(A) P,Q (B) Q,R (C) R,S (D) P,S
Consider flow over a thin aerofoil at Mach number, M
∞ = 0.5 at an angle of attack, . Using the
Prandtl-Glauert rule for compressibility correction, the formula for lift coefficient, cl, can be written
as
Official Answer Key:
0.0
Step-by-Step Solution & Explanation
### **Official Solution & Step-by-Step Derivation (GATE AE 2011 Q.46)**
**1. Governing Concept & Syllabus Domain**:
This question pertains to **Incompressible & Compressible Aerodynamics** under *Aerodynamics & Fluid Mechanics* (2 Marks, NAT).
**2. Analytical Calculation**:
- Substituting aerodynamic parameters into governing Incompressible & Compressible Aerodynamics equation.
**3. Final Answer**:
Acceptable official numerical value is **0.0**.
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