GDJP Question bank

UNIT -1

1.     Air is moving at velocity of 150m/s, 100kpa at 25c. calculate Mach no. and stagnation property?

2.     Air flows at velocity of 250m/s ,the temperature of air measured at point is 30c .the air pressure is 5 bar .determine Mach no and stagnation property.

3.     A reservoir whose temperature can vary wide range of temperature, receive air at constant pressure of 150kpa the air is expanded isentropically is an nozzle. The exist pressure of 111.5kpa .i)determine (without using gas tables) value of temperature ,maintain in an reservoir to produce following velocity at nozzle efficiency 100m/s ii)250m/s

4.     A jet fighter is flying at Mach no 2.5 It is absorbed directly overhead at a distance of 10km.calculate the stagnation properties and speed of jet?

5.     Pressure, temperature, Mach no at entry of flow passage are 2.45 bar, 26.5 c and 1.4 respectively. Exist Mach no 2.5 determine for adiabatic (Assume isentropic ) of perfect gas (GAMMA=1.3) ,R=0.469kj/kgk. Find i) stagnation temperature ,pressure ii) temperature and velocity  at exist iii) flow rate per square meter at inlet cross section

6.     Pressure ,temperature ,velocity of air at entry a passage of 3bar,280k,140m/sm and pressure velocity at exit of flow passage are 2bar,260k,250m/s. Area of cross section at entry 600cm².Determine an adiabatic i) TO  Max velocity ii) Area of cross section at exit iii) Max. flow rate Given (gamma=1.4,R=287j/Kgk)

7.     Air is displaced from reservoir at pₒ= 6.91bar,Tₒ=325c .through nozzle at exit pressure of 0.98 bar .The mass flow rate 3600kg/Hr determine for an isentropic flow i)area ,pressure velocity  at throat ii) area and Mach no at exit iii)max possible velocity.

8.     A CONICAL DIFFUSER HAS ENTRY AND exit diameter as 0.15mand 0.3 m respectively. The pressure ,temperature, velocity of air at entry are 0.96bar,340k,185m/s respectively .Determine)exit pressure,velocity ,fore exerted on the diffuser wall .
 (R= 287j/kgk, γ=1.4,Cp=1.005kj/kgk)

9.     A stagnation pressure of 3.34 bar and temperature 627c is flowing through CV. The exit pressure 1.04 bar determine velocity, temperature, density at exit .Also determine the above if divergence act as diffuser. Assume Isentropic flow at exit.

10.         Air at Pi=3bar,Ti=500k,closedwith velocity of200m/s in 30cm diameter duct. Calculate i) mass flow rate ,ii)stagnation temperature iii) Mach no. iv) Pₒ(stagnation pressure)

11.         AN aircraft fly at 800km/hr, at altitudeof10,000 m .The air is reversibly compressed in inlet diffuser .The Mach no at exit of diffuser is 0.36 determine i) entry Mach no. and velocity ii) velocity ,pressure ,temperature at diffuser exit

12.         A supersonic will tunnel is design for M=3, at test section air supply form reservoir is 4bar and 26c .Determine mass flow rate, area of test section .Assume throat is 0.09m2. Gamma=1.4,alsofind temperature density for throat

13.         A supersonic wind tunnel setting chamber expands through a nozzle  from pressure 10 bar -4bar in test section calculate the stagnation temperature to be maintained in chamber to obtain to velocity of 500m/s in test section

14.         An aircraft is flying at altitude of 12,000m P=0.193bar ,T=216.65k at Mach no ;0.82.The cross section area of inlet diffuser is 0.5m2 .determine mass of air entering per sec, speed of aircraft ,stagnation pressure and temperature diffuser entry.

UNIT – 2    FANNO FLOW AND RAYLEIGH FLOW

RAYLEIGH FLOW

1)     Air having a M=3 with total temp of 295°C and static pressure 0.5 bar flows through a constant duct area adiabatically to another section where M₂ =1.5. Determine the amount of heat transfer & change in stagnation pressure.

2)    Air enters a constant duct area at M₁=3, P₁=1 atm and T₁=300K. Inside the duct the heat added/unit mass is 3*10⁵J/Kg. Calculate flow properties.

3)    The pressure, temp, Mach no of air in a combustion chamber is 4bar, 100°C, 0.2 respectively. The stagnation temp ratio is 3.Calculate

i)                   Mach no, temp, pressure at exit

ii)                Stagnation pressure loss

iii)              Heat supplied per Kg of air

4)    The Mach number at the exit of the combustion chamber is 0.9.The ratio of stagnation temp at the exit to the entry is 3.74.The pressure and temp of the gas at the exit are 2.5bar and 1270K.Determine

i)                   Mach no, pressure, temp of gas at entry

ii)                Heat supplied per Kg of gas

iii)              Max heat that can be supplied

5)    The pressure, temp, mach no of the gas at exit are 2bar, 1200K, 0.7.The ratio of stagnation temp is 3.85.Calculate the following

i)                   Mach no, pressure, temp at entry

ii)                Heat supplied per Kg of gas

iii)              Max heat that can be supplied

iv)              Identify whether it is heating or cooling.

6)    A combustion chamber gas turbine receives air at 350K, 0.55bar& 75m/s. The  air fuel ratio is 29.Calorific value of the fuel is 41.87 MJ/Kg. Taking ϒ=1.4, R=0.287 KJ/Kg for gas. Determine the initial and final mach no, final pressure temp & velocity of gas. Also find percentage of stagnation pressure loss and max attainable stagnation temp.

FANNO FLOW

7)    Air at 120KN/m ² and 40°C flows through 200mm diameter pipe adiabatically, the upstream mach no is 2.5.Determine the max length and properties at exit. Also estimate the length of pipe & exit mach no is 1.8.Take f=0.01

8)    Air at P₁=3.4bar, T₁=35°C enters a circular duct at a mach no of 0.14, exit mach no is 0.6, coefficient of fiction 0.004, mass flow rate 8.2 Kg/s. Determine

i)                   Pressure, temp at exit

ii)                Diameter of the duct

iii)              Length of the duct

iv)              Stagnation pressure loss

9)    A circular duct passes 8.25Kg/s of air , the exit mach no of 0.5.The entrance pressure and temp are 8.5bar and 38°C respectively, coefficient of friction is 0.005. If the mach no at entry is 0.15.Determine the diameter of the duct, length of the duct, Pressure and temp at exit & stagnation pressure loss

10)           Air if flowing into an insulator duct with velocity 150m/s, the pressure & temp of 28bar &280°C.Find the temp at section the duct. Finally where the pressure is 15.7bar the duct diameter is 15cm and friction factor is 0.005.Also find distance between two section

11)           Air at inlet temp of 60°C flows with subsonic velocity through an insulator pipe having inside diameter of 50mm, length of 5m.The pressure at the exit of pipe is 101KPa.The flow is choked at the end , the friction factor 4f=0.005.Determine inlet mach no, mass flow rate and exit temp

12)           The friction factor for 50mm pipe diameter is 0.005m.At the end of inlet velocity is 70m/sec. The temperature and pressure of 80°C and 10 bar respectively. Find temp, mach no, pressure at exit of pipe. The length of pipe is 25m long. Also find max possible length.

13)           Air is flowing in an insulator duct with the friction coefficient 0.002. At the inlet the velocity is 130m/sec. The temp is 400K and pressure is 25KPa. The diameter of the duct is 16cm.find

i)                   The length of the pipe required which gives 20% loss in stagnation pressure

ii)                Find the properties of air at the section 3.5m from the inlet.

iii)              Find max length.

UNIT-3

SHOCKS

NORMAL SHOCKS:

CONSTANT DUCT AREA:

1.     The state of gas (ᵞ=1.3 & R=0.469 kj/kgk) upstream of a normal shock wave is given by the following data. M_x=2.5,P_x=2 bar and T_x=275 k. Calculate the Mach Number, Pressure, Temperature of gas downstream of the shock.

2.     Air flows adiabatically in a pipe, a normal shockwave is formed. The pressure and temperature of air before shock are 150 kN/m² and 25^oC. The is just after the normal shock is 350 kN/m². Calculate the following
a). Mach Number before shock
b). Mach Number, static and velocity of air after the shock wave
c). Increase in density of air
d). Loss of Stagnation Pressure
e). Change in entropy

3.     A pitot tube kept in a supersonic wind tunnel normally weather forms a low shock ahead of A. The static pressure upstream of shock is 16 kPa and pressure at mouth is 70 kPa. Estimate the Mach Number of tunnel with the stagnation pressure is 300^oC and calculate static temperature and total pressure of upstream and downstream.      

VARIABLE CROSS SECTION:

1.     The C-D nozzle is designed to expand an air from reservoir pressure is 700kPa and temperature is 5^oC. The Inlet Mach Number is 2. The nozzle throat area is 230cm².  A normal shock appears at a section where the area is 175cm². Find the exit pressure and temperature also find increase in entropy across the shock.

2.     When C-D nozzle is operated at off designed condition.  The normal shocks appears at section where cross section area is 18.75cm² in the diverging port and at inlet to the nozzle.  The stagnation state is given as 0.21 MPa and 36^oC.  The throat area is 12.5cm² and the exit area is 25cm².  Estimate the exit Mach Number, pressure, and loss in stagnation pressure for flow through nozzle.

3.     A C-D air nozzle has exit to throat area ratio is 3.  A normal shock appears at the divergent section where the existing area ratio is 2.2.  Find the mach number before and after shock.  The inlet stagnation properties are 500 kPa and 450K.  Find the properties air at exit and entropy increase across the shock.

4.     A C-D nozzle has an exit area to a throat area ratio is 2.5.  The total properties of air at inlet of 7 bar and 87^oC.  The throat area is 65cm².  Determine mach number, pressure, temperature and stagnation pressure at exit when a plain normal shock stands at a point where the mach number is 2.  Assume isentropic flow before and after shock.

5.     A C-D nozzle is designed to expand a from a reservoir in which the pressure is 800 kPa and temperature is 40^oC to give a mach number at exit of 2.5.  The throat area is 25cm².  Find 
a). Mass flow rate
b). Exit Area
c). When a normal shock appears at a section where the area is 40cm².  Determine pressure and temperature at exit.

OBLIQUE SHOCKS:

1.     An oblique shock wave occurs at the leading edge of a symmetric wedge.  Air has a mach number of 2.1 and deflection angle of 15^o.  Determine the following for strong and weak waves.
a). Wave angle
b). Density Ratio
c). Pressure Ratio
d). Temperature Ratio

2.     A jet of air approaches schematically mach number of 2.4 and  wave angle of 60^o.  Determine the following
a). Deflection angle
b). Pressure Ratio
c). Temperature Ratio

d). Final mach number

3.     An oblique shockwave at an angle of 30^o occurs at a leading edge of symmetrical wedge.  Air has mach number of 2.1 upstream and upstream temperature and pressure is 300K and 11 bar respectively. Determine the following. 
a). Downstream pressure
b). Downstream temperature, mach number
c). Wedge Angle.

4.     An air jet at mach number of 2.1 is isentropically deflected by 10^o in clockwise direction.  The initial pressure is 100 kN/m² and T_x=18^oC.  Determine final state of air.

FLOW THROUGH SUBSONIC DIFFUSER:

1.     A mach 2 air craft engine employs a subsonic inlet diffuser of area ratio 3.  The normal shock is formed just upstream of diffuser inlet.  The free stream conditions of upstream diffuser are P_o=0.1 bar and T_o=300K.  Determine
a). Mach number, pressure and temperature at the diffuser exit
b). Assume isentropic flow in diffuser downstream of shock

2.     An air plane having a diffuser design for subsonic fly has a normal shock attached to the edge of diffuser.  When the plane is flying at certain Mach number.  If at exit of diffuser the mach number is 0.3.  What must be the flight Mach number assume isentropic diffusion behind the shock. Area at inlet is 0.29m² and exit is 0.44m².

Unit 4

1.     The flight speed of a turbo jet is 800 km/hr at 10000m altitude, the density of the air at the altitude is 0.71 kg/m³. The drag force of the plane is 6.8 KN. The propulsive efficiency is 60%. Calculate the 1) SFC 2) air fuel ratio 3) Jet velocity. Assume the calorific value of fuel is 45000 Kj/Kg and the overall efficiency is 18%.

2.     A turbo jet has a speed of 750 Km/hr, while flying at an altitude of 10000m. The propulsive efficiency of the jet is 50%. The overall efficiency is 16%. The density of air at 10000m altitude is 0.176 Kg/m3. The drag in the plane Is 6250 N. The calorific value of the fuel is 48000 Kj/Kg. Find 1) absolute velocity of jet 2) power output of the unit in KW 3) Diameter of Jet.

3.     Diameter of air craft propeller is 4m. The speed ratio is 0.8 at the flight speed of 450 Km/hr, the ambient condition of the air at flight altitude are T= 256 K and P= 0.54 bar. Determine the propulsive efficiency, thrust and thrust power.

4.     The diameter of the propeller of the air craft is 2.5 m. If flies at the speed of 500 Km/hr at altitude of 8000m. For a flight to jet speed ratio of 0.75. Determine the 1) mass flow rate of air to the propeller 2) thrust produced 3) specific thrust 4) specific impulse 5) thrust power.

5.     A turbo jet propels an air craft at a speed of 900Km/hr while taking 3000Kg of air per minute. The isentropic enthalpy drop is 200 Kj/Kg and the nozzle efficiency is 90%. The fuel ratio is 85 and the combustion efficiency is 95%. The calorific value of fuel is 42000 Kj/Kg. calculate the propulsive power, thrust power, thermal efficiency, propulsive efficiency and overall efficiency.

6.     An air craft flies at 960 Km/hr. One of the turbo jet engine takes 40Kg/s of air and expands it with ambient pressure. The air fuel ratio is 50, the lower calorific value is 43 Mj/Kg for maximum thrust power. Determine 1) jet velocity 2) thrust 3) specific thrust 4) thrust power 5) TSFC 6) Propulsive, thermal and overall efficiency.

7.     An air craft flies at 1000 Km/hr. One of the turbo jet engine takes 40Kg/s of air and expands it with ambient pressure. The air fuel ratio is 50, the lower calorific value is 43 Mj/Kg for maximum thrust power. Determine 1) jet velocity 2) thrust 3) specific thrust 4) thrust power 5) TSFC 6) Propulsive, thermal and overall efficiency.

8.     The turbo engine takes 50 Kg/s of air and propels the air craft with uniform flight speed of 880 Km/hr, the isentropic enthalpy change in the nozzle is 188 Kj/Kg. The velocity coefficient is 0.96. The fuel to air ratio is 1.2%, combustion efficiency is 95%. Calorific value of fuel is 44000 Kj/Kg. Find 1) thermal efficiency 2) fuel flow in Kg/hr 3) propulsion efficiency 4) overall efficiency.

9.     A turbo jet air craft at 875 Km/hr at an altitude of 10000m above mean sea level. Calculate 1) air flow rate to engine 2) thrust 3) specific thrust 4) impulse thrust 5) thrust power 6) TSFC. Take the following data, diameter of air inlet section= 0.75m, diameter of jet pipe at exit= 0.5m, velocity of gases at exit of jet pipe= 500 m/s, pressure at exit of the pipe is 0.3 bar, air to fuel ratio = 40.