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Question Bank 

ME 6604- GAS DYNAMICS AND JET PROP!SION

NIT-"-PART-A #$ Ma%ks&

1) State the difference between compressible fluid and incompressible fluid ?

2) Define stagnation pressure?

3) Express the stagnation enthalpy in terms of static enthalpy and velocity of flow?

) Explain !ach cone and !ach angle?

") Define adiabatic process?

#) Define !ach number?

$) Define %one of action and %one of silence ?

&) Define closed and open system?') (hat is the difference between intensive and extensive properties?

1) Distinguish between !ach wave and normal shoc*?

1) Differentiate +diabatic and ,sentropic process-

2) Differentiate no%%le and diffuser ?

3) (hat is ,mpulse function ?

) Differentiate between adiabatic flow and adiabatic flow ?

") State the expression for d+.+ as a function of !ach number ?

#) /ive the expression for 0.0o and 0.0 for isentropic flow through variable area in terms of 

!ach number ?

$) Draw the variation of !ach number along the length of a convergent divergent duct when

it acts as a a) o%%le b) Diffuser c) 4enturi

&) (hat is choc*ed flow through a no%%le?

') (hat type of no%%le used for sonic flow and supersonic flow?

1) (hen does the maximum mass flow occur for an isentropic flow with variable area?

1) (hat are the assumptions made for fanno flow?

2) Differentiate 5anno flow and 6ayleigh flow?

3) Explain choc*ing in 5anno flow?

) Explain the difference between 5anno flow and ,sothermal flow?

") (rite down the ratio of velocities between any two sections in terms of their !ach number 

in a fanno flow ?

#) (rite down the ratio of density between any two sections in terms of their !ach number in

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a fanno flow?

$) (hat are the three e7uation governing 5anno flow?

&) /ive the expression to find increase in entropy for 5anno flow?

') /ive two practical examples where the 5anno flow occurs?

1) (hat is 6ayleigh line and 5anno line?

1) (hat is mean by shoc* wave?

2) (hat is mean by ormal shoc*?

3) (hat is obli7ue shoc*?

) Define strength of shoc* wave?

") (hat are applications of moving shoc* wave?

#) Shoc* waves cannot develop in subsonic flow? (hy?

$) Define compression and rarefaction shoc*? ,s the latter possible?&) State the necessary conditions for a normal shoc* to occur in compressible flow?

') /ive the difference between normal and obli7ue shoc*?

1) (hat are the properties changes across a normal shoc*?

1) Differentiate 8et propulsion and roc*et propulsion or) differentiate between air breating

and roc*et propulsion?

2) (hat is monopropellant? /ive one example for that?

3) (hat is bipropellant?) 9lassify the roc*et engines based on sources of energy employed?

") (hat is specify impulse of roc*et?

#) Define specific consumption?

$) (hat is weight flow co:efficient?

&) (hat is ,(6?

') (hat is thrust co:efficient?

1) Define propulsive efficiency?

Pa%t - B #"6 Ma%ks&

1- Stating the assumptions used-+n air 8et r ;1-< 6;2&$ =.>g >) at > has sonic

velocity -Determine

1- velocity of sound at > 

2- 4elocity of sound at the stagnation conditions-

3- !aximum velocity of the 8et-

- Staganation enthalpy-

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"- crocco number-

2) 0he pressure< temperature and !ach number at the entry of a flow passage are 2-" bar<

2#-"@ 9 and 1- respectively- ,f the exit !ach number is 2-" determine for adiabatic

flow of perfect gas A ;1-3< 6;-#' >=.>g >)-

3) +ir A ;1-g >) enters a straight axis symmetric duct at 3 >=.>g >

#) a) (hat is the effect of !ach number on compressibility prove for 

 A;1-< o FG . H G cB ; 1 IJ !B I 1. ! I KK- &)

b) Show that for sonic flow the deviation between the compressible and incompressible

flow values of the pressure coefficient of a percent gas A ;1-) is about 2$-" per cent-

&)

$) +ir at stagnation condition has a temperature of & >- Determine the stagnation velocity of 

Sound and the maximum possible fluid velocity- (hat is the velocity of the sound

when the flow velocity is at half the maximum velocity 1#)&) +ir flow through a duct- 0he pressure and temperature at station one are pressure is -$ bar 

and temperature is 39- +t a second station the pressure is -" bar- 9alculate

temperature and density at the second station- +ssume the flow is to be ,sentropic 1#)

nit -$

Pa%t - B #"6 Ma%ks&

1)+ir flowing in a duct has a velocity of 3 m.s

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1) Stagnation pressure and temperature< )

2) 4elocity of sound in the dynamic and stagnation conditions< #)

3) Stagnation pressure assuming constant density- #)

2) + conical diffuser has entry and exit diameters of 1" cm and 3cm respectively-

-0he pressure g.hr determine for isentropic flow

1)0hroat area< pressure3)!aximum velocity- )

") + super sonic wind tunnel settling chamber expands air or 5reon:21 through a

no%%le from a no%%le from a pressure of 1 bar to bar in the test section - calculate the

stagnation temperature to the maintained in the setting chamber to obtain a velocity of 

" m.s in the test section for +ir =.>g >< 5reon :21

< 9p ;-$&" >=.>g >

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$) +ir flowing in a duct has a velocity of 3 m.s g.hr determine for isentropic

flow

1)0hroat area< pressure3)!aximum velocity- )

11) + supersonic wind tunnel settling chamber expands air or 5reon:21 through a no%%le

from a no%%le from a pressure of 1 bar to bar in the test section- 9alculate the

stagnation temperature to the maintained in the setting chamber to obtain a velocity of 

" m.s in the

test section for< +ir =.>g >< 5reon :21 =.>g >-

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(hat is the test section !ach number is each case? 1#)

12) Derive the following relations for one dimensional isentropic flow

d+.+ ;dC.G cB1 :!B) &)

 A p.p ;2.AI1 IA:1 .AI1!B) &)

nit -'

1)+ circular duct passes &-2">g.s of air at an exit !ach number of -"- 0he entry pressure and

temperature are 3-" bar and 3&@9 respectively and the coefficient of friction -"-,f the

!ach number at entry is -1"< determine

,- 0he diameter of the duct < 2)

,,- Length of the duct< )

,,,- Cressure and temperature at the exit<) ,4- Stagnation pressure loss< and )

4- 4erify the exit !ach number through exit velocity and temperature- 2)

2) + gas A ;1-3=.>g>) at p1 ;1bar< 01 ; * enters a 3cm diameter duct

at a !ach number of 2--+ normal shoc* occurs at a !ach number of 1-" and the exit

!ach number is1-

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 b) Diameter of the no%%le throat< and #)

c) Cressure and temperature at the pipe exit- #)

") Show that the upper and lower branches of a 5anno curve represent subsonic and

supersonic flows respectively- Crove that at the maximum entropy point !ach number is

unity and all processes approach this point -Mow would the state of a gas in a flow change

from

the supersonic to subsonic branch ? 1#)

5low in constant area ducts with heat transfer 6ayleigh flow)

#) 0he !ach number at the exit of a combustion chamber is -'- 0he ratio of stagnation

temper ature at exit and entry is 3-$- ,f the pressure and temperature of the gas at exit are 2-"

 bar 

and 1@9 respectively determine a) !ach number g> 1#)

$) 0he conditions of a gas in a combuster at entry are C1;-33bar =.>g>< A ;1- 1#)

&) + combustion chamber in a gas turbine plant receives air at 3" >

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nit-4

1)0he state of a gas A;1-3 is 2-2 -0here is a standing normal

shoc* wave 8ust outside the diffuser entry- 0he flow in the diffuser is isentropic -

Determine at the exit of the diffuser-

1- !ach number < )

2- 0emperature< and )3- Cressure )

- (hat is the stagnation pressure loss between the initial and final states of the flow ? )

3) 0he velocity of a normal shoc* wave moving into stagnant air p;1- bar< t;1$@9 ) is "m.s

-

,f the area of cross: section of the duct is constant determine a) pressure b) temperature c)

velocity of air d) stagnation temperature and e) the mach number imparted upstream of 

the wave

front- 1#)

) 0he following data refers to a supersonic wind

tunnel o%%le throat area ;2cmB

0est section cross: section ;33$-"cmB

(or*ing fluid Qair A ;1-< 9p ;-2&$ >=.>g >)

Determine the test section !ach number and the diffuser throat area if a normal

shoc* is located in the test section- 1#)

") + supersonic diffuser for air A ;1-) has an area ratio of -1# with an inlet !ach number of 

2- design value)- Determine the exit !ach number and the design value of the pressure

ratio across the diffuser for isentropic flow- +t an off: design value of the inlet !ach number 

2-$) a normal shoc* occurs inside the diffuser -Determine the upstream !ach number and

area ratio at the section where the shoc* occurs< diffuser efficiency and the pressure ratio

across the diffuser- Depict graphically the static pressure distribution at off design- 1#)

#) Starting from the energy e7uation for flow through a normal shoc* obtain the

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following relations or) prandtl F meyer relation 9x 9y ;a B !x !y ;1 1#)

5low with obli7ue shoc* waves

$) +ir approaches a symmetrical wedge A ;1"@) at a !ach number of 

2--Determine for the strong and wea* waves a) wave angle b) pressure ratio c)

density ratio<

d) temperature ratio and e)downstream !ach number 4erify these values using

/as tables for normal shoc*s- 1#)

&) + gas A ;1-3) at p1 ;3" !bar< 01; 3" > and !1;1-" is to be isentropically expanded

to 13& !bar- Determine a) the deflection angle =.>g- 0a*e 9p;1-">=.*g< A ;1-- 1#)

13) + combustion chamber in a gas turbine plant receives air at 3" > < -"" bar and

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$"m.s- 0he air Ffuel ratio is 2' and the calorific value of the fuel is 1-&$ !=.>g-

0a*ing A ;1- and 6 ;-2&$ >=.>g > for the gas determine

a) 0he initial and final !ach number< )

 b) 5inal pressure

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9ombustion chamber efficiency ;-'$"

0urbine efficiency ;-&1

!echanical efficiency of the power transmission between turbine and compressor ;-'&

Exhaust no%%le efficiency;-'$

Specific impulse ;2" seconds

+ssuming the same properties for air and combustion gases calculate

 A 5uel Fair ratio< 2)

 A 9ompressor pressure ratio< )

 A 0urbine pressure ratio< )

 A Exhaust no%%les pressure ratio -0he calorific

value of the fuel used is !=.>g -0he properties of the combustion gases are same as

those of 

air A ;1-< 6;2&$=.>g > )- 0he velocity of air at the diffuser exit is negligible

9alculate

a) the efficiency of the ideal cycle< b) flight speed c) air flow rate d) diffuser pressure ratioe) fuel Fratio f)no%%le pressure ratio g) no%%le 8et !ach number h) propulsive efficiency

i) and thrust- +ssume the following valuesRD ;-'

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flow rate of "->g.s -,f the heat of reaction of the propellants is #">=.>g of 

the propel at mixture determineQ

a) Cropulsion efficiency and propulsion power< #)

 b) Engine output and thermal efficiency

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1") riefly explain the construction and wor*ing of 

+- 6oc*et engine #)

- 6am8et engine #)

9- Culse8et engine )

1#) (ith the help of a neat s*etch describe the wor*ing of a ram8et engine- Depict

the various thermodynamic process occurring in it on h:s diagram- (hat is the effect

of flight !ach number on its efficiency? 1#)

1$) Explain with a neat s*etch the wor*ing of a turbo:pump feed system used in a

li7uid propellant

roc*et? 1#)