You are expected to develop the code in Fortran, C++, or Matlab, Python, or any other...

 

1. 
   

Transcribed Image Text:

You are expected to develop the code in Fortran, C++, or Matlab, Python, or any other engineering programming language to calculate and draw the nozzle contour. General Specifications: The rocket is to discharge into an atmosphere where the ambient pressure is 101 kPa. Assume the propellant properties are equivalent to air, y= 1.40 and R = 287 J/(kg. K) The stagnation temperature TO of the propellant is 1000 K, and the stagnation pressure p0 of the propellant is 1476.62 kPa The velocity of the fluid as it crosses the exit plane must be uniform and parallel to the axis of the nozzle The overall expansion ratio Aexit/Athroat must be large enough so that the specific kinetic energy of the exhaust gas, 0.5 Ve2 is as high as possible percentage of specific stagnation enthalpy, h0, of the propellant. What is the best percentage of this kinetic energy of exhaust with respect to the specific stagnation enthalpy, h0, of the propellant? The geometry downstream of throat should not exhibit sharp corners of large turning angles (this is required to avoid boundary layer separation due to sharp corners) The length of the nozzle should be as short as possible subject to above constraints You must show explicit calculations of the exit speed, Ve, and the exit Mach number, Me. 1 point (b) You must show a sample calculation of the characteristic lines that you use to create the nozzle contour 1 point. Use suitable number of characteristics (25 - 50) to produce a smooth contour. (c) A table of the nozzle contour coordinates. The x, y coordinates must be expressed in SI units, as measured from the axis of the nozzle as a function of axial distance, x. from the throat section. Also, plot the contour as A/Athroat vs. x/L, where A is the local nozzle area, Athroat is the throat area, and L is the length of the nozzle. 2 points (d) Provide smooth graphics of your final nozzle shape with few characteristic lines. No need to show all the characteristic lines. 1 point Program listing must be presented. 2 points Provide plots and discussions of nozzle flow property variations, such as Mach number, temperature, and pressure through the nozzle. Compare and contrast the property variations for MLN and Aerospike. You are expected to develop the code in Fortran, C++, or Matlab, Python, or any other engineering programming language to calculate and draw the nozzle contour. General Specifications: The rocket is to discharge into an atmosphere where the ambient pressure is 101 kPa. Assume the propellant properties are equivalent to air, y=1.40 and R = 287 J/(kg. K) The stagnation temperature TO of the propellant is 1000 K, and the stagnation pressure p0 of the propellant is 1476.62 kPa The velocity of the fluid as it crosses the exit plane must be uniform and parallel to the axis of the nozzle The overall expansion ratio Aexit/Athroat must be large enough so that the specific kinetic energy of the exhaust gas, 0.5 Ve2 is as high as possible percentage of specific stagnation enthalpy, h0, of the propellant. What is the best percentage of this kinetic energy of exhaust with respect to the specific stagnation enthalpy, h0, of the propellant? The geometry downstream of throat should not exhibit sharp corners of large turning angles (this is required to avoid boundary layer separation due to sharp corners) The length of the nozzle should be as short as possible subject to above constraints You must show explicit calculations of the exit speed, Ve, and the exit Mach number, Me. 1 point (b) You must show a sample calculation of the characteristic lines that you use to create the nozzle contour 1 point. Use suitable number of characteristics (25 - 50) to produce a smooth contour. (c) A table of the nozzle contour coordinates. The x, y coordinates must be expressed in SI units, as measured from the axis of the nozzle as a function of axial distance, x. from the throat section. Also, plot the contour as A/Athroat vs. x/L, where A is the local nozzle area, Athroat is the throat area, and L is the length of the nozzle. 2 points (d) Provide smooth graphics of your final nozzle shape with few characteristic lines. No need to show all the characteristic lines. 1 point Program listing must be presented. 2 points Provide plots and discussions of nozzle flow property variations, such as Mach number, temperature, and pressure through the nozzle. Compare and contrast the property variations for MLN and Aerospike. You are expected to develop the code in Fortran, C++, or Matlab, Python, or any other engineering programming language to calculate and draw the nozzle contour. General Specifications: The rocket is to discharge into an atmosphere where the ambient pressure is 101 kPa. Assume the propellant properties are equivalent to air, y= 1.40 and R = 287 J/(kg. K) The stagnation temperature TO of the propellant is 1000 K, and the stagnation pressure p0 of the propellant is 1476.62 kPa The velocity of the fluid as it crosses the exit plane must be uniform and parallel to the axis of the nozzle The overall expansion ratio Aexit/Athroat must be large enough so that the specific kinetic energy of the exhaust gas, 0.5 Ve2 is as high as possible percentage of specific stagnation enthalpy, h0, of the propellant. What is the best percentage of this kinetic energy of exhaust with respect to the specific stagnation enthalpy, h0, of the propellant? The geometry downstream of throat should not exhibit sharp corners of large turning angles (this is required to avoid boundary layer separation due to sharp corners) The length of the nozzle should be as short as possible subject to above constraints You must show explicit calculations of the exit speed, Ve, and the exit Mach number, Me. 1 point (b) You must show a sample calculation of the characteristic lines that you use to create the nozzle contour 1 point. Use suitable number of characteristics (25 - 50) to produce a smooth contour. (c) A table of the nozzle contour coordinates. The x, y coordinates must be expressed in SI units, as measured from the axis of the nozzle as a function of axial distance, x. from the throat section. Also, plot the contour as A/Athroat vs. x/L, where A is the local nozzle area, Athroat is the throat area, and L is the length of the nozzle. 2 points (d) Provide smooth graphics of your final nozzle shape with few characteristic lines. No need to show all the characteristic lines. 1 point Program listing must be presented. 2 points Provide plots and discussions of nozzle flow property variations, such as Mach number, temperature, and pressure through the nozzle. Compare and contrast the property variations for MLN and Aerospike. You are expected to develop the code in Fortran, C++, or Matlab, Python, or any other engineering programming language to calculate and draw the nozzle contour. General Specifications: The rocket is to discharge into an atmosphere where the ambient pressure is 101 kPa. Assume the propellant properties are equivalent to air, y=1.40 and R = 287 J/(kg. K) The stagnation temperature TO of the propellant is 1000 K, and the stagnation pressure p0 of the propellant is 1476.62 kPa The velocity of the fluid as it crosses the exit plane must be uniform and parallel to the axis of the nozzle The overall expansion ratio Aexit/Athroat must be large enough so that the specific kinetic energy of the exhaust gas, 0.5 Ve2 is as high as possible percentage of specific stagnation enthalpy, h0, of the propellant. What is the best percentage of this kinetic energy of exhaust with respect to the specific stagnation enthalpy, h0, of the propellant? The geometry downstream of throat should not exhibit sharp corners of large turning angles (this is required to avoid boundary layer separation due to sharp corners) The length of the nozzle should be as short as possible subject to above constraints You must show explicit calculations of the exit speed, Ve, and the exit Mach number, Me. 1 point (b) You must show a sample calculation of the characteristic lines that you use to create the nozzle contour 1 point. Use suitable number of characteristics (25 - 50) to produce a smooth contour. (c) A table of the nozzle contour coordinates. The x, y coordinates must be expressed in SI units, as measured from the axis of the nozzle as a function of axial distance, x. from the throat section. Also, plot the contour as A/Athroat vs. x/L, where A is the local nozzle area, Athroat is the throat area, and L is the length of the nozzle. 2 points (d) Provide smooth graphics of your final nozzle shape with few characteristic lines. No need to show all the characteristic lines. 1 point Program listing must be presented. 2 points Provide plots and discussions of nozzle flow property variations, such as Mach number, temperature, and pressure through the nozzle. Compare and contrast the property variations for MLN and Aerospike.

Answer rating: 100% (QA)

The example will calculate the nozzle contour and plot the nozzle flow properties For a realworld ap... View the full answer