I.H. Ibrahim

Simulations of the linear plasma synthetic jet actuator utilizing a modified Suzen-Huang model

The linear plasma synthetic jet actuator (L-PSJA) is a unique form of flow control device which harnesses the interaction of induced flows from two linear plasma actuators to form an upward jet. Since each injection can be manipulated in intensity, the synthetic jet has thrust vectoring properties. Our study simulates the L-PSJA by utilizing a modified Suzen-Huang (S-H) model that accounts for drift and diffusive properties in the surface charge. The results of the present model show that the centreline velocity is closer to the experimental values found in literature as compared to the default form of S-H modelling. Thrust vectoring simulations were also performed to demonstrate the feasibility of flow directional variation in the L-PSJA.

Boundary Condition Modifications of the Suzen-Huang Plasma Actuator Model

The accuracy of the Suzen-Huang (S-H) model is improved by altering the boundary condition of the dielectric surface above the lower electrode. For the equation governing the electric field, we introduce a ‘dielectric shielding’ condition at the same region, which results in a spread of the electric field strength along the dielectric surface. For the equation governing the surface charge density, we introduce boundary conditions that modify the behavior of the charge density variable in the S-H model. The conditions represent a fitting procedure by adding the features of propagation and dissipation in a one-dimensional Fokker-Plank equation. The equation is initiated by a normal distribution function centered at the leading edge of the lower electrode. These modifications improved model results by about 50% when comparing the maximum induced velocity value with experimental results. Furthermore, charge density growth is propagating in a similar manner to that obtained by charge transport models.

Effects of the scalar parameters in the Suzen‐Huang model on plasma actuator characteristics

Purpose – For the past decade, plasma actuators have been identified as a subset in the realm of active flow control devices. As research into plasma actuators continues to mature, computational modelling is needed to complement the investigation of the actuators. This paper seeks to address these issues.Design/methodology/approach – In this study, the Suzen‐Huang model is chosen because of its ability to simulate both the charge density and Lorentz body force. Its advantages and limitations have been identified with a parametric study of two constants used in the modelling: the Debye length (λD) and the maximum charge density value (ρc* ). By varying the two scalars, the effects of charge density, body force and induced velocity are examined.Findings – The results show that the non‐dimensionalised body force (Fb*) is nonlinearly dependent on Debye length. However, a linear variation of Fb* is observed with increasing values of maximum charge density. The optimized form of the Suzen‐Huang model shows bett...

Applying Taguchi's off-line quality control method and ANOVA on the maneuverability of the F-5E intake

Air quality leading up to the compressor face of a fighter aircraft determines the engine performance considerably. A deficiency in the quality could lead to flutter or stall in the engines. In this study, two statistical methods; the Taguchi Method (TM) and the Analysis of Variance (ANOVA) are used to evaluate airflow quality through the intake via fighter aircraft maneuvers. The three factors associated directly with aircraft maneuverability are the Mach number (M), Angles of Incident (@a) and Sideslip (@b). Desirable air quality can be described as having high pressure recoveries as well as low distortion at the Aerodynamic Interface Plane (AIP). The intake studied is the port side F-5E duct. Results show that an increase in the Mach number affects the streamwise diffusion of the fluid more than the changing the angles of attack and sideslip, resulting in lower pressure recovery. The secondary flow formation in the streamwise direction is unable to dissipate and increases in strength with increasing Mach number. The curvature in the z-axis is more pronounced than that existing in the x-axis, leading to the formation of more adverse pressure gradients forming and hence greater secondary flow strength. This results in a more distorted flow leading to the AIP. This observation is in tandem with the values of the DC (60) readings obtained. The F-5Es Taguchis Method results show that Mach number had the greatest effect on pressure recovery, and AOA affected distortion most considerably. Results from ANOVA show that Factors A, B and C and Interactions AC and BC affect the distortion of airflow. However, Factor B or the angle of attack affects this distortion most significantly.

Utilizing the L-PSJA for controlling cylindrical wake flow

Purpose – The cylindrical wake flow is an important part of many engineering applications, including wake turbulence, acoustic noise, and lift/drag forces on bodies. The suppression of von Karman vortex street (VKS) is an important goal for flow control devices. The paper aims to discuss these issues. Design/methodology/approach – The linear plasma synthetic jet actuator (L-PSJA) is utilized as a flow control device to suppress the VKS formation. Different configurations of the device is studied numerically. Findings – Of the 12 configurations that were investigated, five configurations were able to suppress the formation of the VKS. Originality/value – For the first time, the L-PSJA has been shown (through numerical simulations) to be able to suppress VKS.

Numerical Investigation on Flow Separation Control of Low Reynolds Number Sinusoidal Aerofoils

The paper presents a computational analysis of the characteristics of a NACA 634- 021 aerofoil incorporated with sinusoidal leading-edge protuberances at Re = 14,000. The protuberances are characterized by an amplitude and wavelength of 12% and 50% of the aerofoil chord length respectively. An unsteady Reynolds Average Navier Stokes (RANS) analysis of the full-span aerofoils was carried out using Transition SST (Shear Stress Transport) turbulence model across five different angles-of-attack (AOA). Comparisons with previous experimental results reported good qualitative agreements in terms of flow separation when the aerofoils are pitched at higher AOAs. Results presented here comprised of near-wall flow visualizations of the flow separation bubble at the peaks and troughs of the protuberances. Additionally, results indicate that the aerofoil with leading-edge protuberances displayed distinctive wall shear streamline and iso-contour characteristics at different span-wise positions. This implies that even at a low Reynolds number, implementations of these leading-edge protuberances could have positive or adverse effects on flow separation.

Effect of Leading Edge Protuberance on Thrust Production of a Dynamically Pitching Aerofoil

The paper presents a computational analysis of the characteristics of a NACA 634-021 aerofoil modified by incorporating sinusoidal leading-edge protuberances at Re = 14,000. The protuberances are from the tubercles of the humpback whale flipper with leading edge acting as passive-flow control devices that improve performance and manoeuvrability of the flipper. They are characterized by an amplitude and wavelength of 12% and 50% of the aerofoil chord length respectively. Three-dimensional CFD on the modified aerofoil oscillating about a point located on the centreline at quarter-chord has been performed with the frequency and amplitude of oscillation being 4Hz and 10 deg respectively. In addition to the lift and thrust coefficients, near wall flow visualisations and the shedding of vortices during oscillations are presented to illustrate the unsteady flow features on the performance of the oscillating flipper. The results show an improvement in the thrust production when compared to previous studies on similar symmetric aerofoil without the leading edge modifications.

Flight Maneuverability Characteristics of the F-16 CFD and Correlation with its Intake Total Pressure Recovery and Distortion

Abstract: An investigation of the intake pressure recovery and distortion effects observed by varying maneuverability conditions of the F-16 was conducted. Simulations were conducted by using the Partial Differential Equation (PDE) calculator COMSOL. Results were validated with the classical s-duct, showing reasonable qualitative agreement. The maneuverability of the F-16 was defined using three factors: Mach number, angle of attack and angle of sideslip. The values of pressure recovery and distortion were measured at the Aerodynamic Interface Plane (AIP). Two sets of simulations were conducted: one to compare the sea level and 10 km altitude flight condition and another to correlate the required air intake and engine performance with pressure recovery and distortion at the AIP. Results of the flight level comparisons showed a slightly higher pressure recovery at 10 km of 0.985 compared to sea level condition value of 0.983. Distortion coefficient values for both conditions were however similar at 0.120. The difference in recovery values is attributed to the less viscous characteristic of air at higher altitude. This resulted in lower secondary flow formations in the intake. With increasing altitude, temperature plots showed a linear decreasing profile throughout the intake. The study provides a platform to a future model which couples the results obtained from the intake as inlet conditions to turbine flow simulations. This would provide a more holistic understanding of the propulsion system in the F-16.

Effect of the angle of attack on the YF-16 inlet

A k-e numerical analysis of a fuselage-shielded inlet similar to that of the YF-16 was conducted. The simulations were conducted at Mach 0.6 at -10° to 40° angles of attack. Results were validated with experimental values obtained in the 90° bending pipe and circular S-duct. Pre-entry separation effects were simulated by a flat plate and offset by the diverter height. Results showed that the pressure recoveries resembled that of the YF-16 inlet, where negligible losses were up to 30° angle of attack (AOA). The increase in the angle of attack also resulted in the diminishing strength of the twin swirls.

Swirl flow analysis of the port side F-5E intake using CFD

Swirl represents a form of energy loss, as the energy is used in accelerating the flow in the angular direction and does not contribute to engine thrust. Significant swirl flow in resulting in the fighter aircraft intakes during manoeuvres can result in an undesirable engine performance. The CFD study conducted using FLUENT shows that the F-5E intake is more sensitive to the negative angles of attack, especially for AOA at −10°, where bulk swirls were seen forming. In the results of Taguchi and ANOVA, both statistical model point to the angle of attack play the most significant role in determining a more undesirable swirl coefficient.