To investigate the optimal flight trajectory for hypersonic vehicle and Turbine-Based Combined Cycle (TBCC) engine to enhance mission efficacy, a performance calculation model of a variable-cycle TBCC engine is developed utilizing the method of airframe/engine integrated performance analysis. Employing the Radau pseudospectral method for trajectory optimization, the study of the hypersonic vehicle involves the angle of attack, the scramjet engine equivalence ratio, and the core driven fan stage (CDFS) stator vane angle as control variables. Two optimization scenarios were considered: minimizing climb time and maximizing range. The findings indicate that by optimally adjusting the angle of attack and the scramjet engine’s equivalence ratio, the vehicle’s climb time can be reduced by up to 33.68 %, and the total flight duration by 8.07 %. Moreover, optimizing these variables, along with CDFS stator vane angle, can decrease fuel consumption during climb and potentially extend cruising distance by 8.56 %, increasing overall range by 3.63 %.

Optimizing fuel distribution strategies in scramjets is crucial for efficient supersonic combustion, thermal management, and performance enhancement. Existing research has overlooked the potential to enhance combustion performance while reducing heat load through fundamental fuel distribution strategies. This study investigates the effect of core-to-mainstream fuel distribution on the combustion performance of a core burning scramjet combustor under hypersonic inflow conditions (Mach 2.8, total temperature 1400 K, and total pressure 1.5 MPa) through experiments and numerical simulations. The findings demonstrate that the core burning scramjet combustor achieves self-sustained combustion of kerosene fuel. A fuel distribution strategy with a low core equivalence ratio and high mainstream equivalence ratio enhances combustion efficiency and total pressure recovery coefficient. The detailed flow field analysis reveals low wall heat load characteristics, with high-temperature regions concentrated downstream of the core burning chamber and at the center of the main combustion chamber. By integrating these findings, this research advances the understanding of fuel distribution strategies for the core burning scramjet combustor and offers practical guidance for optimizing combustion performance and thermal management in hypersonic propulsion systems.

This paper presents a comprehensive study on the control law design for the turbine-to-ramjet mode transition in an adaptive-cycle turbine-based combined cycle (TBCC) engine, aiming to mitigate the persistent “thrust gap” challenge. An integrated conceptual configuration of a hypersonic vehicle with a parallel-duct TBCC system, which replaces the conventional turbofan with a three-bypass adaptive cycle engine (ACE), is proposed. High-fidelity performance models for both the ACE and the scramjet are developed, with a Kriging surrogate model employed to accelerate the computationally intensive ACE simulations during the transition. A co-optimization framework is established, defining a comprehensive performance index that balances thrust tracking accuracy and control smoothness under rigorous intake-engine flow matching constraints. Using sequential quadratic programming (SQP), the control schedules for the ACE’s variable geometries are optimized. Comparative analyses reveal that the ACE, with its flexible bypass management and multiple adjustable mechanisms, can actively adapt its airflow demand to match the restricted intake supply. Consequently, the optimized ACE-based TBCC reduces total airflow fluctuation during the Mach 3–3.5 transition from 106% (conventional turbofan baseline) to 42.5%, while maintaining required thrust. This work quantitatively demonstrates the superior flow-handling capability of adaptive cycle technology, providing a viable and effective solution for ensuring stable and efficient mode transition in future hypersonic TBCC propulsion systems.