Research Summary for May 02, 2002
A. VISCOUS TRANSONIC AIRFOIL WORKSHOP CASES:
 
    NACA 0012:
Table 1. NACA 0012 cases in the Viscous Transonic Airfoil Workshop
free-stream Mach
angle of attack [deg.] (corrected)
Re Number (based on chord length)
Data type
comments
 0.700
1.49
9 x 106
 Cp, CL, CD
 flow attached, slightly supersonic near the leading-edge upper surface
 0.550
8.34
 9 x 106
Cp, CL, CD
 flow has a supersonic bubble well forward on the airfoil upper surface and is slightly separated at the foot of the shock. Angle of attack value is about 1 deg below the maximum lift value.
 0.799
2.26 
 9 x 106
 Cp, CL, CD
 A shock wave exists on the airfoil upper surface at about x/c=0.5, which is strong enough to cause significant boundary layer separation. Relatively "severe" case compared to the others.

In addition to the above,  the paper gives the following cases for NACA 0012 airfoil:

RAE 2822: Test Cases 6 and 10 of the AGARD Report (Table 4).

Table 2. RAE 2822 cases in the Viscous Transonic Airfoil Workshop

free-stream Mach
angle of attack [deg.] (uncorrected)
Re Number (based on chord length)
Data type
comments
 0.725
2.92
6.5 x 106
 Cp, CL, CD, Cf, 
B-L velocity profiles, 
B-L displacement thickness, and 
B-L momentum thickness		 Case 6 of the AGARD report. Various angle of attack and mach number corrections made by the authors presenting CFD results. Transonic flow. A shock exists at around x/c=0.6. Cf distribution indicates no flow separation. 
 0.750
3.19
 6.2 x 106
  Cp, CL, CD, Cf, 
B-L velocity profiles, 
B-L displacement thickness, and 
B-L momentum thickness		 Case 10 of the AGARD report. Various angle of attack and mach number corrections made by the authors presenting CFD results. Transonic flow. One of the most difficult RAE 2822 cases of the AGARD report. Shock wave causes a significant amount of boundary layer separation. 

B. AGARD REPORT CASES:

        NACA 0012:  Wall corrections are applied to the pressure data of the cases shown in Table 3. Re number (based on chord length)  change between 1.85 x 106 and 4.05 x 106 .  Aerodynamic force &  moment coefficients and Cp distributions are the available data. 

Table 3. NACA 0012 cases in the AGARD report.

free-stream Mach
angle of attack (alpha) [deg.] for force and moment coef. measurements (Cx, Cz, Cm) 
angle of attack (alpha) [deg.] for pressure measurements (Cp)
0.3
-2 = <  alpha = < 14
0 / 4
0.4
-2 = <  alpha = < 12
-
0.5
-2 = <  alpha = < 10
0 / 2 / 4 / 6 / 8 
0.6
-2 = <  alpha = < 9
4
0.7
-2 = <  alpha = < 9
0 / 4
0.75
-2 = <  alpha = < 7
0 / 1 / 2 / 3 / 4
0.8
-2 = <  alpha = < 6
0
0.83
-2 = <  alpha = < 6
0

In addition to the above cases, the data (Cp, Cx, Cz, and Cm) acquired in another tunnel at 0 deg. angle of attack for a range of Mach numbers (0.49<Mach<0.93) are also presented in the AGARD report. For these cases no wall corrections are made.

        RAE 2822: No corrections are applied to the data. Available data are the same as the ones given in Table  2.

Table 4. RAE 2822 cases in the AGARD report.

case number
free-stream Mach
angle of attack [deg.] (uncorrected)
Re Number (based on chord length)
1
0.676
2.40
5.7 x 106
2
0.676
-2.18
5.7 x 106
3
0.600
2.57
6.3 x 106
6
0.725
2.92
6.5 x 106
7
0.725
2.55
6.5 x 106
8
0.728
3.22
6.5 x 106
9
0.730
3.19
6.5 x 106
10
0.750
3.19
6.2 x 106
11
0.730
3.19
2.7 x 106
12
0.745
3.19
2.7 x 106
13
0.740
3.19
2.7 x 106