Gas Dynamics and Jet Propulsion - 2 Marks - All 5 Units

080120037 - Gas Dynamics and Jet Propulsion

P.A. COLLEGE OF ENGINEERING AND TECHNOLOGY

PALLADAM ROAD, POLLACHI - 642 002

DEPARTMENT OF MECHANICAL ENGINEERING

GAS DYNAMICS AND JET PROPULSION

TWO MARK QUESTIONS AND ANSWERS

ACADEMIC YEAR 2012 - 2013

Prepared By

Prof. C. Sowmya Dhanalakshmi. M.E., MISTE, (Phd)

P. A. College of Engineering and Technology, Mechanical Department 1


UNIT - 1 COMPRESSIBLE FLOW - FUNDAMAENTALS

PART-A

1. What is the basic difference between compressible and incompressible fluid flow?

Compressible flow In Compressible flow1. Fluid velocities are appreciable compared with the velocity of sound

1. Fluid velocities are small compared with the velocity of sound

2. Density is not constant 2. Density is constant1. Compressibility factor is greater than

one3. Compressibility factor is one

2. Write the steady flow energy equation for an adiabatic flow of air.In an adiabatic flow q = 0. Therefore energy equation becomes, h1 + c1

2/2 + gZ1 = h2 + c22/2 + gZ2 + W1

Adiabatic energy equation is h0 = h + 1/2c22

3. Explain the meaning of stagnation state with example.The state of fluid attained by isentropically decelerating it to zero velocity at zero elevation is referred as stagnation state.E.g. Fluid in a reservoir or in a settling chamber

4. Distinguish between static and stagnation pressures.In stagnation pressure state the velocity of the flowing fluid is zero whereas in the static pressure, the fluid velocity is not equal to zero

5. Differentiate between the static and stagnation temperatures.The actual temperature of the fluid in a particular state is known as static temperature whereas the temperature of the fluid when the fluid velocity is zero at zero elevation known as stagnation temperatureTo = T+c2/2Cp

6. What is the use of Mach number?Mach number is defined as the ratio between the local fluid velocity to the velocity of sound. Mach number M=c/a. It is used for the analysis of compressible fluid flow problems. Critical mach number is a dimensionless number at which fluid velocity is equal to its sound velocity. Mcritical = (c/a) – 1

7. What is Crocco number?It is a non-dimensional fluid velocity which is defined as the ratio of fluid velocity to its maximum fluid velocity, Cr=c/cmax

8. Write down the relationship between stagnation and static temperature interms of the flow, mach number for the case of isentropic flow.T0/T = [1+ (γ -1)/2] M2

9. Give expression of P/P0 for an isentropic flow through a duct.The expression is P/P0 = 1/{[1+ (γ -1)/2] M2} γ-1

10. Name the velocities that are used in expressing the fluid velocities in non-dimensional form.



Local velocity of sound, stagnation velocity of sound, Maximum velocity of sound, critical velocity of sound

11. What are the different regions of compressible flow?Incompressible regionSubsonic regionTransonic regionSupersonic regionHypersonic region

12. Define M* and give the relation between M and M*

It is a non-dimensional mach number and is defined by the ratio between the local fluid velocity to its critical velocity of sound, M* = c/a*

13. A plane travels at a speed of 2400Km/hr in an atmosphere of 5 degree, find the Mach angle?C=2400/3.6 = 666.67T=278KM=c/√ γRT=1.9947α=sin-1(1/M) = 30.0876°

14. Define Mach angle and Mach wedge.Mach angle is formed when an object is moving with supersonic speed. The wave propagation and changes are smooth. When an object is moving with hypersonic speed the changes are abrupt is shown in figure. Hence for a supersonic flow over two-dimensional object “mach wedge” is used instead of “mach cone”.

15. What is meant by isentropic flow with variable area?A steady one dimensional isentropic flow in a variable area passages is called “variable area flow”. The heat transfer is negligible and there are no other irreversibilities due to fluid friction.

16. Define Mach cone.Tangents drawn from the source point on the spheres define a conical surface referred to as Mach cone.

17. What is characteristic Mach number?M* = [M2(γ-1)/2+ M2(γ-1)]1/2

18. If an aeroplane goes to higher altitudes maintaining the same speed what will happen to the Mach number?At higher altitude the sound velocity ‘a’ will decrease and hence M will increase. Therefore, M is not a constant.

19. Define open and closed system.In open system both mass and energy transfer takes place. But in closed system mass transfer does not occur; only energy transfer takes place.

20. What is the difference between intensive and extensive properties?Intensive properties: These are independent on the mass of the system. Ex:



Pressure and TemperatureExtensive properties: These are dependent on the mass of the system. Ex: Total volume, Total energy

21. Distinguish between Mach wave and normal shock?Mach wave: The lines at which the pressure difference is concentrated and which generate cone are called mach lines or mach wavesNormal shock: A shock wave is nothing but a steep finite pressure wave. When the shock wave is right angle to the flow, it is called normal shock

22. Define zone action and zone of silence.The region inside the Mach cone is called the zone of action an the region outside the Mach cone is termed as the zone of silence.

23. Define adiabatic process.In an adiabatic process there is no heat transfer between the system and the surrounding, Q=0

24. What is meant by transonic flow?If the fluid velocity is close to the speed of sound that type of flow is called as transonic flow. Mach number is between 0.8 and 1.2

25. What is meant by hypersonic flow?In hypersonic flow, fluid velocity is much greater than sound velocity. Mach number is always greater than 5

26. What is the difference between nozzle and diffuser?Nozzle is a device which increases the velocity and decreases the pressure of working substance. Diffuser is a device which increases the pressure and decreases the velocity of the working substance.

PART-B

1. An air jet at 300 K has sonic velocity. Determine the following:Velocity of sound at 300 K, Velocity of sound at stagnation conditions, Maximum velocity of jet, Stagnation enthalpy and Crocco number. Take γ = 1.4, R=287 J/kgK

ANSWER: Page number 1.42 (2)

2. Derive an expression for the energy equation.

ANSWER: Page number 1.14(section 1.16)

3. The pressure, temperature and fluid velocity of air at the entry of a flow passage are 3 bar, 280 K and 140 m/s. The pressure, temperature and velocity at the exit of a low passage are 2 bar, 260K and 250 m/s. The area of cross section at entry is 600 cm2. Determine for an adiabatic flow, the stagnation temperature, maximum velocity, mas flow rate and area of cross section at exit. Take γ = 1.4, R=287 J/kgK

[ANSWER: Page number 1.56 (6)



4. A gas flows in a duct of 40 cm diameter at inlet pressure of 3 bar, temperature of 450 K and a velocity of 180 m/s. Calculate at inlet, the mass flow rate, stagnation temperature, Mach number and stagnation pressure values assuming the flow as compressible and incompressible.


5. An aircraft is flying at an altitude of 11,000 meters, at 800 km/hr. The air is reversibly compressed in an inlet diffuser. The inlet temperature is 216.65 K and pressure is 0.226 bar. If the Mach number at the exit of the diffuser is 0.35, calculate the following. Entry Mach number, velocity, pressure and temperature of air at the diffuser exit.


6. An aircraft is flying at an altitude of 10,000 meters. The inlet Mach number is 0.82, temperature is 223.15 K and pressure is 0.246 bar. The cross sectional area of the inlet diffuser before the low pressure compressor stage is 0.45 m2. Calculate the following: The mass of air entering the compressor per second, the speed of the air craft and stagnation pressure at diffuser entry and stagnation temperature at diffuser entry.


7. Argon is stored in a reservoir at 280 K. Determine stagnation enthalpy and stagnation velocity of sound for γ = 1.65 and the molecular weight of argon is 39.94, if the argon at a temperature of 150 K flowing at a velocity of 300 m/s, find the Mach number and Mach angle.


8. Air (γ = 1.4, R=287 J/kgK) at an inlet mach number of 0.2 enters a straight duct at 400K and expands isentropically. If the exit Mach number is 0.8, determine the following: Stagnation temperature, critical temperature, static temperature at exit and area ratio A1/A2


9. The pressure, temperature and Mach number at the entry of a flow passage are 2 bar, 275 K and 1.3 respectively. If the exit Mach number is 2.4, determine the velocity of sound at stagnation condition, the maximum velocity, the temperature and pressure at exit and Mach number M1

* M1* and M2

*

Take γ = 1.3, R=0.460 kJ/kgKANSWER: Page number 1.85 (13)

10. In a settling chamber air is maintained at a temperature of 400 K and a pressure of 6 bar. Calculate the following: Stagnation enthalpy, stagnation velocity of sound, maximum velocity, critical velocity of fluid and critical velocity of sound


11. The air moving at a velocity of 150 m/s. The static conditions are 100 kPa and 25 ˚C. Calculate the Mach number and stagnation properties verify the values with table values.


12. An aircraft flies at a velocity of 700 kmph in an atmosphere where the temperature is 75



kPa and temperature is 5˚C. Calculate the Mach number and stagnation properties.ANSWER: Page number 1.97 (4)

13. A steam of air flows with a velocity of 250 m/s in a duct of 10 cm diameter. Its temperature and pressure at that point are 5˚C and 40 kPa. What will be its stagnation pressure and temperature? What is the mass flow rate?


14. The following data refers to the entry and exit of a passage where isentropic flow occurs: Entry:p1 = 207 kPa, T1 = 300 K, M1=1.4 Exit: M2 = 2.5, Assuming ideal gas, determine velocity of sound at stagnation condition, maximum velocity and temperature and pressure at exit.


15. The pressure, temperature and Mach number at th entry of a flow passage are 2.45 bar, 26.5˚C and 1.4 respectiively. If the exit Mach number is 2.5, determine the stagnation temperature, temperature and velocity of a gas at exit and the flow rate per square metre of the inlet cross section for adiabatic flow of a perfect gas (γ = 1.3, R=0.460 kJ/kgK).


16. Air (γ = 1.4, R=287 J/kgK) enters a straight axis symmetric duct at 300K, 3.45 bar and 150 m/s and leaves it at 277 K, 2.058 bar and 260 m/s. The area of cross section at entry is 50 cm2. Assuming adiabatic flow determine stagnation temperature, maximum velocity, mass flow rate and area of cross section at exit.


UNIT - 2 FLOW THROUGH VARIABLE AREA DUCTS

PART-A

1. Differentiate Adiabatic and Isentropic process.Adiabatic process:In a process there is no heat transfer from the fluid to surroundings or from the surroundings to the fluid.Isentropic process:In a isentropic entropy remains constant and it is reversible .During this process there is no heat transfer from the fluid to surroundings or from the surroundings to the fluid. Therefore an isentropic process can be stated as reversible adiabatic process.



S = constant, Q =0

2. Differentiate nozzle and diffuser ?Nozzle:It is a device which is used to increase the velocity and decrease the pressure of fluids.Diffuser:It is a device which is used to increase the pressure and decrease the velocity of fluids.

3. What is Impulse function ?The sum of pressure force ( pA ) and impulse force ( þAc² ) gives Impulse function (F)F = pA + þac²

4. Differentiate between adiabatic flow and diabatic flow ?Diabatic flow :Flow in a constant area duct with heat transfer and without friction is known as diabatic flow (Rayleigh flow)Adiabatic flow:Flow in a constant area duct with friction and without heat transfer is known as adiabatic flow (Fanno flow).

5. State the expression for dA/A as a function of Mach number ?dA/A =dp/þc² [ 1-M² ]

6. Give the expression for T/To and T/T* for isentropic flow through variable area interms of Mach number ?To/T =1+[_-1/2]M²To/T = 1

7. Draw the variation of Mach number along the length of a convergent divergent duct when it acts as a (a) Nozzle (b) Diffuser (c) Venturi

8. What is chocked flow through a nozzle?The mass flow rate of nozzle is increased by decreasing the back pressure. The maximum mass flow conditions are reached when the throat pressure ratio achieves critical value. After that there is no further increase in mass flow with decrease in back pressure .This condition is called chocking. At chocking condition M=1.

9. What type of nozzle used for sonic flow and supersonic flow?Constant area duct nozzle is used for sonic flow and divergent nozzle is used for supersonic flow.

10. When does the maximum mass flow occur for an isentropic flow with variable area?Mass flow rate will be maximum at throat section where the Mach number is one.

11. Give the expression for To/T and T/Y* for isentropic flow through variable area in terms of Mach number T/To = 1/[(1+(γ-1)/2)/M2] T/T* = (γ-1)/ [(1+(γ-1)/2)/M2]

12. Sketch the isentropic and adiabatic expansion process in P-V and T-S diagram.



13. Represent the adiabatic flow through a diffuser on T-S diagram. Label the different states, the initial and final points.

14. What will happen if the air flowing through a nozzle is heated?When the flowing air is heated in a nozzle, the following changes like increase inair velocity, increase in temperature and enthalpy, increase in pressure and increase in entropy will occur.

15. Write the Fliegner’s formula.Mmax/A* x √To/Po = 0.0404

16. Write the equation for efficiency of the diffuser.Diffuser efficiency = static pressure rise in actual process/ static pressure rise in ideal processP2-P1/P2’-P1

17. What is impulse function and give its uses?Impulse function is defined as the sum of pressure force and inertia force. Impulse function F=Pressure force ρA + inertia force ρAc2. Since the unit of both the quantities are same as unit of force, it is very convenient for solving jet propulsion problems. The thrust exerted by the flowing fluid between two sectons can be obtained by using change in impulse function.

18. What is chocked flow? When the back pressure is reduced in a nozzle, the mass flow rate will increase. The maximum mass flow conditions are reached when the back pressure is equal to the critical pressure. When the back pressure is reduced further, the mass flow rate will not change and is constant. The condition of flow is called ‘chocked flow’.

19. State the necessary conditions for chocked flow to occur in a nozzle. The necessary conditions for this flow to occur in a nozzle is the nozzle exit pressure ratio must be equal to the critical pressure ratio where the mach number M=1.

20. Give the difference between nozzle and venture.

Nozzle VenturiThe flow is accelerated continuously.(mach number and velocity increases continuously)

The flow is accelerated upto M=1 and then mach number is decreased

Used to increase velocity and mach number

Used for flow measurement (discharges)

Generally convergent portion is short Convergent and divergent portions are equal

21. What is normal shock?When the shock waves are right angles to the direction of flow and the rise in pressure is abrupt are called normal shock waves.



22. What is meant by normal shock as applied to compressible flow?Compression wave front being normal to the direction of compressible fluid flow. It occurs when the flow is decelerating from supersonic flow. The fluid properties jump across the normal shock.

23. Shock waves cannot develop in subsonic flow? State the reason.Shocks are introduced to increase the pressure and hence it is a deceleration process. Shocks are possible only when the fluid velocity is maximum.

24. Define strength of a shock wave.Strength of a shock wave is defined as the ratio of increase in static pressure across the shock to the inlet static pressure.Strength of shock = (Py – Px)/Px

25. Calculate the strength of shock wave when normal shock appears at M=2.M=2, γ=1.4, Py/Px = 4.5Strength of shock = 3.5/4.5

26. Draw the shape of the nozzle for the expansion of air from 1 Mpa to 700 kPa.

PART-B

17. 1. Air is discharged from a reservoir at Po =6.91bar and To =325°c through a nozzle to an exit pressure of 0.98 bar .If the flow rate is 3600Kg/hr determine for isentropic flow:1) Throat area, pressure,and velocity, 2) Exit area,Mach number3) Maximum velocity.


2. A conical diffuser has entry and exit diameters of 15 cm and 30cm respectively . The pressure ,temperature and velocity of air at entry are 0.69bar,340 k and 180 m/s respectively . Determine1) The exit pressure2) The exit velocity3) The force exerted on the diffuser walls.Assume isentropic flow,_ γ=1.4,Cp =1.00 KJ Kg-K..


3. An air nozzle is to be designed for an exit Mach number of 3.5. the stagnation conditions for the isentropic flow are 800 kPa and 240˚C. Estimate pressure, temperature, velocity and area at throat and exit for a mas flow rate of 3.5 kg/s.


4. A diffuser has exit to throat area ratio of 1.5 to 1. The inlet Mach number is 0.8 The initial pressure and temperature are 1 bar and 15˚C. Assuming the flow to be isentropic,



calculate the exit pressure, temperature and Mach number for air.ANSWER: Page number 2.94 (3)

5. A supersonic diffuser, diffuses air in an isentropic flow from a Mach number of 3 to a Mach number of 1.5. the static conditions of air at inlet are 70 kPa and -7˚C. If the mass flow rate of air is 125 kg/s, determine stagnation conditions, area at the throat and exit and static conditions of air at exit.


6. An air enters an isentropic diffuser with a mach number of 3.6 and is decelerated to a mch number of 2. The diffuser passes a flow of 15kg/s. The initial static pressure and temperature of the air are 1.05 bar and 40˚C. Assuming γ=1.4, calculate the inlet area, total pressure and total temperature at inlet, exit area, total pressure, total temperature and static pressure.


7. A thrust chamber pressure of a rocket nozzle is 350 bar and the nozzle throat section area is 6 cm2. If the mach number at the nozzle exit is 5.2, calculate the thrust developed by the rocket.


8. Air enters the nozzle from a large reservoir at 7 bar and 320˚C. the exit pressure of the nozzle is 0.94 bar and mass flow rate is 3500 kg/h. Calculate the following for isentropic flow. Throat area, throat pressure, throat velocity, exit area, exit mach number, maximum velocity


9. The pressure, temperature and velocity of air at the entry of a diffuser are 0.7 bar , 345 K and 190 m/s respectively. The entry diameter of diffuser is 15 cm and exit diameter is 35 cm. Determine the following. Exit pressure, exit velocity and force exerted on the diffuser walls. Assuming isentropic flow and take γ=1.4, cp=1005 J/kgK


10. A supersonic wind tunnel settling chamber expands air or Freon-21 through a nozzle from a nozzle from a pressure of 10 bar to 4bar in the test section . calculate the stagnation temperature to the maintained in the setting chamber to obtain a velocity of 500 m/s in the test section for,1) Air ,Cp =1.025 KJ/Kg K, Cv =0.735 KJ/Kg K2) Freon -21 ,Cp =0.785 KJ/Kg K ,Cv= 0.675 KJ/Kg K.What is the test section Mach number is each case?




UNIT - 3 FANNO AND RAYLEIGH FLOW

PART-A

1. What are the consumption made for fanno flow?One dimensional steady flow.Flow takes place in constant sectional area.There is no heat transferThe gas is perfect with constant specific heats.

2. Differentiate Fanno flow and Rayleigh flow?Rayleigh flow:Flow in a constant area duct with heat transfer and without friction is known as Rayleighs flow.Fanno Flow:Flow in a constant area duct with friction and without heat transfer is known as Fanno flow.

3. Explain chocking in Fanno flow?



In a fanno flow, subsonic flow region, the effect of friction will increase the velocity and Mach number and to decrease the enthalpy and pressure of the gas.In supersonic flow region, the effect of friction will decrease the velocity and Mach number and to increase the enthalpy and pressure of the gas.In both cases entropy increases up to limiting state where the Mach number is one(M=1) and it is constant afterwards. At this point flow is said to be chocked flow.

4. Explain the difference between Fanno flow and Isothermal flow?

Fanno Flow Isothermal FlowFlow in a constant areaduct with friction andwithout heat transfer isknown as fanno flow.

Flow in a constant area duct withfriction and the heat transfer isknown as isothermal flow.

Static temperature is notconstant

Static temperature remainsconstant

5. Write down the ratio of velocities between any two sections in terms of their Mach number in a fanno flow ?[1+[_-1/2] M1²]½C2/C1=M1/M2[1+[_-1/2] M2²]½

6. Write down the ratio of density between any two section in terms of their Mach number in a fanno flow?Þ2/ Þ1= M1/M2 [1+ [_-1/2] M1²]½[1+ [_-1/2] M1²]½

7. What are the three equation governing Fanno flow?Energy equation, continuity equation and equation of state.

8. Give the expression to find increase in entropy for Fanno flow?S2-S1R =_n M1/M2 [1+ [_-1/2] M1²]_ +1/2(_ -1)[1+ [_-1/2] M1²]_ +1/2(_ -1)

9. Give two practical examples where the Fanno flow occurs?Flow in air breathing enginesFlow in refrigeration and air conditioningFlow of fluids in long pipes.

10. What is Rayleigh line and Fanno line?Rayleigh line: Flow in a constant duct area with heat transfer and without friction is described by a curve is known as Rayleigh line.Fanno Line: Flow in a constant duct area without heat transfer and with friction is described by a curve is Fanno line

11. What are the assumptions of Fanno flow?One dimensional steady flowFlow takes place in constant sectional area



There is no heat transferThe gas is perfect with constant specific heats

12. Write down expression to find increase in entropy for Fanno flow.(S2-s1)/R = ln M1/M2 [

13. Define fanning’s coefficient of skin frictionIt is the ratio between wall shear stress and dynamic headF = wall shear stress/dynamic head

14. Define oblique shock. Also mention where it occurs.The shock wave which is inclined at an angle to the two dimensional flow direction is called as oblique shock. When the flow is supersonic, the oblique shock occurs at the corner due to the turning of supersonic flow.

15. Define Fanno line.The locus of the state which satisfy the continuity and energy equation for a frictional flow is known as fanno line.

16. Define isothermal flow with friction.A steady one dimensional flow with friction and heat transfer in a constant area duct is called isothermal flow with friction.

17. Give the applications of isothermal flow with friction.In long ducts where sufficient time is available for the heat transfer to occur and therefore the temperature may remain constant.

18. State the assumptions made to derive the equations for isothermal flow.One dimensional flow with friction and heat transferConstant area ductPerfect gas with constant specific heats and molecular weightsIsothermal flow

19. Give the assumptions made in Rayleigh flowOne dimensional flow without friction and heat transferConstant area ductPerfect gas with constant specific heats and molecular weightsAbsence of body forces.

20. Write the continuity equationC1/c2 = ρ2/ρ1

PART-B

18. 1. A combustion chamber in a gas turbine plant receives air at 350 K ,0.55bar and 75



m/s.The air –fuel ratio is 29 and the calorific value of the fuel is 41.87 MJ/Kg. Taking γ=1.4 and R =0.287 KJ/kg K for the gas determine.a) The initial and final Mach numbersb) Final pressure ,temperature and velocity of the gasc) Percent stagnation pressure loss in the combustion chamberd) The maximum stagnation temperature attainable


2. Obtain an equation representing the Rayleigh line . Draw Rayleigh lines on the h-s and p-v planes for two different values of the mass flux.


3. The conditions of a gas in a combuster at entry are: P1=0.343bar ,T1 = 310K C1=60m/s.Detemine the Mach number ,pressure ,temperature and velocity at the exit if the increase in stagnation enthalpy of the gas between entry and exit is 1172.5KJ/Kg.Take Cp=1.005KJ/KgK, γ=1.4


4. The pressure, temperature and Mach number of the gas at exit are 2 bar, 1200˚C and 0.7 respectively. The ratio of stagnation temperature at exit to entry is 3.85, calculate the following. Mach number, pressure and temperature of the gas at entry, the heat supplied per kg of gas, the maximum heat supplied and state is it a cooling or heating process.


5. The condition of a gas in a combustion chamber at entry are T1=375 K. p1=0.5 bar, c1=70m/s. The air-fuel ratio is 29 and the calorific value of the fuel is 42 MJ/kg. Calculate the initial and final Mach number, final pressure, temperature and velocity of the gas, percentage of stagnation pressure loss and maximum stagnation temperature. Take γ=1.4 and R =0.287 KJ/kg K


6. Given diabatic flow(Rayleigh flow) of dry air having of some section a Mach number equal to 3 and a stagnation temperature of 300 K, while the static pressure is 0.5 bar. For some other section where mach number is 1.5. Find stagnation temperature, stagnation pressure, static pressure and amount of heat transferred that caused the reduction in Mach number.


7. The stagnation temperature of air is raised from 85˚C to 376˚C in a heat exchanger. If the inlet mach number is 0.4, determine the final mach number and percentage drop in pressure.


8. Air is heated in a frictionless duct from an initial static properties of P1=110 kPa and T1=300 K. Calculate the amount of heat necessary to check the flow at exit of the duct when the inlet mach number is 2.2 and when it is 0.22




9. Show that c1/c2=ρ2/ρ1 = [(M1/M2)2x (1+γ M22/1+γ M1

2)]ANSWER: Page number 3.116 (6)

10. Prove the stagnation pressure ratioANSWER: Page number 3.115 (3)

11.

UNIT - 4 NORMAL SHOCK

PART-A

1. What is mean by shock wave ?A shock wave nothing but a steep finite pressure wave. The shock wave may be described as a compression wave front in a subsonic flow field across which there is abrupt change in flow properties.

2. What is mean by Normal shock?When the shock wave at right angle to the flow it is called normal shock.

3. What is oblique shock?When the shock wave is inclined at an angle to the flow it is called oblique shock.

4. What are applications of moving shock wave ?It is used in Jet engines, Shock tubes, Supersonic wind tunnel and Practical admission turbines

5. Shock waves cannot develop in subsonic flow? Why?In subsonic flow the velocity of fluid is less than the velocity of sound .Due to this reason,



deceleration is not possible in subsonic flow so shock waves cannot develop in subsonic flow.

6. Define compression and rarefaction shock? A shock wave which is at a higher pressure than the fluid into which it is moving is called a compression wave.The shock wave which is at a lower pressure than the fluid into which it is moving is called a expansion shock wave or rarefaction shock wave.

7. State the necessary conditions for a normal shock to occur in compressible flow?1. The compression wave is to be at right angle to the compression flow2. Flow should be supersonic

8. Give the difference between normal and oblique shock?In Normal Shock, the wave is right angle to the Flow and its is a one dimensional flowIn oblique shock, Shock wave is inclined at an angle to the flow and it is a two dimensional flow.

9. What are the properties change across a normal shock ?1. Stagnation pressure decreases2. Stagnation temperature remains const3. Static pressure and temperature increase

10. What is Prandtl – Meyer relation?It is the basis of other equation for shock waves. It gives the relationship between the gas velocities before and after the normal shock and the critical velocity of sound.

11. Define strength of shock wave.It is defined as the ratio of difference in downstream and upstream shock pressures to upstream shock pressure. It is denoted by گ(Py-Px)/Px

12. Is the flow through a normal shock an equilibrium one.No. Since the fluid properties like pressure, temperature and density are changed during normal shock.

13. Calculate the strength of the shock waves when normal shock appears at M=2.Strength of shock = (Py-Px)/Px

For, Normal shocks table for Mx=2 and γ=1.4, Py/Px = 4.5Therefore, strength = 4.5 – 1 = 3.5

14. Show the normal shock in h-s diagram with the help of Rayleigh line and Fanno line.

15. Write down the static pressure ratio expression for a normal shock.Py/Px = (2γ/γ+1) x Mx

2 – [(γ-1)/γ+1)]

16. What are expansion wave?A wave which is at a lower pressure than the fluid in to which it is moving is called an expansion wave or refraction wave.



17. What are compression wave?A wave which is at a higher pressure than the fluid in to which it is moving is called compression wave.

PART-B

19. 1. The state of a gas (γ=1.3,R =0.469 KJ/Kg K) upstream of a normal shock is given by the following data:Mx =2.5, px= 2bar ,Tx =275K calculate the Mach number ,pressure,temperature and velocity of the gas downstream of the shock; check the calculated values with those give in the gas tables.


2. The ratio of th exit to entry area in a subsonic diffuser is 4.0 .The Mach number of a jet of air approaching the diffuser at p0=1.013 bar, T =290 K is 2.2 .There is a standing normal shock wave just outside the diffuser entry. The flow in the diffuser is isentropic. Determine at the exit of the diffuser.a) Mach numberb) Temperaturec) PressureWhat is the stagnation pressure loss between the initial and final states of the flow?


3. Starting from the energy equation for flow through a normal shock obtain the following relations (or) prandtl – meyer relation Cx Cy =a* ² M*x M*y =1


4. Air is entering into supersonic wind tunnel at nozzle throat area of 200 cm2 and test cross sectional area of 330 cm2. If the normal shock is located in the test section, find the following.Test section Mach number and Difuser throat area.


5. The pressure, temperature and velocity of a normal shock wave moving into stagnant air are 1 bar, 20˚C and 520 m/s respectively. If the area of cross section of the duct is constant, calculate at exit the pressure, temperature, velocity, stagnation temperature and Mach number.


6. Air flows adiabatically in a pipe. A normal shock wave is formed. The pressure and temperature of air before the shock are 150 kN/m2 and 25˚C respectively. The pressure just after the normal shock is 350 kN/m2, calculate the mach number, static temperature and velocity o air after the shock wave, increase in density of air, loss of stagnation pressure of air and change in entropy.




7. A convergent nozzle is designed to expand air from a reservoir in which the pressure is 800 kPa and temperature is 40˚C to give a Mach number at exit of 2.5. the throat area is 25 cm2 find the mass flow rate, exit area, when a normal shock appears at a section where the area is 40 cm@, determine the pressure and temperature at exit.


8. A convergent divergent nozzle is designed to expand air from a reservoir in which the pressure is 700 kPa and temperature is 5˚C and the nozzle inlet mach number is 0.2 the nozzle throat area is 46 cm2 and the exit area is 230 cm2. A normal shock appears at a section where the area is 175 cm2. Find the exit pressure and temperature. Also find the increase in entropy across the shock.


UNIT - 5 JET PROPULSION

PART-A

1. Differentiate jet propulsion and rocket propulsion (or) differentiate between air breathing and rocket propulsion?

Jet propulsion Rocket propulsionOxygen required forcombustion purpose is takenfrom the atmosphere

Oxygen is filled in a tank in the rocketengine itself and used for combustion purpose

Altitude limitation No altitude limitationFlight speed always less than jetvelocity.

Flight speed can be greater than jet velocity

Reasonable efficiency Low efficiency expect at extremely high flight speed

Trust decreases with altitude Trust improves slightly with altitude

2. What is monopropellant? Give one example for that?The liquid propellant both the fuel and oxidizer in a single chemical is known as aMono propellant. It is stable at normal ambient conditions and liberates thermal chemical energy on heating.Example: Nitroglycerine and Nitro methane



3. What is bipropellant?If the fuel and oxidizer are different from each other in its chemical nature, the propellant is called the bipropellant.Example: Liquid oxygen –gasoline and Hydrogen peroxide – hydrazine

4. Classify the rocket engines based on sources of energy employed?On the basis of source of energy employed rocket engine is classified as:Chemical rocket enginesSolar rocket enginesNuclear rocket enginesElectrical rocket engines

5. What is specifying impulse of rocket?The thrust developed by unit weight flow rate of the propellant is known as specific impulse. Isp =F/Wp

6. Define specific consumption?The propellant consumption rate per unit thrust is known as specific propellant consumption. SPC =Wp/F

7. What is weight flow co-efficient?It is the ratio of propellant flow rate to the throat force. Cw =Wp/poA*

10. What is IWR?IWR (impulse to weight ratio) is the ratio of total impulse of the rocket to the total weight of the rocket. IWR = I total/Wtotal

11. What is thrust co-efficient?It is the ratio of the thrust to the thrust force. Cf = F/po A*

18. Define propulsive efficiency?It is ratio of the propulsive power to the power output of the engineηp =propulsive power/power output of the engine.

19. What is thrust or drag?The force which propels the aircraft towards at an given speed is called as thrust or propulsive force. This thrust mainly depends on the velocity of gases at the exit of the nozzle.

20. Define Effective Speed ratio.The ratio of flight speed to jet velocity is known as effective speed ratio. Σ = u/cj

21. Define specific thrust.The thrust developed per unit mass flow rate is known as specific thrust (Fsp) (Fsp) = F/m.

22. What is thrust specific fuel consumption(TSFC)?It is defined as the ratio of fuel consumption rate per unit thrust.

23. Define specific impulse.



The thrust developed per unit weight flow rate is known as specific impulse.Isp = F/W

24. What are the main parts of Ramjet engine?The main parts of Ramjet engine are,Supersonic diffuser, subsonic diffuser, combustion chamber and discharge nozzle.

25. Give the expression for the thrust developed b a turbojet engine.Thrust F = m.cj - m.

a u

26. Define overall efficiency.It is the ratio of propulsive power to the power input to the engine.ηo = Propulsive power / power input to the engine.

27. What is the type of compressor used in turbo jet? Why?Rotary compressor is used in turbojet engine due to its high thrust and high efficiency.

28. What is turboprop unit?Turboprop engine is very similar to turbojet engine. In this type, a turbine which is used to drive the compressor and propeller.

29. What is thrust augmentation?To achieve better take-off performance, additional fuel is burnt in the tail pipe between the turbine exhaust section and entrance section of the exhaust nozzle. This is called as thrust augmentation

30. Why ramjet engine does not require a compressor and a turbine?In ramjet engine due to supersonic and subsonic diffuser, the static pressure of air is increased to ignition pressure. So there is no need of compressor and turbine.

31. What is scram jet?A supersonic combustion ramjet engine is known as scramjet.

PART-B

20.1. A ramjet engine operates at M=1.5 at an altitude of 6500m.The diameter of the inlet

diffuser at entry is 50cm and the stagnation temperature at the nozzle entry is 1600K.The calorific value of the fuel used is 40MJ/Kg .The properties of the combustion gases are same as those of air (γ=1.4, R=287J/Kg K ). The velocity of air at the diffuser exit is negligible,Calculate (a) the efficiency of the ideal cycle, (b) flight speed (c) air flow rate (d) diffuser pressure ratio (e) fuel –ratio (f)nozzle pressure ratio (g) nozzle jet Mach number (h) propulsive efficiency (i) and thrust. Assume the following values: D =0.90,B =0.98,j=0.96.Stagnation pressure loss in the combustion chamber =0.002




2. Explain the working principle of ram jet engine with neat sketch.ANSWER: Page number 5.2

3. Explain the working principle of turbo jet engine with neat sketchANSWER: Page number 5.9

4. Explain the working principle of pulse jet engine with neat sketchANSWER: Page number 5.7

5. Explain the working principle of turbo prop engine with neat sketchANSWER: Page number 5.14

6. An aircraft takes 45 kg/s of air from the atmosphere and flies at as speed of 950 kmph. The air fuel ratio is 50 and the calorific value of the fuel is 42 MJ/kg. For maximum thrust power, find jet velocity, specific thrust, propulsive efficiency, overall efficiency, thrust, thrust power and thermal efficiency.


7. A turbo engine operates at an altitude of 3500 m above the sea level and an aircraft speed of 520 kmph. If the inlet diffuser efficiency of the engine is 0.86, compressor efficiency is 0.75, velocity of air at compressor entry is 95 m/s, temperature rise through the compressor is 240 K, find the pressure rise through the inlet diffuser, pressure ratio developed by the compressor, power required by the compressor per unit flow rate of air and air standard efficiency..


8. A turbo jet engine operates at an altitude of 11 km and at a speed of 900 kmph. The engine has the following data.Stagnation temperature at the turbine inlet = 1500 KTemperature drop in the turbine =205˚CCalorific value of the fuel = 42 MJ/kgTurbine efficiency = 0.92Compressor efficiency = 0.76Combustion chamber efficiency = 0.95Exhaust nozzle efficiency = 0.93


9. A turbo jet propels an aircraft at a speed of 900 km/h while taking 3000 kg of air per minute. The isentropic enthalpy drop in the nozzle is 200 kJ/kg and the nozzle efficiency is 90%. The air fuel ratio is 85 and the combustion efficiency is 95%. The calorific value of the fuel is 42,000 kJ/kg, calculate propulsive power, thrust efficiency, propulsive efficiency.


10. Calculate the thrust and specific thrust of a jet propulsion unit whose data are as follows:Total head isentropic efficiency of the compressor = 80%Total head isentropic efficiency of the turbine = 85%Total pressure ratio including combustor pressure loss = 4 : 1



Combustion efficiency = 98%Mechanical Transmission efficiency = 99%Nozzle efficiency = 90%Maximum cycle temperature = 1000 KAir flow rate = 220 N/sFor gases cp = 1153 J/kgK and γ=1.3Ambient temperature and pressure are 15˚C and I bar. Neglect the weight of fuel.


11. A turbo jet has a speed of 750 km/h while flying at an altitude of 10000 m. The propulsive efficiency of the jet is 50% and the over all efficiency of the turbine plant is 16%. The density of air at 10,000 m altitude is 0.73 kg/m3. The drag on the plane is 6250 N. Calorific value of the fuel is 48,000 kJ/kg, Calculate the absolute velocity of the jet, diameter of the jet and power output of the unit in kW.



Gas Dynamics and Jet Propulsion - 2 Marks - All 5 Units

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Transcript of Gas Dynamics and Jet Propulsion - 2 Marks - All 5 Units