Center of rotation estimation for rocket nozzle by infrared reflective makers and image sequences

The determination of an accurate center of rotation of rocket motor nozzle or other object to be measured is of great interest across a wide range of applications,such as rocket,missile,robotics,industry,spaceflight,aviation and human motion analysis fields,particularly for clinical gait analysis.A new approach was proposed to estimate the moving objects' instantaneous center of rotation and other motion parameters.The new method assumes that the two segment of object to be measured are rigid body which rotates around a center of rotation between each other relatively.The center of rotation varies with time in the global coordinate system but is fixed in the local coordinate system attached to each segment.The models of rocket motor nozzle and its movement were established.The arbitrary moving object's corresponding to motion equations were deduced,and the least square closed-form solutions of the object's motion parameters were figured out.It is assumed that the two high speed CCD cameras mounted on the 750 nm infrared(IR) filter are synchronized and calibrated in advance.The virtual simulation experiment using 3D coordinates of markers was conducted by synchronized stereo image sequences based on 6-DOF motion platform and the experimental results prove the feasibility of our algorithm.The test results show that the precision of x,y,z component on center of rotation is up to 0.14 mm,0.13 mm,0.15 mm.

Investigation on the mechanism of aeroelastic hazard during ground test of rocket nozzle

Side loads and aeroelastic stability of rocket nozzle were studied by solving Navier-Stokes equation coupled with structural equation of motion.The computation was implemented at different total pressure inlet conditions,and flow phenomena of free shock separation(FSS) and restricted shock separation(RSS) were captured.At certain total pressure inlet conditions,it was found that both kinds of separations existed in nozzle flow filed,while RSS exhibited combined space asymmetry and time unsteady characteristics.The corresponding asymmetric circumferential pressure distribution,strong pressure fluctuation in separation region and large range of displacement of shock wave all led to severe side loads.Besides,for flexible nozzles,the low pressure gradient in separation region might reduce structure stability at nozzle exit,resulting in large local deformation.It was also found that aeroelasticity exhibited buffeting characteristic due to the asymmetric separation,resulting in reduction of aeroelastic stability,even structure destruction.Moreover,aeroelasticity might amplify side loads and aggravate its growth rate.However,with increment of inlet pressure,nozzle aeroelastic stability was also increased when a full flow was nearly reached.

THE RESEARCH ON HIGH-SPEED GAS JET OF ROCKET NOZZLE UNDERWATER

In this paper, a study of the high-speed gas jet of a rocket nozzle underwater was carried out using commercially available CFD software FLUENT with it’s user-defined-function. The volume of fluid technique based on finite volume method was adopted to solve the time-dependent multiphase flow including a compressible phase, and the PISO algorithm was included. The computed results show that this problem was calculated successfully. The gas bubble behind the nozzle, and the wave structure existing in highly compressed gas in water were captured accurately.

Analysis and control of the compaction force in the composite prepreg tape winding process for rocket motor nozzles

In the process of composite prepreg tape winding,the compaction force could influence the quality of winding products.According to the analysis and experiments,during the winding process of a rocket motor nozzle aft exit cone with a winding angle,there would be an error between the deposition speed of tape layers and the feeding speed of the compaction roller,which could influence the compaction force.Both a lack of compaction and overcompaction related to the feeding of the compaction roller could result in defects of winding nozzles.Thus,a flexible winding system has been developed for rocket motor nozzle winding.In the system,feeding of the compaction roller could be adjusted in real time to achieve an invariable compaction force.According to experiments,the force deformation model of the winding tape is a time-varying system.Thus,a forgetting factor recursive least square based parameter estimation proportional-integral-differential（PID）controller has been developed,which could estimate the time-varying parameter and control the compaction force by adjusting the feeding of the compaction roller during the winding process.According to the experimental results,a winding nozzle with fewer voids and a smooth surface could be wounded by the invariable compaction force in the flexible winding system.

Control of the Asymmetric Flow in Rocket Nozzles

In some rocket nozzle flows, the existence of the transition from FSS to RSS and the occurrence of asymmetric flow are known in previous researches. As a result, the transition causes excessive side-loads that may damage the nozzle. Thus, it is important to investigate the method in order to control the asymmetric flow separation. In the present study, the relationship between the asymmetric separation and the rate of change of the pressure ratio with time was investigated from the point of view of the transition from FSS to RSS in the supersonic nozzle experimentally. Further, change of the flow separation by using step and cavity, and the possibility of the control was demonstrated. As a result, it was shown that the method using a cavity was effective for the control of the separation pattern.

Estimation of Exhaust Gas Temperature of the Rocket Nozzle using Hybrid Approach

In this paper the theoretical model is built for ZEpHyR(ZARM Experimental Hybrid Rocket) main engine which is being developed at ZARM institute,Bremen,Germany.The theoretical model is used to estimate the temperature of exhaust gas.The Conjugate Gradient Method(CGM) with Adjoint Problem for Function Estimation iterative technique is used to solve the Inverse Heat Conduction Problem(IHCP) to estimate the heat flux and internal wall temperature at the throat section of the nozzle.Bartz equation is used to calculate the convective heat transfer coefficient.The exhaust gas temperature is determined using the estimated heat flux,the wall temperature at internal surface of nozzle and the heat transfer coefficient.The accuracy of CGM iterative scheme to solve the IHCP is also investigated and its results are presented.

Numerical simulation of multi-phase combustion flow in solid rocket motors with metalized propellant

Multi-phase flowfield simulation has been performed on solid rocket motor and effect of multi-phases on the performance prediction of the solid rocket motor(SRM) is investigation.During the combustion of aluminized propellant,the aluminum particles in the propellant melt and formliquid aluminum at the burning propellant surface.So the flow within the rocket motor is multi phase or two phase because it contains droplets and smoke particles of Al2O3.Flowsi mulations have been performed on a large scale motor,to observe the effect of the flowfield onthe chamber and nozzle as well.Uniform particles diameters and Rosin-Rammler diameter distribution method that is based on the assumption that an exponential relationship exists betweenthe droplet diameter,dand mass fraction of droplets with diameter greater thandhave been used for the si mulation of different distribution of Al2O3 droplets present in SRM.Particles sizes in the range of 1-100μm are used,as being the most common droplets.In this approachthe complete range of particle sizes is dividedinto a set of discrete size ranges,eachto be defined by single streamthat is part of the group.Roe scheme-flux differencing splitting based on approxi mate Riemann problem has been used to si mulate the effects of the multi-phase flowfeild.This is second order upwind scheme in which flux differencing splitting method is employed.To cater for the turbulence effect,Spalart-All maras model has been used.The results obtained show the great sensitivity of this diameters distribution and particles concentrations to the SRMflowdynamics,primarily at the motor chamber and nozzle exit.The results are shown with various sizes of the particles concentrations and geometrical configurations including models for SRM and nozzle.The analysis also provides effect of multi-phase on performance prediction of solid rocket motor.

Computational Study of Performance Characteristics for Truncated Conical Aerospike Nozzles

Aerospike nozzles are advanced rocket nozzles that can maintain its aerodynamic efficiency over a wide range of altitudes. It belongs to class of altitude compensating nozzles. A vehicle with an aerospike nozzle uses less fuel at low altitudes due to its altitude adaptability, where most missions have the greatest need for thrust. Aerospike nozzles are better suited to Single Stage to Orbit （SSTO） missions compared to conventional nozzles. In the cur- rent study, the flow through 20% and 40% aerospike nozzle is analyzed in detail using computational fluid dy- namics technique. Steady state analysis with implicit formulation is carried out. Reynolds averaged Navier-Stokes equations are solved with the Spalart-AUmaras turbulence model. The results are compared with experimental results from previous work. The transition from open wake to closed wake happens in lower Nozzle Pressure Ratio for 20% as compared to 40% aerospike nozzle.

Computational study on flow through truncated conical plug nozzle with base bleed

Conical plug nozzle and truncated conical plug nozzle are advanced rocket nozzles suitable for use as altitude compensating nozzles.In this study flow through the conical plug and truncated conical plug nozzles are numerically simulated to first validate with experimental data and then to compare the performance when a base bleed is introduced.The numerical analysis has considered two-dimensional axisymmetric models.Reynolds-averaged NavierStokes equations are solved with two equation shear stress transport k-ω turbulence model.For the validation of the plug nozzle,flow features and wall pressure along the length of the nozzle is taken for different nozzle pressure ratios.For the validation of truncated plug nozzle,flow features and base pressures at various nozzle pressure ratios are compared.The base bleed is taken as 2%of the inlet mass flow rate.The comparison of results shows that the introduction of base bleed helps to compensate for the loss of thrust due to conical plug nozzle truncation.

Multi-phase flow effect on SRM nozzle flow field and thermal protection materials

Multi-phase flow effect generated from the combustion of aluminum based composite propellant was performed on the thermal protection material of solid rocket motor(SRM) nozzle.Injection of alumina(Al2O3) particles from 5% to 10% was tried on SRM nozzle flow field to see the influence of multiphase flow on heat transfer computations.A coupled,time resolved CFD(computational fluid dynamics) approach was adopted to solve the conjugate problem of multi-phase fluid flow and heat transfer in the solid rocket motor nozzle.The governing equations are discretized by using the finite volume method.Spalart-Allmaras(S-A) turbulence model was employed.The computation was executed on the different models selected for the analysis to validate the temperature variation in the throat inserts and baking material of SRM nozzle.Comparison for temperatures variations were also carried out at different expansion ratios of nozzle.This paper also characterized the advanced SRM nozzle composites material for their high thermo stability and their high thermo mechanical capabilities to make it more reliable simpler and lighter.