The structural and aerodynamic performance of a a low aspect ratio SiC/SiC CMC High Pressure Turbine 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 are 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.
The density of Ceramic Matrix Composite (CMC) materials is approximately 1/3 the density of metals currently used for High Pressure Turbine (HPT) blades. A lower density, and consequently lower centrifugal stresses, increases the feasibility of shrouding HPT blades. Shrouding HPT blades improves aerodynamic efficiency, especially for low aspect ratio turbine blades. This paper explores aerodynamic and structural issues associated with shrouding HPT rotor blades. Detailed Navier-Stokes analysis of a rotor blade showed that shrouding improved blade row aerodynamic efficiency by 1.3%, when the clearance was 2% of the blade span. Recessed casings were used. Without a shroud the depth of the recess equaled the clearance. With a shroud the recess depth increased by the shroud thickness, which included a knife seal. There was good agreement between the predicted stage efficiency for the unshrouded blades and the experimentally measured efficiency. Structural analysis showed a strong interaction between stresses in the shroud and peak stresses at the hub of the blade. A thin shroud of uniform thickness only moderately increased maximum blade stress, but there were very high stresses in the shroud itself. Increasing shroud thickness reduced stresses in the shroud, but increased blade stresses near the hub. A single knife seal added to the thin shroud noticeably decreased maximum shroud stress, without increasing maximum blade stress. Maximum stresses due to pressure loads and combined pressure and centrifugal loads were nearly the same as the maximum stresses for individual pressure or centrifugal loads. Stresses due to a 100K temperature difference across the blade walls were much lower than centrifugal or pressure load stresses.
Through thickness, hoop, and spanwise component stresses were calculated for two Ceramic Matrix Composite (CMC) vane configurations. The analyses are for the first stage vane of a High Pressure Turbine. One configuration is for a vane with trailing edge ejection, and the other has no trailing edge ejection. The effects of analyzing separate pressure and thermal loads, as well as combining these loads, are examined. For the case without trailing edge ejection the effects of variations in the stiffness modulus are given. Results are discussed for the midspan region as well as for the entire span. Pressure loads were determined assuming a mainstream gas and coolant pressure of 50 atm. Thermal loads were determined assuming a gas temperature of 2141°K(3394°F), and a maximum Environmental Barrier Coating temperature of 1756°K(2700°F). The desired maximum CMC temperature was 1589°C(2400°F).
Issues associated with using SiC/SiC Ceramic Matrix Composite (CMC) materials for High Pressure Turbine (HPT) rotor blades are explored. SiC/SiC materials have higher temperature capability than current HPT superalloys. The strength versus temperature characteristics of SiC/SiC CMCs differs from that of superalloys. Stress analyses were done for a NASA specified notional single aisle aircraft engine blade to be available in the N+3 time frame, (beyond 2030). Stacking, the relative position of hub and tip sections, depends on both pressure and centrifugal forces, and material density. The effect of blade stacking on blade stresses is examined. The change in stresses as the rotation rate varies is examined. The change in engine weight, and thus fuel consumption, due to changes in engine size as the rpm changes is discussed. SiC/SiC CMC materials are generally not isotropic. The effect on stresses and strains of a directional variation in Young’s modulus is examined. Shrouding metallic HPT rotor blades is not common. Shrouding SiC/SiC CMC rotor blades may be feasible due to the lower density, and thus lower centrifugal loads, of SiC/SiC blades. The increase in stresses due to shrouding a SiC/SiC blade is discussed.
The structural and aerodynamic performance of a a low aspect ratio SiC/SiC CMC High Pressure Turbine 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 are 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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