The performance deterioration of a high-speed axial compressor rotor due to surface roughness and airfoil thickness variations is reported. A 0.025 mm (0.001 in.) thick rough coating with a surface finish of 2.54-3.18 rms μm (100-125 rms μin.) is applied to the pressure and suction surface of the rotor blades. Coating both surfaces increases the leading edge thickness by 10 percent at the hub and 20 percent at the tip. Application of this coating results in a loss in efficiency of 6 points and a 9 percent reduction in the pressure ratio across the rotor at an operating condition near the design point. To separate the effects of thickness and roughness, a smooth coating of equal thickness is also applied to the blade. The smooth coating surface finish is 0.254-0.508 rms μm (10-20 rms μin.), compared to the bare metal blade surface finish of 0.508 rms μm (20 rms μin.). The smooth coating results in approximately half of the performance deterioration found from the rough coating. Both coatings are then applied to different portions of the blade surface to determine which portions of the airfoil are most sensitive to thickness/roughness variations. Aerodynamic performance measurements are presented for a number of coating configurations at 60, 80, and 100 percent of design speed. The results indicate that thickness/roughness over the first 2 percent of blade chord accounts for virtually all of the observed performance degradation for the smooth coating, compared to about 70 percent of the observed performance degradation for the rough coating. The performance deterioration is investigated in more detail at design speed using laser anemometer measurements as well as predictions generated by a quasi-three-dimensional Navier-Stokes flow solver, which includes a surface roughness model. Measurements and analysis are performed on the baseline blade and the full-coverage smooth and rough coatings. The results indicate that adding roughness at the blade leading edge causes a thickening of the blade boundary layers. The interaction between the rotor passage shock and the thickened suction surface boundary layer then results in an increase in blockage, which reduces the diffusion level in the rear half of the blade passage, thus reducing the aerodynamic performance of the rotor.

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The effect of adding roughness and thickness to a transonic axial compressor rotor

The performance deterioration of a high speed axial compressor rotor due to surface roughness and airfoil thickness variations is reported. A 0.025 mm (0.001 in.) thick rough coating with a surface finish of 2.54-3.18 RMS microns (100-125 RMS microinches) is applied to the pressure and suction surface of the rotor blades. Coating both surfaces increases the leading edge thickness by 10% at the hub and 20% at the tip. Application of this coating results in a loss in efficiency of 6 points and a 9% reduction in the pressure ratio across the rotor at an operating condition near the design point. To separate the effect of thickness and roughness, a smooth coating of equal thickness is also applied to the blade. The smooth coating surface finish is 0.254-0.508 RMS microns (10-20 RMS microinches), compared to the bare metal blade surface finish of 0.508 RMS microns (20 RMS microinches). The smooth coating results in approximately half of the performance deterioration found from the rough coating. Both coatings are then applied to different portions of the blade surface to determine which portions of the airfoil are most sensitive to thickness/roughness variations. Aerodynamic performance measurements are presented for a number of coating configurations at 60%, 80%, and 100% of design speed. The results indicate that thickness/roughness over the first 10% of blade chord accounts for virtually all of the observed performance degradation for the smooth coating, compared to about 70% of the observed performance degradation for the rough coating. The performance deterioration is investigated in more detail at design speed using laser anemometer measurements as well as predictions generated by a quasi-3D Navier-Stokes flow solver which includes a surface roughness model. Measurements and analysis are performed on the baseline blade and the full-coverage smooth and rough coatings. The results indicate that coating the blade causes a thickening of the blade boundary layers. The interaction between the rotor passage shock and the thickened suction surface boundary layer then results in an increase in blockage which reduces the diffusion level in the rear half of the blade passage, thus reducing the aerodynamic performance of the rotor.

The Effect of Adding Roughness and Thickness to a Transonic Axial Compressor Rotor

Thesis (Ph. D.)--Massachusetts Institute of Technology, Dept. of Aeronautics and Astronautics, 2003.

Probabilistic aerothermal design of compressor airfoils

Aircraft fan and compressor blade leading edges suffer from atmospheric particulate erosion that reduces aerodynamic performance. Recontouring the blade leading edge region can restore blade performance. This process typically results in blades of varying chord length. The question therefore arises as to whether performance of refurbished fans and compressors could be further improved if blades of varying chord length are installed into the disk in a certain order. To investigate this issue the aerodynamic performance of a transonic compressor rotor operating with blades of varying chord length was measured in back-to-back compressor test rig entries. One half of the rotor blades were the full nominal chord length while the remaining half of the blades were cut back at the leading edge to 95% of chord length and recontoured. The rotor aerodynamic performance was measured at 100, 80, and 60% of design speed for three blade installation configurations: nominal-chord blades in half of the disk and short-chord blades in half of the disk; four alternating quadrants of nominal-chord and short-chord blades; nominal-chord and short-chord blades alternating around the disk. No significant difference in performance was found between configurations, indicating that blade chord variation is not important to aerodynamic performance above the stall chord limit if leading edges have the same shape. The stall chord limit for most civil aviation turbofan engines is between 94-96% of nominal (new) blade chord.

The Effect of Variable Chord Length on Transonic Axial Rotor Performance

Especially at transonic flow conditions the leading edge shape influences the performance
of a fan profile. At the same time the leading edge of a fan profile is highly affected
by erosion during operation. This erosion leads to a deformation of the leading edge
shape and a reduction of the chord length. In the present experimental and numerical
study, the aerodynamic performance of an original fan profile geometry is compared to
an eroded fan profile with a blunt leading edge (BLE) and a chord length reduced by
about 1%. The experiments are performed at a linear fan blade cascade in the Transonic
Cascade Wind Tunnel of DLR in Cologne. The inflow Mach number during the tests is
1.25 and the Reynolds number 1.5�106. All tests are carried out at a low inflow turbulence
level of 0.8%. The results of the investigation show that losses are increased over
the whole operating range of the cascade. At the aerodynamic design point (ADP) the
losses raise by 25%. This significant loss increase can be traced back to the increase of
the shock losses at the leading edge. The change in shock structure is investigated and
described in detail by means of particle image velocimetry (PIV) measurements and
Schlieren imaging. Additionally, the unsteady fluctuation of the shock position is measured
by a high-speed shadowgraphy. Then the frequency range of the fluctuation is
obtained by a Fourier analysis of the time resolved shock position. Furthermore, liquid
crystal measurements are performed in order to analyze the influence of the leading edge
shape on the development of the suction side boundary layer. The results show that for
the original fan blade the transition occurs at the shock position on the blade suction side
by a separation bubble whereas the transition onset is shifted upstream for the fan blade
with the BLE.

L. Reid

Papers

Experimental Evaluation of the Effects of a Blunt Leading Edge on the Performance of a Transonic Rotor