Abstract
Abstract The structural and aerodynamic performance of a low aspect ratio SiC/SiC ceramic matrix composite (CMC) high pressure turbine (HPT) blade was determined. The application was a NASA notional single aisle aircraft engine to be available in the N + 3, beyond 2030, time frame. The notional rpm was maintained, and to satisfy stress constraints, the annulus area was constrained. This led to a low span blade. For a given clearance, low span blade is likely to have improved efficiency when shrouded. The efficiency improvement due to shrouding was found to strongly depend on the axial gap between the shroud and casing. Axial gap, unlike clearance or reaction, is not a common parameter used to correlate the efficiency improvement due to shrouding. The zero clearance stage efficiency of the low aspect ratio turbine was 0.920. Structural analyses showed that the rotor blade could be shrouded without excessive stresses. The goal was to have blade stresses less than 100 MPa (14.5 ksi) for the unshrouded blade. Under some not very restrictive circumstances, such as blade stacking, a one-dimensional radial stress equation accurately predicted area averaged Von Mises stress at the blade hub. With appropriate stacking, radial and Von Mises stresses were similar.
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