Effectiveness of synthetic jet actuators for separation control on an airfoil

The aerodynamic performance of an airfoil could be improved by controlling flow separation using active flow control techniques. In this study, a synthetic jet actuator (SJA) based on piezoelectric diaphragm has been developed. The selection of the SJA was due to their advantages in being lightweight, no external air supply required, simple system assembly, fast time response, low power consumption, easy installation, low cost and relatively small in size. Basically, the performance of the SJA depends on the specification and configuration of jet orifice, cavity, and oscillating membrane. The parameters studied include waveform signal, frequency, voltage, cavity and orifice physical characteristics. Final design and geometry of the SJA were determined based on these parameters. The SJA design with the best performance has been developed to generate sufficient air jet velocity to control flow separation. The experimental results measured by a hot-wire anemometer show that the maximum jet velocity obtained by the SJA with circular and slot orifice were 41.71 m/s and 35.3 m/s at an applied frequency of 900 Hz and 1570 Hz respectively. Next, the selected SJA was embedded into the wing with NACA 0015 airfoil and placed at 12.5% chord from the leading edge. Wind tunnel testing was conducted for stationary and oscillating airfoil conditions, with and without the SJA. The unsteady aerodynamic loads were calculated from the surface pressure measurements of 30 ports along the wing chord for both upper and lower surfaces. The airfoil was tested at various angles of attack at a free-stream velocity of up to 35 m/s corresponding to a Reynolds number of 1.006 x 106. Specifically for an oscillating airfoil, the reduced frequency, k, was varied from 0.02 to 0.18. The results of an airfoil with SJA showed that the CLmax and stall angle increased up to 13.94% and 29% respectively. Based on the results obtained, the SJA has an excellent capability to control the flow separation with delaying the stall angle, increasing the maximum lift, reducing the drag and delaying the intense nose down pitching moment

Static analysis of unsteady aerodynamics wake of simplified helicopter model via simulation work

Computational tools have led and helped researchers in providing advanced results, notably in rotorcraft research, as flow around the helicopter is dominated by complex aerodynamics and flow interaction phenomena. This research work aimed to evaluate the aerodynamic computational results on a simplified model helicopter when the model was subjected to the angles of attack 0°, -5°, -15°, and -20°, respectively. The study also examined the unsteady flow behaviour on the three-dimensional elliptical shape of a fuselage equipped with a rotor hub of the single rotor blade. The computational domain for the aerodynamic flow field was created within the size of 7 m (length) x 5 m (width) x 5 m (height). Results showed that an increase in the angle of attack in the rotor component caused additional drag of about 34% to 45% whilst the fuselage component contributed about 55% to 65% to drag increment. Also, a significant value of total pressure from -235 Pa to 250 Pa demonstrated along the simplified model helicopter distinctly showed that the complexity of geometry caused adverse pressure. The findings of this research work could potentially improve the understanding of complex flow surrounding the helicopter that has always baffled the aerodynamicists

Study the orifice effects of a synthetic jet actuator design

The synthetic jet actuator is an active flow control device that is used to improve the aerodynamic performance on working surfaces such as wings, helicopter blades and ground vehicles. The performance of synthetic jet actuator depends on the design of the orifice and cavity, and the oscillating driver. Piezoelectric diaphragm was used as an oscillating driver because of its small size and easier installation. The focus of this project is to study the effects of orifice size and shape for a synthetic jet actuator design. The effects were studied on circular and rectangular shapes, and different sizes of orifice. Meanwhile, the configurations of the cavity are fixed. Experiments were performed to determine the maximum pulse jet velocity and turbulence intensities of the jet coming out of the orifice, driven by the Piezoelectric diaphragm at different frequencies, at constant input voltage of 2V. The experiment mainly involved the measurement of the exit pulse jet velocity using a hot-wire anemometer. The results demonstrated that the circular orifice produced higher maximum pulse jet velocity and smaller sizes orifices, both circular and rectangular, results in higher velocity jets

Superaugmented pitching motion of UTM CAMAR UAV using advanced flying handling qualities

This paper focused of a robust flight control system (FCS) for a small UAV. The main objective of this design is to ensure the small UAV can fly safely in severe gusty conditions. The Superaugmentation FCS consisted of Stability Augmentation System (SAS) and Command Stability Augmentation System (CSAS) was developed in UTMLST to improve the dynamic characteristics of the longitudinal stability of UAV; i.e UTM Camar. A combination of the variable stability technique along with advanced flying and handling qualities (FHQ) requirements are used to reduce the gust effect on the aircraft or UAV. The results obtained from the simulation studies showed that the superaugmented aircraft can be operated in severe gust environments than augmented aircraft. The result from here has reduced strain on the elevator activity in both extreme and calm weather conditions. Moreover, the superaugmentation FCS in the longitudinal axis meets the requirements of the level 1 handling qualities specification in flight phase

The effects of reynolds number on flow separation of Naca Aerofoil

The purpose of this study is to investigate the flow separation above UTM 2D Airfoil at three different Reynolds numbers which are 1 × 106, 1.5 × 106 and 2 × 106 using pressure distribution method and flow visualization. The experiment was conducted in UTM-LST (Low Speed Tunnel). The pressure distribution is done on three different wing span, which are 40%, 50% and 70%m of span and was measured and plotted to observe the flow characteristic at angle of attack from 0° to 35° for all three different Reynolds numbers. The flow visualization method was done at 10m/s, 20m/s and 30m/s airspeed from 0° to 18°. It is concluded that the Reynolds number of 1 × 106 separates at 16° Reynolds number of 1.5 × 106 separates at 18° and Reynolds number of 2 × 106 separates at 20°

Wind tunnel experiments on a generic sharp-edge delta wing UAV model

Delta wing is a triangular shape platform from a plan view. Delta wing can be applied to aircraft development as well as UAV. However, the flow around delta wing is very complicated and unresolved to date. On the upper surface of the wing, vortex is developed which need more studies to understand this flow physics. This paper discusses an experiment study of active flow control applied on the sharp-edged generic delta wing UAV. This paper focuses on the effect of rotating propeller on the vortex properties above a generic 550 swept angle model. The model has an overall length of 0.99 meter and the experiments were performed in Universiti Teknologi Malaysia Low Speed wind tunnel sized of 1.5 x 2.0 meter2. In this experiment, the experiments were conducted at a speed of 18 m/s. In order to differentiate the effect of propeller size on the vortex system, the experiment was carried out in three stages, i.e., experimental without propeller called as clean wing configuration and followed by the experiment with propeller diameter of 13”. The final experiment was the experiment with propeller diameter of 14”. During the experiments, two measurement techniques were employed; steady forces and surface pressure measurements. The experimental data highlights an impact of propeller size on the coefficients of lift, drag, and moment and vortex system of the delta-shaped UAV. The results obtained indicate that the lift is increased particularly at high angle of attack. The results also show that vortex breakdown is delayed further aft of the wing when propeller rotating at about 5000 RPM

Micro Air Vehicle: Technology Review and Design Study

This paper presents the technology review and design study now carrying-out at UTM on the micro air vehicles (MAVs). There are several existing MAVs in the market with various configurations and built for a designated function. These vehicles may carry visual, acoustic, chemical or biological sensors for such missions as traffic management, hostage situation surveillance, rescue operations and a few more. In this report, all existing technology for MAV airframes and systems was reviewed. After reviewing all the configurations, the concept of rotary wing MAV was selected for our design study. The concept was selected mainly due to its stability during manoeuvres. The components selection and system integration required for building the MAV were also discussed. The components were arranged in such a way that the centre of gravity of the MAV is located at the centreline of the body. This is to stabilise the vehicle statically and dynamically

Control of vibrational power input to semi-infinite anisotropic beam

A computational method has been developed to realize an efficient and accurate user friendly computational program called RocketCalculator, which integrates carefully assessed methodologies and offer flexibilities of the rocket configurations and velocities up to Mach number 3.0 and angle of attack of 25 degrees. The RocketCalculator is capable of analyzing the configurations of wing-alone, bodyalone, wing-body combination, and wing-body-tail combination of rocket. USAF Datcom Method has been chosen as the analysis method and the programming language is Microsoft Visual Basic. The result would be displayed in form of the corresponding lift (CL ), normal force (CN ) and drag (CD ) coefficients at certain Mach number and angle of attack. Experimented data for several models have been taken out from available sources to validate the program output. Comparisons of the program output and experimental results generally show good agreement with average error of less than 10

Development of a computer program for rocket aerodynmic coefficients estimation

A computational method has been developed to realize an efficient and accurate user friendly computational program called RocketCalculator, which integrates carefully assessed methodologies and offer flexibilities of the rocket configurations and velocities up to Mach number 3.0 and angle of attack of 25 degrees. The RocketCalculator is capable of analyzing the configurations of wing-alone, bodyalone, wing-body combination, and wing-body-tail combination of rocket. USAF Datcom Method has been chosen as the analysis method and the programming language is Microsoft Visual Basic. The result would be displayed in form of the corresponding lift (CL ), normal force (CN ) and drag (CD ) coefficients at certain Mach number and angle of attack. Experimented data for several models have been taken out from available sources to validate the program output. Comparisons of the program output and experimental results generally show good agreement with average error of less than 10%