Active Flow Control (AFC) experiments performed at the Caltech Lucas Adaptive Wall WindTunnel on a 12%-thick, generic vertical tail model indicated that sweeping jets emanating from the trailing edge (TE) of the vertical stabilizer significantly increased the side force coefficient for a wide range of rudder deflection angles and yaw angles at free-stream velocities approaching takeoff rotation speed. The results indicated that 2% blowing momentum coefficient (C µ ) increased the side force in excess of 50% at the maximum conventional rudder deflection angle in the absence of yaw. Even C µ = 0.5% increased the side force in excess of 20% under these conditions. This effort was sponsored by the NASA Environmentally Responsible Aviation (ERA) project and the successful demonstration of this flow-control application could have far reaching implications. It could lead to effective applications of AFC technologies on key aircraft control surfaces and lift enhancing devices (flaps) that would aid in reduction of fuel consumption through a decrease in size and weight of wings and control surfaces or a reduction of the noise footprint due to steeper climb and descent.

This study provides quantitative insight into the formation of vortex breakdown and the onset of global instability in a turbulent swirling jet. A water jet is guided through a rotating honeycomb that imparts the rotational motion, passed through a contraction, and discharged into a large water tank. The flow states evolving at increasing swirl are mapped out via time-resolved particle image velocimetry. The experimental results scale properly with the swirl number based on the axial momentum flux when the commonly used boundary-layer approximations are omitted. The instantaneous velocity field reveals that vortex breakdown occurs intermittently at a wide range of swirl numbers before it appears in the mean flow. At this intermittent state, the evolving breakdown bubble oscillates heavily between two streamwise locations where the vortex core is subcritical. Upon further increasing the swirl, the breakdown oscillations decay and a region of reversed flow appears in the mean flowfield. The formation of this socalled axisymmetric breakdown state is accompanied by a supercritical-to-subcritical transition of the inflowing vortex core. The reversed flow region is found to grow linearly with increasing swirl until the flow undergoes a supercritical Hopf bifurcation to a global single-helical mode, and vortex breakdown adopts a spiral shape. The global mode shape is extracted from the particle image velocimetry snapshots by means of proper orthogonal decomposition and Fourier analysis. The present experiment reveals that, at gradually increasing swirl, the jet first transitions to an axisymmetric breakdown state that remains globally stable until a critical swirl number is exceeded. This sequence of flow states agrees well with the transient formation of vortex breakdown observed in laminar flows.Nomenclature jA sat j = saturation amplitude at limit-cycle oscillation a i = ith temporal proper orthogonal decomposition mode D = nozzle diameter E tot = total coherent energy Fx; r; m; f = complex Fourier coefficient f = frequency f s = measurement acquisition sampling frequency i = index of particle image velocimetry snapshot m = azimuthal wave number N = number of particle image velocimetry snapshots psd = power spectral density pdf = probability density function P RF = probability of reversed flow P VB = probability of vortex breakdown Re = Reynolds number; Eq. (1) r core = radial position where w x is equal to 0 r crit = Benjamin's critical radius; Eq. (9) S = swirl number; Eq. (2) S crit = critical swirl number for Hopf bifurcation; Eq. (11) S VB = minimum swirl number for vortex breakdown St = Strouhal number, fD=U m T = duration of measurement series t = time U cl = axial mean velocity at jet centerline U m = bulk velocity at zero swirl V = mean velocity vector, (U; V; W) v = instantaneous velocity vector, u; v; w v = coherent velocity vector, (ũ;ṽ;w) v 0 = fluctuating velocity vector, (u 0 ; v 0 ; w 0 ) x = position vector in Cartesian coordinates, (x; y; z) x = position vector in cylindrical coordinates, (x; r; ) x...

The application of active flow control on a vertical tail of a typical twin engine aircraft was investigated. Sweeping jets installed into the rudder surface were used and their effect was assessed by force measurements, flow visualization and local pressure distributions. The airfoil forming the tail is a NACA 0012 with a rudder using 35% of its chord. The tests were carried out at the Lucas Wind Tunnel at the California Institute of Technology at representative Reynolds numbers of up to Re=1.5 million. Multiple flap deflections and spanwise actuator configurations were tested resulting in an increase of up to 50-70% in side force depending on the free stream velocity and momentum input.

Experiments aimed at delaying flow separation through discrete jets pointing in the direction of streaming and sweeping side to side along the span were conducted on a V-22 airfoil with and without deflected trailing-edge flaps. The results indicated substantial drag reduction and lift increase at moderately low inputs of mass and momentum. Additional experiments were carried out on a semispan V-22 wing/nacelle combination, and they too provided an increase in lift-to-drag ratio L=D of approximately 60% (although active flow control was applied to the wing only). The effectiveness of the sweeping jets on reducing the download force acting on a V-22 full-span powered model in hover was also examined. A 29% reduction in download was realized using the embedded sweeping jets, corresponding approximately to a 2000 lb increase in hover lift. NomenclatureA slot = total area of the nozzle exits A wing = area of the wing= jet momentum L=D = lift-to-drag ratio Re = Reynolds number r = air density U jet = jet velocity U 1 = freestream velocity = angle of attack

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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 μ) input. The results indicated that a collective C μ 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 μ 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. Nomenclature

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