本研究以涡轮导叶为研究对象，采用压力敏感漆（PSP）实验测量的方式，在保证开孔率一致的前提下，设定密度比（DR）1.4和3.0模拟飞机低空、高空飞行时发动机温比，系统研究了不同质量流率（MFR）下孔径扩大对于前缘气膜冷效的影响。实验结果表明:在DR=1.4和3.0条件下，扩大孔径均导致前缘面平均冷却效率下降，但吸力面与压力面呈现显著差异。吸力面在两种密度比下，面平均冷却效率随孔径增大呈先降后升趋势（DR=1.4:降幅6%，增幅3%；DR=3.0:降幅17%，增幅4%）；压力面在DR=1.4时冷却效率随孔径增大持续降低（降幅13%–16%），而在DR=3.0时表现为先升后降（增幅10%，降幅11%）。研究揭示了高/低密度比工况下孔径与气动条件的非线性耦合机制，为空天发动机前缘冷却结构的孔径优化设计提供了实验依据。

This study investigates the impact of hole diameter enlargement on the leading-edge film cooling effectiveness of turbine guide vanes through pressure-sensitive paint (PSP) experiments under fixed porosity conditions. By simulating engine thermal environments during low-altitude (DR=1.4) and high-altitude flights (DR=3.0) with varying mass flow ratios (MFR), systematic measurements reveal that increased hole diameters consistently reduce area-averaged cooling effectiveness across both density ratios. However, distinct surface-specific behaviors emerge: The suction surface exhibits a non-monotonic response, with initial cooling efficiency reductions (6% at DR=1.4; 17% at DR=3.0) followed by partial recovery (3% and 4% improvements, respectively). Conversely, the pressure surface demonstrates divergent trends—monotonic degradation at DR=1.4 (13%-16% reduction) versus an initial enhancement (10% improvement) succeeded by deterioration (11% reduction) at DR=3.0. These findings elucidate the nonlinear interactions between geometric scaling and aerodynamic conditions under different operational regimes, providing critical experimental evidence for adaptive hole diameter optimization in next-generation aerospace engine cooling systems.