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

080120037 080120037 - Gas Dynamics and Jet Propulsion Propulsion

P.A. COLLEGE OF ENGINEERING ENGINEERING AND TECHNOLOGY PALLADAM ROAD, POLLACHI 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

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080120037 080120037 - Gas Dynamics and Jet Propulsion Propulsion UNIT - 1 COMPRESSIBLE FLOW - FUNDAMAENTALS PART-A

1.

What What is the basic basic diffe differenc rencee between between comp compressi ressible ble and and incompre incompressib ssible le fluid fluid flow? flow? Compressible flow 1. Fluid velocities are appreciable compared compared with the velocity of sound 2. Density is not constant Compressibili ibility ty factor factor is greater greater than 1. Compress one

In Compressible flow 1. Fluid velocities are small compared with the velocity of sound 2. Density is constant 3. Compressibility factor is one

2.

Write Write the steady steady flflow ow energ energyy equati equation on for an adiab adiabatic atic flow of air. air. In an adiabatic flow q = 0. Therefore energy equation becomes, h1 + c12/2 + gZ1 = h2 + c22/2 + gZ 2 + W1 Adiabatic energy energy equation is is h 0 = h + 1/2c22

3.

Expla Explain in the the me mean aning ing of of sta stagn gnati ation on stat statee with with exam exampl ple. e. The state of fluid attained by isentropically decelerating 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.

Distin Distingu guish ish bbetw etween een stat static ic and and stagn stagnati ation on pres pressur sures. es. In stagnation pressure state the velocity of the flowing fluid is zero whereas in the static pressure, pressure, the fluid velocity is not equal to zero

5.

Differ Differen entia tiate te betwee betweenn the static static and and stagnat stagnation ion temp tempera eratur tures. es. 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 temperature temperature To = T+c2/2Cp

6.

What What is is the the use of of Mac Machh num numbe ber? r? Mach number is defined as the ratio between the local fluid velocity to the velocity of sound sound.. Ma Mach ch numb number er M=c/a M=c/a.. It is used used for the analy analysis sis of comp compres ressib sible le fluid fluid flow flow problems. problems. Critical mach number is a dimensionless dimensionless number at which fluid velocity is equal to its sound velocity. M critical = (c/a) – 1

7.

What is is Cr Crocc occo nu numbe ber? r? It is a non-dimensional fluid velocity which is defined as the ratio of fluid velocity to its maximum fluid velocity, C r =c/c =c/cmax

8.

Write Write down down the relat relation ionsh ship ip betwe between en stagn stagnati ation on and and static static temp tempera eratur turee inter interms ms of the the flow, mach number for the case of isentropic flow.











080120037 080120037 - Gas Dynamics and Jet Propulsion Propulsion 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 region Subsonic region Transonic region Supersonic region Hypersonic region 12 Define M* and give the relation relation between 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 2400Km/hr in an atmosphere of 5 degree, degree, find the Mach . angle? C=2400/3.6 C=2400/3.6 = 666.67 T=278K M=c/√ γRT=1.9947 α=sin-1(1/M) = 30.0876 ° 14 Define Mach angle and Mach wedge. . Mach Ma ch angle angle is forme formedd when when an obj objec ectt is mo movin vingg with with supe superso rsonic nic spee speed. d. The The wave wave propagation and changes are smooth. When an object is moving with hypersonic speed the change changess are abrupt abrupt is shown shown in figur figure. e. Hence Hence for a supe superso rsonic nic flow flow over over twotwodimensional dimensional object “mach wedge” is used instead of “mach cone”. 15 What is meant by isentropic isentropic flow with variable area? . A steady one dimensional dimensional isentropic isentropic flow in a variable area passages 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 characteristic Mach number? * 2 . M = [M (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. increase. Therefore, M is not a constant.











080120037 080120037 - Gas Dynamics and Jet Propulsion Propulsion 20 What is the difference between intensive and extensive properties? properties? . Intensive properties: properties: These are independent independent on the mass of the system. Pressure and Temperature Extensive properties: properties: These are dependent on the mass of the system. Ex: Total volume, Total energy

Ex:

21 Distinguish between Mach wave and normal shock? . Mach Ma ch wave: wave: The The lines lines at which which the the pres pressur suree diffe differe rence nce is concen concentra trated ted and and which which generate cone are called mach lines or mach waves Normal 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, 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 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. substance. PART-B

1.

An air air jet jet at 300 300 K has has sonic sonic veloc velocity ity.. Determi Determine ne the the follow followin ing: g: 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: ANSWER: Page number 1.42 (2)

2.

Derive Derive an expre expressi ssion on for for the energ energyy equa equatio tion. n. ANSWER: ANSWER: Page number 1.14(section 1.16)

3.

The press pressure, ure, tempe temperatur raturee and fluid fluid vel velocity ocity of of air at at the entry entry of of a flow flow passage passage are are 3 bar, 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 cm 2. Determine for











080120037 080120037 - Gas Dynamics and Jet Propulsion Propulsion ANSWER: Page number 1.56 (6) [ANSWER:

4.

A gas flows flows in in a duct duct of 40 cm diame diameter ter at inle inlett pressu pressure re of 3 bar, bar, tempe temperat rature ure of 450 450 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. ANSWER: ANSWER: Page number 1.63 (7)

5.

An aircr aircraf aftt is flyin flyingg at an alti altitu tude de of 11,0 11,000 00 mete meters rs,, at 800 800 km/h km/hr. r. The air is reve revers rsib ibly ly 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. [ANSWER: Page number 1.74 (9)

6.

An aircr aircraf aftt is flyin flyingg at an alti altitu tude de of 10,0 10,000 00 meter meters. s. The The inle inlett Ma Mach ch numbe numberr is 0.82, 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 m 2. 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 temperature at diffuser entry. [ANSWER: [ANSWER: Page number 1.77 (10)

7.

Argon Argon is stored stored in a rese reservo rvoir ir at 280 K. Deter Determin minee stagnat stagnation ion entha enthalpy lpy and stag stagnat nation ion 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. ANSWER: Page number 1.80 (11)

8.

Air (γ (γ = 1.4, R=287 R=287 J/kg J/kgK) K) at an inle inlett mach num numbe berr of 0.2 enter enterss a straig straight ht duct duct at 400K 400K and expands expands isentrop isentropical ically. ly. If the exit Mach Mach num number ber is 0.8 0.8,, dete determin rminee the following: following: Stagnation temperature, temperature, critical temperature, static temperature at exit and area ratio A 1/ A2 ANSWER: Page number 1.82 (12)

9.

The pres pressur sure, e, tempe temperat rature ure and and Mach Mach number number at the entry entry of a flow flow passag passagee are 2 bar, bar, 275 K and 1.3 respectively. respectively. If the exit Mach number is 2.4, determine determine the velocity of sound at stagnation condition, the maximum velocity, the temperature and pressure at exit and Mach number M 1* M1* and M2* Take γ = 1.3, R=0.460 kJ/kgK ANSWER: 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. . Calcula Calculate te the following following:: Stagna Stagnation tion enth enthalpy alpy,, stagnati stagnation on velocity velocity of sound, sound, maximum maximum velocity, critical velocity of fluid and critical velocity of sound ANSWER: Page number 1.88 (14)

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.











080120037 080120037 - Gas Dynamics and Jet Propulsion Propulsion ANSWER: ANSWER: Page number 1.97 (4)

13 A steam of air flows with a velocity velocity of 250 m/s in a duct of 10 cm diameter. diameter. Its temperature ˚ . and pressure at that point are 5 C and 40 kPa. What will be its stagnation pressure and temperature? temperature? What is the mass flow rate? ANSWER: ANSWER: Page number 1.100 (6)

14 The following data refers to the entry and exit of a passage where isentropic flow occurs: . Entry:p1 = 207 kPa, T 1 = 300 K, M1=1.4 Exit: M 2 = 2.5, Assuming ideal gas, determine velocity of sound at stagnation condition, condition, maximum velocity and temperature temperature and pressure at exit. ANSWER: ANSWER: Page number 1.102 (7)

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). ANSWER: ANSWER: Page number 1.105 (8)

16 Air (γ = 1.4, R=287 J/kgK) J/kgK) enters a straight axis symmetric symmetric duct at 300K, 3.45 bar and 150 150 . 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, temperature, maximum velocity, mass flow rate and area of cross section at exit. ANSWER: Page number 1.108 (9)

UNIT - 2 FLOW THROUGH VARIABLE AREA DUCTS PART-A

1.

Differ Differen entia tiate te Adia Adiaba batic tic and and Ise Isentr ntrop opic ic proce process. ss. Adiabatic process: process: In a proc proces esss ther theree is no heat heat tran transf sfer er from from the the flui fluidd to surr surrou ound ndin ings gs or from from the the surroundings surroundings to the fluid. Isentropic process: In a isentropic entropy remains constant and it is reversible .During this process there is













080120037 080120037 - Gas Dynamics and Jet Propulsion Propulsion 2.

Diffe Differe rent ntia iate te noz nozzl zlee and and diff diffus user er ? 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 What is Impu Impuls lsee fun funct ctio ionn ? The sum of pressure force ( pA ) and impulse force ( þAc² ) gives Impulse function (F) F = pA + þac²

4.

Differ Differen entia tiate te betwe between en adiab adiabati aticc flow flow and dia diaba batic tic flow 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 flow:Flow in a constant area area duct with friction and without without heat transfer transfer is known as adiabatic flow (Fanno flow).

5.

State State the the expr express ession ion for for dA/A dA/A as as a funct function ion of of Mach Mach numb number er ? dA/A =dp/þc² [ 1-M² ]

6.

Give the the expres expression sion for for T/To and and T/T* T/T* for isen isentrop tropic ic flow throu through gh varia variable ble area area interm intermss of Mach number ? To/T =1+[_-1/2]M² To/T = 1

7.

Draw the the variati variation on of Mach Mach numb number er along along the the len length gth of a converg convergent ent diverg divergent ent duct duct when when it acts as a (a) Nozzle (b) Diffuser (c) Venturi

8.

What What is is choc chocke kedd flow flow thr throu ough gh a noz nozzl zle? e? The mass flow rate of nozzle is increased 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 What type type of noz nozzle zle uuse sedd for soni sonicc flow flow and and supers supersoni onicc flow? flow? Constant area duct nozzle is used for sonic flow and divergent nozzle is used for supersonic supersonic flow. 10 When does the maximum mass flow occur for an isentropic flow with variable area?











080120037 080120037 - Gas Dynamics and Jet Propulsion Propulsion 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 process P2-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 ρAc 2. 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 Venturi The The flow flow is acce accele lera rate tedd cont contin inuo uous usly ly.. The flow is accelerated upto M=1 and then t hen (mac (machh numb number er and and velo veloci city ty incr increa ease sess mach number is decreased continuously)












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 = (P y – Px)/Px 25 Calculate the strength of shock wave when normal shock appears at M=2. . M=2, γ=1.4, Py/Px = 4.5 Strength 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 discharged from a reservoir reservoir at Po =6.91bar and To To =325°c through 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 area,Mach number 3) Maximum velocity. ANSWER: Page number 2.88 (1)

2.

A conical conical diffu diffuser ser has has entry entry and exit exit diame diameters ters of 15 15 cm and and 30cm 30cm respect respectivel ivelyy . The pressure ,temperature ,temperature and velocity of air at entry are 0.69bar,340 k and 180 m/s respectively . Determine 1) The exit pressure 2) The exit velocity 3) The force exerted on the diffuser walls. Assume isentropic isentropic flow,_ γ=1.4,Cp =1.00 KJ Kg-K.. Kg-K.. ANSWER: Page number 2.107 (8)











080120037 080120037 - Gas Dynamics and Jet Propulsion Propulsion rate of air is 125 kg/s, determine stagnation conditions, area at the throat and exit and static conditions of air at exit. [ANSWER: Page number 2.97 (5) 6.

An air ente enters rs an isent isentropi ropicc diffuser diffuser with with a mach mach num number ber of 3.6 3.6 and is is decelera decelerated ted to a mch mch numb number er of 2. The The diffu diffuse serr passe passess a flow flow of 15kg/ 15kg/s. s. The The initi initial al static static pressu pressure re and and temperature of the air air are 1.05 bar and and 40˚C. Assuming γ=1.4, calculate calculate the inlet area, area, total pressure and total temperature at inlet, exit area, total pressure, total temperature and static pressure. pressure. ANSWER: Page number 2.107 (6) [ANSWER:

7.

A thrust thrust chambe chamberr pressure pressure of of a rocket rocket nozzle nozzle is is 350 bar bar and and the nozz nozzle le throat throat section section area area 2 is 6 cm . If the mach number at the nozzle exit is 5.2, calculate the thrust developed by the rocket. [ANSWER: ANSWER: Page number 2.86 (18)

8.

Air ente enters rs the nozz nozzle le from from a large large reservo reservoir ir at 7 bar and and 320˚C. 320˚C. the exit exit pres pressur suree 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 ANSWER: Page number 2.56 (8)

9.

The pressu pressure, re, tempera temperature ture and and velocit velocityy of air air at the the entry entry of a diffuse diffuserr are 0.7 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 isentropic flow and take γ=1.4, c p=1005 J/kgK ANSWER: Page number 2.60 (9)

10 A supersonic wind tunnel settling chamber 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 K 2) Freon -21 ,Cp =0.785 KJ/Kg K ,Cv= 0.675 KJ/Kg K.












UNIT - 3 FANNO AND RAYLEIGH FLOW PART-A

1.

What What are are the the consum consumpti ption on mad madee for for fann fannoo flflow? ow? One dimensional steady flow. Flow takes place in constant sectional area. There is no heat transfer The gas is perfect with constant specific heats.













080120037 080120037 - Gas Dynamics and Jet Propulsion Propulsion and it is constant afterwards. At this point flow is said to be chocked flow. 4.

Expla Explain in the diffe differe rence nce betwe between en Fann Fannoo flow and and Is Isoth other ermal mal flow? flow? Fanno Flow

5.

Isothermal Flow

Flow in a constant area Flow in a constant area duct with duct with friction and friction and the heat transfer is without heat transfer is known as isothermal flow. known as fanno flow. Static temperature is not Static temperature remains constant constant Write Write down the ratio ratio of veloc velocitie itiess between between any any two section sectionss in terms terms of their their Mach Mach number number in a fanno flow ? [1+[_-1/2] M1²]½ C2/C1=M1/M2 [1+[_-1/2] M2²]½

6.

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

7.

What What are are the the thre threee equa equatio tionn gove governi rning ng Fan Fanno no flow flow?? Energy equation, continuity equation and equation of state.

8.

Give Give the expre expressi ssion on to find find in incre crease ase in in entrop entropyy for Fann Fannoo flow? flow? S2-S1











080120037 080120037 - Gas Dynamics and Jet Propulsion Propulsion 13 Define fanning’s coefficient coefficient of skin friction . It is the ratio between wall shear stress and dynamic head F = 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 dimensional flow with friction and heat transfer in a constant area duct is called isothermal 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 dimensional flow with friction and heat transfer Constant area duct Perfect gas with constant constant specific heats and molecular weights











080120037 080120037 - Gas Dynamics and Jet Propulsion Propulsion ANSWER: Page number 3.74 (5)

2.

Obtain Obtain an equa equation tion repres representi enting ng the Rayle Rayleigh igh line line . Draw Draw Rayleigh Rayleigh lines lines on the the h-s and ppv planes for two different values of the mass flux. ANSWER: Page number 3.115 (1)

3.

The The con ondi ditition onss of a gas in a com ombbuste ster at en entr tryy are: P1=0 1=0.34 .343b 3baar ,T1 ,T1 = 310 10K K C1=60m/s.Detemine the Mach number ,pressure ,temperature and velocity at the exit if the the increase in stagnatio tion entha thalpy of the gas between entry try and exit is 1172.5KJ/Kg.Take 1172.5KJ/Kg.Take Cp=1.005KJ/KgK, Cp=1.005KJ/KgK, γ=1.4 [ANSWER: Page number 3.31 (3)

4.

The press pressure, ure, temper temperatur aturee and Mach Mach number number of of the gas gas at exit exit are are 2 ba bar, r, 1200˚C 1200˚C and 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 supplied and state is it a cooling or heating process. ANSWER: Page number 3.35 (4)

5.

The The cond condititio ionn of a gas in a comb combus ustition on cham chambe berr at entr entryy are are T 1=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 percentage of stagnation pressure loss and maximum stagnation temperature. Take γ=1.4 and R =0.287 KJ/kg K [ANSWER: Page number 3.44 (6)

6.

Given Given diabatic diabatic flow(Ra flow(Raylei yleigh gh flow) flow) of dry dry air having having of of some some section section a Mach Mach number number equal equal to 3 and a stagnation temperature of 300 K, while the static pressure is 0.5 bar. For some











080120037 080120037 - Gas Dynamics and Jet Propulsion Propulsion 11.

UNIT - 4 NORMAL SHOCK PART-A

1.

What What is me mean an by shoc shockk wav wavee ? A shock wave nothing nothing but a steep finite finite pressure wave. wave. The shock wave may be











080120037 080120037 - Gas Dynamics and Jet Propulsion Propulsion 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 State the necessa necessary ry conditio conditions ns for for a norma normall shock shock to occu occurr in comp compressi ressible ble flow? flow? 1. The compre compression ssion wave wave is to be at right right angle angle to the compres compression sion flow flow 2. Flow Flow sho shoul uldd be be ssup upers erson onic ic

8.

Give Give the the diffe differen rence ce betw between een norm normal al and and obl obliqu iquee shock shock?? In Normal Shock, the wave is right angle to the Flow and its is a one dimensional flow In oblique shock, Shock wave is inclined at an angle to the flow and it is a two dimensional dimensional flow. What What are are the the prope properti rties es chan change ge acros acrosss a norm normal al shock shock ? 1. Stagn Stagnati ation on press pressure ure dec decrea reases ses 2. Stagnatio Stagnationn temp temperat erature ure remains remains const const 3. Static Static pres pressure sure and tem tempera perature ture increase increase

9.

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. 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














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 ,pressure,temperature and velocity of the gas downstream of the shock; check the calculated calculated values with those give in the gas tables. ANSWER: Page number 4.66 (1)

2.

The ratio ratio of th exit exit to en entry try area area in a sub subson sonic ic diffuser diffuser is is 4.0 .The .The Mach Mach numbe numberr of a jet jet of air approaching 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 number b) Temperature Temperature c) Pressure What is the stagnation pressure loss between the initial and final states of the flow? ANSWER: Page number 4.58 (8)












8.

A converg convergent ent diver diverge gent nt noz nozzle zle is desig designe nedd to ex expan pandd air from a reser reservoi voirr in which the 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 cm 2 and the exit area is 230 cm 2. A normal shock appears at a section where the area is 175 cm 2. Find the exit pressure and temperature. Also find the increase in entropy across the shock. ANSWER: Page number 4.74 (4)

UNIT - 5 JET PROPULSION











080120037 080120037 - Gas Dynamics and Jet Propulsion Propulsion Nuclear rocket engines Electrical rocket engines 5.

What What is is spec specififyi ying ng imp impul ulse se of of roc rocke ket? t? The thrust developed by unit weight flow rate of the propellant propellant is known as specific impulse. impulse. Isp =F/Wp

6.

Defi Define ne sp specif ecific ic cons consum umpt ptio ion? n? The propellant consumpti consumption on rate per unit thrust is known as specific propellant consumption. consumption. SPC =Wp/F

7.

What What is weig weight ht flow flow co-e co-eff ffic icie ient nt?? 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*












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 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. propeller. 29 What is thrust augmentation? . To achieve better take-off performance, 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 augmentation 30 Why ramjet engine does not require a compressor and a turbine?












5.

Explain Explain the the worki working ng prin principl ciplee of turbo turbo prop prop engi engine ne with with ne neat at sketch sketch ANSWER: ANSWER: Page number 5.14

6.

An aircr aircraft aft takes takes 45 kg/s kg/s of air from from the atmo atmosph sphere ere and and flies flies at as speed speed of 950 950 kmph. kmph. The air fuel ratio is 50 and the calorific value of the fuel is 42 MJ/kg. For maximum thrust power, power, find jet velocity velocity,, specific specific thrust, thrust, propuls propulsive ive efficien efficiency, cy, overall overall efficienc efficiency, y, thrust, thrust, thrust power and thermal efficiency. ANSWER: Page number 5.43 (3)

7.

A turb turboo engi engine ne opera operate tess at an altit altitud udee of 3500 3500 m abov abovee the the sea sea leve levell and and an airc aircra raft ft 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.. ANSWER: Page number 5.49 (5)

8.

A turbo turbo je jett engin engin

tes tes at an alti altitu tude de of 11 km and and at a

d of 900 900 kmph. kmph. Th













080120037 080120037 - Gas Dynamics and Jet Propulsion Propulsion 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/m 3. 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 output of the unit in kW. ANSWER: Page number 5.99 (6)

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

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