Measurement technique for the determination of the deformation of a flexible thermal protection ring inside solid fuel booster rocket
A new generation of the solid booster rockets of the Ariane 5 launcher is being under construction to improve stability and controlling characteristics of the vehicle. One crucial element of solid rocket boosters is the separator ring that permits the division of the total quantity of solid fuel into well defined, independent stages. Up to now, these separator rings introduced some unwanted perturbations in the operation of the launcher, thus they are replaced by a flexible element, the characteristics of which are to be investigated. The flexible thermal protection ring is able to change its form due to the gas flow inside the booster. The form is changed partially due to deformation induced by the flow, and by erosion effects due to the burning fuel. At the present investigation the flow induced deformation is measured. The measurements were carried out in a facility that does not take into account the effects of burning, thus optical access was possible. An optical image treatment based technique was developed and applied for the determination of the deformation of the thermal protection ring for various flow velocities. The method was found to be accurate, robust and reliable, thus it can be used as a reference measurement technique for the calibration of other type of nonoptical sensors that can be used later also in fuel burning conditions. This technique is discussed in this publication in detail. Keywords—Fluid structure interaction, optical technique, Ariane 5
Cold flow investigation of the flow acoustic coupling in solid propellant boosters
This paper deals with an analysis of the flow instabilities in the Ariane 5 solid propellant booster. They may occur as vortices shed from partiallyburned inhibitors can excite one of the acoustic modes of the cavity, leading to a flow-acoustic coupling. The fundamental aspects of axial flow instabilities promoted by two obstructions in a cylindrical duct ended by a nozzle and their interference with the acoustic field are firstly investigated. Resonant pressure fluctuations occur at several velocities coinciding to the existence of an integer number of vortices between the obstacles, with no more than 4 vortices observed. After this, the flow-acoustic coupling is studied in a setup respecting the geometrical and dynamic similarity with the real booster at a 1/15* scale with radial flow simulation and sonic condition at the nozzle. Vortex shedding is characterized by a constant Strouhal number of 0.6 but its frequency remains too far from the acoustic ones for flow-acoustic coupling to be possible. The use of porous walls for simulating the radial flow affects the acoustic field and the improvement of the side injection technique is under way at VKI. The nature of the flow in the combustion chamber is complex since it generates radially from the burning surface and then develops longitudinally to reach the exhaust nozzle where it interacts with a very complex geometrical shape. The three segments are separated by two inhibitor rings which, during combustion, may protrude into the flow-field giving rise to a possible vortex shedding (figure 2). So, flow instabilities may occur as vortices shed from partially-burned propellant inhibitors can excite one of the acoustic mode of the cavity, leading to a resonant flow acoustic coupling. Excitation of acoustic modes by flow instabilities could result in significant pressure fluctuations in the combustion chamber, with the consequent adverse effect upon stability of the combustion itself and upon rocket performances. This exhibition of pressure and thrust oscillations is already recorded in the literature for rockets of similar dimensions (U.S. Space Shuttle and Titan SRM)''.
Parietal vortex shedding as a cause of instability for long solid propellant motors - Numerial simulations and comparisons with firing tests
This paper describes a new mechanism leading to vortex shedding instabilities in long (large L/D) solid propellant motors. This mechanism is termed "parietal vortex shedding" and has been discovered thanks to numerical simulations of the unsteady, 2D, compressible, Navier-Stokes equations. It seems to involve hydrodynamic instabilities of the mean flow velocity profiles, corresponding to injection induced internal flow (so called Culick or Taylor profiles), that couple with the acoustic frequencies of the chamber. Although this mechanism is found to be very powerful, it seems to need some background noise to feed it. The presented simulations can explain observed instabilities in configurations, without segmentation or without protruding inhibitor rings, of a simplified subscale setup of the Ariane 5 MPS P230 solid propellant motor. Detailed comparisons are proposed and the influence of the propellant combustion response and of the turbulence of the flow field are analyzed by means of recently developed models. INTRODUCTION-OBJECTIVES not precisely known if inhibitor rings are or not destroyed or completely slack; on the other hand non segmented motors have also given rise to a same type of instability as vortex shedding : this is the case of one of the configuration (LP3D) of the LP3 set up presented in a preceding paper. As it is known ' ' , classical linear acoustic balance computations do not give reliable stability predictions in complex internal geometries (such as in the P230 motor), and then an effort was carried out to perform full numerical simulations of the unsteady, compressible internal flow fields. On the other hand, motor internal flows are mostly non-observable, and without numerical simulations it is not possible to describe the path of the aerodynamic instability development. The object of the present paper is to show hydrodynamic instabilities (which drive pressure oscillations) in configurations without shear flows induced by inhibitor rings, and to explain how these instabilities occur by means of numerical simulations. LP3D AND LP3E TEST CASES This work is part of the overall research effort, supported by CNES, accompanying the development of the Ariane 5 P230 MPS solid propellant motor (program ASSM for Aerodynamics of Segmented Solid Motors) and makes use of experimental results obtained during the combustion stability assessment program carried out for BPD and CNES, by delegation of ESA. In this scope, it is the continuation of earlier works about numerical simulations in solid propellant rocket motors ". The Ariane 5 motor, as other large segmented motors (U.S. Space Shuttle and Titan SRM) has been reported to exhibit pressure and thrust oscillations. Until recently, it was believed that such instability was exclusively due to the segmented design : inhibitor rings induce shear layers and vortex shedding driven oscillations. Nevertheless it is * Research scientist, Energetics Dept. 1 Project manager, Energetics Dept., Member AIAA Copyright © 1996 by the American Institute of Aeronautics and Astronautics, Inc., All rights reserved. The test cases are based on two configurations of the LP3 motor. The LP3 motor, which has been presented in reference , is a simplified 1/15 subscale set-up of the Ariane 5 P230 motor used to test several inter-segment arrangements. The configurations of interest here, called LP3D and LP3E, have no prominent obstacles at the mid chamber point, see figure 1. Both configurations have a cylindrical chamber of length 1632 mm and inner diameter 203 mm. At ignition, the main propellant burning surface is cylindrical, inner diameter 90 mm, with a chamfered surface near the aft end. The supersonic outlet nozzle is submerged, with a throat diameter of 56.5 mm, then the total length of the motor is 1650 mm. LP3D and LP3E have a small forward segment of 225 mm length, cylindrical for LP3D (169.5