Metacomp Technologies participated in the 7th AIAA CFD Drag Prediction Workshop held at the AIAA AVIATION 2022 Conference in Chicago, IL, USA. The aim of the workshop was to predict the effect of shock-induced separation on the variation of lift and pitching moment with increasing angle-of-attack at transonic conditions. The workshop focused on the NASA Common Research Model (CRM) Wing-Body configuration, which consists of a contemporary supercritical transonic wing and a fuselage representative of a wide-body commercial transport aircraft. The CRM is designed for a cruise Mach number of 0.85 and a corresponding design lift coefficient of CL=0.5. This study used a half geometry model that did not include the tail, nacelle, pylon or sting-support system.
Compressible RANS simulations were conducted using the one-equation SA model for turbulence closure, adding the rotational-curvature correction (RC) and quadratic constitutive relations (QCR) to the base SA model. CFD predictions of lift, drag, and pitching moment coefficients were compared to wind tunnel data from the NASA Langley National Transonic Facility . The flow was simulated at a Mach number of 0.85 and a Reynolds number of 20 million with a reference temperature of -250◦ F. Each angle-of-attack used a different geometry that included the aero-elastic deflections measured in the European Transonic Wind Tunnel (ETW) for that specific set of conditions.
Figure 2: Pressure coefficient contours for α = 2.23˚ (CL driver run) and α = 4.25˚
The first simulation was run using the CL driver to automatically adjust the angle-of-attack to achieve the desired CL of 0.50. This simulation converged to an angle-of-attack of 2.23 degrees. Subsequently, an angle-of-attack sweep was performed from 2.75˚ to 4.25˚ in increments of 0.25˚. Figure 2 shows the pressure coefficient contours over the wing-body model from two representative solutions at angles-of-attack of 2.23˚ (CL driver run) and 4.25˚. Figure 3 shows skin friction contours near the wing tip with “oil-flow” streamlines, showing the increasing levels of separation at that location. Shock-induced separation was observed from α=2.75˚ onwards. Figure 4 shows the CL-CD and CL-CM curves over the range of angles-of-attack. Note that a shift in CM was applied to match the test data, as guided by the workshop. The CL-CM curve shows that the experimentally-observed pitch break was captured by CFD++.
Figure 3: Skin friction contours near the wing tip for α = 2.75˚ and α = 4.25˚
Figure 4: CL-CD and CL-CM curves over the range of angles-of-attack