The analysis of flow on round-edged delta wings

The flow around three-dimensional aircraft wings, including delta-wings is very complicated. Much experimental and numerical work has been performed to discover its complexity. To date, all numerical calculations on delta wings have been carried out for either fully laminar or fully turbulent boundary layers. The transition status of the boundary layer is considered unknown despite several efforts to identify transition from laminar to turbulent flow. One such study, called the International Vortex Flow Experiment – 2 (VFE-2), has been carried out by an international group and mainly focuses on the boundary layers on delta wings. The data from the VFE-2 experimentals potentially provide the location of transition on the upper and lower surfaces of the wing to guide associated numerical studies. The effects of Reynolds number, Mach number, angle of attack and the leading edge bluntness are also investigated. Almost all delta wing studies to date have involved tests on wings with sharp leading edges and these have led to the conclusion that the flows are relatively independent of Reynolds number. In fact, most real wings have finite leading edge radii. Hence, the flow separation is no longer fixed at the leading edge, thus making the flow dependent on Reynolds number. This particular aspect has been studied extensively by the VFE-2 team. As part of the VFE-2 project, Glasgow University constructed a delta wing with four different sets of leading edges. Small-, medium- and large-radius edges and a pair of sharp leading edges were constructed in order to compare results from four delta wing configurations. In the current study experiments were carried out on these wings in the 2.65 metre by 2.04 metre, closed circuit, Argyll Wind tunnel of Glasgow University. The models were mounted on a specially designed sting support structure that allowed them to be pitched around a constant centre of rotation throughout the experiments. Tests were conducted at speeds of 20.63 m/s and 41.23 m/s representing Reynolds numbers of 1 x 106 & 2 x 106 respectively, based on the mean aerodynamic chords of the wings. The tests were conducted in three phases. In the first phase, steady and unsteady forces and moments on all wings were measured at an angle of attack that varied from α =100 to 250. The forces and moments were captured at two sampling rates; i.e., 100 Hz and 8000 Hz. The second test series captured flow visualization data on the four wings. In these experiments, a mixture of Ondina oil and paraffin was combined with Dayglo powder and applied to the surfaces of the delta wings. The images of the flow topology on the wings were recorded. The final series of experiments involved Particle Image Velocimetry measurements. A stereo-PIV arrangement was applied in this experiment and two CCD-Cameras were positioned outside the test section for image capture. The current study has identified interesting features of the interrelationship between the conventional leading edge primary vortex and the occurrence and development of the inner vortex on the round-edged delta wings. The inner vortex was first identified and verified by the VFE-2 team. The effects of Reynolds number, angle of attack and leading-edge radii on both vortices are discussed in detail. The steady balance data have shown that the normal force coefficients are sensitive to leading edge bluntness at moderate angles of attack but are less so at high angles of attack. In relation to this, the flow visualization images have also shown that the primary vortex origin is located further aft on the wing at higher leading edge bluntness. This impacts on the strength of the inner vortex which remains a significant flow feature until the primary vortex approaches the apex. The lateral extent of the inner vortex is very dependent on the primary vortex at the leading edge; i.e. the weakening of the primary vortex has positive effects on the inner vortex. Particle Image Velocimetry shows that the increase in leading edge bluntness significantly decreases the swirl magnitude of the primary vortex. The results obtained from the current investigation provide considerable insight into the effects of Reynolds number, angle of attack and bluntness on the flow structures experienced by delta wings, with rounded leading edges. This work will, therefore, inform and guide future investigations of delta wing flows and has the potential to impact on future wing design

Computational Fluid Dynamic Simulation (CFD) and Experimental Study on Wing-external Store Aerodynamic Interference

The main objective of the present work is to study the effect of an external store to a subsonic fighter aircraft. Generally most modern fighter aircraft is designed with an external store installation. In this project a subsonic fighter aircraft model has been manufactured using a computer numerical control machine for the purpose of studying the effect of the external store aerodynamic interference on the flow around the aircraft wing. A computational fluid dynamic (CFD) and wind tunnel testing experiments have been carried out to ensure the aerodynamic characteristic of the model then certified the aircraft will not facing any difficulties in stability and controllability. In the CFD experiment, commercial CFD code is used to simulate the interference and aerodynamic characteristics of the model. Subsequently, the model together with an external store was tested in a low speed wind tunnel with test section sized 0.45 m×0.45 m. Result in the two-dimensional pressure distribution obtained by both experiments are comparable. There is only 12% deviation in pressure distribution found in wind tunnel testing compared to the result predicted by the CFD. The result shows that the effect of the external storage is only significant at the lower surface of the wing and almost negligible at the upper surface of the wing. Aerodynamic interference is due to the external storage were mostly evidence on a lower surface of the wing and almost negligible on the upper surface at low angle of attack. In addition, the area of influence on the wing surface by store interference increased as the airspeed increase.

Computational Fluid Dynamic Simulation (CFD) and Experimental Study on Wing-external Store Aerodynamic Interference of a Subsonic Fighter Aircraft

The main objective of the present work is to study the effect of an external store on a subsonic fighter aircraft. Generally most modern fighter aircrafts are designed with an external store installation. In this study, a subsonic fighter aircraft model has been manufactured using a computer numerical control machine for the purpose of studying the effect of the aerodynamic interference of the external store on the flow around the aircraft wing. A computational fluid dynamic (CFD) simulation was also carried out on the same configuration. Both the CFD and the wind tunnel testing were carried out at a Reynolds number 1.86×105 to ensure that the aerodynamic characteristic can certify that the aircraft will not be face any difficulties in its stability and controllability. Both the experiments and the simulation were carried out at the same Reynolds number in order to verify each other. In the CFD simulation, a commercial CFD code was used to simulate the interference and aerodynamic characteristics of the model. Subsequently, the model together with an external store was tested in a low speed wind tunnel with a test section sized 0.45 m×0.45 m. Measured and computed results for the two-dimensional pressure distribution were satisfactorily comparable. There is only a 19% deviation between pressure distribution measured in wind tunnel testing and the result predicted by the CFD. The result shows that the effect of the external storage is only significant on the lower surface of the wing and almost negligible on the upper surface of the wing. Aerodynamic interference due to the external store was most evident on the lower surface of the wing and almost negligible on the upper surface at a low angle of attack. In addition, the area of influence on the wing surface by the store interference increased as the airspeed increased

Experimental and simulation Studies of a Two Seater Light Aircraft

This paper presents the aerodynamic studies carried out on a three-dimensional aircraft model. The test model is a 15% scaled down from a two-seater light aircraft that close to the Malaysian made SME MD3-160 aircraft. The aircraft model is equipped with control surfaces such as flaps, aileron, rudder and elevator and it is designed for pressure measurement testing and direct force measurement using a 6-components balance system. This aircraft model has been tested at two different low speed tunnels, at Universiti Teknologi Malaysia tunnel sized 1.5 x 2.0 meter2 test section, and at Institute Aerodynamic Research, National Research Council of Canada sized 3 x 2 meter2 tunnel. The speed during testing at UTM and IAR/NRC tunnels was up to 70 meter/second, which is corresponds to Reynolds numbers of 1.3 x 106.The longitudinal and lateral directional aerodynamic characteristics of the aircraft such as coefficients of pressure, forces (lift, drag, side) and moments (roll, pitch and yaw) have been experimentally measured either using direct force measurement or pressure measurement method. The data reduction methods include the strut support interference factor using dummy image and the blockage correction have been applied in this project. The results showed that for the undeployed flap configuration, the stalling angle of this aircraft is 160 at CLMax = 1.05 measured by UTM - LST, compared to CLMax =1.09 at stalling angle 150 by IAR- NRC. Beside the experimental study, simulation also be performed by using a commercial Computational Fluid Dynamics (CFD) code, FLUENT Version 5.3. Experimental works at UTM and IAR – NRC tunnel show that the aerodynamic characteristics of this light aircraft are in a good agreement with each other. Simultaneously, the aerodynamic forces obtained from experimental works and CFD simulations have been compared. The results proved that they are agreeable especially at a low angle of attack

The effect of edge profile on delta wing flow

This paper presents flow measurements on four delta wing configurations which are differentiated by their leading edge profiles; sharp-edged, small, medium, and large radius. The experiments were performed as a part of the European Vortex Flow Experiment-2 campaign. Tests were conducted at speeds of 20.63 m/s and 41.23 m/s representing Reynolds numbers of 1 × 10<sup>6</sup> and 2 × 10<sup>6</sup>, respectively. In this paper, oil flow visualization data are presented for the four wings together with particle image velocimetry results for the large radius wing. The study has identified interesting features of the interrelationship between the conventional leading edge primary vortex and the occurrence and development of the inner vortex on the round-edged delta wings. The effects of Reynolds number, angle of attack, and leading-edge radii on both vortex systems are discussed in detail

Tomography systems and sensor application for sharp-edged delta wing analysis

This paper presents the results on wind tunnel testing above non-slender sharp-edged delta wing under pitching motion. Above the sharp-edged delta wing the flow topology is very complex, disorganized and unresolved till date. The primary vortex onset is occurred at the wing apex of the sharp leading-edge delta wing and it develops from the leading edge of the wing to the trailing edge. There are several factors that influenced the vortex properties above the wing such as angle of attack, Reynolds number, Mach number, leading-edge bluntness and flow control techniques. The main objective of this study is to show the flow control technique effects known as the blower, above the sharp-edged non-slender delta wing on the flow topology. The experiments were carried out in the wind tunnel at Reynolds number of 0.8×106 with the speed of 25m/s. A generic delta wing model was fabricated in UTM with a sweep angle of 55˚, the model was designed to be installed to UTM Aerolab external strain gauge. During the experiments, the blowers were placed at three different positions 15%, 50% and 70% from the Apex of the wing and these locations are named as location I, II and III. In order to measure the vortex, a measurement technique called as surface pressure measurement was employed on the wing. The experiments were divided into two phases. The first phase was the clean wing configuration where the experiment was performed without the flow control. The final experiment was the experiments with the flow control at three different locations. The results have shown that the location of blower has influenced the flow characteristics above the wing. The result obtained shows that the blower at location I has an impact at higher angle of attack while with the blower at location II and III, the blower has significantly increased the primary vortex in size and at the high angle of attack the vortex breakdown is considerably delayed

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°

Air pressure sensor using fiber bragg gratings (FBG as air pressure sensors on generic UTM-LST half model)

This work was performed to investigate the feasibility of using Fiber Bragg Gratings (FBGs) strain sensor in detection of air pressure on aeroplane model known as Generic UTM Half-Model. The FBGs was attached on the surface of the aeroplane model where its position is as near as possible to the location of static conventional pressure sensor. Then, the sensing performance was tested inside UTM Low Speed Tunnel (UTM-LST) with the wind speed set at 30 ms^(-1), 40. ms.^(-1), and 50 ms^(-1). The direction of wind was arranged to be in perpendicular to the FBG and the position of wing model was varied at angle of 0°, 5°, 10°, 15°, and 20°. The measured pressure coefficient, Cp based on Bragg wavelength shift was compared with Static FKPS 30DP Pressure Measuring Module data. The results reveal that the shift in Bragg wavelength was found to increase linearly from angle 0° until 10° and after that the wavelength shift become saturated. The pressure coefficient obtained by FBGs has well agreed with the value obtained by pressure coefficient of pressure sensor module at low angle of attack from 0° to 10°

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