#NACA SUPERSONIC AIRFOIL GENERATOR SERIES#
#NACA SUPERSONIC AIRFOIL GENERATOR FREE#

Pressure and shear are in units of force per unit area (N/m2)Net unbalance creates an aerodynamic force Shear stress, tw, (friction) on surfaceActs tangentially to surface High ARLow ARFirst airplane designed for sustained flight at Mach 2Very sharp leading edge on wings (razor sharp leading edges, thickness 3.4 %)Designed to minimize wave drag at supersonic speedsVery poor low-speed aerodynamic performanceSuch wings tend to stall at low angles of attack, CLmax is only about 1.15Vstall (full of fuel) ~ 198 MPHVstall (fuel empty) ~ 152 MPHVstall proportional to W1/2EXAMPLE: F-104 LOCKHEED STARFIGHTERДЄspect RatioMean Chamber Line: Set of points halfway between upper and lower surfacesMeasured perpendicular to mean chamber line itselfLeading Edge: Most forward point of mean chamber lineTrailing Edge: Most reward point of mean chamber lineChord Line: Straight line connecting the leading and trailing edgesChord, c: Distance along the chord line from leading to trailing edgeCamber: Maximum distance between mean chamber line and chord lineMeasured perpendicular to chord lineAIRFOIL NOMENCLATUREДЄerodynamic forces exerted by airflow comes from only two sourcesPressure, p, distribution on surfaceActs normal to surface NACA FOUR-DIGIT SERIESFirst set of airfoils designed using this approach was NACA Four-Digit SeriesFirst digit specifies maximum camber in percentage of chordSecond digit indicates position of maximum camber in tenths of chordLast two digits provide maximum thickness of airfoil in percentage of chordД®xample: NACA 2415Airfoil has maximum thickness of 15%of chord (0.15c)Camber of 2% (0.02c) located 40%back from airfoil leading edge (0.4c) NACA 2415INFINITE VERSUS FINITE WINGS Slots and slats permit a new boundary layer to start on the main wing portion, eliminating the detrimental effect of the initial adverse gradient. Leading Edge DevicesLeading edge devices such as nose flaps, Kruger flaps, and slats reduce the pressure peak near the nose by changing the nose camber. As the aerofoil is designed in such a way that its upper surface is longer than the bottom, and because the molecules that hit the leading edge must meet again at the trailing edge, the ones that travel on the upper surface do so with greater velocity than the lower Drag Reduction And lift Augmentation MethodsДЇlap systemLeading edge devicesMulti element airfoilsCirculation controlLaminar flow controlwingletsFlap is an element attached to the aileron of the wing sectionIt is always possible to reduce stall speed by increasing wing area Flap systemsДЇlaps change the airfoil pressure distribution, increasing the camber of the airfoil and allowing more of the lift to be carried over the rear portion of the section This theory describes how the shape of the aerofoil produces a pressure difference which generates lift. The most common one is the Longer Path Theory. Many theories have been developed on how a wing generates lift. Reynolds No, Boundary Layer Transition and surface roughnessДЄn airfoil designed for minimum drag and uninterrupted flow of the boundary layer is called a laminar airfoil.Laminar Flow Airfoils **The tendency for flow to leak around the wing tipsgenerally cause streamlines over the top surface ofthe wing to veer to the wing root and streamlinesover the bottom surface veer to the wing tips.Д®ndplates (winglets) at end of wing reduces tip vortex If planes height is not changing then: Lift = WeightДЄp and c are independent of CL =L/(1/2 V2Ap)ДЄ10Camber line is straight for symmetrical aerfoilДЄs angle of attack increasesstagnation point moves downstream along bottom surface, causing an unfavorable pressure gradient at the nose*.

#NACA SUPERSONIC AIRFOIL GENERATOR FREE#

$X Requires LS-DYNA MPP R8.0.Forces on airplane atlevel speed and constantheight and speed.Lift force is the component of R that is perpendicular tofree stream velocity, and drag is the component of R parallel to the free stream velocity. $ Copying for non-commercial usage allowed if Bonhaus, A HIGHER ORDER ACCURATE FINITE ELEMENT METHODДЇOR VISCOUS COMPRESSIBLE FLOWS, PhD thesis, Virginia Polytechnic Institute Test case, the flow will be considered inviscid, no boundary layer will be developed on theД. Potential normal shock waves on the upper camber and lower camber. Over the airfoilвЂ™s surface, the flow will accelerate and become locally supersonic resulting in Transonic flow occurs when there is mixed subsonic and supersonic local flow in the sameДЇlow field (typically with free stream Mach numbers from M = 0.7 or 0.8 to 1.3). In this test case, the Mach number is high enough to reach the so called transonic flow. The NACA 0012 is a symmetrical airfoil frequently used for benchmarking test cases.

#NACA SUPERSONIC AIRFOIL GENERATOR SERIES#

The four digit series define the profile by describing its maximum camber, the location of the maximum camber and the airfoilвЂ™s maximum thickness. The NACA airfoils are airfoil shapes for aircraft wings developed by the National Advisory Committee for Aeronautics (NACA, former NASA) in the 1940s.

Naca supersonic airfoil generator