Aerodynamics II Posted on January 14, 2019 by aerovijay Aerodynamics is a branch of dynamics that deals with the motion of air and other gaseous fluids and with the forces acting on bodies in motion relative to such fluids. The qualities of an object that affect how easily it is able to move through the air. An example of a study of aerodynamics is to determine the potential speed of a vehicle based on the way its shape will move through air. Supersonic aerofoil region in transonic flow and linearized supersonic flow, (2D cases) No fundamental differences between two cases In linear supersonic flow the characteristics (Mach wave) are straight lines with a variable slope In the transonic case, both families of characteristics are present (expansion and compression waves). For both cases disturbances must propagate away from the wing in the far field What happens if you reduce the curvature of the aerofoil leading edge? Advanced the shock wave formation Delay the shock wave formation It will have no effect on shock wave formation None of these Newtonian models for Hypersonic momentum balance : Density ratio across shock is large Tangential components conserved Normanl components conserved None of these For supersonic wing increasing the aspect ratio : Lift-curve slope increases Supersonic drag increased Trailing-edge vortex drag increases None of these According to the slender body theory the fuselage Does not contribute to lift but contributes to pitching moment von Karman give is pointed at both ends Lift of body pointed at both ends is not zero None of these Slender body theory in supersonic flow Cd is dependent of Mach number for pointed nose of a slender body Cd is independent of Mach number for pointed nose and trailing end Drag is zero for pointed nose and trailing end Drag is zero for pointed nose and drag is normal on trailing edge Supersonic Delta wing Forward shock moves inboard and intersects rear shock causes flow seperation Delta wing with supersonic leading edge, reduce wave drag Delta wing with subsonic leading edge, large wave drag None of these Supersonic Swept wings Sweep delays drag rise and reduces peak drag Sweep reduces CL Both are correct None of these In a compressible Laminar Boundary Layer BL Pr=1 and zero pressure gradient yields h0=const. Re analogy is valid in stagnation point Pr=1 and h0=cte for adiabatic wall and adverse pressure gradient None of these Recuperation factor, r, in a laminar compressible BL Tp0, hot wall Tp>Tf implays qp<0, Cold wall r is a fraction of the U2/2Cp recovered at the wall None of these BL Laminar, for given wall heat flux qw and M, calculation sequence for Cf and Tw: Get T* from Monagan reference temperature method Get Cf from Reynolds analogy Both are correct None of these Maintaining natural laminar flow on the attachment line Increasing swept wing angle Reducing sweep leading edge radius Cooling has large impact on CF than on TS None of these The value of the Bulk Modulus of elasticity for an incompressible fluid is Zero Unity Very low Infinity The compressible fluid flow outside the boundary layer is isentropic because: thermal energy viscosity and diffusion effects are zero fluid elements are in an adiabatic and reversible process Both are correct None of these A flow of Inviscid fluid: It is irrotational It may be rotational and at constant entropy is a potential flow Steady and is not isentropic is rotationa None of these For a supersonic flow the Mach nº increases if : Flow is isentropic and area increases Flow is adiabatic with friction in a pipe (A=constant) It enters a convergent duct Heating the Inviscid pipe flow An Inviscid 1D flow with heat transfer: Mach increases only if M2 ⦤γ with heating Mach increases with heating Mach increases with heating only in subsonic Mach increases with cooling For a normal shock wave: Stagnation pressure decreases Mach is always supersonic after the shock Stagnation temperature decreases Infinitesimal shock corresponds to Mach wave An Inviscid 1D flow with heat transfer: Mach increases only if M2 ⦤γ with heating Mach increases with heating only in subsonic Mach increases with heating Mach increases with cooling In a shock tube: p and Mach are continuous across contact discontinuity p and U are continuous across contact discontinuity Mach of the flow behind shock can not be M>1 entropy s is constant through expansion and contact discontinuity The compressible steady 2D potential equations : Sound speed is constant Is always hyperbolic Satisfies only mass and energy Characteristics depend on deflection angle and Mach angle The Prandtl-Glauert rule Relates CM of the same airfoil and different angle of attack Cp is the same at points y’ and, y=βy' Same airfoil may have the same Cp at two angles of attack Relates Cp from M=0 to M=1 Linearized potential flow Airfoil affline related may have the same Cp at M=0 and M>0 Prandtl-G relates the drag coefficients for M>Mcritic The information of angle of attack is in the profile shape The CL from Gothert and P-G rules are always equal for M>0 Supercritical aerofoils : The shock waves are developed further aft Have highly cambered aft section Both are correct None of these You need to choose the climb speed that gets you the most altitude in the shortest horizontal distance. Which speed are you flying? Vx Vy None of these The value of prandtl number from water range from 1 to 4 20 to 45 5 to 10 100 to 500 Where are self-aligning pitot tubes used? Bomber planes Commercial airliners High angle of attack fighter aircraft General aviation What is the difference between the actual free stream pressure and the measured static pressure called? Static error Static defect Free stream error Dynamic pressure The mechanism of shock wave heating can be understood from an analysis of the discontinuities in pressure and other physical quantities that happen at the_______ Oblique wave Wing tip Wing root Shock The oblique shock strength is______ High Moderate Low None of these Two pipes of diameters d1 and d2 converge to form a pipe of diameter 2d. If the liquid flows with a velocity of v1 and v2 in the two pipes, what will be the flow velocity in the third pipe? v1 + v2 v1 + v2/2 v1 + v2/4 2(v1 + v2) Speed of sound in air is______ 343ms-1 343ms-2 341ms-1 250ms-1 The different probes that disturb the airflow is called ______ Air data prope Data prope Instrusive prope Static prope Below a diagram that gives a vectorical visual representation of the movement of a body or a fluid is called as_________. Line graph Pie chart Bar graph Hodograph The Prandtl Number approximates ___________ Thermal diffusivity to momentum diffusivity Shear stress to thermal diffusivity Thermal diffusivity to kinematic viscosity Momentum diffusivity to thermal diffusivity Which one of the following is the correct relation between compressibility β and Bulk Modulus k β = k β = 1/k β = 2k β = k/2 While the upstream and downstream flow directions areunchanged across a normal shock, they are different for flow across an Normal shock wave Supersonic flow Oblique shock wave Subsonic flow Rayleigh line represents the locus of states with same impulse pressure and mass velocity. True False The adiabatic flow through a constant area duct where the effect of friction is considered as_______. Fanno flow Supersonic flow Subsonic flow None of these The depth of a trapezoidal channel section is 2m, base width of 3m and has a side slope of 1H:2V. Calculate n if the bed slope is 1 in 1000. 0.012 0.013 0.014 0.015 Calculate the velocity of flow through a channel having depth of 1.2m and specific energy equal to 1.24m. 0.6 m/s 0.7 m/s 0.8m/s 0.9m/s The specific energy of a channel section is 1.01m and the velocity of flow is 0.5m⁄s, calculate the depth of flow 0.8m 1.0m 1.2m 1.4m Supersonic wings After the aircraft goes supersonic the Aerodynamic Center (AC) moves from where to where in reference to the Mean Aerodynamic Chord (MAC)? It shifts forward from 50% MAC to 25% MAC It shifts aft from 25% MAC to 30% MAC It shifts aft from 25% MAC to 50% MAC The AC never shifts by definition Momentum is the measurement of _____ in motion? Energy Work Mass Time Finish the formula: Momentum = Mass * _________ Velocity Acceleration Power Work Which of the following is a unit of measurement for momentum? N J kg m/s N m/s What do we call it when two or more moving objects exert forces on each other for a short period of time? Momentum Collision Friction Energy The continuity equation is based on the principle of Conservation of mass Conservation of momentum Conservation of energy Conservation force Which of the following is true about adiabatic flow in constant area duct with friction? It has constant G It has constant ho It follows Fanno line Above all Bow occurs at what Mach number? Mach number = 1 Mach number greater than 1 Mach number is negative Mach number = 0 A boundary over which physical conditions undergo changes is called _______ Shock front Shear front Contact front Cap front Rayleigh line represents the locus of states with same impulse pressure and mass velocity True False Speed of sound in air depends on the Chemical condition Pitch Physical condition Area If a liquid enters a pipe of diameter d with a velocity v, what will it’s velocity at the exit if the diameter reduces to 0.5d? v 0.5v 2v 4v For M=0. Cp is known over the wing profile. Use Prandtl-Glauert To know Cp for M=0 With the same t, apply PG rule With the same T, and different AR apply PG rule With different AR, and apply PH rule None of these Two pipes, each of diameter d, converge to form a pipe of diameter D. What should be the relaion between d and D such that flow velocity in the third pipe becomes double of that in each of the two pipes? D = d D = 2d D = 2d D = 4d A shock wave carries_______ Pressure Heat Energy Temperature Non-Adiabatic flow through a constant area duct where the effect of heat addition or rejection is considered as________ Rayleigh flow Fanno flow Non-compression flow Adiabatic-compression flow In a two dimensional flow, the component of the velocity along the X-axis and the Y-axis are u = ax + by and v = ax – by. For what condition will the flow field be continuous? Impossible Possible if a = b Possible if a = 2b Possible for all values of a and b In a two dimensional flow, the component of velocity along the X-axis and the Y-axis are u = ax^2 + bxy and v = bxy +ay^2. The condition for the flow field to be continuous is independent of the constants (a; b) but dependent on the variables (x; y) independent of the variables (x; y) but dependent on the constants (a; b) independent of both the constants (a; b) and the variables (x; y) dependent on both the constants (a; b) and the variables (x; y)