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)''.
Vortex Shedding in a Large Solid Rocket Motor Without Inhibitors at the Segment Interfaces
The internal acoustic response of the Titan IV Solid Rocket Motor Upgrade (SRMU) is analyzed using pressure oscillation time histories measured during four static e ring tests. Pressure oscillations for other large solid rocket motors are caused by vortex shedding at annular inhibitors and acoustic feedback resulting from impingement of the vortices on other inhibitors or on the solid rocket motor nozzle. The SRMU does not have inhibitors, but it is shown that acoustic feedback also dee nes its pressure oscillations. Vortices are shed around the cavity between the center and aft segments and impinge on the nozzle entrance. The frequencies of the pressure oscillations vary about that of the motor fundamental acoustic mode and generally agree with a simple empirical relationship that has been used to model acoustic feedback. The effect of the SRMU pressure oscillations on dynamics and control of the Titan IV system is minor. I. Introduction A NEW solid rocket motor design called the Solid Rocket Motor Upgrade (SRMU) will be integrated with the Titan IV system beginning in late 1996. The SRMU has a lower inert weight and higher propellant weight than the current Titan IV solid rocket motor, and therefore, augments system performance. The SRMU has three segments (forward, center, and aft) and is the only large solid rocket motor in production that does not have annular inhibitors at the segment interfaces to prevent combustion and to support the propellant grain during burning. A. SRMU Static Firing Tests Five static e ring tests were conducted for qualie cation of the SRMU. The test articles included one preliminary qualie cation motor (called PQM19) and four qualie cation motors (called QM1, QM2, QM3, and QM4 ). The tests were conducted vertically, nozzle down at the Phillips Laboratory 1-125 Firing Complex, Edwards Air Force Base. 1 The test cone guration and the attachments of the solid rocket motor to the test stand simulated e ight. The physical differences between the e ve test motors were minor and are believed to have had a negligible effect on combustion stability. The static e ring tests were conducted at propellant bulk temperatures in the range 36.5‐ 937F (2.5‐ 33.97C) to assess the effect on motor performance. The static e ring test instrumentation included Condec gauges (CCC-406) of absolute pressure on the forward closure (measurement PCF1 ) and on the aft dome (measurement PCA1), and a Kistler gauge (202M122) of oscillatory pressure (measurement PD4 ) on the forward closure. 2 Measurements PCF1 and PCA1 were recorded by a Neff digital acquisition system, while measurement PD4 was recorded by a Metraplex FM system. The Neff system applied a 100-Hz low-pass e lter and recorded measurements PCF1 and PCA1 at 250 samples/s.