Leo Veldhuis

Measurement of the body force field of plasma actuators

A novel technique is proposed and investigated for the estimation of the body force field resulting from the operation of a dielectric barrier discharge plasma actuator. The technique relies on the measurement of the spatio-temporal evolution of the induced velocity field using high-speed particle image velocimetry (PIV). The technique has the advantage of providing spatial distribution of the body force vector field. A full Navier–Stokes term decomposition is applied on the evolving field along with additional closure norms in order to decouple the pressure gradient and body force terms. Results are compared with load-cell measurements of the direct reaction force and also momentum balance calculations based on the PIV field. Agreement between the different methods is observed. The data can easily be incorporated in computational flow solvers and also be used for validation and calibration of numerical plasma models.

Flow Separation Control on Airfoil With Pulsed Nanosecond Discharge Actuator

An experimental study of flow separation control with a nanosecond pulse plasma actuator was performed in wind-tunnel experiments. The discharge used had a pulse width of 12 ns and rising time of 3 ns with voltage up to 12 kV. Repetition frequency was adjustable up to 10 kHz. The first series of experiments was to measure integral effects of the actuator on lift and drag. Three different airfoil models were used, NACA-0015 with the chord of 20 cm, NLF-MOD22A with the chord of 60 cm and NACA 63-618 with the chord of 20 cm. Different geometries of the actuator were tested at flow speeds up to 80 m/s. In stall conditions the significant lift increase up to 20% accompanied by drag reduction (up to 3 times) was observed. The critical angle of attack shifted up to 5–7 degrees. The relation of the optimal discharge frequency to the chord length and flow velocity was proven. The dependence of the effect on the position of the actuator on the wing was studied, showing that the most effective position of the actuator is on the leading edge in case of leading edge separation. In order to study the mechanism of the nanosecond plasma actuation experiments using schlieren imaging were carried out. It shown the shock wave propagation and formation of large-scale vortex structure in the separation zone, which led to separation elimination. PIV diagnostics technique was used to investigate velocity field and quantitative properties of vortex formation. In flat-plate still air experiments small scale actuator effects were investigated. Measured speed of flow generated by actuator was found to be of order of 0.1 m/s and a span-wise nonuniformity was observed. The experimental work is supported by numerical simulations of the phenomena. The formation of vortex similar to that observed in experiments was simulated in the case of laminar leading edge separation. Model simulations of free shear layer shown intensification of shear layer instabilities due to shock wave to shear layer interaction.

Numerical Study on Control of Tollmien-Schlichting Waves Using Plasma Actuators

A numerical investigation on the use of plasma actuators for transition control is presented. The numerical framework involves the solution of the full unsteady 2D incompressible Navier Stokes equations using a finite volume formulation. The set of equations is formulated by solving for the perturbations in the flow while a mean laminar boundary layer flow is considered fixed and superimposed. The effect of the plasma actuator is represented as an imposed unsteady body force distribution derived from experimental measurements. Furthermore, an adaptive control system based on the filtered-x LMS algorithm is implemented directly into the flow solver. The control system uses pressure signals at the wall in order to compute the frequency, phase and amplitude of the plasma body force which minimizes the intensity of the propagating TS waves. Results show large reductions in wave amplitude for both single and multi frequency cases.

Experimental Study on the Body Force Field of Dielectric Barrier Discharge Actuators

An experimental investigation on thrust and body force of Dielectric Barrier Discharge (DBD) /plasma actuators aimed at low power flow control applications is presented. A parametric study on thrust is conducted for a wide range of geometrical configurations as well as several electrical operational conditions. Direct measurements of the induced thrust are taken using a highly sensitive load cell. Simultaneous readings of current and voltage are also performed, providing the power consumption. Furthermore a novel technique for determination of the spatial distribution of the body-force is proposed, developed and tested. The technique involves the use of a high-speed PIV system to resolve all terms of the Navier-Stokes equation representation of the flow field including body force. Results reveal the existence of an explicit relation between voltage, thrust and consumed power. Furthermore the influence of the geometrical configuration of the actuator on the thrust is shown. The body force obtained with the proposed technique agrees well with the thrust measurements.

Experimental investigation of propeller induced ground vortex under headwind condition

During the ground operation of propeller aircraft, the interaction between the ground and the flow field produced by the propeller may lead to the generation of a vortex originating on the ground. Such ground induced vortex results in non-uniform inflow into the propeller disk. Furthermore, the vortex may raise debris from the ground that may damage the blades, decreasing their life time. For this phenomenon, the flow field near the ground is of interest because it provides information on the strength of the vortex before it is being ingested by the propeller. To obtain insight into the origins and the development of the ground vortex, velocity measurements have been performed in the flow field produced by a scaled propeller model using a three-component PIV system on a plane parallel to the ground and in the proximity of it. The velocities in the test plane suggest that a horseshoe vortex is formed above the ground which is built to help understand the flow topology of ground vortices. The vortex/no-vortex domain boundary which is relied upon the presence of node and saddle points in the time averaged flow is presented. The effect of varying the thrust coefficient of the propeller has been assessed under headwind conditions. The averaged strength of the maximum wall normal vortex is compared at different thrust coefficients. The results show that average levels of vortex strength increases up to a maximum and then decreases to achieve a plateau. The pressure under the propeller is measured and the mean value of pressure gradient magnitude keeps increasing as thrust coefficient increases. Vortex wandering is observed at all test conditions in both the lateral and the longitudinal directions for all test conditions. The influence of the ground induced non-uniform inflow upon the propeller performances is found to be negligible.

Aerodynamic and Aeroacoustic Effects of Pylon Trailing Edge Blowing on Pusher Propeller Installation

The aerodynamic and aeroacoustic effects of pylon trailing edge blowing on the propulsive performance and noise emissions of a propeller installed in a pusher configuration were studied in a wind tunnel. A propeller model and a pylon equipped with a trailing edge blowing system were installed in the large low-speed facility of the German-Dutch wind-tunnels (DNW-LLF). Particle image velocimetry measurements of the flow field downstream of the pylon confirmed a wake re-energization obtained through blowing, with a momentum deficit recovery of 80% compared to the unblown case. For the symmetric inflow conditions considered, the effect of pylon installation on the propulsive performance was found small. Increases in thrust and torque of 1% up to 6% were measured at high and low thrust settings, which was comparable to the measurement variability. Acoustic data obtained using out-of-flow microphones confirmed the strong interaction effects resulting from the installation of the upstream pylon, with an increase in noise levels due to the presence of the pylon of up to 12 dB at a medium propeller thrust setting. The application of pylon trailing edge blowing successfully eliminated the installation effects, resulting in noise levels equal to those of the isolated propeller over the entire axial directivity range. At higher thrust settings the change in blade angle of attack due to the pylon wake impingement is smaller, and the steady blade loads are larger compared to the unsteady loads experienced during the wake passage. Consequently, in this operating regime the propeller noise emissions were dominated by steady sources for all but the most upstream observer positions.

Plasma Assisted Aerodynamics for Transition Delay

This work involves the numerical investigation of Dielectric Barrier Discharge (DBD) actuators used as wave cancellation devises for transition delay of subsonic flows. The operation of the DBD/plasma actuator is described based on a first-principles electrodynamic model. Several cases of different geometrical and electrical configurations are studied. In all cases the electrodynamic model indicates the formation of a weak near-wall body-force component. The research concept involves the utilization of this body-force through a wave superposition method for dampening unsteady Tollmien-Schlichting waves in a laminar boundary layer. To study this effect, the output body-force distribution of the plasma model is planned to be coupled with a Direct Numerical Simulation study of the transition process.

Performance Analysis of Wake and Boundary-Layer Ingestion for Aircraft Design

This paper presents conceptual studies to evaluate the performance of the propulsor and its associated vehicle in the configurations of wake ingestion and boundary-layer ingestion. A power conversion analysis uses the power balance method to elaborate the power-saving mechanism of wake ingestion, showing that the Froude’s propulsive efficiency as a figure of merit should be separated from the power conversion efficiency in these configurations. The body/propulsor interaction occurring in the boundary-layer ingestion configuration is qualitatively analyzed to clarify its influence on the performance of the integrated vehicle. The results suggest that the minimization of power consumption should be used as a design criterion for aircraft using boundary-layer ingestion.

Experimental Study and Numerical Simulation of Flow Separation Control with Pulsed Nanosecond Discharge Actuator

Active flow separation control with a nanosecond pulse plasma actuator, which is essentially a simple electrode system on the surface of an airfoil, introducing lowenergy gas discharge into the boundary layer, with little extra weight and no mechanical parts, was performed in wind-tunnel experiments on various airfoil models. In stall conditions the significant lift increase up to 30% accompanied by drag reduction (up to 3 times) was observed. The critical angle of attack shifted up to 5–7 degrees. Schlieren imaging show the shock wave propagation and formation of large-scale vortex structure in the separation zone, which led to separation elimination. The experimental work is supported by numerical simulations of the phenomena. The formation of vortex similar to that observed in experiments was simulated in the case of laminar leading edge separation. Model simulations of free shear layer show intensification of shear layer instabilities due to shock wave to shear layer interaction. The mechanism of flow control by nanosecond plasma discharge is based on extra vorticity created by the shock wave, which is produced from the layer of the hot gas. This hot gas in generated during the fast thermalisation process, in which up to 60% of the discharge energy is converted to heat in less than 1 µs [1]. This phenomenon gives an opportunity for nanosecond discharge actuator to be effective at high velocities [2, 3]. The current work continues studying the performance of nanosecond plasma actuator. A series of wind tunnel experiments was carried out with different actuator layouts at flow velocities up 80 m/s at various airfoils with chords up to 1.5 m and spans up to 5 m. A numerical model was developed to prove the shock wave mechanism of actuator operation. 2. Experiment In the present work, a linear actuator was used [4]. The actuator consisted of a base layer of insulator attached onto the surface of the airfoil, a covered electrode, an interelectrode layer of insulation and an exposed electrode. In the majority of the cases, exposed electrode was ground, and the high-voltage electrode was covered one. High-voltage nanosecond pulses were provided by three different nanosecond pulsers, which were capable of producing pulses of up to 50 kV with rising time of 3-15 ns and duration from 10 to 50 ns at repetition frequencies up to 10 kHz. Low-speed experiments was carried out in open jet wind tunnel using the NACA0015 airfoil with the chord of 20 cm and span about 75 cm. The tunnel was equipped with an

Spatial-temporal and modal analysis of propeller induced ground vortices by particle image velocimetry

During the ground operation of aircraft, there is potentially a system of vortices generated from the ground toward the propulsor, commonly denoted as ground vortices. Although extensive research has been conducted on ground vortices induced by turbofans which were simplified by suction tubes, these studies cannot well capture the properties of ground vortices induced by propellers, e.g., the flow phenomena due to intermittent characteristics of blade passing and the presence of slipstream of the propeller. Therefore, the investigation of ground vortices induced by a propeller is performed to improve understanding of these phenomena. The distributions of velocities in two different planes containing the vortices were measured by high frequency Particle Image Velocimetry. These planes are a wall-parallel plane in close proximity to the ground and a wall-normal plane upstream of the propeller. The instantaneous flow fields feature highly unsteady flow in both of these two planes. The spectral analysis is co...