Solid Propellant Rocket Engine Research Articles

Design of 2-DOF Thrust vector control system for Rockets and Missiles

This paper outlines the development and implementation of a Thrust Vector Control (TVC) system tailored specifically for solid propellant rocket engines, aimed at enhancing the portability of hybrid and liquid propellant rocket systems. The proposed system ensures trajectory stability and rapid response to flight emergencies by effectively addressing external perturbations like wind. Key components are Gyroscopic (GYRO) sensors, good micro-controller, and Servo motors, collectively referred to as Thrust vectoring components, which dynamically adjust thrust direction opposite to the rocket's trajectory to maintain stability. Drawing from an extensive literature review analyzing existing TVC system designs, with a focus on thrust characteristics and engine combustion timings, the design integrates considerations of geometric properties, kinematics, forces, energy requirements, safety, cost, and control methods, aligning with both mechanical and avionic design principles. The anticipated outcome is an advancement in rocket propulsion technology, characterized by improved angular speed control, trajectory linearity, and rapid response capabilities.

Моделювання інжекції газу в надзвукову частину сопла при газодинамічному регу-люванні вектора тяги

Solid-propellant rocket engines are simple in design, highly reliable, and able to store the propellant for a long time without its degradation. Their main feature is that the propellant is a mixture of a solid fuel and a solid oxidizer, thus ensuring a uniform combustion and a stable discharge of the combustion products. However, the combustion rate cannot be controlled, and the combustion cannot be stopped or restarted. This calls for efficient methods of thrust vector control. Gas-dynamic methods, such as a gas injection into the supersonic nozzle area, offer a required flight path control without complex high-power mechanical systems. The importance of this study lies in improving the accuracy and efficiency of rocket flight control, which is critical for today’s space and defense tasks. The numerical simulation of gas-dynamic control systems, in particular by an asymmetric gas injection, allows one to obtain detailed data on the flow behavior and optimize the design and operating conditions of the system. This study is concerned with a full-scale solid-propellant rocket engine with a gas-dynamic thrust vector control system based on the use of asymmetric forces that occur on the nozzle wall when the supersonic flow interacts with the injected transverse jets. To simulate the process in the Ansys Fluent software package, a geometric model of a nozzle with an asymmetric injection of the chamber gas into the supersonic area was developed. The injection flow rate was controlled by moving the valve flap. The simulation was carried out taking into account the temperature dependence of the main thermophysical gas parameters with consideration for dissociation processes by way of data approximation. The approximation was performed using piecewise polynomial functions. Nozzle flow patterns were obtained. The calculated results were compared with experimental test data and shown to be in satisfactory agreement with the lateral force measured during the fire bench tests of the prototype. From a practical point of view, the results obtained may be directly used to improve existing thrust vector control systems and develop new ones. This will improve rocket navigation accuracy, flight stability, and maneuverability, which is critical for complex space and defense tasks.

Effects of Thermal Damage on Impact Response Characteristics of High-Energy Propellants.

Thermal damage due to microstructure changes will occur in propellants under thermal stimulation. It can significantly affect the sensitization, combustion, and other properties of the propellant, which, in turn, affects the impact safety of the solid propellant rocket engine. A new component which uniformly heats the sample was designed to conduct the Lagrange test and EFP impact test at different temperatures. The thermal decomposition and damage characteristics of the propellant during the heating process were quantitatively analyzed. Additionally, the effects of ambient temperature on impact initiation and detonation growth of the high-energy propellant were elucidated at a mesoscopic level. The results showed that the porosity of the specimen increased by 0.89% under the thermomechanical mechanism, which was mainly characterized by interfacial de-bonding between the AP and the binder. The increase in thermal damage changed the hot spot reaction rate and significantly affected the growth process of propellant impact initiation. A method was proposed to systematically calibrate the reaction rate model for the propellant at different temperatures. The theoretical model parameters of the high-energy propellant at two typical temperatures were calibrated in this way. The critical shell thicknesses computed using LS-DYNA, which, for 20 and 70 °C, were obtained as 15 and 20 mm, respectively.

ANALYSIS OF FOREIGN EXPERIENCE OF DUAL USE SOLID ROCKET MOTORS

In countries, such as the United States, China, India, France, Israel and many others, solid-propellant rocket engine technologies are used not only in military missiles, but also in launch vehicles of the rocket and space industry. Such an approach of dual application of technologies makes it possible to introduce and develop new innovative solutions to increase the degree of unification of related technologies. One of the most important factors influencing the efficiency of using solid propellant rocket engines in space programs is the cost of solid rocket propellants and propulsion system components. For this reason, a high degree of unification of solid fuel products for various purposes is aimed at ensuring a high utilization of the industry’s production capacities, which can significantly reduce prices for materials, fuel and the design of propulsion systems as a whole. The analysis confirmed that the creation of launch vehicles based on rocket engines of solid fuel for military missiles naturally leads to a reduction in the time and costs for the development of such products, a reduction in technical risks and an increase in the reliability of the technologies used.

Parametrical analysis of the strength of the nozzle of a solid fuel rocketer

The paper presents an approach to solving the problem of designing a solid propellant rocket engine nozzle using a design feature in the form of a carbon fiber insert. The design task is to select the optimal parameters of the plate thickness (plate shape) that provide the necessary strength and stability of the structure with a minimum weight. During the design process, a parametric analysis of a carbon fiber insert in the solid propellant rocket nozzle was carried out. By varying the thickness of the plate, an optimal design scheme is found that meets the given safety and stability factors. Parametric analysis of an insert plate made of CM includes modeling of the main weight and strength parameters: determination of the stress-strain state of the structure, values of natural frequencies, determination of the buckling margin, determination of the mass of the solid propellant rocket motor nozzle. Analysis of the bearing capacity of the solid propellant rocket motor nozzle with an insert plate made of CM was carried out using the finite element method using the SolidWorks Simulation software package. During the parametric analysis, two variants of the solid propellant rocket engine nozzle with and without an insert plate were considered. According to the results of the parametric analysis of the solid propellant rocket nozzle, its geometric dimensions were determined and the mass of the structure was minimized.

Reliability Analysis of a Three-Engine Simultaneous Pouring Control System Based on Bayesian Networks Combined with FMEA and FTA

Pouring is an important process in the production of solid propellant rocket engines, and usually, the cost of a solid propellant rocket engine is extremely high. Therefore, pouring production with high reliability is very important. The pouring of three engines of solid propellant rocket engines simultaneously can greatly improve its production efficiency. However, it makes the system more complex and redundant. For a multi-state system, it is difficult to make an accurate evaluation of system reliability. Aiming at the redundancy of multiple engines and acousto-optic combined control in the three-engine simultaneously slurry pouring alarm control system with dissimilar redundant alarm units, a reliability analysis method is proposed based on the combination of Failure Mode Effect Analysis (FMEA) and Fault Tree Analysis (FTA). The control system is divided into several redundant states according to the alarm function, and then the Bayesian Networks method is used for reliability evaluation and calculation. Finally, the reliabilities of systems with dissimilar redundancy degrees are obtained. The tangible results of this research work are as follows: (1) The research results obtained by applying the FMEA method laid a foundation for the establishment of a fault trees model for analyzing the reliability of the control system using the FTA method, in addition, which can guide the maintenance and fault identification of the control system during engineering application. (2) The calculated value of the reliability of the control system is 0.999989, and the mean time between failures is MTBF is 5 × 104 by using the fault tree analysis method, which proves that the designed three-engine simultaneous pouring system is very reliable. (3) Based on the calculation and comparison of the Bayesian Networks of redundant three-engine pouring control systems, the circuit diagram of the improved control system is identified.

Вплив шорсткості поверхні камери енергетичної установки на низькочас-тотні автоколивання холодного робочого газу

Dynamic processes in the combustion chamber have a significant effect on the characteristics of the working processes of solid-propellant rocket engines (LPREs). Pressure jumps and a sharp increase in the local temperature of the combustion products in non-stationary engine operation modes can lead to overrating values of operating parameters and a failure of the LPRE combustion chamber structure. The dynamic processes in the LPRE combustion chamber develop in a complex interconnection of a large number of physical and chemical processes that occur in the gas-dynamic part of the working space of the engine chamber and often lead to self-oscillating modes of engine operation. This is evidenced by numerous data on LPRE fire tests. This paper presents the results of a numerical study of the effect of the LPRE chamber inner surface roughness on LPRE operating parameter low-frequency self-oscillations. The study was made using up-to-date computer simulation means and analysis. Low-frequency (up to 1,000 Hz) oscillations in an LPRE combustion chamber were studied for a power plant test chamber in cold operation with the use of two different approaches to numerical modeling of the dynamics of in-chamber processes: the development and study of a 3D model of the dynamic system of combustion chamber structure – combustion products using the finite element method and the development and study of an axisymmetric 2D model of engine chamber gas flow using the finite volume method. The study revealed a self-oscillatory flow regime caused by combustion product vorticity and acoustic feedback due to vortices colliding with the chamber components or the LPRE nozzle. It was shown that accounting for the wall roughness increased gas vorticity in the gas–solid dynamic interaction zone and the chamber gas oscillation amplitude (on the average, by a factor of 2.5 at a maximum wall roughness height of 56 ?m). The calculated gas flow pattern in the vorticity zones of the chamber and the low-frequency gas pressure oscillation parameters are in qualitative agreement with the experimental ones.

Performance Evaluation of Hybrid Propellant Rocket Engine

In this article, the potential of the hybrid propellant rocket engine technique is discussed. it is a great dream to flight engineers to design and fabricate a rocket engine with more reliability and nil rate of combustion instability probably for long range applications. Solid propellant rocket engine and liquid propellant rocket engine has a considerable rate of advantages but combustion instability is a great threaten in both the engines. Hybrid propellant rocket engine has a very less rate of combustion instabilities when compared to both solid and liquid propellant engines. The ultimate aim of every engine design is to prevent occurrence of this combustion instability and maintain reliable operation during entire span of the combustion process. This article is mainly focused on occurrence of various type of combustion instabilities in hybrid propellant rocket engine and its preventing methods. The performance of the hybrid propellant rocket engine was analysed under various aspects and its unique characteristics were presented in detail.

Algorithm for simulation of vibration effects during track tests of aircraft and missile technology

Track high-speed ground tests of products of aviation and rocket technology make it possible to experimentally confirm the aerodynamic and strength characteristics of new aircraft or their components. Experimental installation Rocket rail track 3500, located on the territory of the Federal State Enterprise Scientific Test Range of Aviation Systems named after L. K. Safronov (FKP GkNIPAS named after L.K. Safronov), is a complex hydro-gas-dynamic engineering structure. The main part of this installation is a rail track placed on a special concrete base, which is based on piles to eliminate the influence of viscoelastic ground vibrations. Between the rails, taking into account the profile of the track, a hydrodynamic multi-level tray is made, designed for hydraulic braking of the stored track sled. The movable rocket track sled rests on slip shoes covering the top rail head. On the track sleigh, solid propellant rocket engines and the object of study are placed. The article describes the approach to the calculation of determining the dynamic loads acting on the elastic structure of the track sled with the test object. The design is represented by a schematic spatial model of elastic beams in the form of rods, plates, pipes with equivalent mass and stiffness, interconnected by elastic links. A model of non-stationary aerodynamic forces is formulated taking into account viscosity for numerical calculations of the flow around the sled structure. Programs have been developed and modeling of the aerodynamic interaction in the case of a supersonic air flow around the structure of a rocket sled has been carried out. The results of vibrational accelerations of elements, components of the sled and the test object, depending on the speed of movement, were obtained by calculation. The forms and frequencies of natural free vibrations of the sled structure are determined, and the densities of the vibration acceleration spectra are calculated.

Числове моделювання поступальних і обертальних коливань рдтп у стапелі під час вогневих стендових випробувань

This article dwells on results of firing bench testing of the solid-propellant rocket engine (SPRE), fastened to the thrust-measuring assembly stand. It is shown that when engine enters the steady-state mode of operation, plane (forward and rotation) vibrations of the SPRE can take place in the assembly stand due to the sudden pattern of thrust generation and displacement of the center of mass of the vibrating system from the engine axis. These vibrations distort measured values of engine thrust and pattern of its change versus time. The purpose of this work is to simulate the oscillating processes of the engine atop the assembly stand to single out in the distorted values of the measured thrust the components related to the processes in the engine and components, which are introduced into the thrust measurement by the oscillating processes in the system “assembly stand – engine”. Model of vibrating system is suggested, which consists of two rigidly connected bodies, containing elastic links, enabling forward and rotary motion and limited by the rigidity of the links. Mathematical model of the vibrating system is developed. Internal forces and moments acting in oscillatory system are defined. Method of numerical simulation of plane vibrations within the limits of the developed model is suggested. Plane vibrating motion and elastic force curve (curve based on force sensor readings) were simulated in thrust-measuring system for different cases of thrust curve and values of vibrating system parameters. Resonance condition was simulated and mutual influence of elastic parametrical link between forward and rotary vibrations was established. Impact of thrust-measuring system rigidity on peak values of force sensor readings was found out. Elastic force vibrations in thrust-measuring system with vibrating system parameters were simulated including variant of thrust change versus time, implemented during firing bench tests of one of the SPRE. It is shown that registered simulation results recreate thrust measurement results in pattern and values obtained by the force sensor during the firing bench tests, and owing to this, it was concluded that oscillating process parameters, assumed in the model, meet the actual ones. It is concluded that simulation provides objective interpretation of the thrust curve, reliable and comprehensive analysis of engine run during firing bench tests, more detailed and exact design of the assembly stand. Key words: vibrating system, plane vibrations, forward vibrations, rotary vibrations, resonance, thrust measurement.

Парирование нештатных ситуаций при пусках исследовательских баллистических ракет с созданием резервных условий входа объекта испытаний в атмосферу

The article is devoted to solving the problem of determining and implementing backup conditions for the research ballistic missiles entry with the test objects in the inflight emergency situations. A test object is understood as a descent vehicle entering the atmosphere and moving at the speed close to the first cosmic velocity. In emergency situations during launch and on the active trajectory leg, backup entry conditions are making it possible to carry out the guaranteed flight testing, albeit with less efficiency, but with exception of the launch emergency completion and complete failure in fulfilling the flight test tasks. Detailed study of this problem is quite suitable for practical implementation and goes far beyond the scope of this work, since it requires design studies, including aerodynamic, dynamic, thermal and strength calculations, as well as the development of additional onboard and prelaunch algorithms. This paper notes only certain main features in decision making to terminate or continue the flight after the emergency situations for а research ballistic missile with liquid rocket engines and solid propellant rocket engines. Mathematical apparatus of the external ballistic design is used to solve the problem.

МЕТОДИКА ОПРЕДЕЛЕНИЯ ХАРАКТЕРИСТИК ВИБРОПРОЧНОСТИ КОНСТРУКЦИИ ПРИ ВЫСОКОСКОРОСТНЫХ ТРЕКОВЫХ ИСПЫТАНИЯХ АВИАЦИОННОЙ ТЕХНИКИ

Achieving the speed limits by ballistic aircraft is a priority, relevant and new technical challenge for the present. Track tests of such products are one of the final stages, confirming their practical performance and efficiency. Testing of products of aviation and rocket technology at the "Rocket Rail Track 3500" FKP "GkNIPAS" named after L.K. Safronov allow to accurately simulate the real conditions in which they are operated. The experimental setup includes a two-rail track, made on a special base, which excludes the unacceptable deflection of the R65 rail, with its fixing every 0.5 m by a special embedment design. The rail track has an acceleration section with an angle of attack of 2º and a length of 2500 m and a deceleration section. The test object is placed on the missile track carriage in such a way as to exclude the impact on it of shock waves reflected from the elements of the carriage and relief. Solid propellant rocket engines are installed on the carriage, which provide the necessary thrust. The sliding bearings of the rocket carriage cover the head of the rails. The thrust of the starting rocket engines provides the necessary acceleration to achieve the maximum values of the required test speed. Track high-speed tests of special equipment objects are accompanied by intense vibration and shock loading of the structure. As the speed of products increases above 600 m/s, as tests have shown, the amplitude of elastic oscillations of the structure can reach the maximum values permissible from the strength conditions. Experimental and theoretical study of vibration and shock effects on the structure of a track carriage with test objects under the conditions of an existing rail track is an urgent task for practice.

THERMAL PROTECTION SOLUTION FOR SOLID FUEL ROCKET ENGINES

This paper presents a new approach to the thermal protection of solid-fuel rocket engines. The method of thermal protection using coatings not only reduces the risk of damage to the engine casing but also increases the engine’s efficiency in working. Parameters such as thrust and the mass per square meter of the engine casing are used to compare the differences between engines with and without thermal protection coating. Numerical calculation applied to the cruise engine of the rocket 9M39. Calculation results show that the thickness of the thermal protective coating and the thickness of the combustion chamber have been optimized regarding rocket weight. The survey results showed that the mass per square meter of the engine casing decreased by 2.5 times, and the engine’s impulse total increased by 1.13% when using the thermal protective coating. These results are used as data to develop methods to improve the performance of solid propellant rocket engines.

Gas-dynamic losses in the flow part of the channel charge of a solid-propellant rocket engine

Modern methods of computational fluid dynamics have been used to study flows with mass supply in axisymmetric channels of solid-propellant rocket motor charges. The studies were carried out with the aim of increasing the accuracy of predicting the intraballistic parameters for performing engineering calculations. The analysis of changes in intraballistic characteristics in the flow part of the channel charge in the classical and nozzleless solid propellant rocket motors is carried out for different rates of mass supply from the combustion surface. For an isobaric combustion chamber, a characteristic velocity profile is shown along a tubular charge path and a charge with sudden expansion. It is shown that for a steady cosine profile of axial velocity after sudden expansion of the channel, a length of more than three gauges is required. For a high-speed combustion chamber, a comparison of the axial velocity profile is carried out depending on the flow velocity under different conditions of mass supply. The tendency of the influence of flow velocity on the character of the profile is noted. It is shown that at a Mach number over 0.5, an increase in the mass flow rate from the combustion surface provides a less filled velocity profile tending to cosine. The differences between the pressure losses in the flow passage of the channel charge, calculated in the axisymmetric approximation, and the losses determined using gas-dynamic functions, are shown. An increase of the mass flow rate in the charge channel of a nozzleless solid-propellant rocket motor leads to a decrease in pressure losses at Mach number over 0.8.

How Can Ablative Cooling in A Solid Propellant Rocket Engine Be Accurately Modelled in Order to Optimise Performance?

Abstract: This essay investigates the use of ablative cooling in solid propellant rocket engines. It begins by exploring the mechanisms by which ablation occurs. It then demonstrates how heat transfer to an ablator can be modelled, and how this can be used to find ablator recession rate and hence the necessary ablator thickness for a rocket engine. It does so by considering simplified mathematical models. These are then compared to the more complex models that have recently been developed. The different variables involved and how they might be used or calculated are discussed. The next section of the essay ties this theoretical knowledge and modelling into practical engineering use by considering the impact ablation has on performance. MATLAB is used to demonstrate how an expanding throat diameter of the nozzle can decrease thrust and specific impulse, and that this can greatly decrease payload capacities. Other variables involved in creating a thrust profile for a solid propellant rocket engine are considered. Finally, the essay will look at how to choose an optimal ablator. It goes through the universally desirable characteristics and uses the Space Shuttle SRMs as an example of disadvantages that may not initially be considered when selecting an ablator

АНАЛИЗ СОВРЕМЕННОГО СОСТОЯНИЯ И ПЕРСПЕКТИВЫ РАЗВИТИЯ РАКЕТНЫХ ДВИГАТЕЛЕЙ ДЛЯ ИССЛЕДОВАНИЯ ДАЛЬНЕГО КОСМОСА

The article analyzes the current state and development prospects of rocket engines for space exploration. Currently, for flights to other planets, not to mention the stars, the use of liquid-propellant and solid-propellant rocket engines is becoming increasingly unprofitable, although many rocket engines have been developed. Thus, to reach manned planes even the nearest planets, it is necessary to develop rocket launchers on engines operating on principles different from chemical propulsion systems. The most promising in this regard are electric, laser and nuclear rocket engines, as well as hybrid rocket engines such as solid-state nuclear rocket engine (NRE) and electric jet engine (ERE) or gas-phase NRE and ERE.

The Effect of Flight Overload on the Intra-Chamber Process in a Solid-Propellant Rocket Engine

The Effect of Flight Overload on the Intra-Chamber Process in a Solid-Propellant Rocket Engine

ДИНАМИЧЕСКОЕ ПОВЕДЕНИЕ РАКЕТНОГО ДВИГАТЕЛЯ ТВЕРДОГО ТОПЛИВА В ПРОЦЕССЕ ЭКСПУАТАЦИИ (ОБЗОР)

One of the problems that has been solved for more than 80 years is the dynamic behavior both during the development of a rocket engine and during operation, these are oscillations that occur in the combustion chamber. Gas vibrations and mechanical vibrations of the rocket engine elements can cause vibration loads, which, under certain conditions, leads to resonance phenome-na. This can cause engine failures. An analysis of the behavior of a rocket engine as a dynamic system with an assessment of frequency interactions over the entire time of its operation has not been completely resolved today. Various options for studying the dynamic behavior of a rocket engine, algorithms for determining natural oscillation frequencies are considered. The analysis of existing approaches for solving the problem of determining the dynamic behavior of the rocket engine was carried out. In various works, the mechanical vibrations of the engine housing or its elements, such as a sliding nozzle, are calculated using various methods. In a number of works, the structure is considered as a model of discrete masses, where the elements are connected through the stiffness and viscosity coefficients. In other cases, fluctuations of the gas flow during combustion in the combustion chamber are considered, methods of numerical simulation of the process are developed that take into account the features of vor-tex formation and instability of the gas flow, as well as dependence on the shape of the charge. However, the joint problem has not been solved; in the presented works, the mutual influence of the vibrations of the rocket engine case with the fuel and gas flow during operation is not considered. To get a complete picture of the dynamic loads experienced in a solid propellant rocket engine, this interaction must be taken into account.

Validation of RANS Turbulence Models for the Conjugate Heat Exchange Problem

This paper addresses problems of mathematical modeling of heat exchange processes in the pre-nozzle volume of a solid propellant rocket engine with a charge with starlike cross-section and a recessed hinged nozzle. Methods of mathematical modeling are used to solve the quasi-stationary spatial conjugate problem of heat exchange. An analysis is made of the influence of RANS turbulence models on the flow structure in the flow channels of the engine and on the computed heat flow distributions over the surface of the recessed nozzle. Methods of mathematical modeling are used to solve the quasi-stationary spatial conjugate problem of heat exchange. Results of validation of RANS turbulence models are presented using well-known experimental data. A comparison of numerical and experimental distributions of the heat-transfer coefficient over the inlet surface of the recessed nozzle for the engine with a cylindrical channel charge is made for a primary choice of turbulence models providing a qualitative agreement between calculated and experimental data. By analyzing the results of numerical modeling of the conjugate problem of heat exchange in the combustion chamber of the solid propellant engine with a starlike channel, it is shown that the SST $k-\omega$ turbulence model provides local heat-transfer coefficient distributions that are particularly close to the experimental data.

Anomalous Features of Dynamics of Intrachamber Processes in a Low-Thrust Solid Propellant Rocket Engine with a Multi-Grain Propellant Charge

Anomalous Features of Dynamics of Intrachamber Processes in a Low-Thrust Solid Propellant Rocket Engine with a Multi-Grain Propellant Charge