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Master Thesis Report Literature Review 1 − τs(k−1)(1 − ()loss) ηBLI = k τs − 1 (2.40) Substituting equation 2.39 into equation 2.38 results in an expression which relates the thermal effi- ciency of a non ideal Brayton cycle to the inlet recovery ratio. It can be seen that the inlet recovery ratio has a strong negative effect on the overall thermal efficiency. It should be pointed out that the loss in thermal efficiency is more significant than the gain in propulsive efficiency [4]. As a result, the overall impact is a reduction in the overall efficiency of the propulsion system. A related experiment which was conducted by NASA on the influence of wall boundary layer on the performance of an axial flow fan rotor [5] had shown that fan peak efficiency ηH reduces as much as 2.5% when a 1inch spoiler was used. At the peak efficiency flow quantity coefficient Q1 , the 8 nD3 measured displacement boundary layer thickness was found to be 0.069inch and 0.060inch for the inner and outer surface respectively. The flow quantity coefficient is defined as ratio of the volumet- ric flow rate Q1 to the product of the rotation speed n multiplied by the cube of the diameter D. At the low blade loadings,(that is at high value of Q1 ), the thickest boundary layer reduced the nD3 efficiency approximately 8 %. It was concluded that the overall loss in efficiency may possibly be re- duced by decreasing the blade pitch angle in the boundary layer to conform to the upstream velocity profile. Essentially, the increase in the boundary layer thickness leads to a reduction of the axial velocity, resulting in an increase in the flow incidence angle. Therefore, the blade pitch angle should be reduce to better match the flow direction. The results of this experiment is illustrated in figure 2.10. Figure 2.10: Plot of total pressure rise efficiency versus flow quantity coefficient [5] In another related experiment, Kimzey [6] modelled the influence of inlet total pressure distortion on the stability of a XC-1 compressor, by varying the distortion in a 60◦ sector, located on the bottom of the inlet. The results are shown in figure 2.11. In the figure, N is the compressor rotor speed, P is 13PDF Image | Aerodynamic Design of the NASA Rotor 67 for Non Uniform Inflow
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