内转式进气道拥有较好的来流捕获能力和较高的压缩效率，但此类进气道在初步设计时无法进行局部型面调节，难以改善激波/边界层干扰导致的流动分离、二次流等复杂流动结构对进气道性能造成的不利影响，仍具有较大的优化设计空间。目前，针对高超声速内转式进气道开展气动优化设计面临着型面复杂多变、设计变量规模较大、流场求解精度要求高等难题。为此，采用基于离散伴随的梯度类优化方法，对带斜楔前体内转式进气道开展了气动优化设计。优化结果表明，进气道内外压缩段型面的起伏变化显著改变了内部激波结构，减小了壁面压力梯度，进而弱化了流向涡；隔离段内激波串与附面层的干扰强度显著减弱，抑制了低能流区的扩张。相比于初始构型，优化构型在设计工况下出口处的总压恢复系数提升了8.767%，流量系数提升了0.163%，增压比提升了0.763%，阻力降低了1.658%，进气道的气动性能得到了一定改善。

Despite its relatively good inlet capture ability and high air compression efficiency， the inward turning inlet cannot be locally adjusted in the preliminary design， and improvement in the adverse effects of complex flow structures such as flow separation and secondary flow caused by shock/boundary layer interference is difficult. Therefore， performance optimization is necessary. The current design optimization of the hypersonic inward turning inlet faces many challenges such as the complex profile， large-scale design variables， and the high accuracy requirement of the flowfield numerical simulation. The gradient optimization method based on discrete adjoint is adopted to carry out the aerodynamic design optimization for the inward turning inlet with a wedge forebody. The optimization results show that the undulating shape of the inner and outer compression sections significantly changes the internal shock structure， reduces the wall pressure gradient， and thereby abates the streamwise vortex. In addition， the interference intensity between the shock train and the boundary layer in the isolation section is significantly weakened， inhibiting the expansion of the low energy flow region. Under the design condition， compared with the initial configuration， the aerodynamic performance of the optimized configuration is significantly improved. The total pressure recovery coefficient， the flow coefficient and the pressure ratio at the exit are increased by 8.767%， 0.163%， and 0.763%， respectively， while the drag is reduced by 1.658%.