Higher productivity of single-phase flow simulations thanks to GPU acceleration

Higher productivity of single-phase flow simulations thanks to GPU acceleration
NACA 2412 wing test case

Case setup

A flow over a simple wing with a NACA 2412 airfoil was simulated. The wings parameters were defined as follows:

  • chord – c = 1m,
  • taper ratio – ctip/croot = 1,
  • wing span – b = 6m – span,
  • wing area – S = 6m2.

A symmetry boundary condition was used so flow over half of the wing was simulated. The computational domain was 21m x 10m x 12m. Inlet boundary had the half-cylinder shape so the different angles of attack could be simulated without changes to the mesh. Mesh had 3401338 (3.4M) cells.

Figure 1. Mesh visualization
Figure 1. Mesh visualization
Figure 2. Zoomed part of the mesh where the wing connects to the symmetry plane
Figure 2. Zoomed part of the mesh where the wing connects to the symmetry plane

Simulations were run for different angles of attack from 0 to 45 degrees with a 5 degree step. In a single simulation 500 steps was done using SIMPLE algorithm using both OpenFOAM and SpeedIT FLOW.

Following boundary conditions were used:

  • Fixed velocity value on the inlet and lower boundaries,
  • Zero gradient pressure on the outlet and upper boundaries,
  • Slip condition on the side boundaries.

Following numerical schemes were used:

  • Gauss linear for gradient,
  • Gauss upwind for divergence,
  • Gauss linear corrected for laplacian,
  • Linear interpolation.

Results

Results from OpenFOAM (OF) and SpeedIT Flow (SITF) were compared. In Fig. 3 a comparison between lift and drag coefficients for different angles of attack computed with OpenFOAM and SpeedIT Flow is shown. The coefficients are nearly identical up to a angle of 35 degrees. A small differences for higher angles may be caused by the flow separation behind the wing. Flow over the wing is shown in Fig. 4. For 40 degrees angle of attack a large recirculation zone can be seen. In Fig. 5 a lift to drag ratio is shown. The results obtained with OpenFOAM and SpeedIT Flow are nearly identical for the whole range of investigated angles of attack.

Figure 3. Lift and drag coefficients computed by OpenFOAM (OF) and SpeedIT Flow (SITF) after 500 steps of the SIMPLE algorithm for different angels of attack
Figure 3. Lift and drag coefficients computed by OpenFOAM (OF) and SpeedIT Flow (SITF) after 500 steps of the SIMPLE algorithm for different angels of attack
Figure 4. Visualization of pressure field on the plane of symmetry and wing and streamlines colored with velocity magnitude with different angle of attack: left – 0 deg., center – 20 deg., right – 40 deg.
Figure 5. Lift to drag ratio calculated with OpenFOAM (OF) and SpeedIT Flow (SITF) after 500 steps with SIMPLE algorithm for different angles of attack
Figure 5. Lift to drag ratio calculated with OpenFOAM (OF) and SpeedIT Flow (SITF) after 500 steps with SIMPLE algorithm for different angles of attack


Acceleration

For the solution of the NACA 2412 wing we used following hardware:

  • CPU: 2x Intel(R) Xeon(R) CPU E5649 @ 2.53GHz (24 threads),
  • GPU: NVIDIA Quadro K6000 12GB RAM,
  • RAM: 96GB
  • OS : Ubuntu 12.04.4 LTS 64bit

Times to solution and acceleration for each angle of attack is given in the Fig. 6. The sum of times of simulations are:

  • OF: 26965 s,
  • SITF: 7722 s.

which gives an acceleration of 3.5x.

Figure 6. Comparisons of time to solution for calculations done with OpenFOAM (OF) and SpeedIT Flow (SITF)
Figure 6. Comparisons of time to solution for calculations done with OpenFOAM (OF) and SpeedIT Flow (SITF)

Validation

Comparison of numerical results obtained with OpenFOAM and SpeedIT Flow is shown in Fig. 7. For lower angles of attack there is a good agreement between the results. For the 40 degrees angle of attack there are some differences caused by the formation of the separation of the flow.

Figure 7. Comparison of numerical results obtained for OpenFOAM (line) and SpeedIT Flow (dots). Upper row – results on a veritcal line 1m behind wing, lower row – pressure distribution along wing section, for different angles of attack: left – 0 deg., center – 20 deg., right – 40 deg.