Solid-propellant Motors Research Articles

TÍNH TOÁN MÔ PHỎNG DÒNG CHẢY TRONG BUỒNG ĐỐT ĐỘNG CƠ NHIÊN LIỆU RẮN VỚI THUỐC PHÓNG HỖN HỢP HAI LỚP

This article presents a method of calculating the simulation of flow in solid fuel combustion chamber, which applied to the two-layer mixed solid propellant motor, using simulation software Ansys Fluent. The regression of burning surface of the propellant is simulated using dynamic meshing tool combined with user-defined functions (UDF), written in C++ programming language. The author proposes a method to calculate the average fire rate for mixed solid fuel, thereby simplifying the simulation process and still get results corresponding to the experimental results. The calculation thrust-time curve has a profile similar to the experimental results. The maximum thrust achieved during the steady burning phase is 4.98 kN, the correspoding experimental value is 5.0 kN.

Increasing Burning Rate and Motor Thrust by Expandable Graphite Additives

Controlling rocket thrust may be done via propellant burning rate catalysts and enhancers. This paper presents an experimental investigation on increasing the thrust of hybrid and solid motors by adding a small fraction of expandable graphite (EG) within the binder matrix to enhance burning rate. EG is a form of intercalated graphite flakes that upon heating change their appearance to elongated fibers/strings of substantially larger length and volume. The elongated EG strings at the burning surface are hypothesized to conduct heat from the hot surroundings to the bulk, thereby increasing the burning rate. High-speed photography of the surface phenomena of fuel slabs containing EG additive subjected to flame supports the greater effect on burning rate enhancement (up to twofold) for polyester versus hydroxyl-terminated polybutadiene or paraffin wax fuels in hybrid motors. Similar investigation on the burning of ammonium perchlorate–polymer solid propellant strands revealed different surface phenomena and substantial burning rate increase (60% and more) for hydroxyl-terminated polybutadiene versus polyester binder with 5% EG additive. It can be concluded that EG can serve as a novel burning rate and thrust enhancer without deterioration of the mechanical properties of the polymeric fuel/binder for hybrid (including solid fuel ramjet) and solid propellant motors.

TECHNICAL AND ECONOMIC SUBSTANTIATION OF THE AIR LAUNCH AEROSPACE SYSTEM PARAMETERS

The substantiation of the technical and economic parameters of the air launch aerospace system, which consists of a reusable unmanned aerial vehicle and integrated launch vehicle, is given. A combination of a turbofan engine and a ramjet engine is used as the unmanned aerial vehicle propulsion system. The considered vehicle is capable of delivering a payload into low-Earth orbits without the use of a spaceport. We have developed the methodology of the technical and economic substantiation of the parameters of the air launch aerospace system. The results were obtained by searching for the minimum of the objective function, which established the relationship between the technical and economic parameters of the aerospace system. For the objective function solution, the design parameters of the integrated launch vehicle and unmanned aerial vehicle were determined, as well as the limitation of the total acceleration velocity of the aerospace system. The methods used allowed us to determine the velocity constraints provided by the operation of the turbofan engine, ramjet engine, and the three solid propellant motors of integrated launch vehicle stages, as well as maximum dynamic pressure and maximum permissible temperature on the unmanned aerial vehicle surface. We determined the scheme for estimating the cost of launching the spacecraft into Earth orbit by the air launch aerospace system. The result of the substantiation is the determination of the technical and economic parameters of the integrated launch vehicle, unmanned aerial vehicle, and the aerospace system as a whole. The influence of the maximum temperature on the surface of the unmanned aerial vehicle and the specific impulse of the ramjet engine on the parameters of the aerospace system was also evaluated. The substantiation is the first step towards the creation of the Ukrainian aerospace air launch system.

Вплив імпульсних рульових двигунів на динаміку твердопаливної ракети

The object of this study is a rocket with a solid propellant engine launched from a vehicle launcher. The current work studies the oscillations of a rocket airframe in flight under the influence of impulses of the steering motors and the effectiveness of flight control along the trajectory by the steering motors located in the head of the solid rocket. Analytical and numerical methods for determining the frequencies of free vibrations of a rigid body and analytical methods for integrating differential equations of motion are applied. A mathematical model of rocket flexural oscillations has been constructed. Methods for determining the modes and frequencies of natural oscillations have been analyzed. Three lower frequencies and modes of natural oscillations of the rocket were calculated. The discrepancy between the results of the calculation by the computer program and the analytical calculation of the first approximation is 3%. The calculation of oscillations of a rocket with a caliber of 30 mm and a length of 7655 mm showed that torsional and longitudinal oscillations have relatively high frequencies and have little effect on the stability and controllability of the rocket flight. The vibrations of the rocket airframe, which has a gas-dynamic flight control system with the help of rudders and a gas-jet one with the help of steering engines, are studied. Steering motors are small-sized disposable solid propellant motors, which are located in five rows around the perimeter of the rocket airframe at the head. Methods for calculating the forced oscillations of a rocket during the operation of steering engines have been developed. Under the action of the control pulse, transverse oscillations of the rocket airframe y are excited and simultaneously, the rocket rotates as a solid body around the center of mass. Airframe oscillations quickly decay, and rotation changes the pitch or slip angle. To stop the rotation and leave some required angle of inclination or azimuth of the trajectory, a second impulse of force must be given by the steering motor mounted on the opposite side of the hull. The dependence of the change in pitch and slip angles on the duration of time intervals between control pulses was studied. The analysis of the dependence obtained shows that, although the weight of the fuel charge is approximately half the weight of the rocket, the controllability by the steering engines is equally effective at any fuel consumption.

Rapid detection for the surface and hidden crack of solid propellant motor charge using laser arrays scanning thermography (LAsST)

Laser arrays scanning thermography (LAsST) was introduced to detect the surface and hidden cracks of solid propellant motor charge. A series of simulated surface cracks of solid propellant charge was prepared and detected by LAsST. The thermal-wave response signal is reconstructed by truncation pseudo-static matrix reconstruction (TC-PSMR), and then, the thermal image sequence similar to the impulse response is reconstructed. The principle of characteristic extraction algorithms [such as Fast Fourier transform (FFT), principal component analysis (PCA), and partial least squares regression (PLSR)] are introduced and employed to obtain the feature images from the thermal-wave response. The morphological matrix is used to denoise the processed image. The experimental results demonstrated that the defects of the PLSR image and FFT phase image are more obvious, and the image noise after morphological matrix processing is significantly reduced. Meanwhile, the LAsST approach can enable rapid detection for the surface and hidden cracks of solid propellant motor charge.

A comprehensive modeling study of particulate coalescence and breakup mechanisms in solid propellant motor nozzles

Gas-particle flows are modeled to account for the coalescence and breakup of liquid aluminum oxide particles dispersed within the gas phase in solid propellant motors. The one-way coupled population balance equation, which addresses a change in number concentration of particles by the particle to particle and surrounding to particle interactions, is solved using the direct quadrature method of moments (DQMOM). Simulations are performed on the converging-diverging nozzles attached to solid propellant motors and the mass mean diameter at the nozzle exit is compared to previous measurements and a correlation derived from numerous existing measurements data. The results from a series of simulations are presented to assess the effects of chamber pressure, initial particle concentration and size, and nozzle scale on resulting particle size growth.

Impact of lattice inclusion of Cu and Fe ions on thermal decomposition characteristics of ammonium perchlorate

Ammonium perchlorate (AP) is the universal oxidiser in use for all the solid rocket propellant motors used for space exploration due to its high available oxygen content and thermal decomposition without any solid residue. The inclusion of reactive species in AP directly affect the viscoelastic and ballistic properties of the propellant. Variations in lattice configuration of AP change its physical and thermal characteristics dramatically. In the present work AP was doped with Copper perchlorate and Iron perchlorate through co crystallisation. The impact of inclusion of these ionic species in the lattice on the thermal decomposition characteristics of AP was examined. The incorporation affected the physical as well as the ballistic characteristics of the resultant AP. The incorporation of foreign ions into AP crystals significantly changed the crystal morphology. The decomposition temperature decreased vis-a-vis with normal AP. The activation energy remarkably decreased for the doped AP crystals.

Mitigation of irregular burning in a small solid propellant rocket motor

An irregular burning phenomenon arises in some solid propellant motors(SRM), where the time-averaged pressure differs largely from that dedicated by instantaneous area burning. This phenomenon is generally considered the result of the interaction of pressure disturbances within the motor cavity with solid propellant deflagration. Hence, it is a combustion-dependent acoustic phenomenon. Historically, different suppression devices included the classical resonance rods and paddles, and arrays of acoustic resonators and various baffle arrangements were applied. In this paper, irregular burning which arises in a modification of a shoulder-launched rocket is investigated. Several techniques dealing with the irregular burning problem were attempted. In each case, the signal of pressure was analysed in the time domain through time average and AC pressure signal, and in the frequency domain through Fourier transform and power spectral density. The most successful technique was the drilling of holes through the grain, connecting the perforation to the external burning region of the cylinder. Next was the increasing of blocking factor, both methods lead to stabilizing the modified motor.

Experimental investigation of star grains in dual thrust solid propellant motors

One of the great challenges in designing tactical solid missiles is to achieve high acceleration in the boost phase then maintaining constant speed during the sustain phase. This could be achieved by using a dual thrust solid-propellant rocket motor. Many of these tactical motors use a combination of star, tubular or finocyl grains to achieve this profile. The present study uses two tandem star grains with different design parameters and different transition geometry. Previous researches had consistently shown that the main advantage of star grain is the potential higher volumetric loading in addition to high tailorability. The pressure-time curve for the designed grains is calculated using a zero-dimensional internal ballistic module and a small-scale test motor is used to verify the calculated pressure-time curve. Different transition geometries are compared. Tapered transition is shown to give a comparable performance with the sharp transition with the advantage of higher volumetric loading.

Studies of CT Applications to Development of Technologies for Space Systems Using Gaseous Propellant

AVIO is a Company leader in manufacturing of products for space primary propulsion. In particular, AVIO production covers Solid Propellant Motors manufacturing for European launchers ARIANE and VEGA. Also, cryogenic stage made at AVIO are installed on these launchers and then research is also involved in development of systems used for liquid propellant applications. In this research new technologies as Additive Manufacturing are well addressed and in this effort a CT technique used as NDT, but also as metric tools, has been investigated. Results are presented in this paper.

An Experiment on Solid Propellant Combinations; Evaluate the Propellant Properties using Ultrasonic Impulse Measuring System

Impulse force is the key for all high speed aero propulsion system. Solid propellants are uncontrollable once it ignited because it has the ingredients (fuel, oxidizer) for combustion within the chamber in which they are ignited. Since an entire solid propellant motor performance depends on propellant choice, chamber pressure, and nozzle design, the evaluation of propellant properties is highly needed. In the proposed Ultrasonic Impulse measuring system, the solid propellant motor, when fired, produces a thrust force which makes a bending moment in the simply supported beam. A sensitive Ultrasonic sensor will calibrate the bending moment, proportional to the Impulse force. The effective burning time is noted. Thermocouple reading, load cell reading, and the effective burning time value is used for calculating the mass flow rate, effective exhaust velocity, and specific impulse. The different fuel – oxidizer combinations with catalysers, binders are made. The small scaled propellant motor with length 0.12m, Diameter 0.025m, Exit Nozzle Diameter 0.003m, Empty Motor weight 20grams are used. The Recrystallation method is used for propellant preparation. The pyrotechnic ignition is adapted. The Grain test, Rope test are carried out. We have chosen high molecular weight, and low heat of formation value propellants. There are 4 different propellant combinations are tested with the proposed impulse measuring system. The Specific impulse reading of KNO3 + C12H22O11 + Al powder combination has highest value among all the propellants. It has 87seconds of specific impulse. The proposed impulse measuring system gives the reasonable readings with minimum error. We have compared the experimental results with the theoretical results and got the specific impulse of 90 seconds. The heat of formation of potassium nitrate and sucrose combination gives a low heat of formation value. The heat of combustion value and the characteristic velocity of the solid propellant combination (KNO3 + C12H22O11 + Al powder) are decreased.

Experimental Examination of Scalable Under Water Launching Missile System

Experimental Investigations were carried out to study the performance of underwater launching missile system by adapting the electromagnetic launcher concept. There are two energy sources utilized to propel the missile in this proposed research; solid propellant motor and electromagnetic launcher. Electromagnetic launcher propels the missile from underwater to the water surface. Solid propellant motor propels the missile from water surface to the fixed target in the atmosphere. Solid propellants were prepared and the specific impulse of the propellants was also observed by impulse measuring system. 65/35 ratio of fuel-oxidizer (KNO3-Succrose) has the specific impulse of 89.3seconds. Rope test and vertical launch test were conducted for finding out the missile altitude and performance as well. The Electromagnetic launcher used in this research can replace torpedo tubes which are conventionally fitted with compressor supply. Since capacitor banks are being used as power source for the launcher, reusability and simplicity of the launching system are improved remarkably. The electromagnetic launcher generates Lorentz force to accelerate the missile through the water. 220VDC, 5000μFD capacitors connected in parallel produce 363N-m energy. 10grams of missile model made up of aluminium material was tested and found velocity and projectile of the missile.

Numerical and Experimental Investigation of Motor Pressure Effect on Thermochemical Erosion of Graphite Nozzle in Solid Fuel Engines

Using heat shield, especially in throat area has a significant effect on combustion chamber pressure and thermal efficiency of solid fuel engines. Precise prediction of the regression of throat area for different pressures, will lead to optimal design of the motors, specifically for those of long burnout times. In this work, erosion of graphite nozzles employed in solid propellant motors with a specific composite propellant and variable pressures, is investigated. The numerical model utilized includes the Naiver-Stokes equations, chamber gas thermodynamic equations and thermochemical and heat conduction equations for the nozzle surface. In order to validate the numerical results, a cartridge type solid propellant motor with a graphite nozzle is experimentally tested. Using a 3D scanner in the experimental setup, the amount of inner surface regression for variable pressures (60, 90, 120 and 200 bars) is measured. Numerical and experimental results are in a proper conformity with each other. There is a direct relationship between convection heat transfer coefficient and the pressure. The overall erosion is the same for all four engine pressures. The erosion rate increases with increasing pressure. This rate for the fuel is about 0.21 mm/s for every 100 times the pressure up to 300 times. For a pressure higher than 300 times, a significant leakage occurs at the corrosion rate.

Using an Inhibitor to Prevent Plasticizer Migration from Polyurethane Matrix to EPDM Based Substrate

The loss of adhesion between the propellant and insulator is one of the most important problems in solid propellant motors due to migration of plasticizer to interface of propellant and insulator. In this work, the polyurethane (PU) binder containing DOP plasticizer was used as a polymeric matrix and β-cyclodextrin (β-CD) was applied as inhibitor agent to prevent plasticizer migration from the PU matrix into the ethylene propylene diene monomer (EPDM) substrate. To increase the compatibility of β-CD and PU matrix, a derivative of β-CD has been synthesized using toluene diisocyanate (β-CD-TDI). The synthesized derivative was characterized by MALDI-MS and FTIR-ATR analyses. FTIR-ATR results confirmed the formation of bonding between β-CD and the polymeric network while the MALDI-MS results showed that the synthesized derivative contained two β-CD and 7 TDI molecules bonded to β-CD. Investigation of the mechanical properties of PU modified by β-CD-TDI showed a decrease in tensile strength and an increase in elongation at break with increasing β-CD-TDI content. DMTA results showed a decrement in crosslinking density by increasing the β-CD-TDI content. Also, to investigate plasticizer migration, extraction of the DOP plasticizer from samples was performed using dichloromethane solvent and its concentration was measured by gas chromatography. The results of migration evaluation after four months showed that using β-CD as an inhibitor agent in the PU binder could prevent the migration of plasticizer to EPDM substrate.

The Solid–Propellant Motors with Regulated Thrust

The Solid–Propellant Motors with Regulated Thrust

A preliminary multidisciplinary design procedure for tactical missiles

During missile system development, multidisciplinary design procedure is iteratively implemented based on the missile objective and target nature including internal ballistic, warhead function, and airframe configuration. By applying missile preliminary design, a good estimation for different design parameters can be obtained which will be useful through further detail design process. The aim of this paper is to build a preliminary design procedure for an unguided tactical missile that uses single-stage solid propellant motor to deliver a defined payload mass to a desired ground range. Based on data of available similar mature missile systems, two empirical formulas are developed to serve in the initial sizing of the missile with consideration of slenderness ratio, warhead mass, and desired ground range. Two different design concepts are implemented for tubular and star grains with different propellant compositions and chamber filling coefficients while the body-alone airframe configuration is adopted. The results demonstrate the capability of the proposed design procedure in defining the detailed design parameters. The impact of changing the propellant compositions and chamber filling coefficients on the obtained ground range is also explored.

Dynamics of solid propellant motor composite casing under impact pressure

A dynamic behavior of solid propellant motor composite casing under the action of an internal impact pressure is treated. This pressure describes the engine operation. The thin-walled casing consists of the cylindrical shell and two bottoms. These bottoms are truncated hemisphere. The casing is clamped along two edges of the bottoms. A shear, a rotary inertia and stress–strain relations for an orthotropic material are accounted. Semi analytical method is suggested to analyze the structure stress–strain state. The thin-walled casing dynamic is described by large dimension system of the ordinary differential equations.

An experimental and numerical investigation of the effect of geometric parameters on the flexible joint nonlinear behavior for thrust vector control

Solid-propellant motors are used in a variety of space systems. This type of motor requires a guidance system, and a thrust vector control to function properly. One of the most effective methods for thrust vector control can be achieved by utilizing a flexible joint system. The flexible joint is the most widely used device in advanced nozzle system, such as space shuttle boosters, Vega, and Ariane 5. This setup includes multiple layers of elastomers to provide the flexibility, and a few metallic or composite rings as the scaffold of the system. In this study, a nonlinear simulation of the flexible joint is investigated using finite element method. These results are verified by construction and hydrostatic test, and there is a good consistency between predictions and empirical data. To the best of our knowledge, contemporary architecture does not provide the optimum result. Therefore, we propose delicately designed configuration, which can be mentored in a variety of motors, to improve the functionality of the flexible joints. A detailed study on the alteration of geometric parameters has been done to unravel its effect on the behavior of the flexible joint.

Sounding Rockets: analysis, simulation and optimization of a solid propellant motor using Hopsan

Abstract The aim of the present paper is to expose the performance simulation of a small sounding rocket that operates with a solid propellant motor. In order to contribute in the understanding of solid propellant rocket engines, this work analyses the basic functioning of this type of motor and simulates its performance in the experimental rocketry field. In a solid propellant rocket, the combustion is clearly determined by the grain shape, which establishes the burning surface as a function of time. It affects directly the chamber pressure and, along with burn rate, it determines thrust. So, as well as propellant characteristics, burning surface is an essential parameter and the engineer can modify it in order to obtain the desired performance. The simplicity of the solid motor and the large number of possible grain burning shapes make solid propellants an interesting option for a rocket engine design. Therefore, the focus of this study is to build a complete model of solid combustion, with a specific burning shape, that simulates the rocket performance. The motor dimensions have been based on an Aerotech engine used in experimental rocketry. It has served as a model for the first calculations and its real performance has proved to be a good comparison tool for the results obtained. The simulation has been carried out in Hopsan, a multi-domain software developed at the Linkoping University. Once the simulation is completely built, an optimization of the engine has been done in order to improve the performance and to increase the altitude reached by the rocket. Finally, the results obtained with the optimization of the system are compared with different case dimensions. Hence, the comparison gives an idea of how this parameter affects rocket performance.

Energy and Propulsion Optimization of Solid-Propellant Grain of a Hybrid Power Device

A method of distribution of an additional solid-phase component (oxidizer) providing uniformity of grain burning for the purpose of evaluation and optimization of energy and propulsion parameters of hybrid solid-propellant motor is proposed in the paper.