This paper presents a state of the art design for the blade tip injection. The design includes the means to inject high-pressure gas jet directly into a circumferential casing groove formed in the shroud adjacent to the blade tip. The casing groove is positioned over the blade tip and exceeds 30% of the blade axial chord beyond the impeller to both upstream and downstream directions. In order to validate the multi block model used in the tip gap region, main flow characteristics are verified with the experimental data for smooth casing with a design clearance of 0.5% span. Three arbitrary mass flow rates (1.75%, 2.45%, and 4.35% of choked mass flow) have been studied. The results indicate remarkable advantageous effects on the compressor stability margin. Further, compared to classical design for tip injection, the current design can significantly improve the compressor stall margin due to direct injection of flow. An increase of the injected air may enhance the stall margin improvement. Furthermore, results for injection at different angles, shows that the compressor stability margin reaches a maximum when the bleed air in the relative coordinates is aligned with the mean camber line of the blade leading edge. The main objective of this research is to present an improved design for tip injection as well as to determine its effect on the stability enhancement of the compressor. The current research also provides guidelines to an optimum design of tip injection.

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