Abstract: Rotating Detonation Combustors are studied with increasing interest worldwide, and are seen as the most promising of the available pressure gain (stagnation) combustion concepts. The dynamics and s...

Abstract: The specific impulse (Isp) is an important performance parameter that describes energy efficiency of propellant combustion and is intimately related to the rocket engine thrust. In this study, it was possible by using only two variables, i.e., the heat of reaction (Q) and the number of moles of gaseous reaction products per gram of propellant (Ng) calculated according to [H2O-CO2] arbitrary decomposition assumption and constants derived from the ISPBKW code to predict the specific impulse of more than 165 compositions belonging to virtually all classes of propellants such as monopropellants, single-base, double-base, triple-base, and cast modified double-base (CMDB) propellants, pseudo-propellants, composite propellants, liquid mono- and bipropellants, and finally hybrid propellants. Further analysis reveals that for C-H-N-O containing propellants, the specific impulse values estimated using the new method should not deviate more than 5% from the output of the ISPBKW thermochemical code.

Abstract: A 500 N model engine filled with LO2/GCH4 was designed and manufactured. A series of ignition attempts were performed in it by both head spark plug and body spark plug. Results show that the engine can be ignited but the combustion cannot be sustained when head spark plug applied as the plug tip was set in the gaseous low-velocity zone with thin spray. This is mainly because flame from this zone cannot supply enough ignition energy for the whole chamber. However, reliable ignition and stable combustion can be achieved by body spark plug. As the O/F ratio increases from 2.61 to 3.49, chamber pressure increases from 0.474 to 0.925 MPa and combustion efficiency increases from 57.8% to 95.1%. This is determined by the injector configuration, which cannot produce the sufficiently breakup of the liquid oxygen on the low flow rate case.

Abstract: Hybrid rocket propulsion offers inherent safety during handling and launch operations at low cost. This makes it not only attractive for space tourism applications like SpaceShipTwo but also for sounding rockets and for educational activities with students. This has been successfully demonstrated by the HyEnD student project from the University of Stuttgart: On November 8th, 2016 at 10:30 a.m. the hybrid sounding rocket HEROS 3 was launched from the Esrange Space Center to an apogee altitude of 32,300m (106,000 ft). This set a new altitude record for European student and amateur rocketry. Furthermore, this is a world altitude record for hybrid rockets built by students. It was successfully recovered with the drogue and main parachute being released. The rocket was powered by a 10 kN design thrust hybrid rocket engine with a paraffin-based fuel and Nitrous Oxide (N2O) as the oxidizer. The combustion efficiency was verified to be above 97% in ground tests. Liquid burn time in the flight was for 15 s with an additional combustion of gaseous N2O in the blow-down mode of about 10 s. The rocket was 7.5m long with an empty mass of about 75 kg. The rocket structure was made completely from light-weight carbon fibre and glass fibre reinforced plastic.

Abstract: Present paper focuses on the comprehensive survey of Rotating Detonation Engine (RDE) and their research from the basic to the advanced level. In this paper, an abridged archival background of Pulse/Rotating Detonation Engine (PDE/RDE) is briefed. This is followed by a short description of a Continuous Spin Detonation (CSD) and a few essential facts from the prior publications. Furthermore, a summarization of the Continuous Detonation Wave Rocket Engine (CDWRE) concepts is examined. At long last, a detailed numerical investigation and experiment work of RDE is also presented.

Abstract: A dynamic model of a liquid-propellant rocket engine has been developed in this paper, with the future purpose of meeting the more demanding requirements of rocket-engines control forced by the new reusability scenarios. This transientrepresentative modelling approach has been carried out in two phases: initially the purely thermodynamic modelling and subsequently its adaptation to the control framework. The former was tackled by building first a representative simulator of the well-known Vulcain engine. The differential equations considered come from mass, momentum and energy conservation equations and from turbo-pump dynamics. In general, macroscopic behavior at component and system levels is considered. Once this simulator started to provide satisfactory results, it was translated into a state-space model for control by symbolically joining all components’ equations, which led to a set of differential equations capturing system’s global behaviour. Its states consist in mass flows, pressures, temperatures and shaft rotational speeds. The available actuators are five continuously-controllable valves, one binary igniter and one binary starter. That combination of continuous and discrete features forces the definition of a hybrid system in the control sense. The analysis of its dynamic characteristics points to a good controllability of thrust via the gas-generator injection valves, and of mixture ratio via the turbines’ flow-distribution valve.

Abstract: A multi-fidelity framework is established and demonstrated for prediction of combustion instabilities in rocket engines. The major idea is to adapt appropriate fidelity modeling approaches for different components in a rocket engine to ensure accurate and efficient predictions. Specifically, the proposed framework integrates projection-based Reduced Order Models (ROMs) that are developed using bases generated on truncated domain simulations. The ROM training is performed on truncated domains, and thus does not require full order model solutions on the full rocket geometry, thus demonstrating the potential to greatly reduce training cost. Geometry-specific training is replaced by the response generated by perturbing the characteristics at the boundary of the truncated domain. This training method is shown to enhance predictive capabilities and robustness of the resulting ROMs, including at conditions outside the training range. Numerical tests are conducted on a quasi-1D model of a single-element rocket combustor and the present framework is compared to traditional ROM development approaches.

Abstract: This paper considers various methods of intensification of the heat exchange in the regenerative cooling system of a liquid-propellant rocket engine. Experimental data on hydraulic resistance and heat transfer coefficients in different cooling tracts are presented. A natural hydraulic experiment has been performed on the chamber of the liquid-propellant rocket engine, which confirmed the calculations. The efficiency of porous tracts with interchannel transpiration of the coolant and the efficiency of tracts with coplanar channels have been shown.

Abstract: This paper describes one possible approach to a numerical framework for a reduced order tool that aims to simulate combustion instabilities in liquid rocket engines. The numerical framework relies ...

Abstract: This work studies the performance and dry mass of the under development LOX/Ethanol L75 liquid rocket engine. To this end, an object-oriented program written in C++ was developed. The program is intended to be versatile and easily extensible in order to analyze different configurations of liquid rocket engines. The UML (Unified Modeling Language) tool is used to model the architecture of the codes. UML diagrams help to visualize the code structure and the communication between objects, enabling a high degree of abstraction. The cryogenics Vulcain and HM7B engines power cycles along with the staged-combustion SSME engine perform the verification of the codes. Finally, the influence of changes in design parameters on the performance and dry mass of the L75 rocket engine is analyzed.

Abstract: The invention discloses a controllable thrust multi-pulse solid rocket engine. The controllable thrust multi-pulse solid rocket engine comprises a combustion chamber, a control system and a power source. The combustion chamber is composed of an electric control solid propellant, electrodes, a combustion chamber shell, a sensor, a spraying pipe and the like. The electric control solid propellant isan insensitiveness propellant with the electric control effect, combustion can be achieved through electrification, and extinguishment can be achieved through outage. The electric control solid propellant or the electrodes are coupled to a transmission device, and the relative position between the electric control solid propellant and the electrodes can be changed. When all the electrodes make contact with the electric control solid propellant, the electric control solid propellant is powered on and combusted, and when one or more electrodes do not make contact with the electric control solidpropellant, combustion is stopped, and extinguishment is achieved. According to the controllable thrust multi-pulse solid rocket engine, an ignition device of the rocket engine is omitted, the combustion speed of the electric control solid propellant and the mass of the electric control solid propellant consumed for combustion can be accurately controlled, safety is high, thrust of the engine canbe controlled in real time, and repeated starting is conducted; and the controllable thrust multi-pulse solid rocket engine has wide application prospects in the multi-pulse rocket engine field needing real-time thrust control.

Abstract: A novel internal ballistics technique for determining the time-resolved spatially averaged regression rate of the solid-fuel grain in a hybrid rocket engine is presented. This technique makes use o...

Abstract: We present the results of experiments performed with a rotating detonation engine using continuous detonation in an annular combustor to create thrust. Detonation waves propagate in a supersonic and very small region, allowing shortening of the combustor. The combustor of RDE causes high-pressure loss when the propellant is injected, and cooling is necessary due to high heat flux. However, the combustion efficiency of detonation combustion in an annular combustor is the most important, but have not been fully elucidated. In addition, the influence of the injector shape and direct cooling of a rotating detonation combustor require clarification. This paper reports the measurement results of combustor stagnation pressure and thrust, the influence of injector shape on c* efficiency, and the estimate of heat flux. The c* efficiency was 88–100% when we used the convergent or convergent-divergent nozzle and the equivalence ratio was less than 1.0. The shape of the injector influenced wave propagation mode, but the mode did not change the c* efficiency. We estimated time-spatial average heat flux from the terminal temperature, and the heat flux was 8.1 ± 1.8 MW/m2 in no water injection condition. The rocket RDE sled test was successfully performed. The total mass of the rocket RDE system was 58.3 kg, total time averaged thrust was 201 N, the time averaged mass flow rate was 143 g/s, and the specific impulse was 144 s.

Abstract: It is widely recognised that there is a global need for cheaper launchers for small satellites, and there is considerable global effort led primarily by private enterprise to develop these. One way to achieve these lower costs is to use cheaper materials, usually by sacrificing some of the performance benefit. A dual birectional double vortex bipropellant engine has been designed that confines combustion to the core of the rocket engine, thereby maintaining much cooler wall temperatures, enabling the employment of cheaper materials with lower melting temperatures. This paper reports on the design and successful test firing of such an engine with a nominal thrust of 20N, the first time that such an engine has been developed in an educational setting.

Abstract: A numerical analysis is conducted to understand the flowfield inside a launch pad affected by the layouts of clustered rocket engines for the first stage. The Japanese H3 launch vehicle with typica...

Abstract: The invention relates to the field of rocket-engine propellant delivery systems, in particular to a gas-oxygen-kerosene rocket-engine pressurizing delivery system and a small-rocket-engine propellantsupplying system. The pressurizing delivery system comprises an oxygen supplying system, a fuel oil supplying system and a blowing-off system; the oxygen supplying system comprises an oxygen storage tank valve assembly, a first stop valve, a first one-way valve and a second one-way valve; the oxygen storage tank valve assembly, the first stop valve and the first one-way valve are sequentially communicated through a pipe, the second one-way valve is communicated with the blowing-off system, and an outlet of the first one-way valve and an outlet of the second one-way valve are connected with a thrust chamber through the same pipe; the fuel oil supplying system is connected with the blowing-off system to enter the thrust chamber through the pipe. The gas-oxygen-kerosene rocket-engine pressurizing delivery system in the technical scheme is compared with a liquid-oxygen-kerosene rocket propellant delivery system, complex design such as pressurization-system gas-liquid phase transition, delivery-system low-temperature two-phase flow and engine pre-cooling are not required to be considered, and the complex degree and the quality of the system structure are also greatly reduced.

Abstract: A solid rocket or a solid-fuel rocket is a rocket engine that uses solid propellants (fuel/oxidizer). The term "solid fuel" in this context is actually erroneous because solid propellant must contain both a fuel and an oxidizer to support combustion. All rockets used some form of solid or powdered propellant up until the 20th century, when liquid rockets and hybrid rockets offered more efficient and controllable alternatives. In this we consider a composite material Carbon/Epoxy and Kevlar/epoxy for a pole relevant in rocket parts. Its length as 10m, interior breadth as 1m, inside weight as 1000psi or 6.894Mpa, engine packaging is outlined and subjected to basic and modular examination. Fundamentally engine packaging is planned in ANSYS apparatus by considering outline contemplations, assist the model is subjected to stack imperatives and weight parameters as in the constant flight and reproduction examination will be completed and perceptions are made to know the basic functions, loads, stress, strain and proficiency of rocket engine packaging. For the Design and Analysis, Simulation Software ANSYS R 18.0 be utilized and results will be summoned. Layer as stress – strain Distribution is acquired, Static Structural Analysis, Modal Analysis will be done and comes about might be closed.

Abstract: Disclosed is an electrically controlled solid propellant rocket engine adjustable in thrust and capable of being started several times. The engine comprises a combustion chamber, a spring, a movable insulating plate, an electrically controlled solid propellant grain column, a coplanar electrode, a heat insulation layer, a spraying pipe and a wiring column; a grain containing structure of the electrically controlled solid propellant grain column is tubular, and the grain column is pushed by utilizing the elastic force of the spring; the coplanar electrode is in contact with the burning end surface of the electrically controlled solid propellant grain column; the coplanar electrode comprises two identical porous electrodes, the two porous electrodes are connected to the two ends of a power supply respectively, certain voltage is loaded between the two porous electrodes, and the engine can achieve irrigation; the power supply is switched off, and the engine is powered off automatically; the engine can be started several times continuously. By changing the voltage value between the two porous electrodes, the thrust of the engine can be actively controlled in real time. The engine can replace a liquid attitude-orbit control and terminal correction engine and is applied to power devices or auxiliary power devices with engines adjustable in thrust and capable of being started severaltimes for meeting the requirements of aircrafts.

Abstract: Элементы конструкции ракетного двигателя твердого топлива характеризуются сложностью изготовления и последующей сборки. К возможным технологиям, обеспечивающим снижение временных и ресурсных затрат на производство деталей этого двигателя, относится технология лазерного спекания порошков металлополимерных композиций (порошковое 3D-прототипирование), отличительной особенностью которой является послойное создание детали в процессе спекания порошка. Однако внедрение технологии порошкового прототипирования при изготовлении высокоответственных деталей сопряжено с рядом проблем, в том числе с некоторым снижением механических характеристик получаемого материала. Предложена методика расчета теплопрочностных характеристик конструктивных элементов ракетного двигателя твердого топлива с использованием алгоритмов численного моделирования. Проведена оценка изменения коэффициента запаса прочности элемента, выполненного с помощью технологии прототипирования, в зависимости от условий работы двигателя (давления в камере сгорания, температуры конструкции). В результате математического моделирования теплопрочностных характеристик силового каркаса сопла ракетного двигателя твердого топлива показана возможность применения разработанной методики при выборе параметров элементов конструкции таких двигателей.

Abstract: The article deals with the experimental and numerical investigation of thermal and gas dynamics processes in combustion chamber of prospective rocket engine developed in Oles Honchar Dnipropetrovsk...

Abstract: The invention discloses a pre-whirling type pintle injector applied to a double-component liquid propellant rocket engine The pre-whirling type pintle injector applied to the double-component liquidpropellant rocket engine comprises a center rod, an inner nozzle, an outer nozzle and a connecting ring, all of which are rotation bodies The inner nozzle coaxially sleeves the center rod, and a cavity for conveying fuel is formed by the inner nozzle and the center rod The outer nozzle coaxially sleeves the outer nozzle, a cavity for conveying an oxidizing agent is formed by the outer nozzle andthe inner nozzle, and the connecting ring coaxially sleeves the outer nozzle Pre-whirling treatment is conducted before oxidizing agent spraying-out, and the problem that a conventional pintle injector is not high in effective mixing efficiency is effectively solved; meanwhile, through the arrangement of two rows of fuel direct flow injection holes, cooling to a liquid film on the wall face of athrust chamber through propellant is effectively achieved; and the performance and the reliability of the double-component liquid propellant rocket engine can be obviously improved