JNTU
JNTU B.Tech III HIGH SPEED AERODYNAMICS Previous Question Paper (May/June - 2019 - R16)
Part A is compulsory which carries 25 marks. Answer all questions in Part A. Part B consists of 5 Units. Answer any one full question from each unit. Each question carries 10 marks.
PART - A
State the governing equations of inviscid compressible flow.
Answer
The governing equations for inviscid, compressible flow are the Euler equations, which represent the conservation of mass, momentum, and energy. These are expressed as:
where ρ is density, u is the velocity vector, p is pressure, I is the identity tensor, and E is the total energy per unit mass.
Explain Mach number spectrum.
Answer
The Mach number (M) is the ratio of the flow velocity (v) to the local speed of sound (a). The Mach number spectrum categorizes flow regimes based on the compressibility effects and the nature of wave propagation:
- Incompressible Flow (M < 0.3): Density variations are negligible, and the flow is treated as incompressible.
- Subsonic Flow (0.3 < M < 0.8): Compressibility effects become significant, but the flow remains entirely subsonic.
- Transonic Flow (0.8 < M < 1.2): The flow field contains regions of both subsonic and supersonic flow, typically characterized by the appearance of shock waves.
- Supersonic Flow (1.2 < M < 5.0): The flow is entirely supersonic, and disturbances are confined to the Mach cone.
- Hypersonic Flow (M > 5.0): Characterized by high-temperature effects, chemical dissociation, and the formation of thin shock layers close to the body.
For a convergent nozzle, show the variation of mass flow rate for different values of p/p0.
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Write a short note on wave reflections from a free boundary.
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Evaluate strong oblique shock, weak oblique shock solutions to Mach waves.
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What is Prandtl-Meyer angle?
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Which kind of body should be preferred for hypersonic vehicles as per aerodynamic heating point view?
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Show the progression of shock waves with increasing Mach number over an aerofoil with neat sketch during transonic flow.
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Describe in detail about various pressure measuring devices.
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Write the short note on Fringe displacement method.
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PART - B
The pressure temperature and Mach number at the entry of a flow passage are 2.45 bar, 26.50C and 1.4 respectively. If the exit Mach number is 2.5, determine for adiabatic flow of a perfect gas. (γ= 1.3, R = 0.469 kJ/kg K.)
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Derive a general expression for the speed of sound in a compressible gas from first principle.
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Derive the relation between pressure and temperature at inlet and outlet for compressible fluids.
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What is a one dimensional flow? Illustrate some examples of practical flows where you can approximate the flow to be 1-D.
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Air flow is discharged to sea level atmosphere through a sonic nozzle. If the air storage pressure at the reservoir is 40×105 N/m2, determine the pressure, temperature density at the exit of the nozzle. Assume that the reservoir air is at ambient temperature.
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Derive an expression for choked mass flow rate through a converging diverging nozzle in terms of total pressure, total temperature and throat area.
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Derive expressions for the pressure, area and temperature at any section of the nozzle in terms of throat critical conditions.
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Air flowing with a Mach number 2 with a pressure 80 KPa and a temperature 300C passes over a component of an aircraft that can be modelled as a wedge with an included angle of 80 that is aligned with the flow generating an oblique shock wave. Find the pressure acting on the surface of the wedge. Determine the stagnation pressure loss in this case.
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Show a diamond airfoil at an angle of attack α. Draw the shock and expansion wave pattern on it. Draw the pressure profile above and below of it. Discuss the different flow fields.
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In a supersonic flow of air at M1= 2.8 an oblique shock of 420 emerges from a sharp concave corner. Determine: i) Deflection angle ii) Downstream Mach No iii) Static Pressure ratio
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A flat plate placed in the uniform flow of M∞ = 15, and at an angle of attack 30, Calculate L and D per unit area of the plate (ρ∞ = 0.225kg/m3 and T∞ = 157K).
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Explain conical flow method treatment for Rectangular wings.
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Define with neat sketch physical effects characteristics of hypersonic flow.
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If the free-stream Mach number is 3, calculate the pressure coefficient and density coefficient on the surface of a wedge, whose deflection angle is 300, using the Newtonian theory. How does the altitude influence the resultant values of Cp?
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Describe the procedure of getting full scale data from wind tunnel lift characteristics.
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What are various measurement errors encountered during the testing a model and how to capture and minimize these errors?
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Explain in detail about wall interferences and scale effects.
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Describe the methods used for measuring flow angularity in a supersonic wind tunnel.
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