Normal Shock in Variable Duct Areas

In a converging-diverging nozzle, if the back pressure is between the critical pressures designated as point 'a' and point 'b', what phenomenon is expected to occur within the nozzle?

What is the term for a nozzle where the back pressure is higher than the pressure at the exit for ideal isentropic expansion, but the flow has already become supersonic?

In Example 6.1, a converging-diverging nozzle has a throat area of 3 cm squared and an exit area of 9 cm squared. A shock occurs at a location where the cross-section area is 6 cm squared. What is the area ratio (Ax/A*) used to determine the Mach number just before the shock?

Based on the calculations in Example 6.1, what is the approximate Mach number (Mx) just before the normal shock when the area ratio Ax/A* is 2?

After a normal shock occurs in a nozzle, a new, imaginary star area (A*) can be conceptualized for the subsonic flow downstream. In Example 6.1, given a Mach number of about 0.54743 after the shock and an exit area of 9 cm squared, what is the new area ratio Ae/A* used to find the exit conditions?

In Example 6.1, with a stagnation pressure of 4 Bar, what is the calculated exit pressure (Pexit) after the flow passes through a normal shock and exits the nozzle?

For the scenario in Example 6.1, with a stagnation temperature of 308 K, what is the final exit temperature (Texit)?

What is the primary definition of nozzle efficiency (eta) provided in the chapter?

What is the typical range for the efficiency (eta) of a nozzle according to the text?

How is the coefficient of discharge (Cd) defined for a nozzle?

How is the efficiency of a diffuser defined in the text?

In Example 6.2, a nozzle has a back pressure of 2 Bar. The solution procedure starts at the nozzle's exit and progresses towards the entrance. What is the calculated value of the dimensionless group P_exit*A_exit / (P_x0 * A_x*)?

In the analysis for Example 6.2, after determining the exit Mach number, what is the calculated ratio of stagnation pressures across the shock (P0y/P0x)?

Once the stagnation pressure ratio P0y/P0x is found to be 0.5525 in Example 6.2, what is the corresponding upstream Mach number (Mx) of the shock?

In Example 6.2, after finding the pre-shock Mach number Mx to be 2.3709, what is the area (A) where the shock occurs, given a throat area (A*) of 3 cm squared?

For a converging-diverging nozzle, which condition describes an under-expanded nozzle?

What is the relationship between nozzle efficiency (eta) and the velocity coefficient (Vc)?

In the context of the diffuser efficiency formula, η = 2 * (h3 - h1) / U1^2, what do the subscripts 1 and 3 represent?

In Example 6.3, a wind tunnel has a required condition at the test section (point 3) of M = 3.0. What is the nozzle throat area (A*_n) required if the test section area is 0.02 m squared?

In Example 6.3, after a shock occurs in the M=3.0 test section, what is the required diffuser throat area (A*_d) to maintain choked flow in the diffuser?

What is the reason given for why an over-expanded nozzle is considered worse than an under-expanded nozzle for rocket and aviation performance?

In a converging-diverging nozzle with a fixed geometry, what primarily determines the location of a normal shock within the diverging section?

In Example 6.4, a shock is moving at 200 m/sec into a pipe with gas at 350K and k=1.3. What is the upstream Mach number (Mx) used for analysis?

In Example 6.5, a piston moves to create a shock wave, and the pressure is doubled across the shock (Py/Px = 2). What is the Mach number of the flow behind the shock (My), which corresponds to the piston's velocity?

Following Example 6.5, with an initial temperature of 300K and a temperature ratio (Ty/Tx) of 1.2308, what is the calculated velocity of the piston (Uy)?

In Example 6.6, a gas flow is brought to a sudden stop. The upstream gas velocity is 502.25 m/sec and the speed of sound is calculated from the given conditions (k=1.091, R=143, T=350K). What is the static Mach number (Mx') of the upstream flow?

When a shock occurs in a converging-diverging nozzle, what happens to the stagnation pressure of the flow as it passes through the shock?

When a shock occurs in a converging-diverging nozzle, what happens to the stagnation temperature of the flow as it passes through the shock, assuming adiabatic flow?

In the scenario of Example 6.1, what is the back pressure for the critical point 'b', where the flow is supersonic and perfectly expanded to the exit without a shock?

In the scenario of Example 6.1, what is the back pressure for the critical point 'a', where the flow is choked at the throat but remains subsonic throughout the diverging section?

According to the text, what is the mass flow rate for the nozzle in Example 6.1, given the throat area is 3 cm squared and stagnation conditions are 4 Bar and 308K?

What does the text identify as a major reason for the academic nature of locating a shock wave, making it not easily visible in practice?

In Example 6.2, what is the Mach number at the exit of the nozzle, given the dimensionless group P_exit*A_exit / (P0 * A*) is 1.5?

For the wind tunnel in Example 6.3 with a test section Mach number of 3.0, what is the static pressure ratio (P/P0) at this point?

In the same wind tunnel scenario (Example 6.3, M=3.0), what is the Mach number immediately after the normal shock (My)?

In the iterative method described in Example 6.2 for finding a shock location, why is it noted that the initial guess for the shock location (area) should be larger if the resulting exit pressure is too high?

What physical principle allows for the use of the ratio P0 * A* = constant in solving problems like Example 6.2?

In Example 6.1, the Mach number after the shock (My) is found to be 0.54743. What is the isentropic area ratio (A/A*) corresponding to this subsonic Mach number?

What is the pressure ratio (Py/Px) across the shock in Example 6.1, where the upstream Mach number is 2.1972?

The velocity coefficient, Vc, is defined as the ratio of the actual velocity to the ideal velocity. How is it related to the actual and ideal kinetic energies?

What does it mean if a nozzle's back pressure is lower than the critical value needed for continuous isentropic flow (point 'b')?

In the iterative procedure for finding a shock location in Example 6.2, what is the Mach number (Mx) corresponding to a stagnation pressure ratio (P0y/P0x) of 0.55250?

What is the static temperature at the throat (T*) for the nozzle in Example 6.1, where T0 is 308K and k=1.4?

What is the static pressure at the throat (P*) for the nozzle in Example 6.1, where P0 is 4 Bar and k=1.4?

For the moving shock problem in Example 6.6, what is the pressure ratio (Py/Px) when the upstream shock Mach number (Mx) is 2.9222?

The text mentions that for a flow with a shock, the stagnation pressure is constant as well as the stagnation temperature. Is this statement correct?

In the final step of Example 6.1, the exit temperature is calculated to be 299.9K from a stagnation temperature of 308K. Why is the exit temperature not equal to the stagnation temperature?

What is the key word in Example 6.1 that indicates the stagnation conditions are equal to the conditions in the supply tank?

In the perfect gas equation (6.6) for diffuser efficiency, η = 2 * Cp * (T3 - T1) / U1^2, what is the role of Cp?

In Example 6.1, a converging-diverging nozzle has an exit area of 9 cm squared and a throat of 3 cm squared. What is the overall area ratio (A_exit/A_throat) of the nozzle?