PRESENTED BY: VIVEK KUMAR SHRIVASTAV 17AE60R23 DEPARTMENT OF AEROSPACE ENGINEERING DEPARTMENT OF AEROSPACE ENGINEERING INDIAN INSTITUTE OF TECHNOLOGY KHARAGPUR INDIAN INSTITUTE OF TECHNOLOGY KHARAGPUR

SUPER-CRITICAL AIRFOIL DESCRIPTION BENEFITS FLOW PARAMETERS MESH COMPARISON SETUP BOUNDARY CONDITIONS RESULTS CONTENT

SUPER-CRITICAL AIRFOIL Standard wing shapes are designed to create lower pressure over the top of the wing. The camber of the wing determines how much the air accelerates around the wing. As the speed of the aircraft approaches the speed of sound, the air accelerating around the wing will reach Mach 1 and shock waves will begin to form. The formation of these shock waves causes wave drag. Supercritical aerofoils are designed to minimize this effect by flattening the upper surface of the wing.

A supercritical aerofoil is an aerofoil designed, primarily, to delay the onset of wave drag in the transonic speed range. Supercritical aerofoils are characterized by their flattened upper surface, highly cambered (curved) aft section, and larger leading edge radius compared with NACA 6-series laminar airfoil shapes. P.C. Wikipedia Conventional Airfoil Supercritical Airfoil

DESCRIPTION Research aircraft of the 1950s and '60s found it difficult to break the sound barrier, or even reach Mach 0.9, with conventional aerofoils Supersonic airflow over the upper surface of the traditional aerofoil induced excessive wave drag and a form of stability loss called Mach tuck. Due to the aerofoil shape used, supercritical wings experience these problems less severely and at much higher speeds, thus allowing the wing to maintain high performance at speeds closer to Mach 1.

BENEFITS Supercritical aerofoils feature four main benefits. they have a higher drag divergence Mach number. they develop shock waves further aft than traditional aerofoils. they greatly reduce shock-induced boundary layer separation. their geometry allows for more efficient wing design

FLOW PARAMETERS AIRFOIL = SC(2)-0410 & NACA 0010 AOA = 4 0 Pressure=101325N/m 2 Temperature=300kK Density=1.177Kg/m 3 Chord Length=1m

MESH COMPARISON SC20410NACA0010

SETUP Solver Time = Density Based Solver Time = Transient Energy Equation = On Viscous = Standard K-e, standard wall function Material = Ideal Gas

BOUNDARY CONDITIONS

0.7 MACH Mach Number Variation Static Pressure Variation NACA 0010 Mach Number Variation Static Pressure Variation SC20410

0.8 MACH Mach Number Variation Static Pressure Variation NACA 0010 Mach Number Variation Static Pressure Variation SC20410

0.85 MACH Mach Number Variation Static Pressure Variation NACA 0010 Mach Number Variation Static Pressure Variation SC20410

0.95 MACH Mach Number Variation Static Pressure Variation NACA 0010 Mach Number Variation Static Pressure Variation SC20410

OBTAINED VALUE OF FORCES FOR AIRFOILS

Variation Of C D With Mach Number

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Variation Of C D With Mach Number