NASA Rotor 67 Validation Studies

Transcription

1 NASA Rotor 67 Validation Studies ADS CFD is used to predict and analyze the performance of the first stage rotor (NASA Rotor 67) of a two stage transonic fan designed and tested at the NASA Glenn center (Strazisar et al. 1989, Strazisar et al. 2004). Its design pressure ratio is 1.63, at a mass flow rate of kg/sec. The NASA Rotor 67 has 22 blades with tip radii of 25.7 cm and cm at the leading and trailing edge, respectively, with constant tip clearance of 1.0 mm. The hub to tip radius ratio is at the leading edge (TC = 0.6% span) and at the trailing edge (TC = 0.75% span). The design rotational speed is 16,043 RPM, and the tip leading edge speed is 429 m/sec with a tip relative Mach number of Figure 1: NASA Rotor 67 configuration (Strazisar et al. 1989)

2 The NASA Rotor 67 blade geometry was put into (x,y,z) blade sections coordinates at different span-wise locations along the blade height into the ADS CFD Grid Generator Wand as shown below. Figure 2: NASA Rotor 67 blade geometry sections An O-H-H-grid with fine tip mesh type 2 (recommended option) with 561,152 elements is used in the current study. It has 41 cells upstream of the rotor, 41 cells downstream of the rotor, 130 cells around the blade, 65 cells in the blade-to-blade direction, and 65 cells in the radial direction as shown in the below figure. Figure 3: ADS Wand blade custom mesh options selection for NASA Rotor 67

3 Figure 4: NASA Rotor 67 blade mesh

4 Only one blade passage was modeled for the current simulation using the ADS CFD Code Leo, Code Leo is a high accuracy Density based code employing the cell vertex scheme with the Wilcox s original two-equation K-Omega turbulence modeling (with wall functions and integration to the wall options). Code Leo also employs preconditioning to speed up convergence for low speed flow problems as well as multi-grid convergence acceleration scheme for structured mesh and residual propagation method convergence acceleration for unstructured meshes. For the current NASA Rotor 67 simulation the inlet total pressure and temperature conditions are 101,325 Pa and K, respectively. The simulation was initially run with a near design conditions back pressure of 121,600 Pa with turbulence modeling integration to the wall at 100% design rotational speed of 16,043 RPM. Then, the characteristic performance map was generated using the ADS performance map generation wizard. This is done by changing the back pressure of the baseline simulation with certain percentages multiple times until an adequate resolution of the performance speed line is achieved. Figure 5 below shows the convergence history of the (Inlet/Outlet) mass flow rate through the rotor blade passage as well as the adiabatic efficiency for the baseline simulation. It is seen that the simulation achieved convergence after 3,000 iterations. The simulation converged to a mass flow rate of 33.9 kg/sec with pressure ratio of 1.65 and total adiabatic efficiency of 90.36%.

5 a) Mass flow rate (kg/sec) history and its convergence rate

6 b) Adiabatic efficiency (%) history and its convergence rate Figure 5: NASA Rotor 67 Convergence history

7 The predicted choking mass flow rate was kg/sec compared to the measured choking mass flow rate of kg/sec. The difference between the computed and the measured choking mass flow rate is ~ 0.7%, which is within the accepted expected range for the CFD/Experimental uncertainties. Figure 6 shows the y+ distribution over the NASA Rotor 67 blade surface and it is shown that a y+< 3 is maintained over the blade surface which is the requirement of the two equation K-Omega turbulence modeling for wall integration. Figure 6: NASA Rotor 67 y+ blade surface distribution The predicted characteristic performance map is compared with the measured characteristic performance map in the below figure. The Total-to-Total Pressure Ratio is plotted against the non-dimensional mass flow rate (w.r.t. the choking mass flow rate). It is seen that ADS CFD predictions agree very well with the experimental measurements.

8 Figure 7: Comparison of measured and computed characteristic performance map Then, the Total-to-Total Adiabatic Efficiency is plotted against the non-dimensional mass flow rate (w.r.t. the choking mass flow rate). It is seen that ADS CFD predictions agree very well with the experimental measurements. Figure 8: Comparison of measured and computed total adiabatic efficiency

9 Figure 9 below shows the flow mixed averaged values at the (LE/TE) of the rotor blade surface which is a vital post processing step needed in the design cycle of the rotors. Figure 9: NASA Rotor 67 (LE/TE) mixed average profile vs % Span The relative Mach number contours and static pressure contours at the rotor mid-section and at the tip-section are shown below in Figure 10 and Figure 11, respectively, using Paraview. In Figure 12, the relative Mach number contours and static pressure contours around the rotor blade are also shown. The passage shock wave structure is nicely captured.

10 a) Relative Mach number contours b) Static pressure contours Figure 10: NASA Rotor 67 relative Mach number and static pressure contours at mid-section

11 a) Relative Mach number contours b) Static pressure contours Figure 11: NASA Rotor 67 relative Mach number and static pressure contours at tip-section

12 a) Relative Mach number contours b) Static pressure contours Figure 12: Relative Mach number and static pressure contours around the rotor