09A52102 Aerodynamics-II B.Tech Question Paper : sphoorthyengg.com
Name of the College : Sphoorthy Engineering College
University : JNTUH
Department : AERONAUTICAL ENGINEERING
Subject Code/Name : 09A52102/AERODYNAMICS – II
Year : December 2011
Degree : B.Tech
Year/Sem : III/I
Website : sphoorthyengg.com
Document Type : Model Question Paper
Download Model/Sample Question Paper : https://www.pdfquestion.in/uploads/sphoorthyengg.com/4801-09A52102%20-%20AERODYNAMICS%20-%20II.pdf
Aerodynamics-II Question Paper :
B.Tech III Year I Semester Examinations,
(AERONAUTICAL ENGINEERING)
Time: 3 hours
Related : Sphoorthy Engineering College 09A52103 Aerospace Vehicle Structures-II B.Tech Question Paper : www.pdfquestion.in/4800.html
Max. Marks: 75
Answer any five questions ;
All questions carry equal marks :
1.a) Mathematically define compressible flow. What do you mean by isothermal compressibility, isentropic compressibility?
b) Explain in brief various flow regimes of compressible flow using neat sketches. [7+8]
2. Obtain the energy equation for one dimensional flow given by, Mention its alternative forms [15]
3. Explain the design of high lift to drag hypersonic configurations(Wave riders). [15]
4. Define oblique shock wave. Obtain the relation between flow properties viz., Mach number, pressure, density and temperature ahead and behind the oblique shock wave. [15]
5.a) Obtain the area Mach number relation for a variable area duct when there is an isentropic flow of a calorically perfect gas through it.
b) Consider the subsonic-supersonic flow through a convergent-divergent nozzle. The reservoir pressure and temperature are 10 atm and 300 K respectively. There are two locations in the nozzle where A/A* =6; one in the convergent section and other in the divergent section, calculate Mach number, pressure, Temperature and velocity. [8+7]
6. Write a short notes on
a) Critical Mach number
b) Drag divergence Mach number
c) Super critical airfoil [5+5+5]
7.a) Using neat sketches explain briefly how the supersonic flow past different wing plan forms is evaluated.
b) Write short notes on aerodynamic interaction. [9+6]
8. The geometry of a double wedge airfoil section is as shown in figure
Using linear theory, calculate the lift coefficient, wave drag coefficient, pitching moment coefficient and pressure coefficient on each panel of the airfoil. [15]
R09Code No: 09A52102
SET-2
1.a) Define shock polar using neat sketches.
b) Define Mach reflection using neat sketches. [8+7]
2.a) Define first law of Thermodynamics. Obtain the relations for internal energy and enthalpy for a perfect gas.
b) Briefly explain the fluid models which are used to extract the mathematical equations using neat sketches. [7+8]
3.a) Explain the working of a supersonic wind tunnel using neat sketches.
b) A supersonic wind tunnel is designed to produce flow in the test section at Mach 2.4 at standard atmospheric conditions. Calculate:
i. The exit-to- throat area ratio of the nozzle
ii. Reservoir pressure and temperature [7+8]
4. A rectangular wing having an aspect ratio of 3.5 is flying at M8 = 0.85 at 12 Km. A NACA 0006 airfoil section is used at all span wise stations. What is the airfoil section and aspect ratio for equivalent wing in an incompressible flow? [15]
5. Explain the method of characteristics for a two dimensional irrotational supersonic flow and determine the characteristic lines for the same. [15]
6. Solve the equation using singularity distribution method. Obtain the pressure distribution given an arbitrary configuration. [15]
7. Write a short notes on
a) Aerodynamic heating
b) Mach Number independence
c) Law of hypersonic similarity. [5+5+5]
8.a) Define normal shocks, quasi one dimensional flow using neat sketches.
b) At a given point in the high speed flow over an airplane wing, the local Mach number, pressure and temperature are 0.7, 0.9 atm and 250 K, respectively.
Calculate the values of stagnation pressure, stagnation temperature, characteristic temperature, characteristic Mach number, a* at this point. [4+11]
B.Tech III Year I Semester Examinations, Set III :
1.a) State the fundamental physical principles from the laws of nature.
b) Derive momentum equation considering an appropriate fluid model. [4+11]
2. Consider a wedge with a half angle of 10o flying at Mach 2. Calculate the ratio of total pressures across the shock wave emanating from the leading edge of the wedge. [7+8]
3. In detail, explain the design considerations for supersonic aircraft. What are the major considerations to build a supersonic transport aircraft (SST)? Discuss. [15]
4. Write a short notes on
a) Diffusers
b) Wave reflection from a free boundary
c) Three dimensional shock waves