Next Generation Low/Zero Emission Hypersonic Vehicles

With the rapid expansion of space activities and human exploration, along with the evolving demands of next-generation aviation and the space industry, the need for cost-effective and environmentally friendly propulsion technologies has become increasingly critical. While current advancements remain limited, extensive research is being conducted to overcome existing challenges. Hypersonic propulsion systems, including scramjets, ramjets, and hybrid rocket engines, offer promising, cost-effective solutions for sustainable propulsion. To contribute to the advancement of next-generation propulsion systems, this thesis focuses on hypersonic propulsion technologies, specifically scramjets and hybrid rocket systems. This thesis is structured in two parts: the first part is dedicated to experimental research on hybrid rockets, while the second focuses on numerical modeling of supersonic combustion —a combination of propulsive systems that could contribute to next-generation launch applications or for the sustainable aviation. Part I: Experimental Research on Hybrid Rockets The experimental research focuses on developing a novel solid propellant configuration for hybrid rocket engine applications. Hydroxyl-terminated polybutadiene (HTPB) was combined with boron coated with copper(II) oxide (CuO) to enhance energy output and increase the regression rate. Various fuel compositions were formulated, including pure HTPB, HTPB with 15% and 20% boron, HTPB with 15% boron and 2.5% CuO, and HTPB with 20% boron and 5% CuO. A systematic approach was employed to investigate material properties, thermal behavior, combustion characteristics, and ballistic performance. High-Resolution Scanning Electron Microscopy (HRSEM) was utilized to evaluate the effectiveness of ball milling in coating CuO onto boron. Energy Dispersive Spectroscopy (EDS) mapping validated the uniform distribution of these thermite particles within the HTPB matrix. Thermal analysis, including Thermogravimetric Analysis (TGA), Differential Thermal Analysis (DTG), and Differential Scanning Calorimetry (DSC), was conducted to assess the thermal behavior of boron and boron-doped HTPB at elevated temperatures. Additionally, the catalytic effect of CuO on boron combustion was examined. X-ray Diffraction (XRD) analysis was performed to determine the phase composition and crystal structure of the fuel samples before and after combustion. To analyze combustion instabilities in the hybrid rocket engine, frequency–amplitude and frequency–time analyses were conducted. Ballistic performance was evaluated through lab-scale hybrid rocket engine tests using gaseous oxygen as the oxidizer. The performance of different fuel formulations was assessed by measuring regression rates, pressure and thrust. The combustion stability was analysed and overall performance. Part II: Numerical Investigation of the Supersonic Combustion in Scramjet engines The second part of this thesis examines the influence of cavity geometry and fuel injection ratio on air-fuel mixing, flame stabilization, and combustion efficiency in supersonic flows. The HiFIRE-2 scramjet combustor is used as the model for numerical simulations. A hybrid approach combining Reynolds-Averaged Navier-Stokes (RANS) and Large Eddy Simulation (LES) methods, along with a three-step chemical reaction mechanism, is employed to gain deeper insights into combustion dynamics. The study is conducted in two phases. In the first phase, RANS simulations are performed on the baseline HiFIRE-2 geometry to analyze flow fields, scramjet performance, combustion efficiency, and total pressure loss, with comparisons to experimental data. In the second phase, LES simulations are carried out on modified geometry by varying cavity dimensions and the fuel–air equivalence ratio ϕ. The study investigates the role of fuel injection strategies in enhancing streamwise vorticity, analyzing the contributions of the streamwise and span wise components. Additionally, the interaction of sonic-speed fuel injection (transverse and crossflow) with a high-Mach incoming airstream was analsyed on the combustion and total pressure losses and their effects on the dilatational term ∇.U and baroclinic torque were studied, which play a crucial role in supersonic combustion behavior. Understanding the effects of fuel–air equivalence ratio ϕ, cavity geometry, and flow behavior is essential for optimizing combustion performance. The results provided a valuable insight into shock interactions with fuel injection, vorticity development, heat addition, and boundary layer separation, which collectively influence total pressure loss.