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Gas Turbine Design Axial Flow Compressors

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Gas Turbine Design Axial Flow Compressors ( gas-turbine-design-axial-flow-compressors )

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Chapter 8.0 APPENDIX B NASA Stage 35 NASA Stage 35 consisted of a transonic low-aspect ratio rotor with 36 blades and a stator containing 46 blades (Reid, 1978). At the design speed, the rotor and the stage reached peak adiabatic efficiencies of 87.2% and 84.5% at pressure ratios of 1.875 and 1.842, respectively. No mid-span damper was present on the rotor. Tables B1 and B5 provide the blade geometry at the meanline for the rotor and the stator. Tables B2 through B4 contain experimental flow information for the rotor at the meanline at given speeds. Tables B6 through B8 contain experimental flow information for the stator at the meanline for the corresponding Rotor 35 speeds. The accuracy of the experimental data is reviewed in Table B9. Table B1 Tip Radius Hub Radius Inlet Radius of Meanline Exit Radius of Meanline Chord Length (c) LER tmax/c Solidity (σ) a/c Inlet blade metal angle (β1′) Exit blade metal angle (β2′) Rotor 35 Geometry at Meanline .833 ft 0.5833 ft .7125 ft .7088 ft 0.1825 ft 0.000617 ft .056 1.47 0.6 56.16° 44.26° 71

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