The reservoir pressure and temperature for a convergent-divergent nozzle are 5 atm and
The Mach Number
The exit pressure
The exit Temperature
The exit density
The velocity of flow at exit
The stagnation pressure at exit
The stagnation temperature at exit
Answer to Problem 10.1P
The Mach Number
The exit pressure
The exit Temperature
The exit density
The velocity of flow at exit
The stagnation pressure at exit
The stagnation temperature at exit
Explanation of Solution
Given:
The Reservoir pressure is
The Reservoir temperature is
The ratio of exit area to throat area is
Formula used:
The expression for calculating pressure is given as,
The expression for calculating temperature is given as,
The expression for density is given as,
Here,
The expression for velocity of sound is given as,
Here
The expression for speed of velocity is given as,
Calculation:
Refer to the “isentropic flow properties” for the Mach number at the ratio of exit area to throat area. The Mach number is obtained as,
Refer to the “isentropic flow properties” for the pressure ratio at the ratio of exit area to throat area. The pressure ratio is obtained as,
Refer to the “isentropic flow properties” for the temperature ratio at the ratio of exit area to throat area. The temperature ratio is obtained as,
The pressure at the exit is calculated as,
The stagnation pressure at the exit is calculated as,
The Temperature at the exit is calculated as,
The stagnation temperature at the exit is calculated as,
The Density at the exit is calculated as,
The value of gas constant is in English units is
The velocity at exit is calculated as,
The value of adiabatic constant is
Conclusion:
Therefore, The Mach Number
Therefore, The exit pressure
Therefore, The exit Temperature
Therefore, The exit density
Therefore, The velocity of flow at exit
Therefore, The stagnation pressure at exit
Therefore, The stagnation temperature at exit
Want to see more full solutions like this?
Chapter 10 Solutions
EBK FUNDAMENTALS OF AERODYNAMICS
Additional Engineering Textbook Solutions
Introduction To Finite Element Analysis And Design
Engineering Mechanics: Dynamics (14th Edition)
INTERNATIONAL EDITION---Engineering Mechanics: Statics, 14th edition (SI unit)
Vector Mechanics for Engineers: Statics and Dynamics
Thinking Like an Engineer: An Active Learning Approach (3rd Edition)
Statics and Mechanics of Materials (5th Edition)
- Consider the isentropic flow through a supersonic nozzle. If thetest-section conditions are given by p = 1 atm, T = 230 K, and M = 2,calculate the reservoir pressure and temperature.arrow_forwardThe mass flow of air through a supersonic nozzle is 1.5 lbm/s. The exit velocity is 1,500 ft/sec, and the reservoir temperature and pressure are 1,000 'R and 7 atm , respectively. Calculate the area of the nozzle exitarrow_forwardA Pitot tube inserted at the exit of a supersonic nozzle reads8.92 × 104 N/m2. If the reservoir pressure is 2.02 × 105 N/m2, calculatethe area ratio Ae/A∗ of the nozzle.arrow_forward
- Air enters a converging–diverging nozzle of a supersonic wind tunnel at 150 psia and 100°F with a low velocity. The flow area of the test section is equal to the exit area of the nozzle, which is 5 ft2. Calculate the pressure, temperature, velocity, and mass flow rate in the test section for a Mach number Ma = 2. Explain why the air must be very dry for this application.arrow_forwardConsider a normal shock wave in a supersonic airstream where the pressure upstream of the shock is 1 atm. Calculate the loss of total pressure across the shock wave when the upstream Mach number is (a) M1 = 2.5, and (b) M1 = 4.5. Compare these two results and comment on their implicationarrow_forwardAn ideal isentropic nozzle is attached to an infinite reservoir that has stagnation conditions 3 MPa and 2250 K, and a constant specific heat of 1.2. If the nozzle's static exit pressure is 38.871 kPa, what is the exit static temperature? Also determine the nozzle's exit Mach number, stagnation pressure, and stagnation temperature.arrow_forward
- Air flows through the supersonic nozzle . The inlet conditions are 7 kPa and 420°C. The nozzle exit diameter is adjusted such that the exiting velocity is 700 m/s. Calculate ( a ) the exit temperature, ( b )the mass flux, and ( c ) the exit diameter. Assume an adiabatic quasiequilibrium flowarrow_forwardA normal shock produced by an explosion propagate at a constant velocity of 450 m/s into still air of 100 kPa and 23 °C. The ratio of stagnation pressure upstream the shock to stagnation pressure of the gas flow behind the wave isarrow_forwardA convergent-divergent nozzle with an exit-to-throat area ratio of 1.616has exit and reservoir pressures equal to 0.947 and 1.0 atm, respectively.Assuming isentropic flow through the nozzle,calculate the mass flow through the nozzle, assuming that the reservoir temperature is 288 K and the throat area is 0.3 m2.arrow_forward
- Consider an infinitely thin flat plate with a 1 m chord at an angle of attackof 10◦ in a supersonic flow. The pressure and shear stress distributions onthe upper and lower surfaces are given by pu = 4 × 104(x − 1)2 +5.4 × 104, pl = 2 × 104(x − 1)2 + 1.73 × 105, τu = 288x−0.2, andτl = 731x−0.2, respectively, where x is the distance from the leading edgein meters and p and τ are in newtons per square meter. Calculate thenormal and axial forces, the lift and drag, moments about the leadingedge, and moments about the quarter chord, all per unit span. Also,calculate the location of the center of pressure.arrow_forwardAn aircraft flies with a Mach number Ma1=0.962 at an altitude of 7062 m where the pressure is 46.2 kPa and the temperature is 246.2 K. Calculate the stagnation properties (static temperature, pressure, density), cross-section areas A1 and A2=? at the inlet and outlet of the diffuser. The diffuser at the engine inlet has an exit Mach number of Ma2=0.3. For a mass flow rate of 36,2 kg/s, determine the static pressure rise across the diffuser and the exit area. Solve the problem by making the necessary assumptions and drawing the schematic figure.arrow_forwardAir flows into a converging duct, and a normal shock stands at the exit of the duct. Downstream of the shock, the Mach number is 0.54. If p2/p1 = 2, compute the Mach number at the entrance of the duct and the area ratio A1/A2.arrow_forward
- Elements Of ElectromagneticsMechanical EngineeringISBN:9780190698614Author:Sadiku, Matthew N. O.Publisher:Oxford University PressMechanics of Materials (10th Edition)Mechanical EngineeringISBN:9780134319650Author:Russell C. HibbelerPublisher:PEARSONThermodynamics: An Engineering ApproachMechanical EngineeringISBN:9781259822674Author:Yunus A. Cengel Dr., Michael A. BolesPublisher:McGraw-Hill Education
- Control Systems EngineeringMechanical EngineeringISBN:9781118170519Author:Norman S. NisePublisher:WILEYMechanics of Materials (MindTap Course List)Mechanical EngineeringISBN:9781337093347Author:Barry J. Goodno, James M. GerePublisher:Cengage LearningEngineering Mechanics: StaticsMechanical EngineeringISBN:9781118807330Author:James L. Meriam, L. G. Kraige, J. N. BoltonPublisher:WILEY