The Drags Force at the AirSpeed - Case Study Example

EAT216 COMPUTER AIDED ENGINEERING by Introduction The paper investigates the flow pattern of air over the aerofoil and calculates the drag force at the air speed specified. For validation reasons, it compares drag force with that over a flat plate of similar dimensions calculated using the manual calculation techniques developed in EAT223. Description Helicopters are capable of flying due to aerodynamic forces created when air flows about the airfoil. An aerofoil is any exterior exerting more lift than drag the moment the airplane passing through the air at an appropriate angle. Airfoils are in most cases associated with generation of lift. Aerofoils also assist in control (elevator), stability (fin), and propulsion or thrust (rotor or propeller). Selected aerofoils is NACA 0012 aerofoil model (Pantula 2008). This model simulates the passing around an angled NACA 0012 aerofoil at diverse angles of attack with the SST turbulence framework. The outcomes show good accord with the investigational lift statistics of Ladson and the pressure data (Pantula 2008). Values for the Simulation The reference values of simulations applicable in initializing the recreation and to indicating boundary conditions. In this report, the following set reference values for pressure, velocity, and temperature applicable (Findon 2013). The settings for this study case are: Temperature (T) = 260 K equivalent of = -13.15 C Pressure (P) = 85419 Pa = 85419 N.m-2 Velocity (U) = [241.8220 16.9098 0] m.s-1, from: Angle of attack (alpha) = 4° equivalent to =[cos 4, sin 4, 0] = [0.9976 0.06976 0] Mach Number = 0.75 = 242.41 m.s-1 (Findon 2013). Conversion computation using: gamma = 1.4, R = 8.314472 J·K-1·mol-1 and W = 0.02897 Kg·mol-1 (Findon 2013). Pressure Flow Velocity Flow Screen shot The CFD Module is the blueprint for simulating applications and devices that engages sophisticated fluid pressure models. As is the scenario with every module in the COMSOL commodity Suite, the CFD Module offers ready-made physics links that prearranged to receive system inputs through the graphical user interface (GUI), and to employ the inputs to originate model equations. The precise physics interfaces applicable the CFD Module is equipped to model most conditions of fluid flow, including illustrations of compressible, porous media flows, nonisothermal, two-phase, and non-Newtonian (International Symposium on Shock Waves and Kontis 2012). Summary Table Simulation Conditions NACA 0012 airfoil geometry Blunt trailing edge Incompressible flow _ Re = 6.0e6 For compressible scenario, M=.15 Angles of attack (alpha) investigated: -4 - 20 Completely turbulent Outcome of interest Drag coefficient Lift coefficient Pressure coefficient (ACUSIM Software, Inc. 2010) Boundary conditions Source: ACUSIM Software, Inc. 2010 The objective here is to offer a validation scenario for turbulence NACA 0012 model. Contrasting verification that seeks to institute that a model executed correctly, validation assesses CFD outcomes against data in an attempt to institute a models capability of reproducing physics (Langley Research Center 2014). A large range of nested grids of the similar family is given. For this case, fundamentally incompressible NACA 0012 airfoil model, the data are from investigations (Langley Research Center 2014). For the reasons of this validation, the description of the NACA 0012 airfoil is somewhat changed from the initial definition to allow the airfoil ends at chord=1 with a pointed trailing edge (Langley Research Center 2014). To do this, the exact NACA 0012 formula y= +- 0.6*[0.30*sqrt(x) - 0.14*x - 0.35*x2 + 0.28*x3 - 0.10*x4] is applicable to generate an airfoil amid x=0 and x=1.008930411365 (the T.E. is pointed at this position). Then 1.008930411365 minimizes the airfoil (Langley Research Center 2014). Therefore, the consequential airfoil is a ideal scaled replica of the 0012, with utmost thickness of about 11.9% comparative to its chord (the ideal NACA 0012 has an upper limit thickness of 12% comparative to its blunted chord. However, it has a maximum density of 11.894% comparative to its chord extensive to 1.008930411365) (Langley Research Center 2014). The revised definition is: y= +- 0.59468*[0.29822*sqrt(x) - 0.12713*x - 0.35790*x2 + 0.29198*x3 - 0.10517*x4] Prior to mid 2014, a typo in the original scaled principle was applicable. It formula was y= +- 0.59468*[0.29822*sqrt(x) - 0.12712*x - 0.35791*x2 + 0.29199*x3 - 0.10517*x4]. The typo, there was a small array 10-8 non-closure at the irregular edge (T.E) (Langley Research Center 2014). The Velocity and Pressure Profiles Simulate a steady condition, compressible (transonic) turbulent flow about the NACA 0012 airfoil in simulated 2D. Observe pressure coefficient (Cp) curves and the Cp supply over the airfoil plane (International Symposium on Shock Waves and Kontis 2012). Source: clarkson.edu n.d Center of Pressure c.p. the resultant forces (drag and lift) pressuring at the center create no moment (employ centroid principle to find R location) (Parallel CFD Conference and Chetverushkin 2004). Since the pressure supply over the airfoil transforms considering alpha, the position of the center of pressure ranges with alpha (angle-of-attack) (clarkson.edu n.d). Screenshot The FreeCASE mesh is maximized for an inviscid (viscosity gratis) flow recreation, so it has no mesh clustered secure to the airfoil to determine the viscous border layer (Parallel CFD Conference and Chetverushkin 2004). Hence, emploring the mesh is to aid comprehension of how to import a net and organize a transonic flow reproduction with Caedium, instead of producing a definitive viscous outcome (Findon 2013). Significance of the Velocity and Pressure with Regard to Aerofoil Pressure coefficient It involve defining another coefficient that illustrates the pressure supply over an airfoil plane The draft show cp ranges over both lower and upper surfaces. cp can be deliberated experimentally in the current of air tunnel (Dole and Lewis 2000). The upper surface (suction surface) is usually associated with increased velocity and lower stationary pressure. The lower surface (pressure surface) has a moderately higher static pressure compared to the suction surface. The pressure slope between these two planes leads to the lift force created for a given airfoil (Dole and Lewis 2000). Pressure coefficient: Since the first Cp value is 0 all over, no contours will be displayed until the simulation is execute (Findon 2013). Summary If allocations of pressure coefficient statistics over the lower and upper surfaces vs. chordwise space (x) are accessible The lift coefficient establish as the net region between the lower and upper pressure coefficient contours divided by the chord distance end to end The part lift coefficient (cl) equation is an excellent estimate only for small alphas Pressure coefficients defined Pressure coefficient on the pressure surface Pressure coefficient on the suction surface (SEM Conference on Experimental and Applied Mechanics and Proulx 2011) The SolidWorks Calculation of Drag Force Manual Calculation of The Equivalent Flat Plate Drags Force. The moment the drag force FD, the maximum velocity V, and the liquid density  are calculated during pressure over a body, the drag coefficient is acquired from Where A is usually the frontal region (the region projected on a surface normal to the course of flow) of the aircraft (Chang 2014) However, in some scenarios such as flat planes associated with the flow or aircraft wings, the plan form region applicable instead of the frontal region. Planform part is the region projected on a airplane parallel to the course of flow and standard to the lift force. A similar analysis provides Moment coefficient Drag coefficient The coefficient has c clearly included This term denotes for the force x length measures (clarkson.edu n.d). Summarizing (clarkson.edu n.d). Generic drag coefficient variation 12 Cl-α Experimental statistics are important to aircraft design , according to NACA/NASA data. Cl varies linearly through α–camber alterations αL=0. This linear association breaks down the moment stall occurs(clarkson.edu n.d). Reference List ACUSIM Software, Inc. 2010, Validation of AcuSolve for Aerodynamics of Lifting Bodies. Retrieved February 28, 2015, from www.acusim.com Chang, KH 2014, Motion simulation and mechanism design with solidworks motion 2013. clarkson.edu n.d, AE 429-Aircraft Performance and Flight Mechanics. Retrieved February 28, 2015, from http://people.clarkson.edu/~pmarzocc/AE429/AE-429-3.pdf Dole, CE & Lewis, JE 2000, Flight theory and aerodynamics: a practical guide for operational safety. New York [u.a.], Wiley. Findon, A 2013, Transonic Flow Over the NACA 0012 Airfoil. Retrieved February 28, 2015, from http://www.symscape.com/node/819 International Symposium on Shock Waves, & Kontis, K 2012, 28th International Symposium on Shock Waves. Vol 2 Vol 2. Berlin, Springer. http://dx.doi.org/10.1007/978-3-642-25685-1. Langley Research Center, 2014, 2D NACA 0012 Airfoil Validation Case. Turbulence Modeling Resource Retrieved February 28, 2015, from http://turbmodels.larc.nasa.gov/naca0012_val.html Pantula, RS 2008, Modeling Fluid Structure Interaction over a Flexible Fin Attached to a NACA0012 Airfoil. Western Michigan University; ProQuest. Parallel CFD Conference, & Chetverushkin, BN 2004, Parallel computational fluid dynamics advanced numerical methods : software and applications : proceedings of the Parallel CFD 2003 Conference, Moscow, Russia (May 13-15, 2003). Amsterdam, Elsevier. http://public.eblib.com/choice/publicfullrecord.aspx?p=288815. SEM Conference on Experimental and Applied Mechanics, & Proulx, TA 2011, Experimental and applied mechanics. proceedings of the 2011 Annual Conference on Experimental and Applied Mechanics Volume 6 Volume 6. New York, Springer. http://public.eblib.com/choice/publicfullrecord.aspx?p=763776. Read More