Aerodynamic Assignment

Surface static pressure measurement provides essential information for the assessment of airflow around a model. An airfoil pressure distribution provides information on local flow velocities, loads, and an indication of key flow features such as separation. Integration of normal surface static pressure distribution, provides lift and form drag values.

The static pressure coefficient Cp, is defined as

Cp = (P - P∞)/(½ ρair V2)

Where,

• P is the local static pressure

• ρair is air density

• V is the free stream velocity

• P∞ is the static pressure of the undisturbed stream

Surface static pressure is usually tapped from small bore tubes machined perpendicularly into the surface (Figure 1 & Table 1). These tapings are connected to pressure transducers by flexible pneumatic tubing. When several pressure tapings are monitored, multiple transducers are often used to simultaneously measure individual pressures.

2D Model NACA 23012 Airfoil Technical Data

Chord 0.191 m,    Span 0.457 m

Figure 1: 2D model NACA 23012 airfoil geometry

Table 1: Location of Pressure Tapings

Aerodynamics 2D NACA 23012 Airfoil Laboratory Pressure Distribution Test Results

Aerodynamic 3D NACA 23012 Airfoil Forces & Moments Laboratory Test Results 

Note: For wind tunnel test results see also attached Excel files.

Requirements: 

1. Data analysis is required.

2. Calculate the pressure coefficients from the transducer measurements and wind tunnel free stream dynamic pressure. Ensure you have corrected this data for static pressure drop.

3. Plot CP vs x/c for both upper and lower surfaces of the model at each incidence and then plot the CP vs y/c for both rearward and forward facing surfaces of the model at each incidence.

4. Comment on significant features highlighted by the pressure distribution.

5. Calculate the aerodynamic lift coefficient over the incidence range -4 to +21 degrees. Graphical data presentation of both CL vs α and CD vs α should be included. Tabulated values of force coefficients are needed. State any assumptions made to determine these coefficients. 

6. The numerical methods used to carry out the integration should be identified. Discuss the limitations and accuracy of the numerical and experimental techniques. 

7. Identify a suitable source of published data for the NACA 23012 profile with which to compare with your experimental data. Ensure the published data is at a comparable Reynolds Number.

8. Use XFOIL to determine the pressure distribution around the airfoil under the same conditions as you have tested experimentally. Compare the results with your experimental data and discuss.

9. Use your experimental data to plot CL vs α, CD vs α and CD vs CL2.

10. Compare the values obtained for CL and CD and the rate of change of CL vs α with the 2D values you have calculated previously and discuss the reasons for any differences.

11. Determine the zero lift drag coefficient and span efficiency factor for the wing.

12. Determine the angle of attack at which the wing exhibits the maximum efficiency.

13. State your report conclusions.

Aerodynamic-2D-NACA-23012

Aerodynamic-3D-NACA-23012