Explain how a "Lambda Shock Wave" forms when a shock interacts with
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- vortex generators on the upper surface of a wing will a. decrease the spanwise flow at high Mach numbers b. increase the critical Mach number c. decrease the intensity of the shockwave effects d. increase intensity of the shockwave effects1-derive oblique shock wave relations (compressible aerodynamic) 2-what is the difference between Normal schock wave and oblique schock wave (compressible aerodynamic)Consider a flat plate at an angle of attack α to a Mach 2.4 airflow at 1 atmpressure. What is the maximum pressure that can occur on the platesurface and still have an attached shock wave at the leading edge? At whatvalue of α does this occur?
- und engoes anormal Shock as Showeh in the figure An air Strean with Velocity air Stream with Velocity of (65omls) a Static 5.7 Tind e Mach number and velocitn akter the nornal Shoch wave shock 6 The Static conditions afker the normal9 Loave he Stag na tron conditionns after the normal shock wave O The enthopy clange across the norma Shock wave Pメ- 222 koy 义 Tメー452K 650m/SGive a physical description of the phenomenon of chokingin a converging-nozzle gas fl ow. Could choking happeneven if wall friction were not negligible?Consider a cone at zero angle of attack in a hypersonic flow. (Hypersonic flow is very high-speed flow, generally defined as any flow above a Mach number of 5.) The half-angle of the cone is θc, as shown inthe figure. An approximate expression for the pressure coefficient on the surface of ahypersonic body is given by the newtonian sine-squared law : Cp = 2 sin2 θcNote that Cp, hence, p, is constant along the inclined surface of the cone. Along the base of the body, we assume that p = p∞. Neglecting the effect of friction, obtain an expression for the drag coefficient of the cone, where CD is based on the area of the base Sb.
- Explainwhat do you understand by the term “Compressibility Corrections” for thin airfoils in high subsonic flow and elaboratewhy it is important?Consider a flat plate at α = 20◦ in a Mach 20 freestream. Using straightnewtonian theory, calculate the lift- and wave-drag coefficients. Comparethese results with exact shock-expansion theory.The shock wave photograph of a wedge structure is given. By using this photograph determine the Mach number.
- Air flows in a constant-area duct as shown in the figure. Assume Rayleigh line flow and the air to behave as a perfect gas with constant specific heats. For choked duct, determine: 1. Mach number at inlet (M1) 2. Stagnation temperature at inlet (To1) 3. Critical stagnation temperature (T.") 4. Mass flow rate 5. Heat input to choke the duct (q). q =? Vi = 100 m/s T1 = 320 K P1 = 200 kPa d = 1.5 cmCan the Mach number of a fluid be greater than 1 after a normal shock wave? Explain.Consider a flat plate with a chord length (from leading to trailing edge) of 1 m. The free-stream flow properties are M1 = 3, p1 = 1 atm, and T, = 270 K. Using shock-expansion theory, tabulate and plot on graph paper these properties as functions of angle of attack from 0 to 30° (use increments of 5°): a. Pressure on the top surface b. Pressure on the bottom surface c. Temperature on the top surface d. Temperature on the bottom surface e. Lift per unit span f. Drag per unit span g. Lift/drag ratio (Note: The results from this problem will be used for comparison with linear supersonic theory in Chap. 9.)