John Farnsworth

Active Flow Control at Low Angles of Attack: Stingray Unmanned Aerial Vehicle

Active flow control using fluidic actuators, via arrays of synthetic jet actuators, was used to provide control power for the Stingray unmanned aerial vehicle in the longitudinal (pitch) and lateral (roll) directions at low angles of attack. Using this technique, the pitch and roll moments were altered such that the effect is similar to that of a deflection of conventional control effectors. The control effectiveness of the synthetic jets on the aerodynamic performance of the Stingray unmanned aerial vehicle was investigated experimentally in a wind tunnel. Global flow measurements were conducted, where the moments and forces on the vehicle were measured using a six-component sting balance. The effect of the actuation was also examined on the surface static pressure at two spanwise locations. In addition, a particle image velocimetry technique was used to quantify the flowfield over the model, both the global flowfield as well as the localized interaction domain near the synthetic jet orifice. The synthetic jets were able to alter the local streamlines and displace the boundary layer through the formation of a small quasi-steady interaction region on the suction surface of the Stingray unmanned aerial vehicles wing. Phase-locked particle image velocimetry data were acquired to provide insight into the growth, propagation, and decay of the synthetic jets impulse and their interaction with the crossflow. Furthermore, the changes induced on the moments and forces can be proportionally controlled by either changing the momentum coefficient or by driving the synthetic jets with a pulse modulation waveform. This can lead the way for future development of closed-loop control models.

Open-Loop Dynamics of the Asymmetric Vortex Wake behind a von Kármán Ogive at High Incidence

The asymmetric vortex regime of a von Karman ogive with fineness ratio of 3.5 is experimentally studied at a Reynolds number of 156,000. Both port and starboard plasma actuators are used to introduce fluidic disturbances at the tip of the ogive which are amplified through the flows convective instability and produce a deterministic port or starboard asymmetric vortex state (i.e. side force). Accurate control or manipulation of this asymmetric vortex state holds the potential for increased maneuverability and stability characteristics of slender flight vehicles at high angle of attack. Open-loop experimental tests are used to understand and quantify the vortex dynamics due to actuation inputs. Linear time invariant models provide a suitable model structure to replicate the vortex dynamics and allow for simulation and closed-loop control design. Standard PID control is designed and implemented. A closed-loop simulation shows arbitrary side force tracking with adequate disturbance rejection.

Control of the Stingray UAV at Low Angles of Attack

The control effectiveness of active flow control, via arrays of synthetic jet actuators, on the aerodynamic performance of the Stingray UAV at low angles of attack was investigated experimentally in a wind tunnel. Global flow measurements were conducted, where the moments and forces on the vehicle were measured using a six component sting balance. The virtual shape modification technique, which was previously used on 2-D airfoils, was implemented in the present work on the 3-D Stingray configuration. A Solid obstruction was placed near the leading edge on the suction side of the wing in conjunction with a synthetic jet actuator that was placed farther inboard and was issued normal to the surface. Using this technique, the pitching moment was altered where the effect exceeded that of a 5 o deflection of the elevons. This suggests that synthetic-jet based flow control can be used for longitudinal trim control of the UAV during cruise condition, in lieu of the conventional control surfaces.

Aerodynamic Performance Modification of the Stingray UAV at Low Angles of Attack

Active flow control using fluidic actuators, via synthetic jets and steady blowing jets, was used to provide control power for trimming the Stingray UAV in the longitudinal (pitch) and lateral (roll) directions at low angles of attack. Using this technique, the pitching and roll moments were altered such that the effect is similar to that of a deflection of conventional control effectors in trim. The control effectiveness of the flow control on the aerodynamic performance of the Stingray UAV was investigated experimentally in a wind tunnel. Global flow measurements were conducted, where the moments and forces on the vehicle were measured using a six component sting balance. The effect of the actuation was also examined on the surface static pressure at two spanwise locations. In addition, Particle Image Velocimetry (PIV) technique was used to quantify the velocity vector field over the model, both the global flow field as well as the localized interaction domain near the synthetic jet orifice. The synthetic jets were able to alter the local streamlines through the formation of a quasi-steady interaction region on the suction surface of the Stingray UAV’s wing. Phase locked PIV data was acquired to provide insight into the growth, propagation, and decay of the synthetic jet impulse and its interaction with the cross-flow. The changes induced on the moments and forces can be proportionally controlled by either changing the momentum coefficient or by driving the synthetic jets with a pulse modulation waveform. This can lead the way for future development of closed-loop control models.

Closed-Loop Flow Control of a Forebody at a High Incidence Angle

The flowfield around an axisymmetric forebody at a high angle of attack (40<α<60  deg) produces a significant side force. This side force results from an asymmetric pressure distribution around the body due to an asymmetric vortex configuration. Numerical studies of open-loop control using mass blowing slots near the tip of the model have shown a proportional response of the side force over a range of momentum coefficient amplitudes. From the open-loop simulations, a prediction-error minimization method was employed to formulate a linear time-invariant model, which captured the dynamics of the side force response to different mass flow rates applied to either the port or starboard actuator. Based on the linear time-invariant model, a proportional–integral control law was developed for set-point tracking a prescribed side force. The development of the linear time-invariant model, and corresponding linear time-invariant feedback solution are presented to illustrate the model’s capabilities and limitations. ...

Closed-Loop Flow Control of a Tangent Ogive at a High Angle of Attack

The flowfield around an axisymmetric forebody at a moderate angle of attack (40 � < � < 60 � ) produces a significant side force as the result of an asymmetric pressure distribution around the body resulting from an asymmetric vortex flow state. Numerical studies of open-loop control using mass-blowing slots near the tip of the model have shown proportional control of the side force by varying the momentum coefficient of the blowing slots. From the open-loop simulations, a prediction-error minimization method (PEM) was used to develop a linear time invariant (LTI) model which captures the dynamics of the side force response to different mass flow rates applied to the port or starboard actuator. Based on the model, a PI controller was developed for reference tracking a prescribed side force profile. The development of the LTI model, and corresponding controller simulations of the closed-loop system are presented to illustrate the models capabilities as well as its limitations. The ability to track a prescribed reference signal based on the LTI model and corresponding PI control scheme is shown. The results indicate that the bandwidth of the controller is limited to frequencies below half the convective frequency due to the convective time delay as well as the actuator and sensor placement. Finally, the closed-loop controller simulations are compared to Navier-Stokes feedback simulations.

Active Enhancement of Wind Turbine Blades Performance

The feasibility of using synthetic jet actuators to enhance the performance of wind turbine blades was explored in wind tunnel experiments. Using this technique, the global flow field over the blade was altered such that flow separation was mitigated. This, in turn, resulted in a significant decrease in the vibration of the blade. Global flow measurements were conducted, where the moments and forces on the blade were measured using a six component wall-mounted load cell. The effect of the actuation was also examined on the surface static pressure at two spanwise locations; near the blade’s root and near the tip. In addition, Particle Image Velocimetry (PIV) technique was used to quantify the flow field over the blade. Using synthetic jets, the flow over the blade was either fully or partially reattached, depending on the angle of attack and the Reynolds number. Furthermore, the changes induced on the moments and forces, as well as on the blades vibrations were found to be proportionally controlled by either changing the momentum coefficient, the number of synthetic jets used, or by the driving waveform. Finally, a proof-of-concept closed-loop control system was developed to test the ability of using synthetic jet actuators to restore and maintain flow attachment and reduce the vibrations in the blade during dynamic pitch. The synthetic jets were switched on when the root strain vibration spectrum exceeds a predetermined threshold at a given frequency. While the control system implementation used is simplistic, it demonstrated the ability of synthetic jet actuators to reduce blade’s vibrations (by restoring and maintaining attached flow) during the dynamic motion, analogous to the wind gusts seen in wind turbine operation.

Interaction of a Finite-span Synthetic-jet and Cross-flow over a Swept Wing

An experimental investigation was performed to study the three-dimensional flow structures and interactions of a finite-span synthetic jet with the flow over a finite and swept back wing (cross-sectional profile of the NACA 4421, AR=4, sweep angle of 30 deg), at a Reynolds number of 10 5 and zero angle of attack. Two momentum coefficients were considered, corresponding to two blowing ratios of 0.8 and 1.2. Stereoscopic PIV data was collected around the center jet in the mid-span section. The effect of the momentum coefficient (or blowing ratio) was analyzed based on the three-dimensional flow field using time-averaged and phase-averaged statistics. The study showed that, similar to the interaction of a synthetic jet with the flow over an unswept wing, the flow field in the vicinity of the synthetic-jet orifice becomes highly three-dimensional and is governed by the superposition of two kinds of flow structures: (1) streamwise structures that are associated with the finite span of the jet (edge vortices), and (2) spanwise flow structures that are generated along the orifice due to the vortex pairs that are formed by the synthetic jet. Nevertheless, over a swept wing these three-dimensional flow structures evolve almost immediately downstream to the synthetic-jet orifice and evolve into a very complex 3-D flow field.

Pre-swirl Maneuvering Propulsor: Part 1 Computations

Recent concept studies have demonstrated the potential to utilize a pre-swirl propulsor configuration with adjustable upstream stators to generate propulsor side forces. These studies led to a set of experiments and corresponding computations to validate this concept. Ducted and open pre-swirl propulsors were configured with an upstream stator row and downstream rotor. During normal operation, the upstream stator blades are all situated at the same pitch angle and pre-swirl the flow into the propulsor when generating a roll moment to counter the moment produced by the rotor. By varying the pitch angles of the stator blade about the circumference, it is possible to both generate a mean stator side force and subsequently vary the axial velocity and swirl that is ingested into the inflow. The rotor then generates side forces in response to the inflow. Wind tunnel experiments were conducted to measure the steady, spatially varying stator wake flows for various stator geometric configurations using stereo particle image velocimetry. Computations utilized both potential flow and fully viscous 3-D (Reynolds Averaged Navier-Stokes, RANS) computations to predict the stator forces, velocity field and rotor response. Rotor design space investigations varied blade parameters including blade number, rake, skew and a combination of the two. RANS was used to then validate the final propulsor design with experimental data used to validate the computations. Computational data demonstrated that total side force coefficients on the order of 0.2 could be generated by the propulsor alone with results consistent with recent water tunnel measurements. This amount of control authority exceeds current control surface capabilities at 3 knots for Navy 21” unmanned undersea vehicles.

Aeroelastic Response of a Finite Span NACA 0018 Wing Part 1: Experimental Measurements

A rectangular planform finite span NACA 0018 wing section was machined from three different materials to experimentally quantify the steady aeroelastic response of the wing section. This study was done to provide a baseline model for future flow control experiments and to assist with verification and validation of computational simulations being run in parallel. Measurements from both a six component force/moment balance and a stereo vision system were collected simultaneously to match the wing deformation to the aerodynamic loads. A solid aluminum wing was utilized as a rigid comparison to the flexible plastic materials that were tested. Each of the flexble wings displayed significant wing bending, but negligible wing twisting was observed. As a result the aerodynamic lift and drag profiles showed minimal influence from the aeroelastic response of the wings. Finally, from the force and moment measurements, an estimate of the wing spanwise lift distribution was made and the bending deformation profile was calculated to compare with the stereovision measurements. The estimated bending profiles matched the measured within 1 mm across the wing span of 600 mm for a maximum tip deflection of 25 mm.