Showing papers on "Liquid-propellant rocket published in 2014"

Optimization of Hybrid Upper-Stage Motor with Coupled Evolutionary/Indirect Procedure

Abstract: Upper-stage motors used in small launchers constitute an application where hybrid rocket motors may be competitive. A coupled optimization of motor design and trajectory is needed for such an application due to mission characteristics and motor features. The present article presents a cooperative evolutionary method nested with an indirect approach to perform the coupled optimization of hybrid rocket motor and trajectory for an upper stage. The evolutionary method optimizes the parameters that affect the motor design (e.g., grain geometry) and feed system, whereas the indirect method optimizes the trajectory (i.e., thrust direction and motor switching times) for a given motor and mission. A mission profile based on the Vega launcher is considered and the performance index is the payload inserted into the final orbit. The hybrid rocket motor powers the third and last stage and has a pressurizing feed system that is partially regulated. The characteristics of the first and second solid rocket motor stages a...

Combustion Stability Characteristics of the Project Morpheus Liquid Oxygen / Liquid Methane Main Engine

Abstract: The project Morpheus liquid oxygen (LOX) / liquid methane (LCH4) main engine is a Johnson Space Center (JSC) designed ~5,000 lbf-thrust, 4:1 throttling, pressure-fed cryogenic engine using an impinging element injector design. The engine met or exceeded all performance requirements without experiencing any in- ight failures, but the engine exhibited acoustic-coupled combustion instabilities during sea-level ground-based testing. First tangential (1T), rst radial (1R), 1T1R, and higher order modes were triggered by conditions during the Morpheus vehicle derived low chamber pressure startup sequence. The instability was never observed to initiate during mainstage, even at low power levels. Ground-interaction acoustics aggravated the instability in vehicle tests. Analysis of more than 200 hot re tests on the Morpheus vehicle and Stennis Space Center (SSC) test stand showed a relationship between ignition stability and injector/chamber pressure. The instability had the distinct characteristic of initiating at high relative injection pressure drop at low chamber pressure during the start sequence. Data analysis suggests that the two-phase density during engine start results in a high injection velocity, possibly triggering the instabilities predicted by the Hewitt stability curves. Engine ignition instability was successfully mitigated via a higher-chamber pressure start sequence (e.g., ~50% power level vs ~30%) and operational propellant start temperature limits that maintained \cold LOX" and \warm methane" at the engine inlet. The main engine successfully demonstrated 4:1 throttling without chugging during mainstage, but chug instabilities were observed during some engine shutdown sequences at low injector pressure drop, especially during vehicle landing.

Coupled Heat Transfer Analysis in Regeneratively Cooled Thrust Chambers

Abstract: The coupled hot gas–wall–coolant environment that occurs in regeneratively cooled liquid rocket engines is studied by a computational procedure able to provide a quick and reliable prediction of thrust chamber wall temperature and heat flux as well as coolant flow characteristics, like pressure drop and temperature gain in the regenerative circuit. The coupled analysis is performed by means of a computational fluid dynamics solver of the Reynolds-averaged Navier–Stokes equations for the hot gas flow and by a simplified quasi-two-dimensional approach, which widely relies on semiempirical relations, for the coolant flow and wall structure heat transfer in the cooling channels. Coupled computations of the space shuttle main engine main combustion chamber are performed and compared with available literature data. Results show a reasonable agreement in terms of coolant pressure drop and temperature gain with nominal data, whereas the computed wall temperature peak is closer to hot-firing test data than to the ...

Effect of cooling channel aspect ratio on rocket thermal behavior

Abstract: A tradeoff analysis is performed on a test case representative of the cooling system of 1 MN thrust class oxygen/hydrogen liquid rocket engine. The aim of the analysis is to find the channel aspect ratio that maximizes the heat extracted from the hot gas, for a given coolant pressure drop and hot-gas side wall temperature. The analysis requires many cooling channel flow calculations that are performed by means of a simplified model, referred to as quasi-two-dimensional, and a three-dimensional conjugate heat transfer model based on numerical integration of the Navier–Stokes and Fourier’s equations. Both models are able to describe, with different level of details, the whole cooling device composed by the coolant and the solid domain, which is exposed to the hot gas. The fast quasi-two-dimensional approach is used to select channel geometries showing the same pressure loss. Discussion is made on results obtained with the more accurate three-dimensional model. Results identify, for the selected test case, a...

Parametric Analysis of Cooling Properties of Candidate Expander-Cycle Fuels

Abstract: Flow evolution and heat transfer capability in the cooling system of liquid rocket engines heavily depend on propellant thermophysical properties. Coolant thermophysical property analysis and modeling is therefore important to study the possibility of relying on a regenerative cooling system, whose performance is crucial to determine feasibility and convenience of pump-fed liquid rocket cycles of the expander type. The aim of the present study is to compare the behavior of different liquid fuels for expander-cycle engines. They are light hydrocarbons, binary mixtures of them, and liquefied natural gas, which is a mixture made basically of methane and minor fractions of other light hydrocarbons and nitrogen. A parametric analysis is carried out by a validated numerical solver to compare temperature increase, pressure loss, and heat transfer evolution for the different fuels along the same straight tube and subjected to assigned heat fluxes. Results show that similar engine performance can be obtained by the different candidate expander-cycle fuels, but significant differences can be seen in the flow evolution through the cooling channels.

Tri-propellant rocket engine for space launch applications

Abstract: A tri-propellant rocket engine for space launch applications is disclosed. The tri-propellant rocket engine comprises three main assemblies: an injector, a chamber head, and a chamber.

Development Status of L75: A Brazilian Liquid Propellant Rocket Engine

Abstract: This paper provides an overview of the design and development process of the L75 Liquid Propellant Rocket Engine (LPRE). Being developed at Institute of Aeronautics and Space (IAE), it is the first Brazilian open-cycle liquid rocket engine pressurized by turbopump, designed to deliver 75 kN of thrust in vacuum as a cryogenic upper stage engine using liquid oxygen and ethanol. The Preliminary Design Review (PDR) was accomplished in December 2011 and December 2012 the project received from Agencia Espacial Brasileira (AEB) a financial support, through an agreement with Fundacao de Desenvolvimento da Pesquisa (FUNDEP), to proceeding with the development of models and tests. The main components of the engine, briefly described here, are thrust chamber assembly, gas generator, turbopump assembly, control system and ignition system.

The development of 7500 N variable thrust engine for Chang’E-3

Abstract: To implement orbit maneuvering, rendezvous and docking, and soft landing on planets, the throttling liquid rocket engine is preferable. The 7500 N variable thrust engine applied in the ChangE-3 prober is the first throttling liquid rocket engine in china. It can throttle fast with high precision and continuous variability based on the control commands. The trajectory correction and soft landing on the moon is accomplished. This paper provides the development of the 7500 N variable thrust engine, including the technique scenarios, the key technologies and the tests. The test and flight results show that the system is feasible, its performance is high and it works reliably.

Effects of Orifice Geometry on Gelled Propellants Sprayed from Impinging-Jet Injectors

Abstract: G ELLED propellants are liquid propellants with an addition of one or more gellants that impart non-Newtonian behavior [1]. To obtain high combustion efficiency, fine atomization is necessary for this type of propellant. However, the addition of gellants can cause an increase in gel viscosity, making it more difficult to atomize the propellant into fine drops and reach high combustion efficiency in rocket engines. Impinging-jet injectors have been widely used in liquid rocket engines. The shape of the liquid sheet, breakup length, and drop size formed by the impinging jets for Newtonian fluids have been studied extensively [2–5]; the effect of viscosity on sheet formation and breakup has also been studied at length [6–11]. However, the effects of jet-orifice geometry on spray characteristics have received much less attention. Some researchers [12–15] reported that orifice geometry (e.g., the distance between the two orifices, orifice shape, aspect ratio. and length-to-diameter ratio) strongly affects the breakup process of liquid sheet formed by the collision of two jets. The impinging-jet injector is also the most popular injector type used to atomize gelled propellants. The atomization characteristics of gels in impinging-jet injectors have been studied bymany researchers [16–23]. However, all the impinging-jet injectors used in the aforementioned works had exit orifices with a circular section. Although there are some works that examined the flow and breakup characteristics of elliptical liquid jets [24,25] and rectangular jets [26], the spray characteristics of impinging-jet injectors with orifices of different shapes for gelled propellants have not been reported. This Note reports the experimental observations of flow structure and breakup characteristics of a gelled propellant simulant from impinging-jet injectors. The effect of injector orifice section shape and geometry on the breakup characteristics of the liquid sheet is principally discussed.

Liquid Rocket Analysis (LiRA): Development of a Liquid Bi-Propellant Rocket Engine Design, Analysis and Optimization Tool

Abstract: The need to increase specific impulse of rocket launcher engines has lead design engineers to start development of ever more complex engine cycles starting from simple gas pressure fed engines to the very complex staged combustion cycle engines. Currently a wide range of different cycles does exist. These cycles not only differ in terms of performance, but also in terms of mass, cost, reliability etc., which in general makes it difficult to quickly determine which cycle is best suited for a certain mission or task. For this a tool that is capable of analysing different cycles and the impact of design choices has been developed. The tool, named LiRA, has as goal giving the user better system level understanding of the different possible engine cycles and the functions of the components; it can therefore provide valuable and time saving assistance during design or in analysis and optimisation studies. A modular approach is applied where engine components are sized using a performance, dimension and mass model who make use of corrected ideal rocket theory and empirical relations. This work focuses mainly on the methodology and the construction of the models, and further includes the optimisation of an upper stage and a verification, validation and uncertainty and sensitivity analysis of the tool and the optimisation to assess its accuracy, precision and applicability. The completed program has proved to confirm known trends and known cycle characteristics like the mass savings that can be achieved when using turbo-pump fed engines instead of pressure fed engines for mid- to high-thrust booster applications. Further the superior performance of closed cycles, especially staged combustion cycles has been confirmed, and some cycle specific design choices like the need of bypasses in expander cycles have been explained. The tool however is not complete and should be expanded; the addition of a cost and reliability assessment model is for example strongly recommended. There also remain issues with the accuracy and uncertainty of certain estimates which make that the current version should only be used for comparative studies.

Effects of Injector Design on Combustion Stability in Hybrid Rockets Using Self-Pressurizing Oxidizers

Abstract: Interest in nitrous oxide based hybrid rockets is at an all time high. Nitrous oxide (N2O) is a unique oxidizer because it exhibits a high vapor pressure at room temperature ( 730 psia or 5.03 MPa). Due to this high vapor pressure, liquid nitrous oxide can be expelled from a tank without the use of complicated pumps or pressurization systems required by most traditional liquid rocket systems. This results in weight savings and design simplicity. Additional benefits of nitrous oxide include storability, ease of handling, and relative safety compared to traditional liquid oxidizers. The design and modeling of injectors for use with high vapor pressure propellants such as nitrous oxide is made complicated due to the possibility of two-phase flow. The operating pressures within rocket propellant feed systems can often drop below the vapor pressure for these unique propellants, especially within the injector. Injectors operating under these conditions are likely to exhibit cavitation, resulting in significant vapor formation and limitation of mass flow rate. With the introduction of two-phase flow, a critical flow regime can be observed, where the flow rate is independent of backpressure (similar to choking). For a simple orifice style injector, it has been demonstrated that critical flow occurs when the downstream pressure falls su ciently below the vapor pressure, ensuring bulk vapor formation within the injector element. It has been proposed to leverage the insensitivity of critical mass flow rate to downstream pressure as a means of preventing the occurrence of feed system coupled combustion instabilities in hybrid rockets utilizing nitrous oxide. The Peregrine Sounding Rocket is a hybrid rocket that runs on para n wax and nitrous oxide. Its development is a joint e ort between NASA Ames Research Center, Stanford University, and Space Propulsion Group, Inc. For years, progress of the Peregrine program has been hampered by combustion instability problems. Based upon results from the aforementioned small scale injector experiments, a powerful, yet simple solution to the so-called feed system coupled combustion instability was discovered, the details of which are presented. This work also led to the invention of a new class of rocket propellant injectors designed specifically to decrease the likelihood of this type of combustion instability. An in-depth discussion of the proposed design and operation of this novel injection scheme is included, along with the presentation of some prototype cold flow testing results which served as a successful proof of concept.

Investigation on Heat Transfer in Small Hydrocarbon Rocket Combustion Chambers

Abstract: Low costs and nonhazardous properties draw interest in application of hydrocarbon fuels in liquid rocket engines. Within this work, results of experiments on heat transfer, film cooling, transpiration cooled and convectively cooled fiber-reinforced ceramics conducted at a kerosene/oxygen rocket combustion chamber test facility are presented. The experimental data serves as the base for design and validation of simple-to-use models and correlations, which allow estimates of heat flux and cooling needs for the preliminary design of hydrocarbon fuel rocket engines.

Receptivity of a Cryogenic Coaxial Gas-Liquid Jet to Acoustic Disturbances

Abstract: : An experimental study has been conducted at the Air Force Research Laboratory at Edwards Air Force Base to explore the receptivity of cryogenic coaxial jet flows to transverse acoustic disturbances. The shear coaxial jet flow employed liquid nitrogen in the inner jet and cooled helium in the outer annular jet to represent the nominal fluid dynamical conditions of an oxygen/hydrogen liquid rocket engine injector. The injector flow is submerged in a chamber that experiences a monotonic transverse acoustic resonance characteristic of a rocket chamber in the presence of combustion instability. The coaxial jet is exposed to a variety of acoustic conditions including different frequencies, amplitudes, and locations within the resonant mode shape. High-speed back-lit images were captured to record the behavior of the natural (unforced) and forced coaxial jets. Proper orthogonal decomposition and spectral analysis were used to extract natural and forced modes. Convective modes are extracted, and a new Strouhal number is used to characterize the dominant natural convective mode that is analogous to the preferred mode in free jets. The threshold of receptivity was found for a number of different injector flows and acoustic forcing conditions. The results indicate that the dimensionless frequency plays an important role, and there exists a finite forcing amplitude at which the threshold of receptivity occurs. The receptivity threshold and post receptivity response provides useful insight on the suitability of a given injector design for specific rocket combustion chamber conditions.

Genetic algorithm to optimize the design of main combustor and gas generator in liquid rocket engines

Abstract: A genetic algorithm was used to develop optimal design methods for the regenerative cooled combustor and fuel-rich gas generator of a liquid rocket engine. For the combustor design, a chemical equilibrium analysis was applied, and the profile was calculated using Rao’s method. One-dimensional heat transfer was assumed along the profile, and cooling channels were designed. For the gas-generator design, non-equilibrium properties were derived from a counterflow analysis, and a vaporization model for the fuel droplet was adopted to calculate residence time. Finally, a genetic algorithm was adopted to optimize the designs. The combustor and gas generator were optimally designed for 30-tonf, 75-tonf, and 150-tonf engines. The optimized combustors demonstrated superior design characteristics when compared with previous non-optimized results. Wall temperatures at the nozzle throat were optimized to satisfy the requirement of 800 K, and specific impulses were maximized. In addition, the target turbine power and a burned-gas temperature of 1000 K were obtained from the optimized gas-generator design.

Research on the Combustion of Bio-Derived Fuels in Hybrid Propellant Rocket Engine

Abstract: Paper describes the extension of the research of the combustion of non-conventional bioderived fuels in a small-scale Hybrid Propellant Rocket Engine (HPRE) started in 2010. The long-term goal of the research is the investigation of the combustion of non-conventional bioderived fuels such as paraffin, beeswax, lard with different oxidizers (oxygen, hydrogen peroxide, nitrous oxide), including combustion of above fuels with additives, and obtaining regression rate formulas for listed propellants. The small-scale HPRE, test fixture and instrumentation system have been designed, manufactured, assembled and used for the study and analysis of the combustion. The results of the research of combustion of paraffin were presented at the 51 AIAA Meeting and Exhibit. Paper summarizes the results of the study of combustion of bee’s wax with oxygen and represents regression rate formulas. Experimental data were compared with theoretical results and existing data of other researchers. The paper also describes indirect method and shows preliminary results of the estimation of losses of unburned fuel, based on the measurements of exit temperature of combustion products and calculation of actual values of the equivalence ratios and combustion temperatures in the combustion chamber of HPRE.

Development of a Facility for Combustion Stability Experiments at Supercritical Pressure

Abstract: : Combustion instability in liquid rocket engines can have severe consequences including degraded performance, accelerated component wear, and potentially catastrophic failure High-frequency instabilities, which are generally the most harmful in liquid rocket engines, can be driven by interactions between disturbances associated with transverse acoustic resonances and the combustion process The combustion response to acoustic perturbation is a critical component of the instability mechanism, and is in general not well understood The current paper describes an experimental facility at the Air Force Research Laboratory (AFRL) at Edwards Air Force Base that is intended to investigate the coupling between transverse acoustic resonances and single/multiple liquid rocket engine injector flames Critical aspects of the facility will be described, including the capability to operate at supercritical pressures that are relevant to high-performance liquid rocket engines, accurately-controlled and cryogenically-conditioned propellants, and optical access to facilitate the use of advanced diagnostics The transverse acoustic resonance is induced through the use of carefully-controlled piezo-sirens, allowing monochromatic excitation across a range of amplitudes at a number of discrete frequencies The location of the flame within the acoustic resonance mode shape can also be varied through relative phase control of the two acoustic sources The operating space of the facility, for oxygen and hydrogen operation, will be described Preliminary non-reacting and reacting data will also be presented to demonstrate the quality of operation of this facility It is anticipated that future results generated using this facility will provide both fundamental insight into the acoustic-flame interactions as well as provide a database useful for validating combustion instability models

Effective Modeling of Conjugate Heat Transfer and Hydraulics for the Regenerative Cooling Design of Kerosene Rocket Engines

Abstract: The present work is motivated to develop a unified framework to simulate multi-physical processes which are crucial for trade-off design of liquid rocket thrust chambers among propulsive performance, regenerative cooling, and pressure budget. In this paper, an effective modeling of conjugate heat transfer and hydraulics through the regenerative cooling passage has been performed to quantitatively evaluate detailed cooling designs, including spirally twisted channels and bidirectionally branched circuit, as well as to provide the wall heat flux to a compressible reacting flow solver in an interactively coupled manner. The kerosene fuel used as coolant is modeled by a three-component physical surrogate, and the fluid properties required for calculation of a Nusselt number correlation and empirical resistance coefficients are computed over the entire thermodynamic states from compressed liquid to supercritical fluid using the NIST SUPERTRAPP. The present method has been applied to an actual regeneratively co...

Influence of Hydrogen Temperature on the Acoustics of a Rocket Engine Combustion chamber operated with LOX/H2 at Representative Conditions

Abstract: Since the late 1960s, hydrogen temperature is known to have a significant effect on high frequency combustion instabilities in liquid propellant rocket engines with the propellant combination hydrogen and oxygen. In the framework of experimental investigation of high frequency combustion instabilities at the DLR Institute of Space Propulsion in Lampoldshausen, self-excited combustion instabilities were found in an experimental rocket engine combustor with the propellant combination hydrogen / oxygen at a hydrogen temperature of 95K. In order to analyse the influence of hydrogen temperature on these self-excited combustion instabilities, a hydrogen temperature ramping test run was performed. The hydrogen temperature was varied between 40 and 135K. In agreement with previous test runs, highest amplitudes of the 1T mode were found for hydrogen temperatures around 100K.

Thermo-Structural and Thermo-Fluid Dynamics Analyses Supporting the Design of the Cooling System of a Methane Liquid Rocket Engine

Abstract: The HYPROB Program, developed by the Italian Aerospace Research Centre, has the aim to increase the system design and manufacturing capabilities on liquid oxygen-methane rocket engines. It foresees the designing, manufacturing and testing of a ground engine demonstrator of three tons thrust. The demonstrator baseline concept is featured by 18 injectors and it is regeneratively cooled by using liquid methane. In particular, the cooling system is made by a constant number of axial channels and the counter-flow architecture has been chosen; methane enters the channels in the nozzle region in supercritical liquid condition, is heated by the combustion gases along the cooling jacket and then injected into the combustion chamber as a supercritical gas by means of the injection head.The goal of this paper is to describe the thermo-structural and the thermo-fluid dynamic analyses that have been performed in order to support the design activities aiming at identifying the optimal configuration of the cooling jacket in terms of number of channels, rib height and width. In fact, a fully 3-D model, regarding a single channel, heated by the design input heat flux has been considered in order to perform CFD simulations aiming at describing the thermo-fluid dynamic behavior of methane. The results in terms of convective heat transfer coefficients have been taken into account as inputs for the thermo-structural simulations on the most critical sections of the cooling jacket. The thermo-structural activity has been conducted on the demonstrator by means of a Finite Element Method code taking into account the visco-plastic behavior of the adopted materials. In particular, transient thermal analyses and static structural analyses have been performed using ANSYS code on a 2-D model. These analyses have demonstrated that the cooling jacket can withstand the design goal of 5 thermo-mechanical cycles with a safety factor equal to 4 considering a firing time equal to 30 seconds.Copyright © 2014 by ASME

Dual mode chemical rocket engine, and dual mode propulsion system comprising the rocket engine

Abstract: The invention relates generally to dual mode bipropellant chemical rocket propulsion systems to be used in aerospace applications for 1) orbit raising, orbit manoeuvres and maintenance, attitude control and deorbiting of spacecraft, and/or 2) propellant settling, attitude and roll control of missiles, launchers and space planes. The present invention also relates to a dual mode chemical rocket engine for use in such systems. The engine uses low-hazardous storable liquid propellants and can be operated either in monopropellant mode or in bipropellant mode. The monopropellants used are a low-hazard liquid fuel-rich monopropellant, and a low-hazard liquid oxidizer-rich monopropellant, respectively.

Testing system of liquid rocket supercritical helium supercharging

Abstract: The utility model discloses a testing system of liquid rocket supercritical helium supercharging The testing system comprises a helium tank (1), a normal temperature solenoid valve (2), a pressure reducer (3), an orifice plate front pressure gauge (41), an orifice plate rear pressure gauge (42), a normal temperature supercharging orifice plate (5), a liquid helium storage tank (6), a liquid helium storage tank pressure gauge (7), a liquid helium storage tank thermometer (8), an electronic scale (9), a heating heat exchanger (10), a stop valve (11), a supercharging solenoid valve (12), a displacement pipeline (13), a low temperature supercharging orifice plate (14), a flowmeter (15), a storage tank (16), a storage tank pressure gauge (17), a storage tank thermometer (18), an exhaust solenoid valve (19), and an exhaust orifice plate (20) The testing system of liquid rocket supercritical helium supercharging can check matching performance of a supercritical helium supercharging system, obtains a rule of a relation of the supercritical helium filling amount and the normal temperature supercharging air amount, and reveals difficulties of applying the testing system to a rocket in the future

Combustion Performance and Stability of a Porous Injector Compared with a State-of-the-Art Coaxial Injector

Abstract: Injector behavior is of utmost importance for the performance and stability of liquid rocket engines The injection system has to provide for a highly e�cient mixing and consequent chemical reaction of the propellants at a minimum chamber length This article presents results of the experimental investigation for a new injection concept based on the application of porous materials Porous Injectors represent a novel method of propellant injection in liquid propellant rocket engines Earlier investigations revealed their high potential for gas/liquid propellant combi- nations, eg LOX/H2 The simple and low-cost manufacturing process and the good behavior during throttling of this class of injectors are signi�cant advantages compared with coaxial injectors commonly employed for this propellant combination The present study focusses on a direct experimental comparison of a 168-element porous injector (API) with an injector head equipped with 19 state-of-the-art coaxial injector elements comparable to Ariane 5`s Vulcain II engine The hot �re tests were conducted at the P8 test bench in Lampoldshausen An 80 mm diameter calorimeter combustion chamber built by Airbus DS and identical measurement equipment were used with both injector head con�gurations, allowing a direct comparison of the test data The combustion performance, stability, the wall heat �ux and the axial pressure pro�le were determined for operating conditions typical of the operating regime of Vulcain II This included combustion chamber pressures of up to 115 bar and oxidizer to fuel mixture ratios of 54 up to 72

Device for laser ignition of fuel in gas generator of liquid propellant rocket engine

Abstract: FIELD: power engineeringSUBSTANCE: in a device for laser ignition of a gas generator of a liquid propellant rocket engine comprising zones of combustion and mixing of fuel components, comprising a source of electric energy, a block of excitation with optical fibre, at least one laser spark plug with a focusing lens, installed on a nozzle plate of a combustion chamber, having inner and outer walls, differing by the fact that the laser spark plug is installed at the periphery of the nozzle plate at the angle to the axis of the combustion chamber and is made in the form of a hollow sleeve installed outside of the gas generator connected with a metal bushing with the inner cavity of the gas generator, inside the sleeve there is at least one microchip-laser connected by a vacuum tube with a focusing lens at the end, sealed relative to the metal bushing The angle of setting of the laser plug to a fire bottom makes from 60 to 80 degrees The laser focusing may be arranged in the zone of combustion of fuel components Inside each sleeve there is a damping facility The damping facility is made of material having high heat conductivity The damping facility may be metal rubber Inside each sleeve there is a heat accumulator The heat accumulator is made in the form of a container of cylindrical shape with a central hole, the cavity of which is fully or partially fixed with a heat-accumulating material, and is installed concentrically to the axis of the sleeve The heat-accumulating material is sodium acetate trihydrate The focusing lens is installed inside the metal bushing and is deepened into it relative to the inner surface of the inner wall of the fire bottom of the gas generator The value of deepening of the focusing lens is arranged such that the focusing lens is installed inside the sleeve, for instance, near its bottomEFFECT: increased reliability of an ignition device11 cl, 17 dwg

Experimental investigation of methane in transcritical conditions

Abstract: The use of the Methane as coolant in a regenerative liquid rocket engine (LRE) presents some difficulties since transcritical fluidynamics operating conditions occur in the cooling channels. Transcritical conditions cause large fluid properties variation that strongly influences the coolant performance. The HYPROB program is carried out by CIRA under contract by the Italian Ministry of Research with the main objective to improve National system and technology capabilities on liquid rocket engines for future space applications, with specific regard to LOx/LCH4 technology. Its main objective is to develop an test a LOX/LCH4 demonstrator. In order to match this objective a specific breadboard, the Methane Thermal Properties (MTP) breadboard has been manufactured and test. It is based on an electrical heating of a single representative cooling channel that has the aim to validate numerical methodologies and to improve the understanding of relevant physics of methane thermal properties in transcritical conditions. The experimental test campaign has been succsesfull performed at Maurice J. Zucrow Laboratories in Purdue University and the paper presents the main results.

Cooling channel analysis of a LOX/LCH4 rocket engine demonstrator

Abstract: A computational procedure able to describe the coupled hot-gas/wall/coolant environment that occurs in most liquid rocket engines is presented and demonstrated. The coupled analysis is performed by loose coupling of the two-dimensional axisymmetric ReynoldsAveraged Navier-Stokes equations for the hot-gas flow and the conjugate three-dimensional model for the coolant flow and solid material heat transfer in the regenerative cooling circuit. The latter model is in turn based on the coupled Reynolds-Averaged Navier-Stokes equations for the coolant flow and Fourier equation for the thermal conduction in the solid material. In this study, the thermal behavior of a regeneratively cooled oxygen/methane engine demonstrator is analyzed in detail. Starting from a nominal operative condition of the engine, different levels of channel surface roughness and coolant mass flow rate are considered in order to understand their influence on the heat transfer capability of the cooling system. Results show that the heat transfer can be markedly impaired if the operating parameters undergo rather minor changes with respect to the nominal condition.

Method of gasification process simulation for residual liquid rocket fuel and device for method implementation

Abstract: FIELD: engines and pumps.SUBSTANCE: invention relates to simulation devices and can be used in modelling of processes of gasification of liquid fuel residue in tanks of separated stages (SS) of launch vehicles (LV). The device for simulation of process of gasification of residue of the liquid component of rocket fuel in SS tanks of LV stages contains the experimental unit (EU) in the form of the simulated tank with the pan for gasifiable liquid in tanks, temperature and pressure sensors, cylinders with pre-treated gas, the electropneumatic valve, the logic unit, the electric heater (EH). EU is injected with the heat carrier (HC) in the form of gas flow in the form of pre-treated gas with the pre-set parameters and chemical composition corresponding to the products of combustion of burned fuel in the gas generator chamber, the pre-set interaction conditions in the zone of contact of HC with the liquid surface are provided, temperature and pressure are measured in various points, from the thermal sensor the signal is generated into the logic unit, the signal from the thermal sensor is compared with the pre-set signal for EH switching on or off, EN is switched on or off depending on weight coefficients, deviations of HC current temperature, rates of HC cooling and temperature increase, the stationary mode of average temperature of systems is achieved, HC supply into EU is stopped.EFFECT: invention allows to improve the experimental accuracy of gasification process.3 cl, 2 dwg

Liquid rocket engine with hybrid electric motor driven pump

Abstract: A nuclear thermal rocket with a superconducting electric motor driven boost pump submerged within a tank of liquid hydrogen, where the boost pump is driven by both an electric motor and a turbine. The boost pump can be submerged in liquid hydrogen so that the electric motor operates as a superconducting motor. Also, a turbopump for a rocket engine can include both a turbine and an electric motor to drive the liquid oxidizer and liquid fuel pumps of the turbopump.