Session: GT9.1 - Analytics & Digital Solutions and GT6.2 - Emerging Technologies (includes Wind Energy)

Paper Number: 160384

160384 - Nozzle Flow Separation in Thrust Optimized Contour Bell Nozzle 

In rockets, the nozzle undergoes over-expanded conditions during the transient start-up and shutdown of the launch vehicle. This results in a critical phenomenon known as flow separation which involves complex phenomena creating thrust loss. This study investigates the flow separation behaviour in a thrust-optimized contoured (TOC) bell nozzle with an area ratio of 36 focusing on the transition between Free-Shock Separation (FSS) and Restricted-Shock Separation (RSS) during the rocket engine transient start-up and transient shutdown phases. Numerical methods were employed to analyse the flow field by highlighting the gas-dynamic features and hysteresis effects. The numerical simulations were performed using a two dimensional density-based solver in ANSYS Fluent, with turbulence modelled by the SST K - ῳ model. The separation structures and shock interactions were visualised through numerical Schlieren imaging. The study observed that the transition from FSS to RSS occurred at a nozzle pressure ratio (NPR) of 35.6 with significant differences in transition behaviours during the engine transient start-up and transient shutdown phases. The analysis revealed that critical non-isentropic shock features including internal shocks, reflected shocks, Mach stems, triple point, recirculation bubbles, and the formation of cap-shock patterns, strongly influence RSS formation. This study's key findings include the influence of trapped vortex on shear layer growth, quantifying hysteresis effects and highlighting discontinuities in separation points during transient conditions. Additionally, the results reveal the strong dependence of RSS formation on the nozzle divergent geometry design. This study provides valuable insights into the dynamics of flow separation in TOC nozzles and offers guidance for optimising nozzle designs to improve rocket engine performance.

Presenting Author: G Lalitha Maheswari Indian Institute of space science and technology

Presenting Author Biography: G Lalitha Maheswari
M.Tech Candidate, Department of Thermal and Propulsion,
Indian Institute of Space Science and Technology (IIST), Thiruvananthapuram

G Lalitha Maheswari is currently pursuing a Master of Technology in Thermal and Propulsion at the Indian Institute of Space Science and Technology (IIST), Thiruvananthapuram. She holds a bachelor’s degree in aerospace engineering from Bharath University, Chennai. During her undergraduate studies, she worked on an experimental, analytical, and numerical investigation of various geometric nozzles, focusing on shock structures, thrust performance, and centroidal axis pressure measurements.

Before joining IIST, she gained valuable industry experience as an Illustrator at Boeing and a Graduate Apprentice Trainee at Hindustan Aeronautics Limited (HAL), Bangalore. Her current research focuses on the numerical and experimental investigation of flow separation in thrust-optimized contoured nozzles, specifically emphasizing hysteresis effects, side-load phenomena, and