Prediction And Suppression Methods Of Combustion Instability In Large Aspect Ratio Solid Rocket Motors

Large aspect ratio solid rocket motors with high loading of composite solid propellanthave been widely equipped in space propusion and missile weapon system. Howerver, thesekinds of solid rocket motors are prone to combustion instability during the working process,expecially at the end of burning. Combustion instabitliy can lead to the vibration of motorstructure and reduce the reliability of the missile. In serious conditions, it can result in thefailure of missile mission. Based on large aspect ratio solid rocket motor combustioninstability phenomenon, the key driving and damping characteristics of combustioninstability in solid rocket motors are numerically and experimentally studied in this paper.The research work aims at providing theoretical guidance for the prediction andsuppression of combustion instability in large aspect ratio solid rocket motors. The mainresearch work and conclusions are as follows：Based on the VKI (Von Karman Institute for Fluid Dynamics) experimental motor, theobstacle vortex shedding driven pressure oscillations are numerically studied via the LES(Large Eddy Simulation) method. The fluid structure and pressure oscillation characteristicsare obtained by numerical simulation, and the numerical results satisfy well with theexperimental value. The effects of inhibitor position on vortex acoustic pressure oscillationcharacteristics are studied by continuously changing the inhibitor position. Results indicatethat vortex shedding is a periodic process and its accurate frequency can be numericallyobtained. Inhibitor position has little effect on vortex shedding frequency, but has a greatimpact on pressure oscillation amplitude. Pressure amplitude is much larger when theinhibitor locates at the acoustic velocity anti nodes (1/4L,1/2L,3/4L). The farther theinhibitor to nozzle head, the more the vortex energy will be dissipated by the turbulence.Therefore, the vortex shedding amplitude at the second acoustic velocity anti node near3/4L is larger than that of others.The effect of gas temperature on vortex acoutic pressure oscillation characteristic isnumerically studied via changing the gas temperature. Results show that gas temperaturehas little effect on vortex shedding frequency, but it can greatly change the natural acousticfrequency of chamber. As the vortex shedding frequency departs from the natural acousticfrequency, the vortex acoustic feedback loop is decoupled. Consequently, both the vortexshedding and acoustic amplitudes decrease rapidly. A new type of T burner grain structure is designed. The pressure coupled responsefunction of aluminized AP/HTPB composite solid propellant is experimentally studiedbased on T burner. Two repeated tests with one controlled test are carried out in this study,of which the first two tests use same propellant sample, while the propellant formulation isslightly adjusted in the controlled test. The first two experimental results show that thesample propellant has a high response function value under the conditions of high pressure(~11.5MPa) and low frequency (~140Hz). The controlled test result of response functionvalue is half of previous value, which indicates that propellant formulation has an importantinfluence on the pressure coupled response function. The pressure oscillationcharacteristics along with the acoustic pressure distribution in T burner are also numericallystudied. The numerical results of pressure oscillation frequency and acoustic pressuredistribution are in consistence with experimental resluts.The nozzle damping characteristics and the cavity response on pressure oscillations areevaluated by a pulsed method. Resluts indicate that increasing nozzle throat area to portarea (J) along with reducing chamber length and convergent half angle are of benefit tonozzle damping. Convex nozzle provides more damping than the conical nozzle, which inturn provides more damping than the concave nozzle. Submerged nozzle can reduce nozzledamping, and nozzle damping rate is approximately linear with the submerged nozzlecavity volume. Cavity response analysis shows that head end cavity is contributed tosuppressing combustion instability. Middle cavity can easily induce vortex driven pressureoscillations. Convergent middle cavity is better than divergent middle cavity for the motorstability. After end cavity is harmful to motor stability because the aft end cavity can lead tothe reduction of nozzle throat area to port area (J).Based on the M6solid rocket motor with finocyl grain, the acoustic changingcharacthersitcs during the motor working process are analyzed by FEA (Finite ElementAnalysis) method. LES (Large Eddy Simulation) method is utilized to study the flow fliedand pressure oscillation characteristics of the chamber. Finally, the pressure oscillationmechanism at the end of burning is explained via linear combustion instability theory.Results indicate that the first and second axial acoustic frequencies firstly decrease and thenincrease with the burning time. At the end of buring, the dirving factor of pressure coupled response is approaching to nozzle damping factor. Therefore, the M6motor turns from alinear steady state to linear unsteady state, and the stability is greatly decreased. Thus,combustion instability occurs.