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CFD ANALYSIS OF UAV USING ANSYS

Paper: ASAT-13-AE-1313th International Conference on AEROSPACE SCIENCES & AVIATION TECHNOLOGY, ASAT- 13, May 26 28, 2009, E-Mail: asat@mtc.edu.eg Military Technical College, Kobry Elkobbah, Cairo, Egypt Tel : +(202) 24025292 24036138, Fax: +(202) 22621908

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Wind Tunnel Experiments and CFD Analysis of Blended Wing Body (BWB) Unmanned Aerial Vehicle (UAV)

at Mach 0.1 and Mach 0.3

Wirachman Wisnoe*, Rizal Effendy Mohd Nasir**, Wahyu Kuntjoro***, and Aman Mohd Ihsan Mamat

Abstract: This paper discusses the aerodynamics behavior of a baseline design of a Blended Wing Body (BWB) aircraft developed at MARA University of Technology (UiTM). Two methods of analysis are presented, i.e. Steady-state, three- dimensional Computational Fluid Dynamics (CFD) of the BWB at Mach 0.3 and Wind Tunnel experiments on 1/6 scaled half model of the BWB at Mach 0.1. In both methods of analysis, Lift Coefficient (CL), Drag Coefficient (CD) and Pitching Moment Coefficient (CM) are measured and compared at respective Mach numbers with respect to variation of angle of attack. Pressure contours and Mach number contours are plotted and the turbulence area is predicted, both extracted from CFD analysis. Visualization using mini tuft during wind tunnel tests is also executed to complete the analysis where the stall progression patterns can be clearly observed. The presented BWB UAV design here has achieved an unprecedented capability in terms of sustainability of flight at high angle of attack, low parasite drag coefficient and decent maximum lift coefficient. Some recommendations for future improvement of the BWB are given. Keywords: Aerodynamics, Blended-Wind Body (BWB), Unmanned Aerial Vehicle (UAV), Computational Fluid Dynamics, Wind Tunnel, mini-tuft visualization 1. Introduction Blended Wing Body (BWB) aircraft is a concept where fuselage is merged with wing and tail to become a single entity [1]. BWB is a hybrid of flying-wing aircraft and the conventional aircraft where the body is designed to have a shape of an airfoil and carefully streamlined with the wing to have a desired planform. If the wing in conventional aircraft is the main contributor to the generation of lift, the fuselage of BWB generates lift together with the wing

* Associate Professor, Faculty of Mechanical Engineering, Universiti Teknologi MARA, 40450 Shah Alam, MALAYSIA, wira_wisnoe@yahoo.com ** Lecturer, Faculty of Mechanical Engineering, Universiti Teknologi MARA, 40450 Shah Alam, MALAYSIA, fendy117@yahoo.com *** Professor, Faculty of Mechanical Engineering, Universiti Teknologi MARA, 40450 Shah Alam, MALAYSIA, wkuntjoro@yahoo.com Lecturer, Faculty of Mechanical Engineering, Universiti Teknologi MARA, 40450 Shah Alam, MALAYSIA, aman_korup@yahoo.com

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thus increasing the effective lifting surface area. The streamlined shape between fuselage and wing intersections reduces interference drag, reduces wetted surface area that reduces friction drag while the slow evolution of fuselage-to-wing thickness by careful design may suggest that more volume can be stored inside the BWB aircraft, hence, increases payload and fuel capacity [1][2]. The BWB concept aims at combining the advantages of a flying wing with the loading capabilities of a conventional airliner by creating a wide body in the center of the wing to allow space for passengers and cargo. Especially, for very large transport aircraft, the BWB concept is often claimed to be superior compared to conventional configurations in terms of higher lift-to-drag ratio and consequently less fuel consumption [3]. Since September 2005, UiTM has started a research on BWB of Unmanned Aerial Vehicle (UAV) [4]. The planform of BWB UAV determined during the preliminary can be seen in Fig. 1. Preliminary structural configuration has been analyzed using finite element model [5]. Computational fluid dynamic analysis was also performed on the basic planform at various Mach numbers [6][7] and experimental analysis has been conducted to establish the validation of the CFD simulation [8][9]. This paper will focus on aerodynamic study of preliminary design of BWB planform to be used as a UAV. It consists of Computational Fluid Dynamics (CFD) study of the BWB at 0.1 and 0.3 Mach number, and wind tunnel tests at around 0.1 Mach number. Mach 0.3 represents the cruising phase of the BWB during its flight mission and Mach 0.1 represents the loitering phase (Fig. 2). The aerodynamic characteristics such as lift coefficient, drag coefficient and pitching moment coefficient are obtained and compared. In addition, the CFD visualization of flow and pressure distribution on the surface of the BWB model had been plotted to analyze the flow behavior on the BWB surface at the sub sonic level. On the other hand, visualization using mini tuft is performed in the wind tunnel to observe the quality of flow pattern around the BWB at various angles of attack. 2. CFD Approach, Methodology and Experimental Setup

2.1. CFD Approach, Methodology Development of mathematical models involves derivation of geometrical equations for BWB planform. Some parameters such as wing planform area, sweep angle, taper ratio for each section (body, inner wing and outer wing), span and chord for various spanwise locations must be determined for each configuration. These mathematical models are then translated into three dimensional drawing in CATIA where the sizing of preliminary BWB planform configuration is taken as relative to wing span. Fig. 1 shows the three-view drawing of the BWB proposed during its preliminary design. Second stage involves conversion from three dimensional CATIA CAD model into CFD element in GAMBIT to create the meshing element. Then, the succeeded meshing models were exported to FLUENT for the analysis. The result presented is the simulation for sub sonic flow at Mach number equals to 0.1 and 0.3 corresponding to Reynolds number equals to 4.66 106 and 1.4 107 respectively. Boundary conditions and airflow are simulated in this stage purposely-executed for two reasons; first is determination of pressure distribution on the surface of the BWB UAV that later on leads to calculations of aerodynamics characteristics of BWB such as CL, CD and CM at various angle of attack, second is the visualization of the

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airflow around the BWB using Post Processing to recognize some critical area with possible vortex reduction in the near future. The analytical of aerodynamics characteristics for various angles of attacks using CFD simulation will be conducted in this final stage.

2.2. Experimental Setup The tests were conducted using UiTM Low Speed Tunnel LST-1 (Fig. 3). This wind tunnel has a test section area of 0.5 m x 0.5 m x 1.25 m. It is a suction type tunnel, equipped with a 3-component balance, capable of measuring lift force, drag force, and pitching moment. Hence, a half model of BWB is used for the tests. The BWB planform was obtained from [4]. The half model of this BWB-UAV has been manufactured using CNC machine with a size reduction of 1/6 [10]. Fig. 4 shows the dimension of the half model and Fig. 5 shows the manufactured model. The tests are conducted at 4 different air speeds, i.e. 25 m/s, 30 m/s, 35 m/s, and 40 m/s. The Reynolds number, using base chord length as reference length, is on the order of 105, and the Mach number ranges from 0.07 to 0.11. 3. Results and Discussion

3.1 Lift Coefficient Analysis Fig. 6 shows variation of lift coefficient (CL) for different values of angle of attack (). For each wind tunnel airspeed, the value of CL increases as the angle of attack is increased until its maximum value at around = 35 and decreases afterwards with lower slope. Computational Fluid Dynamics (CFD) results at Mach 0.1 and 0.3 also give the same trend with maximum CL located at = 39 and = 35 respectively. Table 1 summarizes the maximum values of CL obtained at different measurement. It is observed that the value of CLmax increases as the air velocity of the wind tunnel is increased. Hence, the CLmax increases with the increase of Reynolds number (Fig. 7). This explains the difference of values of CLmax between the experiments and the CFD.

Table 1. Maximum values of CL for different airspeeds (or Mach numbers)

v (m/s) M (deg) CL max

25 0.072 34 0.649 30 0.086 35 0.664 35 0.101 34 0.687 40 0.115 34.5 0.749

CFD (M=0.1) 0.1 39 1.043 CFD (M=0.3) 0.3 35 1.031

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The experimental curves presented in Fig. 6 also show clearly a deflection at around = 8. The visualization reveals the flow separation around the wing area around this angle of attack that results in reduction in lifting surface area (Fig. 8 and Fig. 9). The deflection of curves is also observed in CFD results, although it is not very clear. Fig. 8 to Fig. 10 show visualization using mini tuft, taken at wind tunnel airspeed of 35 m/s. In Fig. 8, it can be seen that the flow is still attached to the overall surface at = 7. However, in Fig. 9, at = 8, the flow has almost completely separated from the wing, except around the wing tip. This means that, above this angle of attack, only the body generates the lift. The body will continue generating lift until CLmax and from there; separation starts to occur on the body part. The separation does not occur completely, as there is always part of the body where the flow is still attached. Fig. 10 shows the flow pattern around the body at = 42.

3.2 Drag Coefficient Analysis Fig. 11 shows variation of drag coefficient (CD) versus angle of attack () taken at different air speeds and Mach numbers. It is observed that the variation of drag coefficient is very slow and almost constant