Icing is a phenomenon that poses a serious threat to flight safety. This paper systematically conducts research on engine anti-icing design technology using a newly developed civil turbo-shaft engine as a platform, establishing a complete anti-icing design process for turbo-shaft engines from a forward research and development perspective. It proposes an “mesh screen + hot air anti-icing” composited anti-icing method that combines structure anti-icing with a square-hole curved surface mesh screen and hot-air anti-icing through automatic regulation of bleed air volume via a constant pressure valve. Numerical analysis is performed on the inlet flow field with the mesh screen under icing conditions, the total pressure loss at the inlet, changes in compressor parameters, and the impact of ice chunks on the compressor first rotor blades. The research results indicate that after icing, the total pressure loss at the inlet and the distortion of the compressor inlet flow field increase. Coupled with the temperature distortion effect caused by hot-air anti-icing, compressor efficiency and surge margin decrease, leading to engine performance deterioration. To ensure the anti-icing effectiveness of the engine within the icing envelope, the ground idle condition at −20 ℃ can be used as the design point for hot-air anti-icing, with a bleed air volume of approximately 2.8%. To meet the aircraft's power requirements, performance design should focus on ensuring that the engine retains sufficient power margin under high-altitude icing conditions. To prevent surge during engine acceleration under icing conditions, the control system can adopt a optimized control strategy that automatically reduces the acceleration limit of the gas generator rotor. The research findings have been directly applied to the development of a newly developed civil turbo-shaft engine, providing important technical support for the engine's airworthiness certification.

In this paper, the icing airworthiness test is conducted independently. A newly developed civil turbo-shaft engine is taken as the verification platform. The technical problems such as development of icing spray system and simulation of cloud spray flowfield are solved. Icing tests in icing envelope conditions are completed by using the newly constructed icing test bench. The ice formation mechanism of the intake system is revealed. The ice morphology of the engine intake system is obtained, and the impact of the flight environment on ice morphology is analyzed. The changing rules of total pressure loss, compressor efficiency, gas turbine outlet temperature and other parameters are summarized. The test results show that icing mainly occurs in the intake mesh cover of the engine, and the icing type is mainly manifested as glaze ice and rime ice. There are various forms of ice morphology, and the rime ice covers a large area, with the maximum icing thickness up to 260 mm appearing at temperature −10 ℃, altitude 1.2 km, and velocity 280 km/h. Icing has a significant impact on engine performance. Due to icing, the total intake pressure loss coefficient of the engine reaches up to 13.3%, the compressor efficiency value decreases by up to 10.4%, the gas temperature at gas turbine outlet increases by 41 ℃ and the engine power decreases by 15.8% at temperature −30 ℃, altitude 6 km, and velocity 280 km/h. During the icing test, the engine does not show any abnormal phenomena such as stalling or flame-out, which fully verifies the stability and reliability of the engine under adverse icing conditions.