© Faculty of Mechanical Engineering, Belgrade. All rights reserved FME Transactions (2012) 40, 111-118 111 Received: September 2009, Accepted: June 2012 Correspondence to: Vladeta Zmijanović École Polytechnique, 91128 Palaiseau cedex, France E-mail: vladeta.zmijanovic@polytechnique.edu Vladeta Zmijanović Graduate student Fluid Mechanics and Energetics École Polytechnique Boško Rašuo Full Professor University of Belgrade Faculty of Mechanical Engineering Amer Chpoun Full Professor Université Evry Val d'Essonne Laboratoire de Mécanique et d'Energétique Evry (Paris region), France Flow Separation Modes and Side Phenomena in an Overexpanded Nozzle As a part of an aerodynamics Ecole Polytechnique project, separation modes which occur in supersonic nozzles at overexpanded regimes are numerically investigated and compared with known effects. Different shock generation and reflections in different nozzle types are observed and their impact on the two main separation modes, namely Free and Restricted Shock Separation (FSS & RSS) is explored. ONERA’ experimental thust- optimized-contour (TOC) rocket nozzle was the reference case and it is compared with the corresponding Vulcain 2 nozzle and analogues conical and TIC nozzle contours. Strong lateral forces and side effects on the nozzle wall caused by the RSS and by transition FSS to RSS are depicted in terms of flow unsteadiness and presented in flow charts and diagrams. Additionally, some design criteria for modern thrust optimized nozzles according to these effects are proposed. Keywords: launcher nozzle, TOC, TIC, conical contour, overexpanded regime, free, restricted shock separation. 1. INTRODUCTION The nozzle, an end-element of the propulsive process cycle, represents a critical part of any aerospace vehicle. The task of accelerating and efficiently exhausting combusted and reactive gases according to the delivered thrust represents the main objective of the propulsion system design. As such, propulsive convergent- divergent (C-D) axisymmetric rocket nozzles have evolved since the early period of use till nowadays. From technical and historical point of view, it is possible to sort axisymmetric nozzle types by the shape of their divergent profile as conical, ideal - de Laval, truncated ideal contour (TIC) and sorts of modern thrust optimized contour (TOC) nozzles. In the space launcher engine design it is of substantial importance to optimize and control flow end-effects at the conventionally used nozzles. The launcher nozzle, designed to work in extremely wide range of pressure regimes has critical stress points at its lower and upper limits of operation envelope. With the high ambient pressure and thus, lower nozzle pressure ratio (NPR) than the designed one, exhaust plume is over-expanded and it is being recompressed through the shock reflections to the ambient pressure. The nozzle flow under certain circumstances generates critical side- loads on the walls. This mainly occurs when the flow is separated due to an overexpansion and during the start- up transition processes. Different types of flow separation are possible with the different nozzle types as well within the same nozzle. At largely over-expanded regime boundary-layer is detached from the nozzle wall due to an adverse pressure gradient generating the separation shock. This supersonic and separated flow continues downstream and it may interact with the recompression shocks, asymmetric jet portions and/or possibly present internal shock, which can further lead to the severe lateral loads. These lateral or side-loads of uncontrolled flow separation are frequently present during the transient processes of engine start-up and shut-down operation. As the part of the Ecole Polytechnique applied aerodynamics student-project under the guidance of ONERA, the separation effects and nature of such produced flow field are investigated. 2. STATE OF THE ART OF THE PROBLEM IN BRIEF The presented study is motivated by the flow and performance investigations of the contemporary nozzles used in the space launcher technology. This implies use of the thrust optimized and compressed parabolic nozzles. The main reasons for the use of these nozzles are higher thrust performance on the shorter nozzle divergent section length and thus the smaller mass of the launcher propulsion system. Following the principles of the shockless de Laval ideal nozzle, whose contour may be also numerically obtained in respect to the method of characteristics (MoC) procedure [1], it is possible to further optimize this contour in the direction of thrust to mass ratio. Optimization of the ideal nozzle would consider subsonic and supersonic throat section curvature radius but still with a smooth transition and calculation according to Kliegel, Sauer or another method of sonic line as an incipient point for further MoC calculation process. The calculated divergent supersonic profile with marching characteristic until the completely parallel 0deg and uniform exit flow in the first optimal derivation may be truncated according to the maximum delivered thrust. In the usual design process optimal exit angle of the truncated ideal contour (TIC) nozzle profile is found between 4.5 and 7.5 deg but its naturally