects of a pneumatic SJ system beginning with scale-model wind tunnel tests by Andino et al. [49], who showed that spanwise flow on the rudder was reduced and side force increased by up to 50% using a momentum coefficients of O(1%). In flight tests, Whalen et al. [50] showed that the SJ system provided a 14% increase in side force at the maximum tested rudder deflection and at critical sideslip angles. The reduced effectiveness in flight, relative to subscale testing, was attributed to the effects of Reynolds number on the separated flow state of the tail. On the basis of gas dynamics arguments, Hirsch et al. [51] showed experimentally and computationally that higher temperature air supplies to the actuators result in reduced mass flow rates without any degradation in control authority. More specifically, the mass flow rate is reduced for a given momentum coefficient.VII. ConclusionsRecent decades have witnessed significant advances across the spectrum of fundamental studies and basic research, actuator analysis and development, full-scale wind tunnel testing, and flight tests. By far, the greater emphasis is on active flow control (AFC). While fundamental studies focus on physical mechanisms, others show tremendous potential for drag reduction, control of supersonic jets, transition control, and the control of cavity oscillations. Basic research conducted on airfoils has now advanced to straight and swept wing applications, where traditional and nontraditional methods are employed for flight control, while novel and viable actuation methods are being developed for separation control. In the field of actuation, significant attention has been afforded to sweeping jets, encompassing basic research, computation modeling, wind tunnel campaigns, and full-scale flight tests.More than a hundred years after the demonstration of what we today call AFC, basic research has transitioned from the laboratory to full-scale flight tests. Passive flow control left an indelible mark on 20th-century aerodynamics, it is hoped that this century will see a decisive paradigm shift brought about by AFC.David GreenblattProfessor, Faculty of Mechanical Engineering, Technion–Israel Institute of Technology, 3200003 Haifa, Israel; [email protected]ac.il. Associate Fellow AIAA.Edward A. WhalenFlow Control Actuators Manager, Boeing Research & Technology, Boeing Company, Hazelwood, Missouri 63042. Associate Fellow AIAA.Israel J. WygnanskiProfessor, Aerospace and Mechanical Engineering Department, The University of Arizona, Tucson, Arizona 85721. Fellow AIAA. References [1] Anderson J. 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TopicsAircraft Components and StructureAircraft ControlAircraft DesignAircraft Flight Control SystemAircraft OperationsAircraft Operations and TechnologyAircraft Stability and ControlAircraft Wing DesignAircraftsAirfoilAutomatic Flight Control SystemsFlight Control SurfacesShort Takeoff and LandingTakeoff and LandingWing ConfigurationsWing Planforms KeywordsBoundary Layer SeparationAerodynamic LoadsFlight TestingClosed Loop ControlReynolds Averaged Navier StokesDielectric Barrier DischargePlasma ActuatorCommercial AircraftTailless AircraftSeparated FlowsPDF Received26 March 2019Accepted17 June 2019Published online2 August 2019

Tests were carried out on a 45° swept back wing that is highly unstable in pitch due to its large aspect ratio and sweep-back angle. The tests were carried out at velocities of 20-40 meters/second at a representative free stream Reynolds numbers approaching 10. The wing being based on a NACA0012 airfoil had a round leading edge that prevented the generation of a leading edge vortex prior to tip trailing edge stall, thus its first pitch-up departure was due to flow separation at the tip. The interaction between the tip separation and the leading edge vortex resulted in another non-linear pitch behavior. Finally, at larger incidence angle the leading edge vortex propagated inboard and the trailing edge separated region moved upstream. This wing would have been uncontrollable in pitch, were it not for the introduction of active flow control. In this case a single sweeping jet actuator could extend the trimmed incidence angle by 5° to 7°, thus avoiding the pitch break and almost doubling the usable lift coefficient. Balance results, flow visualization and Pressure Sensitive Paint were used to correlate the quantitative results to flow physics.

Active flow control (AFC) was incorporated into the Stability And Control CONfiguration model (SACCON) in order to trim it in pitch. The combination of AFC and traditional control surfaces enables the model to be stable and controllable. The model can be trimmed in pitch, provided that the significance of AFC is considered to be on par with other parameters affecting the aerodynamic characteristics of a wing. The most important outcome of this exercise implies that the traditional boundary-layer approach with inviscidly determined pressure distributions should no longer be used when AFC is to be implemented. This concept is proven by way of examples associated with the SACCON model that can be trimmed either by large deflection of flaps or by AFC. Proper combination of both parameters requires a minimal input of AFC (e.g., two actuators) in conjunction with a small flap deflection. Thus far, an actuator array located near the flap hinge was most effective but this result may not be universal. Tests were carried out at low speeds (i.e., at Mach number<0.2) and Reynolds numbers based on the outer panel chord of the model that approached 10(6).

Active flow control (AFC) subscale experiments were conducted at the Lucas Wind Tunnel of the California Institute of Technology. Tests were performed on a generic vertical tail model at low speeds. Fluidic oscillators were used at the trailing edge of the main element (vertical stabilizer) to redirect the flow over the rudder and delay or prevent flow separation. Side force increases in excess of 50% were achieved with a 2% momentum coefficient (C mu) input. The results indicated that a collective C mu of about 1% could increase the side force by 30-50%. This result is achieved by reducing the spanwise flow on the swept back wings that contributes to early flow separation near their tips. These experiments provided the technical backdrop to test the full-scale Boeing 757 vertical tail model equipped with a fluidic oscillator system at the National Full-scale Aerodynamics Complex 40-by 80-foot Wind Tunnel, NASA Ames Research Center. The C mu is shown to be an important parameter for scaling a fluidic oscillator AFC system from subscale to full-scale wind tunnel tests. The results of these tests provided the required rationale to use a fluidic oscillator AFC configuration for a follow-on flight test on the Boeing 757 ecoDemonstrator.