SOLUTIONS MANUAL FOR
Aircraft Propulsion
and
Gas Turbine Engines
by
Ahmed El-Sayed
86898.indd 1 8/14/08 10:25:48
, CHAPTER 4
TURBOJET ENGINES
1- The following data refer to the operating conditions for an aircraft turbojet engine:
Barometric pressure 101 kPa
Aircraft speed 500 km/h
Compressor air flow 45 kg/s
Exhaust nozzle area 0.2 m²
Exhaust nozzle pressure 200 kPa
Exhaust gas velocity 450 m/s
Fuel flow 0.8 kg/s
a. Calculate the ram drag, momentum thrust, pressure thrust,
gross thrust, net thrust and specific thrust
b. If the aircraft speed is increased to 600, 700, 800 and 900
km/h, while all other data are, maintained constant, plot the
ram drag, gross thrust and net thrust versus the aircraft
speed
c. If the mass flow rate is proportional to the aircraft speed,
and the speed also changes to be 600, 700, 800, and 900
km/h. calculate the new value of the air mass flow and plot
again the ram drag, gross thrust and net thrust versus the
aircraft speed.
,2- Calculate the specific thrust and the thrust specific fuel consumption of a single spool
turbojet engine having the following peculiarities:
Cruise velocity of 280 m/s at altitude of 7000 m
Intake efficiency of 93%
Compressor pressure ratio of 8:1 and efficiency of 87%
Burner efficiency of 98%
Pressure drop in the combustion chamber of 4 % of the delivery pressure of the
compressor
Turbine inlet temperature of 1200 K and efficiency of 90%
Mechanical efficiency of 99%
Nozzle efficiency of 95%
Fuel heating value 44000 kJ/kg
3- Derive an expression for the ratio between the thrust forces of a turbojet engine fitted
with an afterburner when the afterburner is operative ( Tab ) and inoperative ( T ) with
choked nozzle similar to that in equation (4.33)
4- A single spool turbojet engine powers an aircraft flying at Mach number of 0.5, at an
altitude where Ta =226.7k and Pa =28.7kPa. It has the following data
Air mass flow rate 20 kg/s
Compressor pressure ratio 8
Maximum cycle temperature 1200 K
, Fuel heating value QR 43000kJ/kg
All the components of engine are ideal
Calculate the thrust force, fuel to air ratio, and thrust specific fuel consumption
(TSFC)
5- A simple turbojet engine with a compressor pressure ratio of 8, a turbine inlet
temperature 1200 K and a mass flow of 15 kg/s, when the aircraft is flying at 260 m/s
at an altitude of 7000 m. Assuming the following components efficiencies and I.S.A
conditions, calculate the propelling nozzle area required, the net thrust developed and
the TSFC.
Isentropic efficiency of intake 0.9
Isentropic efficiency of compressor 0.88
Isentropic efficiency of turbine 0.92
Isentropic efficiency of propelling nozzle 0.95
Combustion chamber pressure loss is 6% of compressor delivery pressure
Combustion chamber efficiency 0.98
If the gases in the jet pipe are reheated to 2000 K and the combustion chamber
pressure loss is 3% of the pressure at outlet from turbine, calculate:
The percentage increase in nozzle area required the mass flow is to be unchanged.
The percentage increase in net thrust
( Ta = 242.7 K , Pa = 41.06kPa, QR = 43000 kJ / kg )
6- A turbojet engine is installed to a fighter airplane flying at Mach number of 1.3. The
ambient conditions are 236.21 K and 35.64 kPa. The exhaust conditions are 1800 K
Aircraft Propulsion
and
Gas Turbine Engines
by
Ahmed El-Sayed
86898.indd 1 8/14/08 10:25:48
, CHAPTER 4
TURBOJET ENGINES
1- The following data refer to the operating conditions for an aircraft turbojet engine:
Barometric pressure 101 kPa
Aircraft speed 500 km/h
Compressor air flow 45 kg/s
Exhaust nozzle area 0.2 m²
Exhaust nozzle pressure 200 kPa
Exhaust gas velocity 450 m/s
Fuel flow 0.8 kg/s
a. Calculate the ram drag, momentum thrust, pressure thrust,
gross thrust, net thrust and specific thrust
b. If the aircraft speed is increased to 600, 700, 800 and 900
km/h, while all other data are, maintained constant, plot the
ram drag, gross thrust and net thrust versus the aircraft
speed
c. If the mass flow rate is proportional to the aircraft speed,
and the speed also changes to be 600, 700, 800, and 900
km/h. calculate the new value of the air mass flow and plot
again the ram drag, gross thrust and net thrust versus the
aircraft speed.
,2- Calculate the specific thrust and the thrust specific fuel consumption of a single spool
turbojet engine having the following peculiarities:
Cruise velocity of 280 m/s at altitude of 7000 m
Intake efficiency of 93%
Compressor pressure ratio of 8:1 and efficiency of 87%
Burner efficiency of 98%
Pressure drop in the combustion chamber of 4 % of the delivery pressure of the
compressor
Turbine inlet temperature of 1200 K and efficiency of 90%
Mechanical efficiency of 99%
Nozzle efficiency of 95%
Fuel heating value 44000 kJ/kg
3- Derive an expression for the ratio between the thrust forces of a turbojet engine fitted
with an afterburner when the afterburner is operative ( Tab ) and inoperative ( T ) with
choked nozzle similar to that in equation (4.33)
4- A single spool turbojet engine powers an aircraft flying at Mach number of 0.5, at an
altitude where Ta =226.7k and Pa =28.7kPa. It has the following data
Air mass flow rate 20 kg/s
Compressor pressure ratio 8
Maximum cycle temperature 1200 K
, Fuel heating value QR 43000kJ/kg
All the components of engine are ideal
Calculate the thrust force, fuel to air ratio, and thrust specific fuel consumption
(TSFC)
5- A simple turbojet engine with a compressor pressure ratio of 8, a turbine inlet
temperature 1200 K and a mass flow of 15 kg/s, when the aircraft is flying at 260 m/s
at an altitude of 7000 m. Assuming the following components efficiencies and I.S.A
conditions, calculate the propelling nozzle area required, the net thrust developed and
the TSFC.
Isentropic efficiency of intake 0.9
Isentropic efficiency of compressor 0.88
Isentropic efficiency of turbine 0.92
Isentropic efficiency of propelling nozzle 0.95
Combustion chamber pressure loss is 6% of compressor delivery pressure
Combustion chamber efficiency 0.98
If the gases in the jet pipe are reheated to 2000 K and the combustion chamber
pressure loss is 3% of the pressure at outlet from turbine, calculate:
The percentage increase in nozzle area required the mass flow is to be unchanged.
The percentage increase in net thrust
( Ta = 242.7 K , Pa = 41.06kPa, QR = 43000 kJ / kg )
6- A turbojet engine is installed to a fighter airplane flying at Mach number of 1.3. The
ambient conditions are 236.21 K and 35.64 kPa. The exhaust conditions are 1800 K
