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P3. A Schlieren photograph of the flow around a corner reveals the edges of the expansion "fan" a...
1. A uniform supersonic flow at Mach 2.0, with static pressure of 75 kPa and a...
1. A uniform supersonic flow at Mach 2.0, with static pressure of 75 kPa and a temperature oi 250 К, ехрands aгound a 10 degrees pressure p2, temperature T,, and the fan angle. convex corner. Determine the downstream Mach number M,, Fan angle Мi Mа 10°
1. A uniform supersonic flow at Mach 2.0, with static pressure of 75 kPa and a temperature oi 250 К, ехрands aгound a 10 degrees pressure p2, temperature T,, and the fan angle. convex...
Question 2.8 Refering to the figure below, a supersonic flow with upstream Mach number, M, static...
Question 2.8 Refering to the figure below, a supersonic flow with upstream Mach number, M, static pressure, pi, and static temperature, Ti, as specified in the table below, encounters a corner with a turning angle ore Determine the angle of the oblique shock, ?, the angle of the reflected wave, q, the Mach numbers M2 and M, and the downstream static pressure Ps and static temperature Ty Mi P1 M3 P3 T3 Design Data Value Unit Mach number (M) Static...
M 3. Consider a scenario in which supersonic flow is expanded and turned by 150 through...
M 3. Consider a scenario in which supersonic flow is expanded and turned by 150 through a Prandtl-Meyer expansion wave/fan. Consider the gas to be calorically perfect Air with upstream properties as follows: M1 = 4, P = 20 kPa, T = 250 K. Find: 0-150 M2 (a) freestream Prandtl-Meyer function, Vi. (b) downstream Mach number, M2. (c) downstream static pressure, P2. (d) downstream static temperature, T2.
scouS in a pipe are expre in terms of a Jriction Jactor, J. It is important to frictional forces in pipe flow because t...
scouS in a pipe are expre in terms of a Jriction Jactor, J. It is important to frictional forces in pipe flow because they result in pressure loss which must be accounted for in the piping system. For turbulent flow in pipes, the friction factor is calculated using the Colebrook equation: e/D 2.51 VT where e is the roughness height, D is the pipe diameter and Re is the pipe Reynolds number (dimensionless). For e/D = 0.008 and Re =...
4. A supersonic engine inlet is shown below-with a spike centerbody. Suppose the flight Mach number...
4. A supersonic engine inlet is shown below-with a spike centerbody. Suppose the flight Mach number M1 = 2.5, and the pressure is pı = 50,000 N/m². The half-angle of the spike centerbody is 10°, as shown. For a particular mass flow through the engine, it happens that there is an oblique shock at a, and a normal shock wave standing at b. Before entering the second shock wave, the fluid expands through a Prandtl-Meyer turn, as the skech indicates....
Question 2.10 A two-dimensional wedge-shaped airfoil, with chord, c, consists of straight-line segments with wedge angl...
Question 2.10 A two-dimensional wedge-shaped airfoil, with chord, c, consists of straight-line segments with wedge angles, θ¡et and θaf, at the leading and trailing edges, as defined in the figure and given in the table below. The airfoil operates at an angle of attack, α, and it is moving through air at a supersonic speed, M. The atmospheric temperature and pressure far upstream of the airfoil are T and p as specified in the table The various flow regions are...
GAS DYNAMICS 3 M = 2.2 M, = 22 124 12 (a) (b) Figure 4.55 (a)...
GAS DYNAMICS
3 M = 2.2 M, = 22 124 12 (a) (b) Figure 4.55 (a) One-shock spike, (b) two-shock spike diffuser. Figure 4.56 Supersonic flow past a sharp corner. M OM P12 м. 10° Figure 4.57 Flow through incident and reflected oblique shocks. M = 2.4 P1 = 101 kPa 10° Oblique Shock and Expansion Waves 219 Figure 4.58 Oblique shock reflection from a solid wall. M,=2.2 Po= 100 kPa M2 M 30° Air X Jet boundary Figure 4.59...
An axial flow compressor stage is downstream of an IGV that turns the flow 15° in...
An axial flow compressor stage is downstream of an IGV that turns the flow 15° in the direction of the rotor rotation, as shown in figure 1. The axial velocity component remains constant throughout the stage at Cz = 150 M/S. The rotor rotational speed w = 3000 rpm and the pitchline radius is rm=0.5 m. The rotor relative exit flow angle is B2=-15°. The static temperature and pressure of air upstream of the rotor are T1 = 20°C and...
Determine the solution in the expansion fan. In this problem, we consider the flow of water through a porous medium. Specifically, u, is the water content at depth z and timet measured as a ratio...
Determine the solution in the expansion fan.
In this problem, we consider the flow of water through a porous medium. Specifically, u, is the water content at depth z and timet measured as a ratio of water-filled pore space to total pore-space. (Thus, 0 u 1.) We assume that u satisfies the conservation law and initial condition (12), + ut = 0 t>0 x<0 a(x, 0) = 0 u(r,0)-0 >
In this problem, we consider the flow of water through...
8) (1 point) Honeycomb Flow Straightener -Anti-Turbulence Screens Silencer Diffuser Fan Test Section Inlet The wind...
8) (1 point) Honeycomb Flow Straightener -Anti-Turbulence Screens Silencer Diffuser Fan Test Section Inlet The wind tunnel shown above is; a) open circuit/suction/low speed wind tunnel b) open circuit/blow-down low speed wind tunnel c) closed circuit/suction/supersonic wind tunnel d) closed circuit/blow-down/supersonic wind tunnel 9) For the air wind tunnel shown above in Q8: (6 points) Assuming a circular cross section, the atmospheric pressure is 101.325 kPa, density of air is 1.2 kg/m², and absolute viscosity of air is 1.825*10-Ns/m² (assume...
1. A uniform supersonic flow at Mach 2.0, with static pressure of 75 kPa and a temperature oi 250 К, ехрands aгound a 10 degrees pressure p2, temperature T,, and the fan angle. convex corner. Determine the downstream Mach number M,, Fan angle Мi Mа 10°
1. A uniform supersonic flow at Mach 2.0, with static pressure of 75 kPa and a temperature oi 250 К, ехрands aгound a 10 degrees pressure p2, temperature T,, and the fan angle. convex...
Question 2.8 Refering to the figure below, a supersonic flow with upstream Mach number, M, static pressure, pi, and static temperature, Ti, as specified in the table below, encounters a corner with a turning angle ore Determine the angle of the oblique shock, ?, the angle of the reflected wave, q, the Mach numbers M2 and M, and the downstream static pressure Ps and static temperature Ty Mi P1 M3 P3 T3 Design Data Value Unit Mach number (M) Static...
M 3. Consider a scenario in which supersonic flow is expanded and turned by 150 through a Prandtl-Meyer expansion wave/fan. Consider the gas to be calorically perfect Air with upstream properties as follows: M1 = 4, P = 20 kPa, T = 250 K. Find: 0-150 M2 (a) freestream Prandtl-Meyer function, Vi. (b) downstream Mach number, M2. (c) downstream static pressure, P2. (d) downstream static temperature, T2.
scouS in a pipe are expre in terms of a Jriction Jactor, J. It is important to frictional forces in pipe flow because they result in pressure loss which must be accounted for in the piping system. For turbulent flow in pipes, the friction factor is calculated using the Colebrook equation: e/D 2.51 VT where e is the roughness height, D is the pipe diameter and Re is the pipe Reynolds number (dimensionless). For e/D = 0.008 and Re =...
4. A supersonic engine inlet is shown below-with a spike centerbody. Suppose the flight Mach number M1 = 2.5, and the pressure is pı = 50,000 N/m². The half-angle of the spike centerbody is 10°, as shown. For a particular mass flow through the engine, it happens that there is an oblique shock at a, and a normal shock wave standing at b. Before entering the second shock wave, the fluid expands through a Prandtl-Meyer turn, as the skech indicates....
Question 2.10 A two-dimensional wedge-shaped airfoil, with chord, c, consists of straight-line segments with wedge angles, θ¡et and θaf, at the leading and trailing edges, as defined in the figure and given in the table below. The airfoil operates at an angle of attack, α, and it is moving through air at a supersonic speed, M. The atmospheric temperature and pressure far upstream of the airfoil are T and p as specified in the table The various flow regions are...
GAS DYNAMICS
3 M = 2.2 M, = 22 124 12 (a) (b) Figure 4.55 (a) One-shock spike, (b) two-shock spike diffuser. Figure 4.56 Supersonic flow past a sharp corner. M OM P12 м. 10° Figure 4.57 Flow through incident and reflected oblique shocks. M = 2.4 P1 = 101 kPa 10° Oblique Shock and Expansion Waves 219 Figure 4.58 Oblique shock reflection from a solid wall. M,=2.2 Po= 100 kPa M2 M 30° Air X Jet boundary Figure 4.59...
An axial flow compressor stage is downstream of an IGV that turns the flow 15° in the direction of the rotor rotation, as shown in figure 1. The axial velocity component remains constant throughout the stage at Cz = 150 M/S. The rotor rotational speed w = 3000 rpm and the pitchline radius is rm=0.5 m. The rotor relative exit flow angle is B2=-15°. The static temperature and pressure of air upstream of the rotor are T1 = 20°C and...
Determine the solution in the expansion fan.
In this problem, we consider the flow of water through a porous medium. Specifically, u, is the water content at depth z and timet measured as a ratio of water-filled pore space to total pore-space. (Thus, 0 u 1.) We assume that u satisfies the conservation law and initial condition (12), + ut = 0 t>0 x<0 a(x, 0) = 0 u(r,0)-0 >
In this problem, we consider the flow of water through...
8) (1 point) Honeycomb Flow Straightener -Anti-Turbulence Screens Silencer Diffuser Fan Test Section Inlet The wind tunnel shown above is; a) open circuit/suction/low speed wind tunnel b) open circuit/blow-down low speed wind tunnel c) closed circuit/suction/supersonic wind tunnel d) closed circuit/blow-down/supersonic wind tunnel 9) For the air wind tunnel shown above in Q8: (6 points) Assuming a circular cross section, the atmospheric pressure is 101.325 kPa, density of air is 1.2 kg/m², and absolute viscosity of air is 1.825*10-Ns/m² (assume...
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