Materials and process in manufacturing 9th edition E. Paul DeGarmo, J T. Black, Ronald A. Kohser - Solutions

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Consider an infinitely thin flat plate at an angle of attack a in a Mach 2.6 flow. Calculate the lift and wave-drag coefficients for(a) α = 5°(b) α = 15°(c) α = 30°
Consider a diamond-wedge airfoil such as shown in Fig. 9.27, with a half-angle ε = 10°. The airfoil is at an angle of attack α = 15° to a Mach 3 free stream. Calculate the lift and wave-drag coefficients for the airfoil.
Consider sonic flow. Calculate the maximum deflection angle through which this flow can be expanded via a centered expansion wave.
Consider a circular cylinder (oriented with its axis perpendicular to the flow) and a symmetric diamond-wedge airfoil with a half-angle of 5° at zero angle of attack; both bodies are in the same Mach 5 free stream. The thickness of the airfoil and the diameter of the cylinder are the same. The
The reservoir pressure and temperature for a convergent-divergent nozzle are 5atm and 520°R, respectively. The flow is expanded isentropically to supersonic speed at the nozzle exit. If the exit-to-throat area ratio is 2.193, calculate the following properties at the exit: Me, pe, Te, pe, ue, p0,
A flow is isentropically expanded to supersonic speeds in a convergent-divergent nozzle. The reservoir and exit pressures are 1 and 0.3143atm, respectively. What is the value of Ae/A*?
A Pitot tube inserted at the exit of a supersonic nozzle reads 8.92 x 104 N/m2. If the reservoir pressure is 2.02 x 105 N/m2, calculate the area ratio Ae/A* of the nozzle.
For the nozzle flow given in Prob. 10.1, the throat area is 4 in2. Calculate the mass flow through the nozzle.
A closed-form expression for the mass flow through a choked nozzle is Derive this expression.
Repeat Prob. 10.4, using the formula derived in Prob. 10.5 and check your answer from Prob. 10.4.
A convergent-divergent nozzle with an exit-to-throat area ratio of 1.616 has exit and reservoir pressures equal to 0.947 and 1.0atm, respectively. Assuming isentropic flow through the nozzle, calculate the Mach number and pressure at the throat.
For the flow in Prob. 10.7, calculate the mass flow through the nozzle, assuming that the reservoir temperature is 288 K and the throat area is 0.3 m2.
Consider a convergent-divergent nozzle with an exit-to-throat area ratio of 1.53. The reservoir pressure is 1atm. Assuming isentropic flow, except for the possibility of a normal shock wave inside the nozzle, calculate the exit Mach number when the exit pressure pe is(a) 0.94atm(b) 0.886atm(c)
A 20° half-angle wedge is mounted at 0° angle of attack in the test section of a supersonic wind tunnel. When the tunnel is operating, the wave angle from the wedge leading edge is measured to be 41.8°. What is the exit-to-throat area ratio of the tunnel nozzle?
The nozzle of a supersonic wind tunnel has an exit-to-throat area ratio of 6.79. When the tunnel is running, a Pitot tube mounted in the test section measures 1.448 atm. What is the reservoir pressure for the tunnel?
We wish to design a supersonic wind tunnel which produces a Mach 2.8 flow at standard sea level conditions in the test section and has a mass flow of air equal to 1 slug/s. Calculate the necessary reservoir pressure and temperature, the nozzle throat and exit areas, and the diffuser throat area.
Consider a rocket engine burning hydrogen and oxygen. The total mass flow of the propellant plus oxidizer into the combustion chamber is 287.2 kg/s. The combustion chamber temperature is 3600 K. Assume that the combustion chamber is a low-velocity reservoir for the rocket engine. If the area of the
For supersonic and hypersonic wind tunnels a diffuser efficiency, t\d, can be defined as the ratio of the total pressures at the diffuser exit and nozzle reservoir, divided by the total pressure ratio across a normal shock at the test-section Mach number. This is a measure of the efficiency of the
Consider a subsonic compressible flow in Cartesian coordinates where the velocity potential is given by If the free stream properties are given by V? = 700 ft/s, p? = 1atm, and T? = 519°R, calculate the following properties at the location (x, y) = (0.2 ft, 0.2 ft): M, p, and T.
Using the Prandtl-Glauert rule, calculate the lift coefficient for an NACA 2412 airfoil at 5° angle of attack in a Mach 0.6 free stream. (Refer to Fig. 4.5 for the original airfoil data.)
Under low-speed incompressible flow conditions, the pressure coefficient at a given point on an airfoil is -0.54. Calculate Cp at this point when the free stream Mach number is 0.58, using(a) The Prandtl-Glauert rule(b) The Karman-Tsien rule(c) Laitone's rule
In low-speed incompressible flow, the peak pressure coefficient (at the minimum pressure point) on an NACA 0012 airfoil is —0.41. Estimate the critical Mach number for this airfoil, using the Prandtl-Glauert rule.
For a given airfoil, the critical Mach number is 0.8. Calculate the value of p/p∞ at the minimum pressure point when M∞ = 0.8.
Consider an airfoil in a Mach 0.5 free stream. At a given point on the airfoil, the local Mach number is 0.86. Using the compressible flow tables at the back of this book, calculate the pressure coefficient at that point. Check your answer using the appropriate analytical equation from this
Figure 11.5 shows four cases for the flow over the same airfoil wherein M∞ is progressively increased from 0.3 to MCГ = 0.61. Have you wondered where the numbers on Fig. 11.5 came from? Here is your chance to find out. Point A on the airfoil is the point of minimum pressure (hence maximum M) on
Consider the flow over a circular cylinder; the incompressible flow over such a cylinder is discussed in Sec. 3.13. Consider also the flow over a sphere; the incompressible flow over a sphere is described in Sec. 6.4. The subsonic compressible flow over both the cylinder and the sphere is
Using the results of linearized theory calculate the lift and wave-drag coefficients for an infinitely thin flat plate in a Mach 2.6 free stream at angles of attack of(a) a = 5° (b) a = 15° (c) a = 30°Compare these approximate results with those from the exact shock-expansion theory
For the conditions of Prob. 12.1, calculate the pressures (in the form of p/p∞) on the top and bottom surfaces of the flat plate, using linearized theory. Compare these approximate results with those obtained from exact shock-expansion theory in Prob. 9.13. Make some appropriate conclusions
Consider a diamond-wedge airfoil such as shown in Figure 9.24, with a half-angle ε = 10°. The airfoil is at an angle of attack α = 15° to a Mach 3 free stream. Using linear theory, calculate the lift and wave-drag coefficients for the airfoil. Compare these approximate results with those from
Consider two points in a supersonic flow. These points are lo9cated in a Cartesian coordinate system at (x1, y1) = (0, 0.0684) and (x2, y2) = (0, 0121, 0), where the units are meters. At point (x1, y1): y1 = 639 m/s, v1 = 232.6m/s, p1 = 1atm, T1 = 288K. At point (x2, y2) : u2 = 680 m/s, v2 = 0, p2
Repeat Prob. 9.14 using(a) Newtonian theory(b) Modified Newtonian theoryCompare these results with those obtained from exact shock-expansion theory (Prob. 9.13). From this comparison, what comments can you make about the accuracy of Newtonian and modified Newtonian theories at low supersonic Mach
Consider a flat plat at α = 20o in Mach 20 free stream. Using straight Newtonian theory, calculate the lift and wave-draw coefficients. Compare these results with exact shock-expansion theory.
Consider the incompressible viscous flow of air between two infinitely long parallel plates separated by a distance h. The bottom plate is stationary, and the top plate is moving at the constant velocity ue in the direction of the plate. Assume that no pressure gradient exists in the flow
Assume that the two parallel plates in Prob. 15.1 are both stationary but that a constant pressure gradient exists in the flow direction; i.e., dp/dx = constant.(a) Obtain an expression for the variation of velocity between the plates.(b) Obtain an expression for the shear stress on the plates in
The wing on a Piper Cherokee general aviation aircraft is rectangular, with a span of 9.75m and a chord of 1.6m. The aircraft is flying at cruising speed (141 mi/h) at sea level. Assume that the skin friction drag on the flat plate of the same dimensions. Calculate the skin friction drag:(a) If the
For the case in Prob. 17.1, calculate the boundary-layer thickness at the trailing edge for(a) Completely laminar flow(b) Completely turbulent flow
For the case in Prob. 17.1, calculate the skin friction drag accounting for transition. Assume the transition Reynolds number = 5 x 105.
Consider Mach 4 flow at standard sea level conditions over a flat plate of chord 5 in. Assuming all laminar flow and adiabatic wall conditions, calculate the skin friction drag on the plate per unit span.
Repeat Prob. 17.4 for the case of all turbulent flow.
Consider a compressible, laminar boundary layer over a flat plate. Assuming Pr = 1 and a calorically perfect gas, show that the profile of total temperature through the boundary layer is a function of the velocity profile via where Tw = wall temperature and T0,e and ue are the total temperature and
What is the basic principle of Aircraft landing gear system and design? Briefly explain its objectives, equipments and working.

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Materials and process in manufacturing 9th edition E. Paul DeGarmo, J T. Black, Ronald A. Kohser - Solutions

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