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+{
+ "metadata": {
+  "name": "",
+  "signature": "sha256:6ef501d87a4e150fa73f9d6a23345ef6ffddaecf1a006d99b9d65903d53edab9"
+ },
+ "nbformat": 3,
+ "nbformat_minor": 0,
+ "worksheets": [
+  {
+   "cells": [
+    {
+     "cell_type": "heading",
+     "level": 1,
+     "metadata": {},
+     "source": [
+      "Chapter 6 : Aircraft Propulsion"
+     ]
+    },
+    {
+     "cell_type": "heading",
+     "level": 2,
+     "metadata": {},
+     "source": [
+      "Example 6.1 page : 23"
+     ]
+    },
+    {
+     "cell_type": "code",
+     "collapsed": false,
+     "input": [
+      "import math \n",
+      "\n",
+      "\t\t\t\t\n",
+      "#Input data\n",
+      "eff_com = 0.8 \t\t\t\t#Compressor efficiency\n",
+      "eff_t = 0.85 \t\t\t\t#Turbine efficiency\n",
+      "pr = 4. \t\t\t\t#Pressure ratio including combustion pressure loss(Po2s/Po1)\n",
+      "eff_c = 0.98 \t\t\t\t#Combustion efficiency\n",
+      "eff_m = 0.99 \t\t\t\t#Mechanical transmission efficiency \n",
+      "eff_n = 0.9 \t\t\t\t#Nozzle efficiency \n",
+      "Tmax = 1000. \t\t\t\t#Maximum cycle temperature in K\n",
+      "To3 = Tmax \t\t\t\t#Stagnation temperature before turbine inlet in K\n",
+      "w = 220. \t\t\t\t#mass flow rate in N/s\n",
+      "Cp_air = 1005. \t\t\t\t#Specific heat capacity at consmath.tant pressure in J/kg-K\n",
+      "k = 1.4 \t\t\t\t#Adiabatic consmath.tant for air\n",
+      "Cp_gas = 1153. \t\t\t\t#Specific heat capacity at consmath.tant pressure in J/kg-K\n",
+      "k_gas = 1.3 \t\t\t\t#Adiabatic consmath.tant\n",
+      "To1 = 15.+273 \t\t\t\t#Inlet Stagnation temperature of compressor in K\n",
+      "Po1 = 1. \t\t\t\t#Inlet Stagnation pressure in bar\n",
+      "Poe = Po1 \t\t\t\t#Exit stagnation pressure in bar, Since exit at ambient conditions\n",
+      "g = 9.81 \t\t\t\t#Acceleration due to gravity in m/s**2\n",
+      "\n",
+      "\t\t\t\t\n",
+      "#Calculation  \n",
+      "To2s = To1*(pr)**((k-1)/k) \t\t\t\t#Exit Stagnation temperature of compressor at isentropic process in K\n",
+      "To2 = ((To2s-To1)/eff_com)+To1 \t\t\t\t#Exit Stagnation temperature of compressor in K\n",
+      "Wc = (Cp_air*(To2-To1)) \t\t\t\t#Work given to compressor in J/kg, Cp in J/kg-K\n",
+      "To4 = To3-(Wc/Cp_gas*eff_m) \t\t\t\t#Exit Stagnation temperature of turbine in K\n",
+      "To4s = To3-((To3-To4)/eff_t) \t\t\t\t#Exit Stagnation temperature of turbine at isentropic process in K\n",
+      "Po2 = Po1*pr \t\t\t\t#Exit Stagnation pressure of compressor in bar\n",
+      "Po3 = Po2 \t\t\t\t#Exit Stagnation pressure of combustion chamber in bar, Since the process takes place at consmath.tant pressure process \n",
+      "p1 = (To3/To4s)**(k_gas/(k_gas-1)) \t\t\t\t#Stagnation Pressure ratio of inlet and outlet of turbine  \n",
+      "Po4s = Po3/p1 \t\t\t\t#Stagnation Pressure at outlet of turbine at isentropic process in bar  \n",
+      "pr_n = Po4s/Poe \t\t\t\t#Pressure ratio of nozzle\n",
+      "Toes = To4/((pr_n)**((k_gas-1)/k_gas)) \t\t\t\t#Exit Stagnation temperature of nozzle at isentropic process in K\n",
+      "Toe = To4-((To4-Toes)*eff_n) \t\t\t\t#Exit Stagnation temperature of nozzle in K\n",
+      "Cj = math.sqrt(2*Cp_gas*(To4-Toe)) \t\t\t\t#Jet velocity in m/s\n",
+      "m = w/g \t\t\t\t#Mass flow rate of air in kg/s\n",
+      "F = m*Cj*10**-3 \t\t\t\t#Thrust in kN\n",
+      "Fs = (F*10**3)/m  \t\t\t\t#Specific thrust in Ns/kg, F in N\n",
+      "Is = F/w \t\t\t\t#Specific impulse in sec\n",
+      "\n",
+      "\t\t\t\t\n",
+      "#Output\n",
+      "print 'A)Thrust is %3.3f kN \\\n",
+      "\\nB)Specific thrust is %3.2f Ns/kg'%(F,Fs)\n"
+     ],
+     "language": "python",
+     "metadata": {},
+     "outputs": [
+      {
+       "output_type": "stream",
+       "stream": "stdout",
+       "text": [
+        "A)Thrust is 10.180 kN \n",
+        "B)Specific thrust is 453.94 Ns/kg\n"
+       ]
+      }
+     ],
+     "prompt_number": 1
+    },
+    {
+     "cell_type": "heading",
+     "level": 2,
+     "metadata": {},
+     "source": [
+      "Example 6.2 page : 26"
+     ]
+    },
+    {
+     "cell_type": "code",
+     "collapsed": false,
+     "input": [
+      "import math \n",
+      "\n",
+      "\t\t\t\t\n",
+      "#Input data\n",
+      "u = 800.*(5./18) \t\t\t\t#Flight velocity in m/s\n",
+      "Pe = 60. \t\t\t\t\t#Ambient pressure in kPa\n",
+      "Pn = 300. \t\t\t\t\t#Pressure entering nozzle in kPa \n",
+      "Tn = 200.+273 \t\t\t\t#Temperature entering nozzle in K \n",
+      "m = 20. \t\t\t\t\t#Mass flow rate of air in kg/s\n",
+      "Cp = 1005. \t\t\t\t\t#Specific heat capacity at consmath.tant pressure in J/kg-K\n",
+      "k = 1.4 \t\t\t\t\t#Adiabatic consmath.tant for air\n",
+      "\n",
+      "\t\t\t\t\n",
+      "#Calculation\n",
+      "Te = Tn*(Pe/Pn)**((k-1)/k) \t\t\t\t#Exit temperature of nozzle in K\n",
+      "Cj = math.sqrt(2*Cp*(Tn-Te)) \t\t\t\t#Jet velocity in m/s\n",
+      "F = m*(Cj-u) \t\t\t\t#Thrust in N\n",
+      "P = F*u*10**-3 \t\t\t\t#Thrust power in kW\n",
+      "eff = ((2*u)/(Cj+u))*100 \t\t\t\t#Propulsive efficiency in percent\n",
+      "\n",
+      "\n",
+      "\t\t\t\t\n",
+      "#Output\n",
+      "print 'A)Thrust developed is %3.1f N \\\n",
+      "\\nB)Thrust developed is %3.2f kW \\\n",
+      "\\nC)Propulsive efficiency is %3.3f percent'%(F,P,eff)\n"
+     ],
+     "language": "python",
+     "metadata": {},
+     "outputs": [
+      {
+       "output_type": "stream",
+       "stream": "stdout",
+       "text": [
+        "A)Thrust developed is 7395.4 N \n",
+        "B)Thrust developed is 1643.42 kW \n",
+        "C)Propulsive efficiency is 54.586 percent\n"
+       ]
+      }
+     ],
+     "prompt_number": 2
+    },
+    {
+     "cell_type": "heading",
+     "level": 2,
+     "metadata": {},
+     "source": [
+      "Example 6.3 page : 26"
+     ]
+    },
+    {
+     "cell_type": "code",
+     "collapsed": false,
+     "input": [
+      "import math \n",
+      "\n",
+      "\t\t\t\t\n",
+      "#Input data\n",
+      "Mi = 0.8 \t\t\t\t#Inlet mach number \n",
+      "h = 10000. \t\t\t\t#Altitude in m\n",
+      "pr_c = 8. \t\t\t\t#Pressure ratio of compressor\n",
+      "To3 = 1200. \t\t\t\t#Stagnation temperature at turbine inlet in K\n",
+      "eff_c = 0.87 \t\t\t\t#Compressor efficiency\n",
+      "eff_t = 0.9 \t\t\t\t#Turbine efficiency\n",
+      "eff_d = 0.93 \t\t\t\t#Diffuser efficiency \n",
+      "eff_n = 0.95  \t\t\t\t#Nozzle efficiency \n",
+      "eff_m = 0.99 \t\t\t\t#Mechanical transmission efficiency\n",
+      "eff_cc = 0.98 \t\t\t\t#Combustion efficiency\n",
+      "pl = 0.04 \t\t\t\t#Ratio of combustion pressure loss to compressor delivery pressure \n",
+      "k = 1.4 \t\t\t\t#Adiabatic consmath.tant of air\n",
+      "R = 287. \t\t\t\t#Specific gas consmath.tant in J/kg-K\n",
+      "k_g = 1.33 \t\t\t\t#Adiabatic consmath.tant of gas \n",
+      "Cp_a = 1005. \t\t\t\t#Specific heat capacity at consmath.tant pressure of air in J/kg-K\n",
+      "Cp_g = 1100. \t\t\t\t#Specific heat capacity at consmath.tant pressure of gas in J/kg-K\n",
+      "CV = 43000000. \t\t\t\t#Calorific value in J/kg (AssumE)\n",
+      "\n",
+      "\t\t\t\t\n",
+      "#Calculation\n",
+      "Ti = 223.15 \t\t\t\t#Inlet temperature in K from gas tables\n",
+      "Pi = 26.4 \t\t\t\t#Inlet pressure in kPa from gas tables \n",
+      "ai = math.sqrt(k*R*Ti) \t\t\t\t#Sound velocity in m/s\n",
+      "Ci = ai*Mi \t\t\t\t#Velocity of air in m/s,\n",
+      "u = Ci \t\t\t\t#Flight velocity in m/s, Since it is reaction force of Ci\n",
+      "t1 = 0.886 \t\t\t\t#Ratio of static to stagnation temperature a entry from gas tables at M = 0.8 \n",
+      "To1s = Ti/t1 \t\t\t\t#Stagnation temperature at inlet of compressor at isentropic process in K\n",
+      "To1 = ((To1s-Ti)/eff_d)+Ti \t\t\t\t#Stagnation temperature at inlet of compressor in K\n",
+      "p1 = (To1s/Ti)**(k/(k-1)) \t\t\t\t#Pressure ratio i.e. (Po1s/Pi)\n",
+      "Po1s = Pi*p1 \t\t\t\t#inlet Stagnation pressure of compressor at isentropic process in kPa\n",
+      "Po1 = Po1s \t\t\t\t#Inlet Stagnation pressure of compressor in kPa\n",
+      "Po2 = pr_c*Po1 \t\t\t\t#Exit Stagnation pressure of compressor in kPa\n",
+      "To2s = To1s*(Po2/Po1)**((k-1)/k) \t\t\t\t#Exit Stagnation temperature of compressor at isentropic process in K\n",
+      "To2 = ((To2s-To1)/eff_c)+To1 \t\t\t\t#Exit Stagnation temperature of compressor in K\n",
+      "P_los = pl*Po2 \t\t\t\t#combustion pressure loss in kPa\n",
+      "Po3 = Po2-P_los \t\t\t\t#Exit Stagnation pressure of combustion chamber in kPa\n",
+      "To4 = To3-((Cp_a*(To2-To1))/(eff_m*Cp_g)) \t\t\t\t#Exit Stagnation temperature of turbine in K\n",
+      "To4s = To3-((To3-To4)/eff_t) \t\t\t\t#Exit Stagnation temperature of turbine at isentropic process in K\n",
+      "p1 = (To3/To4s)**(k_g/(k_g-1)) \t\t\t\t#Pressure ratio i.e. (Po3/Po4s)\n",
+      "Po4s = Po3/p1 \t\t\t\t#Stagnation Pressure at outlet of turbine at isentropic process in kPa\n",
+      "Poe = Pi \t\t\t\t#Exit stagnation pressure in kPa, Since exit is at ambient conditions\n",
+      "pr_n = Po4s/Poe \t\t\t\t#Pressure ratio of nozzle\n",
+      "Toes = To4/((pr_n)**((k_g-1)/k_g)) \t\t\t\t#Exit Stagnation temperature of nozzle at isentropic process in K\n",
+      "Toe = To4-((To4-Toes)*eff_n) \t\t\t\t#Exit Stagnation temperature of nozzle in K\n",
+      "Cj = math.sqrt(2*Cp_g*(To4-Toe)) \t\t\t\t#Jet velocity in m/s\n",
+      "Fs = Cj-u \t\t\t\t#Specific thrust in Ns/kg\n",
+      "f = ((Cp_g*To3)-(Cp_a*To2))/((eff_cc*CV)-(Cp_g*To3)) \t\t\t\t#Fuel-air ratio\n",
+      "TSFC = (f/Fs)#*10**5 \t\t\t\t#Thrust specific fuel consumption in kg/s-N x10**-5\n",
+      "\n",
+      "\t\t\t\t\n",
+      "#Output\n",
+      "print 'A)Specific thrust is %3.2f Ns/kg \\\n",
+      "\\nB)Thrust specific fuel consumption is %.3e kg/s-N'%(Fs,TSFC)\n"
+     ],
+     "language": "python",
+     "metadata": {},
+     "outputs": [
+      {
+       "output_type": "stream",
+       "stream": "stdout",
+       "text": [
+        "A)Specific thrust is 575.62 Ns/kg \n",
+        "B)Thrust specific fuel consumption is 3.537e-05 kg/s-N\n"
+       ]
+      }
+     ],
+     "prompt_number": 4
+    },
+    {
+     "cell_type": "heading",
+     "level": 2,
+     "metadata": {},
+     "source": [
+      "Example 6.4 page : 29"
+     ]
+    },
+    {
+     "cell_type": "code",
+     "collapsed": false,
+     "input": [
+      "import math \n",
+      "\n",
+      "\t\t\t\t\n",
+      "#Input data\n",
+      "u = 300. \t\t\t\t#Flight velocity in m/s\n",
+      "Pi = 35. \t\t\t\t#Inlet pressure in kPa\n",
+      "Ti = -40.+273 \t\t\t\t#Inlet temperature in K\n",
+      "pr_c = 10. \t\t\t\t#Pressure ratio of compressor\n",
+      "T3 = 1100.+273 \t\t\t\t#Inlet turbine temperature in K\n",
+      "m = 50. \t\t\t\t#Mass flow rate of air in kg/s\n",
+      "k = 1.4 \t\t\t\t#Adiabatic consmath.tant of air\n",
+      "Cp = 1005. \t\t\t\t#Specific heat capacity at consmath.tant pressure of air in J/kg-K\n",
+      "R = 287. \t\t\t\t#Specific gas consmath.tant in J/kg-K\n",
+      "\n",
+      "\t\t\t\t\n",
+      "#Calculation\n",
+      "ai = math.sqrt(k*R*Ti) \t\t\t\t#Sound velocity at diffuser in m/s\n",
+      "C1 = u \t\t\t\t#Velocity of air in m/s, Since it is reaction force of u\n",
+      "T1 = Ti+(C1**2/(2*Cp)) \t\t\t\t#Temperature at inlet of compressor in K\n",
+      "P1 = Pi*((T1/Ti)**(k/(k-1))) \t\t\t\t#Inlet pressure of compressor in kPa\n",
+      "P2 = pr_c*P1 \t\t\t\t#Exit pressure of compressor in kPa\n",
+      "P3 = P2 \t\t\t\t#Exit pressure of combustion chamber in kPa, Since the process takes place at consmath.tant pressure process \n",
+      "T2 = T1*(P2/P1)**((k-1)/k) \t\t\t\t#Exit temperature of compressor in K\n",
+      "T4 = T3-(T2-T1) \t\t\t\t#Exit temperature of turbine in K\n",
+      "P4 = P3/((T3/T4)**(k/(k-1))) \t\t\t\t#Pressure at outlet of turbine in kPa\n",
+      "Pe = Pi \t\t\t\t#Exit pressure in kPa, Since exit is at ambient conditions\n",
+      "pr_n = P4/Pe \t\t\t\t#Pressure ratio of nozzle\n",
+      "Te = T4/((pr_n)**((k-1)/k)) \t\t\t\t#Exit temperature of nozzle in K\n",
+      "Cj = math.sqrt(2*Cp*(T4-Te)) \t\t\t\t#Jet velocity in m/s\n",
+      "sig = u/Cj \t\t\t\t#Jet speed ratio \n",
+      "eff_prop = ((2*sig)/(1+sig))*100 \t\t\t\t#Propulsive efficiency of the cycle in %\n",
+      "\n",
+      "\t\t\t\t\n",
+      "#Output\n",
+      "print 'A)Temperature and pressure of gases at turbine exit is %3.2f K and %3i kPa \\\n",
+      "\\nB)Velocity of gases is %3.2f m/s \\\n",
+      "\\nC)Propulsive efficiency of the cycle is %3.2f percent'%(T4,P4,Cj,eff_prop)\n"
+     ],
+     "language": "python",
+     "metadata": {},
+     "outputs": [
+      {
+       "output_type": "stream",
+       "stream": "stdout",
+       "text": [
+        "A)Temperature and pressure of gases at turbine exit is 1114.47 K and 311 kPa \n",
+        "B)Velocity of gases is 1020.35 m/s \n",
+        "C)Propulsive efficiency of the cycle is 45.44 percent\n"
+       ]
+      }
+     ],
+     "prompt_number": 5
+    },
+    {
+     "cell_type": "heading",
+     "level": 2,
+     "metadata": {},
+     "source": [
+      "Example 6.5 page : 30"
+     ]
+    },
+    {
+     "cell_type": "code",
+     "collapsed": false,
+     "input": [
+      "import math \n",
+      "\n",
+      "\t\t\t\t\n",
+      "#Input data\n",
+      "n = 2 \t\t\t\t#Number of jets\n",
+      "D = 0.25 \t\t\t\t#Diameter of turbojet in m\n",
+      "P = 3000 \t\t\t\t#Net power at turbojet in W\n",
+      "mf_kWh = 0.42 \t\t\t\t#Fuel consumption in kg/kWh \n",
+      "CV = 49000 \t\t\t\t#Calorific value in kJ/kg\n",
+      "u = 300 \t\t\t\t#Flight velocity in m/s\n",
+      "d = 0.168 \t\t\t\t#Density in kg/m**3\n",
+      "AFR = 53 \t\t\t\t#Air fuel ratio \n",
+      "\n",
+      "\t\t\t\t#Calculatioon\n",
+      "mf = mf_kWh*P/3600 \t\t\t\t#Mass flow rate of fuel in kg/s\n",
+      "ma = AFR*mf \t\t\t\t#Mass flow rate of air in kg/s\n",
+      "m = ma+mf \t\t\t\t#Mass flow rate of gas in kg/s\n",
+      "Q = m/d \t\t\t\t#Volume flow rate in m**3/s\n",
+      "Cj = (Q*4)/(2*math.pi*D**2) \t\t\t\t#Jet velocity in m/s\n",
+      "Ca = Cj-u \t\t\t\t#Absolute Jet velocity in m/s\n",
+      "F = ((m*Cj)-(ma*u))*10**-3 \t\t\t\t#Thrust in kN\n",
+      "eff = ((F*u)/(mf*CV))*100 \t\t\t\t#Overall efficiency in %\n",
+      "eff_prop = ((2*u)/(Cj+u))*100 \t\t\t\t#Propulsive efficiency of the cycle in %\n",
+      "eff_ther = (eff/eff_prop)*100 \t\t\t\t#Efficiency of turbine in %\n",
+      "\n",
+      "\t\t\t\t\n",
+      "#Output\n",
+      "print 'A)Absolute velocity of jet is %3.3f m/s \\\n",
+      "\\nB)Resistance of the plane is %3.4f kN \\\n",
+      "\\nC)Overall efficiency is %3.2f percent \\\n",
+      "\\nD)Efficiency of turbine is %3.3f percent'%(Ca,F,eff,eff_ther)\n"
+     ],
+     "language": "python",
+     "metadata": {},
+     "outputs": [
+      {
+       "output_type": "stream",
+       "stream": "stdout",
+       "text": [
+        "A)Absolute velocity of jet is 845.916 m/s \n",
+        "B)Resistance of the plane is 16.0928 kN \n",
+        "C)Overall efficiency is 28.15 percent \n",
+        "D)Efficiency of turbine is 67.839 percent\n"
+       ]
+      }
+     ],
+     "prompt_number": 6
+    },
+    {
+     "cell_type": "heading",
+     "level": 2,
+     "metadata": {},
+     "source": [
+      "Example 6.6 page : 31"
+     ]
+    },
+    {
+     "cell_type": "code",
+     "collapsed": false,
+     "input": [
+      "import math \n",
+      "\n",
+      "\t\t\t\t\n",
+      "#Input data\n",
+      "u = 900*(5./18) \t\t\t\t#Flight velocity in m/s\n",
+      "ma = 3000./60 \t\t\t\t#Mass flow rate of air in kg/s\n",
+      "dh = 200. \t\t\t\t#Enthalpy drop of nozzle in kJ/kg\n",
+      "eff_n = 0.9 \t\t\t\t#Nozzle efficiency \n",
+      "AFR = 85 \t\t\t\t#Air fuel ratio \n",
+      "eff_cc = 0.95 \t\t\t\t#Combustion efficiency\n",
+      "CV = 42000 \t\t\t\t#Calorific value in kJ/kg\n",
+      "\n",
+      "\t\t\t\t\n",
+      "#Calculation\n",
+      "mf = ma/AFR \t\t\t\t#Mass flow rate of fuel in kg/s\n",
+      "m = ma+mf \t\t\t\t#Mass flow rate of gas in kg/s\n",
+      "Cj = math.sqrt(2*eff_n*dh*10**3) \t\t\t\t#Jet velocity in m/s\n",
+      "sig = u/Cj \t\t\t\t#Jet speed ratio \n",
+      "F = ((m*Cj)-(ma*u))*10**-3 \t\t\t\t#Thrust in kN\n",
+      "Pt = F*u \t\t\t\t#Thrust power in kW\n",
+      "Pp = 0.5*((m*Cj**2)-(ma*u**2))*10**-3 \t\t\t\t#Propulsive power in kW\n",
+      "HS = eff_cc*mf*CV \t\t\t\t#Heat supplied in kW\n",
+      "eff_ther = (Pp/HS)*100 \t\t\t\t#Efficiency of turbine in %\n",
+      "eff_prop = (Pt/Pp)*100 \t\t\t\t#Propulsive efficiency of the cycle in %\n",
+      "eff = (Pt/HS)*100 \t\t\t\t#Overall efficiency in %\n",
+      "\n",
+      "\t\t\t\t\n",
+      "#Output\n",
+      "print 'A)Propulsive power is %3.2f kW \\\n",
+      "\\nB)Thrust power is %3.1f kW \\\n",
+      "\\nC)Propulsive efficiency is %3.3f percent \\\n",
+      "\\nD)Thermal efficiency is %3.2f percent \\\n",
+      "\\nE)Total fuel consumption is %3.3f kg/s F)Overall efficiency is %3.3f percent'%(Pp,Pt,eff_prop,eff_ther,mf,eff)\n"
+     ],
+     "language": "python",
+     "metadata": {},
+     "outputs": [
+      {
+       "output_type": "stream",
+       "stream": "stdout",
+       "text": [
+        "A)Propulsive power is 7543.38 kW \n",
+        "B)Thrust power is 4463.2 kW \n",
+        "C)Propulsive efficiency is 59.168 percent \n",
+        "D)Thermal efficiency is 32.14 percent \n",
+        "E)Total fuel consumption is 0.588 kg/s F)Overall efficiency is 19.016 percent\n"
+       ]
+      }
+     ],
+     "prompt_number": 7
+    },
+    {
+     "cell_type": "heading",
+     "level": 2,
+     "metadata": {},
+     "source": [
+      "Example 6.7 page : 32"
+     ]
+    },
+    {
+     "cell_type": "code",
+     "collapsed": false,
+     "input": [
+      "import math \n",
+      "\n",
+      "\t\t\t\t\n",
+      "#Input data\n",
+      "M = 0.8 \t\t\t\t#Mach number \n",
+      "CV = 42800. \t\t\t\t#Calorific value in kJ/kg\n",
+      "h = 10. \t\t\t\t#Altitude in km\n",
+      "F = 50. \t\t\t\t#Thrust in kN\n",
+      "ma = 45. \t\t\t\t#Mass flow rate of air in kg/s\n",
+      "mf = 2.65 \t\t\t\t#Mass flow rate of fuel in kg/s\n",
+      "\n",
+      "\t\t\t\t\n",
+      "#Calculation\n",
+      "m = ma+mf \t\t\t\t#Mass flow rate of gas in kg/s\n",
+      "a = 299.6 \t\t\t\t#Sound velocity in m/s, from gas tables\n",
+      "T = 233.15 \t\t\t\t#Inlet temperature in K \n",
+      "u = a*M \t\t\t\t#Flight velocity in m/s\n",
+      "Cj = ((F*10**3)+(ma*u))/m \t\t\t\t#Jet velocity in m/s\n",
+      "sig = u/Cj \t\t\t\t#Jet speed ratio \n",
+      "Fs = F*10**3/m \t\t\t\t#Specific thrust in Ns/kg, F in N\n",
+      "TSFC = mf*3600/(F*10**3) \t\t\t\t#Thrust specific fuel consumption in kg/N-hr, F in N\n",
+      "Pt = F*u \t\t\t\t#Thrust power in kW\n",
+      "Pp = 0.5*((m*Cj**2)-(ma*u**2))*10**-3 \t\t\t\t#Propulsive power in kW\n",
+      "HS = mf*CV \t\t\t\t#Heat supplied in kW\n",
+      "eff_ther = (Pp/HS)*100 \t\t\t\t#Efficiency of turbine in %\n",
+      "eff_prop = (Pt/Pp)*100 \t\t\t\t#Propulsive efficiency of the cycle in %\n",
+      "eff = (Pt/HS)*100 \t\t\t\t#Overall efficiency in %\n",
+      "\n",
+      "\t\t\t\t\n",
+      "#Output \n",
+      "print 'A)Specific thrust is %3.2f N/kg \\\n",
+      "\\nB)Thrust specific fuel consumption is %3.4f kg/N-hr \\\n",
+      "\\nC)Jet velocity is %3.3f m/s \\\n",
+      "\\nD)Thermal efficiency is %3.2f percent \\\n",
+      "\\nE)Propulsive efficiency is %3.3f percent F)Overall efficiency is %3.2f percent'%(Fs,TSFC,Cj,eff_ther,eff_prop,eff)\n"
+     ],
+     "language": "python",
+     "metadata": {},
+     "outputs": [
+      {
+       "output_type": "stream",
+       "stream": "stdout",
+       "text": [
+        "A)Specific thrust is 1049.32 N/kg \n",
+        "B)Thrust specific fuel consumption is 0.1908 kg/N-hr \n",
+        "C)Jet velocity is 1275.668 m/s \n",
+        "D)Thermal efficiency is 33.04 percent \n",
+        "E)Propulsive efficiency is 31.976 percent F)Overall efficiency is 10.57 percent\n"
+       ]
+      }
+     ],
+     "prompt_number": 8
+    },
+    {
+     "cell_type": "heading",
+     "level": 2,
+     "metadata": {},
+     "source": [
+      "Example 6.8 page : 34"
+     ]
+    },
+    {
+     "cell_type": "code",
+     "collapsed": false,
+     "input": [
+      "import math \n",
+      "\n",
+      "\t\t\t\t\n",
+      "#Input data\n",
+      "Mi = 0.8 \t\t\t\t#Inlet mach number \n",
+      "h = 10. \t\t\t\t#Altitude in km\n",
+      "To3 = 1200. \t\t\t\t#Stagnation temperature before turbine inlet in K\n",
+      "dTc = 175. \t\t\t\t#Stagnation temperature rise through the compressor in K\n",
+      "CV = 43000. \t\t\t\t#Calorific value in kJ/kg\n",
+      "eff_c = 0.75 \t\t\t\t#Compressor efficiency\n",
+      "eff_cc = 0.75 \t\t\t\t#Combustion efficiency\n",
+      "eff_t = 0.81 \t\t\t\t#Turbine efficiency\n",
+      "eff_m = 0.98 \t\t\t\t#Mechanical transmission efficiency\n",
+      "eff_n = 0.97 \t\t\t\t#Nozzle efficiency \n",
+      "Is = 25. \t\t\t\t#Specific impulse in sec\n",
+      "k = 1.4 \t\t\t\t#Adiabatic consmath.tant of air\n",
+      "R = 287. \t\t\t\t#Specific gas consmath.tant in J/kg-K\n",
+      "Cp = 1005. \t\t\t\t#Specific heat capacity at consmath.tant pressure of air in J/kg-K\n",
+      "g = 9.81 \t\t\t\t#Acceleration due to gravity in m/s**2\n",
+      "\n",
+      "\t\t\t\t\n",
+      "#Calculation\n",
+      "Ti = 223.15 \t\t\t\t#Inlet temperature in K from gas tables\n",
+      "ai = math.sqrt(k*R*Ti) \t\t\t\t#Sound velocity in m/s\n",
+      "Toi = (1+((0.5*(k-1)*Mi**2)))*Ti \t\t\t\t#Stagnation temperature at diffuser inlet in K\n",
+      "To1 = Toi \t\t\t\t#Inlet Stagnation temperature of compressor in K, math.since hoi = ho1 \n",
+      "To2 = dTc+To1 \t\t\t\t#Exit Stagnation temperature of compressor in K\n",
+      "pr_c = (1+(eff_c*((To2-To1)/To1)))**(k/(k-1)) \t\t\t\t#Compressor pressure ratio \n",
+      "f = ((Cp*To3)-(Cp*To2))/((eff_cc*CV*10**3)-(Cp*To3)) \t\t\t\t#Fuel-air ratio, calculation mistake in textbook\n",
+      "dTt = dTc/(eff_m*(1+f)) \t\t\t\t#Temperature difference across turbine\n",
+      "pr_t = 1/((1-(dTt/(To3*eff_t)))**(k/(k-1))) \t\t\t\t#Turbine pressure ratio\n",
+      "To4 = To3-dTc \t\t\t\t#Exit Stagnation temperature of turbine in K\n",
+      "u = ai*Mi \t\t\t\t#Flight velocity in m/s\n",
+      "sig = 1/(((Is*g)/u)+1) \t\t\t\t#Jet speed ratio \n",
+      "Ce = u/sig \t\t\t\t#Exit velocity in m/s\n",
+      "Cj = Ce \t\t\t\t#Jet velocity in m/s, Since Cj is due to exit velociy\n",
+      "Te = To4-(Ce**2/(2*Cp)) \t\t\t\t#Exit temperature in K\n",
+      "Tes = To4-((To4-Te)*eff_n) \t\t\t\t#Exit temperature in K, (At isentropic process)\n",
+      "pr_n = (To4/Te)**(k/(k-1)) \t\t\t\t#Nozzle pressure ratio\n",
+      "ae = math.sqrt(k*R*Te) \t\t\t\t#Exit Sound velocity in m/s\n",
+      "Me = Ce/ae \t\t\t\t#Exit mach number \n",
+      "\n",
+      "print 'A)Fuel-air ratio is %3.5f  \\\n",
+      "\\nB)Compressor, turbine, nozzle pressure ratio are %3.3f, %3.3f, %3.2f respectively \\\n",
+      "\\nC)Mach number at exhaust jet is %3.3f'%(f,pr_c,pr_t,pr_n,Me)\n"
+     ],
+     "language": "python",
+     "metadata": {},
+     "outputs": [
+      {
+       "output_type": "stream",
+       "stream": "stdout",
+       "text": [
+        "A)Fuel-air ratio is 0.02503  \n",
+        "B)Compressor, turbine, nozzle pressure ratio are 4.344, 1.996, 1.53 respectively \n",
+        "C)Mach number at exhaust jet is 0.803\n"
+       ]
+      }
+     ],
+     "prompt_number": 9
+    },
+    {
+     "cell_type": "heading",
+     "level": 2,
+     "metadata": {},
+     "source": [
+      "Example 6.9 page : 36"
+     ]
+    },
+    {
+     "cell_type": "code",
+     "collapsed": false,
+     "input": [
+      "import math \n",
+      "\n",
+      "\t\t\t\t\n",
+      "#Input data\n",
+      "D = 2.5 \t\t\t\t#Diameter in m\n",
+      "u = 500.*(5./18) \t\t\t\t#Flight velocity in m/s\n",
+      "h = 8000. \t\t\t\t#Altitude in m\n",
+      "sig = 0.75 \t\t\t\t#Jet speed ratio \n",
+      "g = 9.81 \t\t\t\t#Acceleration due to gravity in m/s**2\n",
+      "\n",
+      "\t\t\t\t\n",
+      "#Calculation\n",
+      "d = 0.525 \t\t\t\t#from gas tables\n",
+      "A = math.pi*D**2*0.25 \t\t\t\t#Area of flow in m**2 \n",
+      "Cj = u/sig \t\t\t\t#Jet velocity in m/s\n",
+      "Vf = (u+Cj)/2 \t\t\t\t#Velocity of flow in m/s\n",
+      "ma = d*A*Vf \t\t\t\t#Mass flow rate of air in kg/s\n",
+      "F = ma*(Cj-u)*10**-3 \t\t\t\t#Thrust in kN\n",
+      "P = F*u \t\t\t\t#Thrust power in kW\n",
+      "Fs = F*10**3/ma \t\t\t\t#Specific thrust in Ns/kg\n",
+      "Is = Fs/g \t\t\t\t#Specific impulse in sec\n",
+      "\n",
+      "\t\t\t\t\n",
+      "#Output\n",
+      "print 'A)Flow rate of air through the propeller is %3.3f m/s \\\n",
+      "\\nB)Thrust produced is %3.3f kN \\\n",
+      "\\nC)Specific thrust is %3.2f N-s/kg \\\n",
+      "\\nD)Specific impulse is %3.3f sec \\\n",
+      "\\nE)Thrust power is %3.1f kW'%(ma,F,Fs,Is,P)\n"
+     ],
+     "language": "python",
+     "metadata": {},
+     "outputs": [
+      {
+       "output_type": "stream",
+       "stream": "stdout",
+       "text": [
+        "A)Flow rate of air through the propeller is 417.584 m/s \n",
+        "B)Thrust produced is 19.333 kN \n",
+        "C)Specific thrust is 46.30 N-s/kg \n",
+        "D)Specific impulse is 4.719 sec \n",
+        "E)Thrust power is 2685.1 kW\n"
+       ]
+      }
+     ],
+     "prompt_number": 10
+    },
+    {
+     "cell_type": "heading",
+     "level": 2,
+     "metadata": {},
+     "source": [
+      "Example 6.10 page : 37"
+     ]
+    },
+    {
+     "cell_type": "code",
+     "collapsed": false,
+     "input": [
+      "import math \n",
+      "\n",
+      "\t\t\t\t\n",
+      "#Input data\n",
+      "h = 3000. \t\t\t\t#Altitude in m\n",
+      "Pi = 0.701 \t\t\t\t#Inlet pressure in bar\n",
+      "Ti = 268.65 \t\t\t\t#Inlet temperature in K\n",
+      "u = 525*(5./18) \t\t\t\t#Flight velocity in m/s\n",
+      "eff_d = 0.875 \t\t\t\t#Diffuser efficiency\n",
+      "eff_c = 0.79 \t\t\t\t#Compressor efficiency\n",
+      "C1 = 90. \t\t\t\t#Velocity of air at compressor in m/s\n",
+      "dTc = 230. \t\t\t\t#Temperature rise through compressor\n",
+      "k = 1.4 \t\t\t\t#Adiabatic consmath.tant of air\n",
+      "Cp = 1005. \t\t\t\t#Specific heat capacity at consmath.tant pressure of air in J/kg-K\n",
+      "R = 287. \t\t\t\t#Specific gas consmath.tant in J/kg-K\n",
+      "\n",
+      "\t\t\t\t\n",
+      "#Calculation\n",
+      "ai = math.sqrt(k*R*Ti) \t\t\t\t#Sound velocity in m/s\n",
+      "Mi = u/ai \t\t\t\t#Inlet mach number \n",
+      "Toi = (1+((0.5*(k-1)*Mi**2)))*Ti \t\t\t\t#Stagnation temperature at diffuser inlet in K\n",
+      "To1 = Toi \t\t\t\t#Inlet Stagnation temperature of compressor in K, math.since hoi = ho1 \n",
+      "T1 = To1-(C1**2/(2*Cp)) \t\t\t\t#Temperature at inlet of compressor in K\n",
+      "P1 = Pi*((1+(eff_d*((T1/Ti)-1)))**(k/(k-1))) \t\t\t\t#Inlet pressure of compressor in bar\n",
+      "dPc = P1-Pi \t\t\t\t#Pressure rise through inlet diffuser in bar\n",
+      "pr_c = (((eff_c*dTc)/To1)+1)**(k/(k-1)) \t\t\t\t#Pressure ratio of compressor\n",
+      "P = Cp*(dTc) \t\t\t\t#Power required by the compressor in kW/(kg/s)\n",
+      "eff = 1-(1/pr_c**((k-1)/k)) \t\t\t\t#Air standard efficiency\n",
+      "\n",
+      "\t\t\t\t\n",
+      "#Output\n",
+      "print 'A)Pressure rise through diffuser is %3.4f bar \\\n",
+      "\\nB)Pressure developed by compressure is %3.4f bar \\\n",
+      "\\nC)Air standard efficiency of the engine is %3.4f'%(dPc,P1,eff)\n"
+     ],
+     "language": "python",
+     "metadata": {},
+     "outputs": [
+      {
+       "output_type": "stream",
+       "stream": "stdout",
+       "text": [
+        "A)Pressure rise through diffuser is 0.0538 bar \n",
+        "B)Pressure developed by compressure is 0.7548 bar \n",
+        "C)Air standard efficiency of the engine is 0.3942\n"
+       ]
+      }
+     ],
+     "prompt_number": 11
+    },
+    {
+     "cell_type": "heading",
+     "level": 2,
+     "metadata": {},
+     "source": [
+      "Example 6.11 page : 38"
+     ]
+    },
+    {
+     "cell_type": "code",
+     "collapsed": false,
+     "input": [
+      "import math \n",
+      "\n",
+      "\t\t\t\t\n",
+      "#Input data\n",
+      "h = 9500. \t\t\t\t#Altitude in m\n",
+      "u = 800*(5./18) \t\t\t\t#Flight velocity in m/s\n",
+      "eff_prop = 0.55 \t\t\t\t#Propulsive efficiency of the cycle\n",
+      "eff_o = 0.17 \t\t\t\t#Overall efficiency\n",
+      "F = 6100. \t\t\t\t#Thrust in N\n",
+      "d = 0.17 \t\t\t\t#Density in kg/m**3\n",
+      "CV = 46000. \t\t\t\t#Calorific value in kJ/kg\n",
+      "\n",
+      "\t\t\t\t\n",
+      "#Calculation\n",
+      "mf = (F*u)/(eff_o*CV*10**3) \t\t\t\t#Mass flow rate of fuel in kg/s\n",
+      "Cj = ((2*u)/(eff_prop))-u \t\t\t\t#Jet velocity in m/s, wrong calculation in textbook\n",
+      "Ca = Cj-u \t\t\t\t#Absolute Jet velocity in m/s\n",
+      "ma = (F-(mf*Cj))/(Ca) \t\t\t\t#Mass flow rate of air in kg/s\n",
+      "m = ma+mf \t\t\t\t#Mass flow rate of gas in kg/s\n",
+      "f = ma/mf \t\t\t\t#Air fuel ratio\n",
+      "Q = m/d \t\t\t\t#Volume flow rate in m**3/s\n",
+      "Dj = math.sqrt((4*Q)/(math.pi*Cj))*10**3 \t\t\t\t#Diameter of jet in mm, Cj value wrong in textbook \n",
+      "P = ((F*u)/eff_prop)*10**-3 \t\t\t\t#Power output of engine in kW\n",
+      "\n",
+      "\t\t\t\t\n",
+      "#Output\n",
+      "print 'A)Diamter of the jet is %3.1f mm \\\n",
+      "\\nB)Power output is %3.1f kW \\\n",
+      "\\nC)Air-fuel ratio is %3.3f \\\n",
+      "\\nD)Absolute velocity of the jet is %3i m/s'%(Dj,P,f,Ca)\n"
+     ],
+     "language": "python",
+     "metadata": {},
+     "outputs": [
+      {
+       "output_type": "stream",
+       "stream": "stdout",
+       "text": [
+        "A)Diamter of the jet is 461.6 mm \n",
+        "B)Power output is 2464.6 kW \n",
+        "C)Air-fuel ratio is 95.161 \n",
+        "D)Absolute velocity of the jet is 363 m/s\n"
+       ]
+      }
+     ],
+     "prompt_number": 12
+    },
+    {
+     "cell_type": "heading",
+     "level": 2,
+     "metadata": {},
+     "source": [
+      "Example 6.12 page : 39"
+     ]
+    },
+    {
+     "cell_type": "code",
+     "collapsed": false,
+     "input": [
+      "import math \n",
+      "\n",
+      "\t\t\t\t\n",
+      "#Input data\n",
+      "u = 960*(5./18) \t\t\t\t#Flight velocity in m/s\n",
+      "ma = 40. \t\t\t\t#Mass flow rate of air in kg/s\n",
+      "AFR = 50. \t\t\t\t#Air fuel ratio\n",
+      "sig = 0.5 \t\t\t\t#Jet speed ratio, for maximum thrust power\n",
+      "CV = 43000. \t\t\t\t#Calorific value in kJ/kg\n",
+      "\n",
+      "\t\t\t\t\n",
+      "#Calculation\n",
+      "mf = ma/AFR \t\t\t\t#Mass flow rate of fuel in kg/s\n",
+      "m = ma+mf \t\t\t\t#Mass flow rate of gas in kg/s\n",
+      "Cj = u/sig \t\t\t\t#Jet velocity in m/s\n",
+      "F = ((m*Cj)-(ma*u))*10**-3 \t\t\t\t#Thrust in kN\n",
+      "Fs = F*10**3/m \t\t\t\t#Specific thrust in Ns/kg, F in N\n",
+      "Pt = F*u \t\t\t\t#Thrust power in kW\n",
+      "eff_prop = ((2*sig)/(1+sig))*100 \t\t\t\t#Propulsive efficiency of the cycle in %\n",
+      "eff_ther = ((0.5*m*(Cj**2-u**2))/(mf*CV*10**3))*100 \t\t\t\t#Efficiency of turbine in %\n",
+      "eff = (eff_prop/100)*(eff_ther/100)*100 \t\t\t\t#Overall efficiency in %\n",
+      "TSFC = mf*3600/(F*10**3) \t\t\t\t#Thrust specific fuel consumption in kg/Nhr\n",
+      "\n",
+      "\t\t\t\t\n",
+      "#Output\n",
+      "print 'A)Jet velocity is %3.1f m/s \\\n",
+      "\\nB)Thrust is %3.3f kN \\\n",
+      "\\nC)Specific thrust is %3.2f N-s/kg \\\n",
+      "\\nD)Thrust power is %3.2f kW \\\n",
+      "\\nE)propulsive, thermal and overall efficiency is %3.2f, %3.2f and %3.3f respectively \\\n",
+      "\\nF)Thrust specific fuel consumption is %3.4f kg/Nhr'%(Cj,F,Fs,Pt,eff_prop,eff_ther,eff,TSFC)\n"
+     ],
+     "language": "python",
+     "metadata": {},
+     "outputs": [
+      {
+       "output_type": "stream",
+       "stream": "stdout",
+       "text": [
+        "A)Jet velocity is 533.3 m/s \n",
+        "B)Thrust is 11.093 kN \n",
+        "C)Specific thrust is 271.90 N-s/kg \n",
+        "D)Thrust power is 2958.22 kW \n",
+        "E)propulsive, thermal and overall efficiency is 66.67, 12.65 and 8.434 respectively \n",
+        "F)Thrust specific fuel consumption is 0.2596 kg/Nhr\n"
+       ]
+      }
+     ],
+     "prompt_number": 13
+    },
+    {
+     "cell_type": "heading",
+     "level": 2,
+     "metadata": {},
+     "source": [
+      "Example 6.13 page : 40"
+     ]
+    },
+    {
+     "cell_type": "code",
+     "collapsed": false,
+     "input": [
+      "import math \n",
+      "\n",
+      "\t\t\t\t\n",
+      "#Input data\n",
+      "u = 960*(5./18) \t\t\t\t#Flight velocity in m/s\n",
+      "ma = 54.5 \t\t\t\t#Mass flow rate of air in kg/s\n",
+      "dh = 200. \t\t\t\t#Change of enthalpy for nozzle in kJ/kg\n",
+      "Cv = 0.97 \t\t\t\t#Velocity coefficient \n",
+      "AFR = 75. \t\t\t\t#Air fuel ratio \n",
+      "eff_cc = 0.93 \t\t\t\t#Combustion efficiency\n",
+      "CV = 45000. \t\t\t\t#Calorific value in kJ/kg\n",
+      "\n",
+      "\t\t\t\t\n",
+      "#Calculation\n",
+      "mf = ma/AFR \t\t\t\t#Mass flow rate of fuel in kg/s\n",
+      "Cj = Cv*math.sqrt(2*dh*10**3) \t\t\t\t#Jet velocity in m/s\n",
+      "F = ma*(Cj-u) \t\t\t\t#Thrust in kN\n",
+      "TSFC = mf*3600/(F) \t\t\t\t#Thrust specific fuel consumption in kg/Nhr\n",
+      "HS = mf*eff_cc*CV \t\t\t\t#Heat supplied in kJ/s\n",
+      "Pp = 0.5*ma*(Cj**2-u**2)*10**-3 \t\t\t\t#Propulsive power in kW\n",
+      "Pt = F*u \t\t\t\t#Thrust power in kW\n",
+      "eff_p = Pt/(Pp*10**3) \t\t\t\t#Propulsive efficiency of the cycle\n",
+      "eff_t = Pp/HS \t\t\t\t#Efficiency of turbine\n",
+      "eff_o = Pt*10**-3/HS  \t\t\t\t#Overall efficiency\n",
+      "\n",
+      "\t\t\t\t\n",
+      "#Output\n",
+      "print 'A)Exit velocity of the jet is %3.2f m/s \\\n",
+      "\\nB)Fuel rate is %3.4f kg/s \\\n",
+      "\\nC)Thrust specific fuel consumption is %3.5f kg/Nhr \\\n",
+      "\\nD)Thermal efficiency is %3.3f \\\n",
+      "\\nE)Propulsive power is %3.2f kW \\\n",
+      "\\nF)Propulsive efficiency is %3.4f \\\n",
+      "\\nG)Overall efficiency is %3.5f'%(Cj,mf,TSFC,eff_t,Pp,eff_p,eff_o)\n"
+     ],
+     "language": "python",
+     "metadata": {},
+     "outputs": [
+      {
+       "output_type": "stream",
+       "stream": "stdout",
+       "text": [
+        "A)Exit velocity of the jet is 613.48 m/s \n",
+        "B)Fuel rate is 0.7267 kg/s \n",
+        "C)Thrust specific fuel consumption is 0.13840 kg/Nhr \n",
+        "D)Thermal efficiency is 0.274 \n",
+        "E)Propulsive power is 8318.03 kW \n",
+        "F)Propulsive efficiency is 0.6060 \n",
+        "G)Overall efficiency is 0.16574\n"
+       ]
+      }
+     ],
+     "prompt_number": 15
+    },
+    {
+     "cell_type": "heading",
+     "level": 2,
+     "metadata": {},
+     "source": [
+      "Example 6.14 page : 41"
+     ]
+    },
+    {
+     "cell_type": "code",
+     "collapsed": false,
+     "input": [
+      "import math \n",
+      "\n",
+      "\t\t\t\t\n",
+      "#Input data\n",
+      "u = 750*(5./18) \t\t\t\t#Flight velocity in m/s\n",
+      "h = 10000. \t\t\t\t#Altitude in m\n",
+      "eff_p = 0.5 \t\t\t\t#Propulsive efficiency of the cycle\n",
+      "eff_o = 0.16 \t\t\t\t#Overall efficiency\n",
+      "d = 0.173 \t\t\t\t#Density in kg/m**3\n",
+      "F = 6250. \t\t\t\t#Thrust in N\n",
+      "CV = 45000. \t\t\t\t#Calorific value in kJ/kg\n",
+      "\n",
+      "\t\t\t\t\n",
+      "#Calculation\n",
+      "sig = eff_p/(2-eff_p) \t\t\t\t#Jet speed ratio\n",
+      "Cj = u/sig \t\t\t\t#Jet velocity in m/s\n",
+      "Ca = Cj-u \t\t\t\t#Absolute Jet velocity in m/s\n",
+      "ma = F/Ca \t\t\t\t#Mass flow rate of air in kg/s\n",
+      "Q = ma*60/d \t\t\t\t#Volume flow rate in m**3/min\n",
+      "A = Q/(Cj*60) \t\t\t\t#Area of flow in m**2\n",
+      "D = math.sqrt((4*A)/(math.pi))*10**3 \t\t\t\t#Diameter in mm\n",
+      "Pt = F*u \t\t\t\t#Thrust power in W\n",
+      "Pp = (Pt/eff_p)*10**-3  \t\t\t\t#Propulsive power in kW\n",
+      "eff_t = eff_o/eff_p \t\t\t\t#Efficiency of turbine\n",
+      "HS = Pp/eff_t \t\t\t\t#Heat supplied in kJ/s\n",
+      "mf = HS/CV \t\t\t\t#Mass flow rate of fuel in kg/s\n",
+      "AFR = ma/mf \t\t\t\t#Air fuel ratio \n",
+      "\n",
+      "\t\t\t\t\n",
+      "#Output\n",
+      "print 'A)Absolute velocity of the jet is %3.2f m/s \\\n",
+      "\\nB)Volume of air compressed per minute is %3.2f m**3/min \\\n",
+      "\\nC)Diameter of the jet is %3i mm \\\n",
+      "\\nD)Power unit of the unit is %3.3f kW \\\n",
+      "\\nE)Air fuel ratio is %3.1f'%(Ca,Q,D,Pp,AFR)\n"
+     ],
+     "language": "python",
+     "metadata": {},
+     "outputs": [
+      {
+       "output_type": "stream",
+       "stream": "stdout",
+       "text": [
+        "A)Absolute velocity of the jet is 416.67 m/s \n",
+        "B)Volume of air compressed per minute is 5202.31 m**3/min \n",
+        "C)Diameter of the jet is 420 mm \n",
+        "D)Power unit of the unit is 2604.167 kW \n",
+        "E)Air fuel ratio is 82.9\n"
+       ]
+      }
+     ],
+     "prompt_number": 16
+    },
+    {
+     "cell_type": "heading",
+     "level": 2,
+     "metadata": {},
+     "source": [
+      "Example 6.15 page : 42"
+     ]
+    },
+    {
+     "cell_type": "code",
+     "collapsed": false,
+     "input": [
+      "import math \n",
+      "\n",
+      "\t\t\t\t\n",
+      "#Input data\n",
+      "P1 = 0.56 \t\t\t\t#Inlet pressure of compressor in bar\n",
+      "T1 = 260 \t\t\t\t#Temperature at inlet of compressor in K\n",
+      "pr_c = 6 \t\t\t\t#Pressure ratio of compressor\n",
+      "eff_c = 0.85 \t\t\t\t#Compressor efficiency\n",
+      "u = 360*(5./18) \t\t\t\t#Flight velocity in m/s\n",
+      "D = 3 \t\t\t\t#Propeller diameter in m \n",
+      "eff_p = 0.8 \t\t\t\t#Efficiency of propeller \n",
+      "eff_g = 0.95 \t\t\t\t#Gear reduction efficiency \n",
+      "pr_t = 5 \t\t\t\t#Expansion ratio\n",
+      "eff_t = 0.88 \t\t\t\t#Turbine efficiency\n",
+      "T3 = 1100 \t\t\t\t#temperature at turbine inlet in K\n",
+      "eff_n = 0.9 \t\t\t\t#Nozzle efficiency \n",
+      "Cp = 1005. \t\t\t\t#Specific heat capacity at consmath.tant pressure of air in J/kg-K\n",
+      "CV = 40000 \t\t\t\t#Calorific value in kJ/kg\n",
+      "k = 1.4 \t\t\t\t#Adiabatic consmath.tant of air\n",
+      "R = 287 \t\t\t\t#Specific gas consmath.tant in J/kg-K\n",
+      "\n",
+      "\t\t\t\t\n",
+      "#Calculation\n",
+      "P2 = pr_c*P1 \t\t\t\t#Exit pressure of compressor in bar\n",
+      "T2s = T1*(pr_c)**((k-1)/k) \t\t\t\t#Exit temperature of compressor at isentropic proces in K\n",
+      "T2 = T1+((T2s-T1)/eff_c) \t\t\t\t#Exit temperature of compressor in K\n",
+      "Wc = Cp*(T2-T1)*10**-3 \t\t\t\t#Power input to compressor in kJ/kg of air\n",
+      "C1 = u \t\t\t\t#Air velocity in m/s, math.since C1 is resulmath.tant of u\n",
+      "C = C1/eff_p \t\t\t\t#Average velocity in m/s\n",
+      "C2 = (2*C)-C1 \t\t\t\t#Exit velocity from compressor in m/s\n",
+      "Ap = 0.25*math.pi*D**2 \t\t\t\t#Area of propeller passage in m**2\n",
+      "Q = Ap*C \t\t\t\t#Quantity of air inducted in m**3/s\n",
+      "mf = ((T3-T2)*Cp)/((CV*10**3)-(Cp*T3)) \t\t\t\t#Mass flow rate of fuel in kg/s\n",
+      "f = mf \t\t\t\t#Fuel consumption in kg/kg of air\n",
+      "AFR = 1/mf \t\t\t\t#Air fuel ratio\n",
+      "P3 = P2 \t\t\t\t#Exit pressure of combustion chamber in bar, Since process is at consmath.tant pressure \n",
+      "P4 = P3/pr_t \t\t\t\t#Exit pressure of turbine in bar\n",
+      "T4s = T3/((pr_t)**((k-1)/k)) \t\t\t\t#Exit temperature of turbine at isentropic proces in K, wrong calculation\n",
+      "T4 = T3-(eff_t*(T3-T4s)) \t\t\t\t#Exit temperature of turbine in K\n",
+      "Po = (1+f)*Cp*(T3-T4)*10**-3 \t\t\t\t#Power output per kg of air in kJ/kg of air\n",
+      "Pa = Po-Wc \t\t\t\t#Power available for propeller in kJ/kg of air\n",
+      "Pe = P1 \t\t\t\t#Exit pressure in bar, Since exit is at ambient conditions\n",
+      "Tes = T4/((P4/Pe)**((k-1)/k)) \t\t\t\t#Exit temperature of nozzle at isentropic proces in K\n",
+      "Cj = math.sqrt(2*Cp*eff_n*(T4-Tes)) \t\t\t\t#Jet velocity in m/s\n",
+      "Fs = ((1+f)*Cj)-u \t\t\t\t#Specific thrust in Ns/kg, F in N\n",
+      "Pp = ((0.5*P1*10**5*Q*(C2**2-C1**2))/(R*T1))*10**-3 \t\t\t\t#Propulsive power by propeller in kJ/s\n",
+      "Ps = Pp/eff_g \t\t\t\t#Power supplied by the turbine in kW\n",
+      "ma = Ps/Pa \t\t\t\t#Air flow rate in kg/s\n",
+      "Fj = ma*Cj*10**-3 \t\t\t\t#Jet thrust in kN, calculation mistake\n",
+      "Fp = (Pp*eff_p)/u \t\t\t\t#Thrust produced by propeller in kN\n",
+      "\n",
+      "\t\t\t\t\n",
+      "#Output\n",
+      "print 'A)Air fuel ratio is %3.2f \\\n",
+      "\\nB)Thrust produced by the nozzle is %3.3f kN \\\n",
+      "\\nC)Thrust by the propeller is %3.3f kN \\\n",
+      "\\nD)mass flow rate through the compressor is %3.2f kg/s'%(AFR,Fj,Fp,ma)\n"
+     ],
+     "language": "python",
+     "metadata": {},
+     "outputs": [
+      {
+       "output_type": "stream",
+       "stream": "stdout",
+       "text": [
+        "A)Air fuel ratio is 60.90 \n",
+        "B)Thrust produced by the nozzle is 7.168 kN \n",
+        "C)Thrust by the propeller is 33.155 kN \n",
+        "D)mass flow rate through the compressor is 27.44 kg/s\n"
+       ]
+      }
+     ],
+     "prompt_number": 17
+    },
+    {
+     "cell_type": "heading",
+     "level": 2,
+     "metadata": {},
+     "source": [
+      "Example 6.16 page : 45"
+     ]
+    },
+    {
+     "cell_type": "code",
+     "collapsed": false,
+     "input": [
+      "import math \n",
+      "\n",
+      "\t\t\t\t\n",
+      "#Input data\n",
+      "M1 = 1.5 \t\t\t\t#Mach number  \n",
+      "h = 6500. \t\t\t\t#Altitude in m\n",
+      "D = 0.5 \t\t\t\t#Diameter in m \n",
+      "To4 = 1600. \t\t\t\t#Stagnation temperature at nozzle inlet in K\n",
+      "CV = 40000. \t\t\t\t#Calorific value in kJ/kg\n",
+      "k = 1.4 \t\t\t\t#Adiabatic consmath.tant of air\n",
+      "R = 287. \t\t\t\t#Specific gas consmath.tant in J/kg-K\n",
+      "eff_d = 0.9 \t\t\t\t#Diffuser efficiency \n",
+      "eff_cc = 0.98 \t\t\t\t#Combustion efficiency\n",
+      "eff_n = 0.96 \t\t\t\t#Nozzle efficiency \n",
+      "pr_l = 0.02 \t\t\t\t#Pressure ratio i.e. Stagnation pressure loss to Exit presure of compressor\n",
+      "Cp = 1005. \t\t\t\t#Specific heat capacity at consmath.tant pressure of air in J/kg-K\n",
+      "\n",
+      "\t\t\t\t\n",
+      "#Calculation\n",
+      "P1 = 0.44 \t\t\t\t#Inlet pressure of compressor in bar\n",
+      "T1 = 245.9  \t\t\t\t#Temperature at inlet of compressor in K\n",
+      "a1 = 314.5 \t\t\t\t#Sound velocity at compressor in m/s \n",
+      "d1 = 0.624 \t\t\t\t#Density at compressor in kg/m**3 \n",
+      "A1 = 0.25*math.pi*D**2 \t\t\t\t#Area at diffuser inlet in m**2\n",
+      "u1 = M1*a1 \t\t\t\t#Flight velocity in m/s\n",
+      "ma = d1*A1*u1 \t\t\t\t#Mass flow rate of air in kg/s\n",
+      "To2 = T1*(1+(((k-1)/2)*M1**2)) \t\t\t\t#Stagnation temperature at commpressor inlet in K\n",
+      "To1 = To2 \t\t\t\t#Stagnation temperature at commpressor outlet in K, (It is in case of diffuser)\n",
+      "pr_d = ((eff_d*(((k-1)/2)*M1**2))+1)**(k/(k-1)) \t\t\t\t#Pressure ratio of diffuser \n",
+      "P2 = pr_d*P1 \t\t\t\t#Exit pressure of compressor in bar\n",
+      "Po2 = P2 \t\t\t\t#Stagnation pressure at exit of compressor in bar \n",
+      "Po3 = (Po2-(pr_l*Po2)) \t\t\t\t#Stagnation pressure at exit of combustion chamber in bar \n",
+      "Poe = P1 \t\t\t\t#Exit stagnation pressure in kPa, Since exit is at ambient conditions\n",
+      "pr_n = Po3/Poe \t\t\t\t#Pressure ratio of nozzle\n",
+      "p1 = 1/pr_n \t\t\t\t#Inverse of pr_n to find in gas tables \n",
+      "M4s = 1.41 \t\t\t\t#Mach number at turbine exit  from gas tables \n",
+      "T4s = To4/(1+((0.5*(k-1)*M4s**2))) \t\t\t\t#Exit temperature of turbine at isentropic process in K\n",
+      "To3 = To4 \t\t\t\t#Stagnation temperature at inlet turbine in K,\n",
+      "T4 = To3-(eff_n*(To3-T4s)) \t\t\t\t#Exit temperature of turbine in K\n",
+      "C4 = math.sqrt(2*Cp*(To4-T4)) \t\t\t\t#Flight velocity of air in m/s\n",
+      "a4 = math.sqrt(k*R*T4) \t\t\t\t#Sound velocity in m/s\n",
+      "Me = C4/a4 \t\t\t\t#Nozzle jet mach number\n",
+      "f = (Cp*(To3-To2))/(eff_cc*CV*10**3) \t\t\t\t#Fuel air ratio\n",
+      "mf = ma*f \t\t\t\t#Mass flow rate of fuel in kg/s\n",
+      "m = ma+mf \t\t\t\t#Mass flow rate of gas in kg/s\n",
+      "eff_i = (1/(1+((2/(k-1))*(1/M1**2))))*100 \t\t\t\t#Efficiency of the ideal cycle in %\n",
+      "sig = u1/C4 \t\t\t\t#Jet speed ratio \n",
+      "eff_p = ((2*sig)/(1+sig)) \t\t\t\t#Propulsive efficiency in %\n",
+      "F = ((m*C4)-(ma*u1))*10**-3 \t\t\t\t#Thrust in kN\n",
+      "\n",
+      "\t\t\t\t\n",
+      "#Output\n",
+      "print 'A)Efficiency of the ideal cycle is %3i percent \\\n",
+      "\\nB)Flight speed is %3.3f m/s \\\n",
+      "\\nC)Air flow rate is %3.3f kg/s \\\n",
+      "\\nD)Diffuser pressure ratio is %3.4f \\\n",
+      "\\nE)Fuel air ratio is %3.5f \\\n",
+      "\\nF)Nozzle pressure ratio is %3.2f \\\n",
+      "\\nG)Nozzle jet mach number is %3.3f \\\n",
+      "\\nH)Propulsive efficiency is %3.4f percent \\\n",
+      "\\nI)Thrust is %3.3f kN'%(eff_i,C4,ma,pr_d,f,pr_n,Me,eff_p,F)\n"
+     ],
+     "language": "python",
+     "metadata": {},
+     "outputs": [
+      {
+       "output_type": "stream",
+       "stream": "stdout",
+       "text": [
+        "A)Efficiency of the ideal cycle is  31 percent \n",
+        "B)Flight speed is 937.202 m/s \n",
+        "C)Air flow rate is 57.800 kg/s \n",
+        "D)Diffuser pressure ratio is 3.2875 \n",
+        "E)Fuel air ratio is 0.03188 \n",
+        "F)Nozzle pressure ratio is 3.22 \n",
+        "G)Nozzle jet mach number is 1.371 \n",
+        "H)Propulsive efficiency is 0.6696 percent \n",
+        "I)Thrust is 28.630 kN\n"
+       ]
+      }
+     ],
+     "prompt_number": 19
+    },
+    {
+     "cell_type": "heading",
+     "level": 2,
+     "metadata": {},
+     "source": [
+      "Example 6.17 page : 47"
+     ]
+    },
+    {
+     "cell_type": "code",
+     "collapsed": false,
+     "input": [
+      "import math \n",
+      "\n",
+      "\t\t\t\t\n",
+      "#Input data\n",
+      "ma = 18. \t\t\t\t#Mass flow rate of air in kg/s\n",
+      "Mi = 0.6 \t\t\t\t#Inlet mach number \n",
+      "h = 4600. \t\t\t\t#Altitude in m\n",
+      "Pi = 55. \t\t\t\t#Inlet pressure in \n",
+      "Ti = -20.+273 \t\t\t\t#Inlet temperature in K\n",
+      "eff_d = 0.9 \t\t\t\t#Diffuser efficiency \n",
+      "pr_d = 5. \t\t\t\t#Diffuser pressure ratio \n",
+      "T3 = 1000.+273 \t\t\t\t#Inlet turbine temperature in K\n",
+      "Pe = 60. \t\t\t\t#Exit pressure in kPa\n",
+      "eff_c = 0.81 \t\t\t\t#Compressor efficiency\n",
+      "eff_t = 0.85 \t\t\t\t#Turbine efficiency\n",
+      "eff_n = 0.915 \t\t\t\t#Nozzle efficiency\n",
+      "CV = 46520. \t\t\t\t#Calorific value in kJ/kg\n",
+      "Cp = 1005. \t\t\t\t#Specific heat capacity at consmath.tant pressure of air in J/kg-K\n",
+      "k = 1.4 \t\t\t\t#Adiabatic consmath.tant \n",
+      "R = 287. \t\t\t\t#Specific gas consmath.tant in J/kg-K\n",
+      "\n",
+      "\t\t\t\t\n",
+      "#Calculation\n",
+      "Ci = Mi*math.sqrt(k*R*Ti) \t\t\t\t#Velocity of air in m/s,\n",
+      "u = Ci \t\t\t\t#Flight velocity in m/s, Since it is reaction force of Ci\n",
+      "T1 = Ti+(Ci**2/(2*Cp)) \t\t\t\t#Temperature at inlet of compressor in K\n",
+      "P1s = Pi*(T1/Ti)**(k/(k-1)) \t\t\t\t#Inlet pressure of compressor at isentropic process in kPa\n",
+      "P1 = Pi+(eff_d*(P1s-Pi)) \t\t\t\t#Inlet pressure of compressor in kPa\n",
+      "P2 = P1*pr_d \t\t\t\t#Outlet pressure of compressor in kPa\n",
+      "T2s = T1*(pr_d)**((k-1)/k) \t\t\t\t#Outlet temperature of compressor at isentropic process in K\n",
+      "T2 = T1+((T2s-T1)/eff_c) \t\t\t\t#Exit temperature of compressor in K\n",
+      "Wc = Cp*(T2-T1)*10**-3 \t\t\t\t#Workdone on compressor in kJ/kg of air\n",
+      "Pc = ma*Wc \t\t\t\t#Power input in kW\n",
+      "Pt = Pc \t\t\t\t#Power out put of turbine for isentropic process in kW \n",
+      "f = (T3-T2)/((CV*10**3/Cp)-T3) \t\t\t\t#Fuel air ratio\n",
+      "Wt = Wc \t\t\t\t#Workdone by the turbine in kJ/kg of air\n",
+      "T4 = T3-(Wt*10**3/Cp) \t\t\t\t#Exit temperature of turbine in K\n",
+      "T4s = T3-((T3-T4)/eff_t) \t\t\t\t#Exit temperature of turbine at isentropic process in K\n",
+      "P3 = P2 \t\t\t\t#Exit pressure of combustion chamber in kPa, Since the process takes place at consmath.tant pressure process\n",
+      "P4 = P3*(T4s/T3)**(k/(k-1)) \t\t\t\t#Pressure at outlet of turbine in kPa\n",
+      "pr_n = P4/Pe \t\t\t\t#Pressure ratio of nozzle\n",
+      "Tes = T4/(pr_n)**((k-1)/k) \t\t\t\t#Exit temperature of nozzle  at isentropic process in K\n",
+      "Te = T4-(eff_n*(T4-Tes)) \t\t\t\t#Exit temperature of nozzle in K\n",
+      "Cj = math.sqrt(2*Cp*(T4-Te)) \t\t\t\t#Jet velocity in m/s\n",
+      "Ce = Cj \t\t\t\t#Flight velocity in m/s\n",
+      "ae = math.sqrt(k*R*Te) \t\t\t\t#Sound velocity at nozzle in m/s\n",
+      "Me = Ce/ae \t\t\t\t#Nozzle jet mach number\n",
+      "F = ma*(((1+f)*Cj)-u) \t\t\t\t#Thrust in N\n",
+      "P = F*u*10**-3 \t\t\t\t#Thrust power in kW\n",
+      "\n",
+      "\t\t\t\t\n",
+      "#Output\n",
+      "print 'A)Power input of compressor is %3.2f kW \\\n",
+      "\\nB)Power output of turbine is %3.2f kW \\\n",
+      "\\nC)F/A ratio on mass basis is %3.4f \\\n",
+      "\\nD)Exit mach number is %3.3f \\\n",
+      "\\nE)Thrust is %3.2f N \\\n",
+      "\\nF)Thrust power is %3.1f kW'%(Pc,Pt,f,Me,F,P)\n"
+     ],
+     "language": "python",
+     "metadata": {},
+     "outputs": [
+      {
+       "output_type": "stream",
+       "stream": "stdout",
+       "text": [
+        "A)Power input of compressor is 3536.17 kW \n",
+        "B)Power output of turbine is 3536.17 kW \n",
+        "C)F/A ratio on mass basis is 0.0179 \n",
+        "D)Exit mach number is 1.245 \n",
+        "E)Thrust is 9668.83 N \n",
+        "F)Thrust power is 1849.7 kW\n"
+       ]
+      }
+     ],
+     "prompt_number": 20
+    }
+   ],
+   "metadata": {}
+  }
+ ]
+}
\ No newline at end of file