First Systematic Testing Platform for Pressurization Feed System Developed for Liquid Rocket Propellant in China

Beijing Aerospace System Engineering Institute of China Academy of Launch Vehicle Technology (CALT) declared recently that theinstitute has set up a laboratory whichwould operate a newly

Identification of Backflow Vortex Instability in Rocket Engine Inducers

Bayesian estimation is applied to the analysis of backflow vortex instabilities in typical three-and four bladed liquid propellant rocket(LPR)engine inducers.The flow in the impeller eye is modeled as a set of equally intense and evenly spaced 2D axial vortices,located at the same radial distance from the axis and rotating at a fraction of the impeller speed.The circle theorem and the Bernoulli’s equation are used to predict the flow pressure in terms of the vortex number,intensity,rotational speed,and radial position.The theoretical spectra so obtained are frequency broadened to mimic the dispersion of the experimental data and parametrically fitted to the measured pressure spectra by maximum likelihood estimation with equal and independent Gaussian errors.The method is applied to three inducers,tested in water at room temperature and different loads and cavitation conditions.It successfully characterizes backflow instabilities using the signals of a single pressure transducer flush-mounted on the casing of the impeller eye,effectively by-passing the aliasing and data acquisition/reduction complexities of traditional multiple-sensor cross correlation methods.The identification returns the estimates of the model parameters and their standard errors,providing the information necessary for assessing the accuracy and statistical significance of the results.The flowrate is found to be the major factor affecting the backflow vortex instability,which,on the other hand,is rather insensitive to the occurrence of cavitation.The results are consistent with the data reported in the literature,as well as with those generated by the auxiliary models specifically developed for initializing the maximum likelihood searches and supporting the identification procedure.

A New Simulation Method for 3D Propellant Grain Burn Analysis of Solid Rocket Motor

Describes a new computer program (Regress-3D) to simulate the regression of complex 3D grain cavity and calculate the burning surface area. It has a large region of applicability in solid rocket motor design and has made new improvements compared with other available codes. User can easily and rapidly build his initial grain shapes and then obtain geometric information of his design. Considering with the calclulting results, redesigning can be performed as desire until reaching at the satisfied result. Advantages and disadvantages of this method are also discussed.

The properties of Sn-Zn-Al-La fusible alloy for mitigation devices of solid propellant rocket motors

The Al and La elements are added to the Sn9Zn alloy to obtain the fusible alloy for the mitigation devices of solid propellant rocket motors. Differential scanning calorimetry(DSC), metallographic analysis,scanning electron microscopy(SEM), energy dispersive spectroscopy(EDS), tensile testing and fracture analysis were used to study the effect of Al and La elements on the microstructure, melting characteristics, and mechanical properties of the Sn9Zn alloy. Whether the fusible diaphragm can effectively relieve pressure was investigated by the hydrostatic pressure at high-temperature test. Experimental results show that the melting point of the Sn9Zn-0.8Al0·2La and Sn9Zn-3Al0·2La fusible alloys can meet the predetermined working temperature of ventilation. The mechanical properties of those are more than 35% higher than that of the Sn9Zn alloy at-50°C-70°C, and the mechanical strength is reduced by 80% at 175°C. It is proven by the hydrostatic pressure at high-temperature test that the fusible diaphragm can relieve pressure effectively and can be used for the design of the mitigation devices of solid propellant rocket motors.

High-energy Metal Fuels for Rocket Propulsion： Characterization and Performance

A joint international effort to improve solid propellant performance within the framework of a FP7European Project was described.Several metallized solid rocket propellants,of the broad family AP/HTPB/Metal in the ratio 68/14/18,were experimentally analyzed seeking to optimize the delivered specific impulse by identifying the most suitable high-energy fuel.Keeping the same nominal composition,different metallic fuels(including micrometric and nanometric Al,AlH3,and a variety of dual metal compositions)were characterized,tested,and contrasted to a conventional micrometric aluminum(30μm average grain size)certified for space flights.In order to overcome the intrinsic performance limitations of the matrix AP/HTPB,a new matrix consisting of ADN/GAP satisfying also the need for environmentally benign propellant formulation was considered as well.A comparative analysis between the two solid propellant systems in terms of ideal thermochemistry and experimental combustion properties reveals advantages and disadvantages of both.Overall,it is judged worthwhile to develop ADN/GAP propellants,with or without metallic fuels,to enhance the current status of solid rocket propulsion.Controlling morphology and mechanical properties of ADN/GAP compositions and understanding their flame structure and aggregation/agglomeration properties are the main issues still challenging industrial users.

Experimental Investigation on Basic Prototype of Solid Propellant Impulsive Microthrusters

A new type of impulsive microthruster and its measurement system were designed for the aim of testing the performance of a basic prototype of solid propellant impulsive microthruster. Two sets of tests were conducted. The tests show that the ignitor and the main charge of the microthruster match well, the dynamic and static capability of the test and measurement meets the test requirement and the result is creditable. The measured technical characteristics of the microthruster are that the ignition delay time is shorter than 0 3?ms, the total impulse is over 3?N·s, the operational time is shorter than 16?ms and the mass ratio of the thruster is 0 216.

Experimental Investigation of the Flame Propagation and Flashback Behavior of a Green Propellant Consisting of N2O and C2H4

Regarding the research on alternatives for monopropellant hydrazine, several so called green propellants are currently under investigation or qualification. Aside others, the DLR Institute of Space Propulsion investigates a N20/C2I-I4 premixed green propellant. During the research activities, flashback from the rocket combustion chamber into the feeding system has been identified as a major challenge when using the propellant mixture. This paper shows the results of ignition experiments conducted in a cylindrical, optical accessible ignition chamber. During the ignition and flame propagation process, pressure, temperature and high-speed video data were collected. The high speed video data were used to analyze the flame propagation speed. The obtained propagation speed was about 20 rn/s at ignition, while during further propagation of the flame speeds of up to 120 m/s were measured. Additionally, two different porous materials as flame arresting elements were tested： Porous stainless steel and porous bronze material. For both materials Peclet numbers for flashback were derived. The critical Peclet number for the sintered bronze material was around 20, while for the sintered stainless steel the critical Peclet number seems to be larger than 40. Due to the test results, sintered porous materials seem to be suitable as flashback arresters.

Cooling process of liquid propellant rocket by means of kerosene-alumina nanofluid

Heat transfer augmentation of kerosene-alumina nanofluid is studied for the possible use in the regenerative cooling channel of semi cryogenic engine.The basic partial differential equations are reduced to oidinary differential equations which are solved using differential transformation method.Velocity and temperature profiles as well as the skin friction coefficient and Nusselt number are determined.The influence of pertinent parameters such as nanofluid volume fraction,viscosity parameter and Eckert number on the flow and heat transfer characteristics is discussed.Tbe results indicate that adding alumina into the fuel of liquid rocket engine(kerosene)can be considered as the way of enhancing cooling process of chamber and nozzle walls.Nusselt number is an increasing function of viscosity parameter and nanoparticle volume fraction while it is a decreasing Junction of Eckert number.

Numerical Simulation of a Liquid Propellant Rocket Motor

This work presents a numerical simulation of the flow field in a liquid propellant rocket engine chamber and exit nozzle using techniques to allow the results to be taken as starting points for designing those propulsive systems. This was done using a Finite Volume method simulating the different flow regimes which usually take place in those systems. As the flow field has regions ranging from the low subsonic to the supersonic regimes, the numerical code used, initially developed for commpressible flows only, was modified to work proficiently in the whole velocity range. It is well known that codes have been developed in CFD, for either compressible or incompressible flows, the joint treatment of both together being complex even today, given the small number of. references available in this area. Here an existing code for compressible flow was used and primitive variables, the pressure, the Cartesian components of the velocity and the temperature instead of the conserved variables were introduced in the Euler and Navier-Stokes equations. This was done to penult the treatment at any Mach number. Unstructured meshes with adaptive refinements were employed here. The convective terms were treated with upwind first and second order methods. The numerical stability was kept with artificial dissipation and in the spatial coverage one used a five stage Runge-Kutta scheme for the Fluid Mechanics and the VODE (Value of Ordinary Differential Equations) scheme along with the Chemkin II in the chemical reacting solution. During the development of this code simulating the flow in a rocket engine, comparison tests were made with several different types of internal and external flows, at different velocities, seeking to establish the confidence level of the techniques being used. These comparisons were done with existing theortical results and with other codes already validated and well acceptal by the CFD community.

Dynamic interaction between clustered liquid propellant rocket engines under their asynchronous start-ups

A nonlinear mathematical model of the low-frequency dynamics of the clustered multi-engine rocket propulsion system has been developed and the computations of the engine transient processes during the start-ups of the four-engine propulsion system with a shared feed system have been made applied.Based on propulsion system start-up modeling the influence of the connectivity of engines in a cluster on the starting characteristics of individual engines is shown.In particular,an advanced nonlinear mathematical model of the pump cavitation phenomena is a distinctive feature of the mathematical model.The computation results showed that the asynchronous engines start-ups during rocket lift-off lead to severely nonlinear engine transients and clustered engine thrust misbalance.The influence of the rocket engines asynchronous start-ups on the clustered feed system transients depends on many factors,mainly on from the clustered feed system low-frequency dynamics,the magnitude of the disturbance and the phase difference between disturbances acting on different branches of the feed system.The deep lingering dips in the flow rate and pressure transients are possible due to the nonlinear dynamic interaction of the engines.In case of great pressure dips at the pump inlet(up to the pressure of saturated vapors during significant periods of start-up time)the cavitation breakdowns of the pumps of one or more engines from the cluster are possible.This can disrupt the operation of the entire propulsion system and leads to the failure of the launch vehicle mission.

空化诱导轮内流动不稳定性的最大似然估

The article illustrates the application of Bayesian estimation to the identification of flow instabilities,with special reference to rotating cavitation,in a three-bladed axial inducer using the unsteady pressure readings of a single transducer mounted on the casing just behind the leading edges of the impeller blades.The typical trapezoidal pressure distribution in the blade channels is parametrized and modulated in time and space for theoretically reproducing the expected pressure generated by known forms of cavitation instabilities(cavitation auto-oscillations and higher-order surge cavitation modes,n-lobed subsynchronous/synchronous/super-synchronous rotating cavitation).The Fourier spectra of the theoretical pressure so obtained in the rotating frame are transformed in the stationary frame,frequency broadened to better approximate the experimental results,and parametrically fitted by maximum likelihood estimation to the measured auto-correlation spectra.Each form of instability generates a characteristic distribution of sidebands in addition to its fundamental frequency.The identification makes use of this information for effective detection and characterization of multiple simultaneous flow instabilities with intensities spanning over about 20 db down to about 4 db signal-to-noise ratios.The same information also allows for effectively bypassing the aliasing limitations of traditional cross-correlation methods in the discrimination of multiple-lobed azimuthal instabilities from the measurements returned by arrays of equally spaced sensors.The method returns both the estimates of the model parameters and their standard deviations,providing the information needed for the assessment of the statistical significance of the results.The proposed approach represents therefore a promising tool for experimental research on flow instabilities in high-performance turbopumps.

Transient flow characteristics and performance of a solid rocket motor with a pintle valve

The pintle valve is currently the most promising technology among all thrust control methods for solid rocket motors.Pintle structure and working condition play a critical role in the successful operation of a pintle motor.Here,2D transient simulations of a pintle motor using dynamic meshing are performed.Reynolds-averaged Navier–Stokes equations are solved with the implementation of an RNG k–e turbulence model.In cold flow test,emphasis is placed on the effect of pintle structure,and in hot flow test,emphasis is placed on the effect of propellant pressure exponent.Validation is performed first by comparing the present results with available cold-test experimental data.This shows that the transient simulations can provide good predictions for pintle motors with a relative error of less than 2%in terms of the chamber pressure.It can be found that,when the gas supply system is different,the working principles and conditions of pintle motors are different.The feedback process in propellant combustion has a significant impact on its operation and the effect on the pintle motor performance of different pintle structures is achieved by different variations in the equivalent throat area.Finally,the pressure exponent is an important parameter in hot flow test and changes of thrust in hot flow test are not monotonic,because changes in the flow field and motor performance are asynchronous.

IR radiation characteristics of rocket exhaust plumes under varying motor operating conditions

The infrared（IR） irradiance signature from rocket motor exhaust plumes is closely related to motor type,propellant composition,burn time,rocket geometry,chamber parameters and flight conditions.In this paper,an infrared signature analysis tool（IRSAT） was developed to understand the spectral characteristics of exhaust plumes in detail.Through a finite volume technique,flow field properties were obtained through the solution of axisymmetric Navier-Stokes equations with the Reynolds-averaged approach.A refined 13-species,30-reaction chemistry scheme was used for combustion effects and a k-e-Rtturbulence model for entrainment effects.Using flowfield properties as input data,the spectrum was integrated with a line of sight（LOS） method based on a single line group（SLG） model with Curtis-Godson approximation.The model correctly predicted spectral distribution in the wavelengths of 1.50–5.50 lm and had good agreement for its location with imaging spectrometer data.The IRSAT was then applied to discuss the effects of three operating conditions on IR signatures：（a） afterburning;（b） chamber pressure from ignition to cutoff;and（c） minor changes in the ratio of hydroxyl-terminated polybutadiene（HTPB） binder to ammonium perchlorate（AP） oxidizer in propellant.Results show that afterburning effects can increase the size and shape of radiance images with enhancement of radiation intensity up to 40%.Also,the total IR irradiance in different bands can be characterized by a non-dimensional chamber pressure trace in which the maximum discrepancy is less than 13% during ignition and engine cutoff.An increase of chamber pressure can lead to more distinct diamonds,whose distance intervals are extended,and the position of the first diamond moving backwards.In addition,an increase in HTPB/AP causes a significant jump in spectral intensity.The incremental rates of radiance intensity integrated in each band are linear with the increase of HTPB,and the growth rates of radiance intensities in some bands reach up to 50% as HTPB weight increases by 3%.

Regenerative Cooling for Liquid Rocket Engines

Heat transfer in the thrust chamber is of great importance in the design of liquid propellant rocketengines. Regenerative cooling is an advanced method which can ensure not only the proper runningbut also higher performance of a rocket engine. The theoretical model is complicated, it relates to fluiddynamics, heat transfer, combustion, etc... In this papers a regenerative cooling model is presented.Effects such as radiation, heat transfer to environment, variable thermal properties and coking areincluded in the model. This model can be applied to all kinds of liquid propellant rocket engines aswell as similar constructions. The modularized computer code is completed in the work.

Chebyshev super spectral viscosity method for water hammer analysis

In this paper,a new fast and efficient algorithm,Chebyshev super spectral viscosity(SSV)method,is introduced to solve the water hammer equations.Compared with standard spectral method,the method's advantage essentially consists in adding a super spectral viscosity to the equations for the high wave numbers of the numerical solution.It can stabilize the numerical oscillation(Gibbs phenomenon)and improve the computational efficiency while discontinuities appear in the solution.Results obtained from the Chebyshev super spectral viscosity method exhibit greater consistency with conventional water hammer calculations.It shows that this new numerical method offers an altemative way to investigate the behavior of the water hammer in propellant pipelines.

Numerical investigation of low cycle fatigue life for channel wall nozzles

The thermal-structural response and low cycle fatigue life of a three-dimensional(3D)channel wall nozzle with regenerative cooling were numerically investigated by coupling the finite volume fluid-thermal method,nonlinear finite element thermal-structural analysis and local strain methods.The nozzle had a high area ratio(nozzle exit area divided by throat area)under cyclic working loads.Parametric studies were carried out to evaluate the effects of channel structural parameters such as channel width,channel height,liner thickness and rib width.Results showed that the integrated effects of three-dimensional channel structure and load distribution caused serious strain,which mainly occurred at the intersectant regions of liner wall on the gas side and the symmetric planes of channel and rib.The cooling effect and channel structural strength were significantly improved as the channel width and height decreased,leading to substantial extension of the nozzle service life.On the other hand,the successive decrease in liner thickness and rib width apparently increased the strain amplitude and residual strain of channel wall nozzle during cyclic work,significantly shortening the service life.The present work is of value for design of the channel wall nozzle to prolong its cyclic service life.