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.

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.

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.

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%.