Liquid Rocket Propulsion Systems Research Articles

Use of combined porous mesh material in phase separation and filtering devices design for space liquid propellant rocket propulsion systems: current state and problems

The article presents justification for the need to design and develop porous mesh materials to replace the fine-mesh metal single-layer mono-grids used in various technical structures. Evolution of the porous mesh materials composition is described depending on the technical requirements imposed. It is noted that a special role is given to the requirement to the method of connecting the porous mesh material and the main structure of a technical product in terms of reliability, tightness and absence of harmful effects on its internal structure. It is demonstrated that this requirement is successfully implemented in composition of the combined porous mesh material (metal) with edges specially prepared for using the joining method based on the laser pulse welding. Examples are provided of the use of welded structures based on combined porous mesh material in space, aviation and gas industries in the Russian Federation. Current problems of specialized production of the porous mesh materials are considered, and recommendations are provided that should assist in overcoming these.

엔진 유지점검과 정지능력을 고려한 액체로켓 추진시스템의 신뢰도 모형

Purpose: In this paper, we present a reliability model for a liquid rocket propulsion system with the engine out capability.<BR>Methods: Previously, researchers have developed such a reliability model, in which a failure decay function is used to modify time-independent reliability to time-dependent reliability for each engine. In that model, engine reliability at different thrust values was estimated independently, and the check-up time and mission time were fixed. Such a model is not useful for determining the check-up time necessary to achieve the system reliability target. In this paper, we present a system reliability model in terms of the engine check-up time and mission time. We use the cumulative exposure model to relate the lifetime of an engine operating at different thrust values, assuming that the lifetime of each engine follows a general distribution. By using a mixed exponential distribution, we obtain a closed form expression of system reliability and explain how time-independent engine reliability can be transformed to time-dependent engine reliability.<BR>Result: The proposed model is useful for determining the check-up time that provides the target mission reliability.<BR>Conclusion: The reliability of a liquid rocket propulsion system must be expressed explicitly in terms of the engine reliability depending on the check-up time and the mission time if the check-up time must be determined to achieve the system reliability requirement.

Modeling and verification of working process parameters in gas generators for pressurizing fuel tanks of liquid propellant rocket propulsion systems

Modeling and verification of working process parameters in gas generators for pressurizing fuel tanks of liquid propellant rocket propulsion systems

FEASIBILITY ANALYSIS OF LIQUID PROPELLANT SETTLING IN THE TANKS OF A SPACE PROPULSION SYSTEM AFTER FLYING IN ZERO GRAVITY USING A SEPARATING TURN MANUEVER OF THE ORBITAL UNIT

The paper presents results of preliminary design feasibility studies of using inertial settling of liquid propellant in tanks of space liquid rocket propulsion systems and separating from the propellant the ullage gases formed in zero gravity by means of centrifugal forces generated when the orbital unit performs the programmed turn maneuver proposed by the author. It shows that performing propellant settling and ullage gas separation by means of the separating turn maneuver makes it possible to significantly reduce propellant consumption in propulsion systems for attitude control and ullage engines. Key words: space liquid rocket propulsion system, main engine, separation of ullage gases from liquid propellant, separating turn maneuver of the orbital unit.

Feasibility analysis of liquid propellant settling in the tanks of a space propulsion system after flying in zero gravity using a separating turn maneuver of the orbital unit

The paper presents results of preliminary design feasibility studies of using inertial settling of liquid propellant in tanks of space liquid rocket propulsion systems and separating from the propellant the ullage gases formed in zero gravity by means of centrifugal forces generated when the orbital unit performs the programmed turn maneuver proposed by the author. It shows that performing propellant settling and ullage gas separation by means of the separating turn maneuver makes it possible to significantly reduce propellant consumption in propulsion systems for attitude control and ullage engines. Key words: space liquid rocket propulsion system, main engine, separation of ullage gases from liquid propellant, separating turn maneuver of the orbital unit.

Особенности системного подхода к безопасности стендовых испытаний двигательных установок жидкостных ракет

Пояснена необходимость разработки и реализации специальных мероприятий при применении системного подхода к безопасности стендовых испытаний двигательных установок ракет. Определена основная концепция промышленной безопасности при стендовых испытаниях кислородно-водородных двигательных установок. Изложено и обосновано содержание специальных мероприятий по безопасности испытаний. Предложена функциональная модель безаварийного проведения испытаний и определены мероприятия для её эффективного применения. It explains the need to develop and implement special measures for a systematic aproach to the safety of stand tests of rocket propulsion systems. The basic concept of industrial safety during stand tests of oxygen-hydrogen propulsion systems is defined. The content of special measures for test safety is described and justified. A functional model of trouble-free testing is proposed and measures for effective application are defined.

Study of Manufacturing Processes for Liquid Rocket Turbopump Impellers: Test and Analysis

In the context of the cooperative development of the L75 Liquid Rocket Engine (with 75 kN of thrust) and in the frame of a global enlargement of competence in the field of turbomachines for liquid rocket propulsion systems, Institute of Space Propulsion of the German Aerospace Center (DLR; in German, Deutsches Zentrum für Luft- und Raumfahrt), in Lampoldshausen, and the Brazilian Institute of Aeronautics and Space (IAE; in Portuguese, Instituto de Aeronáutica e Espaço) have managed to produce its first flight ready components for pumps in liquid rocket propulsion systems. Among these components was a series of prototyped shrouded impellers for the oxidizer pumps manufactured in different materials and by different fabrication processes executed by four independent Brazilian and German workshops. Non-destructive and destructive testing methods have been applied for impellers materials validation, and for manufacturing processes verification. A special attention was given to the spin tests results and analysis, which included verification at maximum operational speed and burst testing, in order to determine the failure speed. Each tested impeller reached the required specification for the application in the L75 turbopump. Spin test logic and test results are discussed later on. Therefore, the scope of the present study is to investigate several traditional and cutting-edge processes applied for shrouded turbopump impellers manufacturing. A verification and validation of these processes will also be discussed.

GROUND-BASED EXPERIMENTAL TESTING OF ELEMENTS OF AUTOMATION OF PNEUMATIC-HYDRAULIC SYSTEMS OF ROCKET AND SPACE TECHNOLOGY

The design and production of sophisticated technical systems, which include modern rockets and other aircraft, requires their reliability and trouble-free operation. To achieve the required level of reliability of aerospace products, a wide variety of test methods are applied at all stages of the life cycle. One of the most important systems of the launch vehicle is the pneumatic hydraulic power system of the liquid rocket propulsion system. Development of new and improvement of existing methods of control and diagnostics is one way of increasing the design and technological reliability of products of aviation and space technology. The use of functional diagnostics systems for bench and flight tests significantly increases the reliability and efficiency of space rocket technology. Researches are directed on increase of a level of reliability of products of aerospace branch. Application of systems of functional diagnostics is described at bench tests. The results of experimental researches of elements of automatics of pneumatic hydraulic power supply systems of liquid rocket engines are considered. The technique of processing of experimental data of a pulsing-acoustic method of diagnostics with use of the mathematical technology of recognition of images is presented. Deciding rules of recognition of a technical condition of object of diagnosing by results of tests are resulted. The developed method with a high degree of accuracy allows to determine the technical condition of the object of diagnosis as defective or to detect the presence of characteristic defects. Experimental testing and the proposed method of processing the results showed the efficiency of the method.

Экспериментальная оценка невырабатываемых остатков топлива в баках c внутрибаковыми устройствами капиллярного типа для двигательных установок космических аппаратов

The paper presents the results of experimental studies of the processes of emptying the fuel tanks of spacecraft liquid rocket propulsion systems at the final stages of free (undisturbed) flight in tanks with innertank capillary devices (IT CD). On the basis of the analysis of the emptying process, the dimensionless complex structure is formed. The complexes determine the dependence of the residual propellant on the physical properties of the propellant and the parameters of the IT CD. Experimental studies were performed on the weightlessness stand, using the principle of realization of the conditions of reduced weight in tested equipment free fall. The results of the experimental studies have shown that the main criterion determining the value of residual propellant is the Froude number. The criterion equation for determination of residual propellant is obtained. The influence of viscosity and surface tension of the propellant on the value of residual propellant begins to affect at the values of the Froude numbers exceeding 104.

Экспериментальная оценка рациональных параметров предпускового ускорения для обеспечения запуска маршевых жидкостных ракетных двигателей в условиях невесомости

The paper presents the results of experimental studies of liquid fuel deposition processes in liquid rocket propulsion system tanks under the conditions of free, i.e. unperturbed, orbital and suborbital flight under the influence of small prestart G-load created by auxiliary engines before liquid sustainer engine launch. Having conducted the dimension analysis, we formed the structure of dimensionless groups, which determine the dependence of the time required for fuel deposition to the tank intake, on the g level, the degree of tank filling and the physical properties of the propellant. Experimental studies were carried out on a zero gravity test bench, using the principle of implementing below-G conditions at the free fall of the equipment under test, and on a laboratory aircraft at the "Kepler's parabola" flight. The results of experimental studies have shown that the effect of viscosity on the deposition rate is practically absent. At the same time, in experiments it was found that, starting with Bond numbers exceeding 40 ... 50, the dimensionless separation time does not depend on Bond number, the value of the latter being determined by the minimum radius of free gas inclusions, deformed during when ascending during the separation process. Findings of research show that with a considerable duration of the prestart G load pulse, the movement of gas bubbles in the liquid under the action of a G load pulse becomes quasistationary.

Optimization of Liquid Rocket propulsion system for Reusable Orbital Transfer Vehicle

The work is devoted to the investigation of possible ways to increase the efficiency of reusable orbital transfer vehicle (R-OTV) by using a multi-criteria optimization procedure of liquid rocket propulsion system (LRPS) design parameters. The “Federation” spacecraft is considered as an example payload of R-OTV for orbital transfer mathematical modeling, due to its known initial mass of 16,500 kg and other design parameters. Payload mass and specific cost of payload transfer are chosen as optimality criteria.Multiple tasks are solved during the work: mathematical model and corresponding software for determining the optimal design parameters of R-OTV LRPS by mass and cost criteria are developed, R-OTV and LRPS calculation and optimization are conducted, the technical appearance of LRPS is designed.Mathematical modeling is used to solve the presented problem. A developed mathematical model has a modular structure, in particular, it has modules for liquid rocket engine (LRE) mass calculating, OTV (Orbital Transfer Vehicle) mass calculating module, payload mass, design cost and manufacturing cost calculating modules for LRE and OTV and module for calculating the specific cost of payload transfer to target orbit. The intermediate result is used by LRPS project parameters design optimization module, then by technical appearance design module and LRPS units design module.Multiple calculations are conducted with a developed mathematical model. Verification of the model is carried out by developing the technical appearance of real OTVLRPS. Several variants of LRPS for spacecraft “Federation” type OTV are designed for low lunar orbit, OTV parameters are compared with optimized LRPS and with existing LRPS.

Numerical investigation on flow condensation process during interphase mixing in oxygen pipeline of liquid rocket

AbstractThis paper presents a numerical study on the flow condensation of high temperature gas oxygen in subcooled flowing liquid oxygen and the corresponding mixing process in the oxygen pipeline of a liquid rocket propulsion system. A computational fluid dynamics model is established with the interfacial phase change solved by user‐defined function, and flow field details such as phase distribution and flow condensation length required for safety operation of liquid rocket engine are obtained. Experimental validation of numerical model is performed with HCFC123 as the working fluid, and the calculated void fraction and flow condensation length of gas show good agreement with those of experimental data. The numerical result indicates that when the gas oxygen is injected into the pipeline, continuous annular gas film forms in the vertical section and then accumulates in the inner side of the bend section due to the centrifugal force. Rapid condensation of oxygen occurs in the following horizontal section due to the enhanced interphase heat transfer by the transverse relative motion. The condensation length of gas oxygen varies with the inlet gas/liquid flow rate ratio, liquid subcooling, and inlet velocity. The operable gas/liquid ratio ranges for complete gas condensation in the given bend pipe is obtained for various liquid inlet parameters.

Mathematical modeling of a pneumatic pressure regulator for aerospace application

High precision Pressure Regulators with wide operating pressure ranges are required for steady performance of Liquid Rocket Propulsion Systems. A Non-linear, dynamic mathematical model, for a dome loaded pressure regulator which meets this requirement essential to understand its behavior has been developed. The thermodynamic property variations of the working medium are modeled using mass and energy conservation equations. A characteristic regulator equation consisting of operating forces and Coulomb friction and pneumatic damping is used to model the dynamics of the regulator. The simulation is carried out by varying characteristic geometrical parameters such as the size of the damping orifice, poppet step length and regulator downstream volume. The effect of varying friction force on the start transient of the regulator is studied. Experiments by varying the same parameters as the simulation are carried out to validate the model. The simulation results show good concurrence with the experiments. This model serves as a tool for optimising the performance of the regulator at a particular flow rate and regulated pressure by arriving at appropriate geometrical parameters.

Review of engineering solutions applicable in tests of liquid rocket engines and propulsion systems employing hydrogen as a fuel and relevant safety assurance aspects

The review presents the author's papers on specific features of experimental development of oxygen-hydrogen liquid rocket engines (LRE), namely the 11D56, 11D57, RD0120, KVD1, and a number of propulsion units and power plants, as well as compares some data on propulsion development activities with relevant data obtained abroad. Also has been shown a role of model studies, component-level tests of engine units and systems, including those performed at simulated flight conditions, and integrated tests in support of experimental development of advanced engines and propulsion systems designed for rocket upper stages. There have been considered techniques and equipment intended to ensure safety of ground testing of rocket engines and power plants involving the use of effective diagnostic systems and emergency protection systems.

Materials and Manufacturing Technology for Aerospace Application

This paper gives an overview of current work in materials and manufacturing technology for aerospace application. Finding the best material to use for a particular application is not simple and it depends on many factors including mission requirements like performance and safety, design requirements, strength to density ratio, operating temperature, and material technology prospects like current state of affordable materials/processes technologies. Materials with high specific strength have long been popular with the aerospace industry, as aerospace vehicle made from such materials provide the required strength with less weight, thereby increasing payload and reducing operating cost. Typical examples are polymer matrix composite and aluminum-lithium alloys for aerospace structures. In liquid rocket propulsion systems, improved high-temperature capability offers the greatest performance payoff, with improved mechanical strength and lower weight also being important. For spacecraft, materials with improved resistance to radiation and atomic oxygen are required. Since before a material can be used in an aerospace system, it must be qualified for use, materials qualification procedure is also presented with an example of shared data.

Dynamic Modeling and Analysis of a High Pressure Regulator

Pressure regulator is a common device used to regulate the working pressure for plants and machines. In aerospace propulsion systems, it is used to pressurize liquid propellant rocket tanks at specified pressure for obtaining the required propellant mass flow rate. In this paper, a generalized model is developed to perform dynamic analysis of a pressure regulator so that constant pressure at outlet can be attained. A nonlinear mathematical model of pressure regulator is developed that consists of dynamic equation of pressure, temperature, equation of mass flow rate, and moving shaft inside regulator. The system of nonlinear and coupled differential equations is numerically simulated and computation of pressure and temperature is carried out for required conditions and given design parameters. Valve opening and mass flow rate are also found as a function of given inlet pressure and time. In the end, an analytical solution based on constant mass flow rate assumption is compared with nonlinear formulation. The results demonstrate a high degree of confidence in the nonlinear modeling framework proposed in this paper. The proposed model solves a real problem of liquid rocket propulsion system. For the real system under consideration, inlet pressure of regulator is decreased linearly from 150 bar to 60 bar and outlet pressure of nearly 15 bar is required from pressure regulator for the complete operating time of 19 s.

Performance Evaluation of Small-scale Liquid Pump using a Radial Turbine with H&lt;sub&gt;2&lt;/sub&gt;O&lt;sub&gt;2&lt;/sub&gt; Gas Generator

A gas-turbine-driven pump system was designed and tested. Pump systems are generally driven by an electric motor, but a gas turbine is occasionally used as the power source for high power and performance; for example, the turbopump of a liquid rocket propulsion system. For research on turbopumps, a gas generator, turbine, and pump are developed. After each component is developed, the performance should be confirmed with a link test. In the case of a turbopump, an axial turbine is applied to generate huge torque, but a radial inflow turbine was used in this research. A radial inflow turbine can be obtained easily and mass-produced. System analysis was conducted using a link test with all components at once. The turbine generated shaft power under steady operation of the gas generator, and the pump performance was measured using a stepped closing valve at the pump exit. Turbine heat loss was considered and a slip factor was applied to the impeller design to modify performance. This research shows the feasibility of developing a pump system operated by a radial inflow turbine and its application to a small turbopump for a hybrid rocket propulsion system and an air-independent system.

Methodology for Designing a Conceptual Model of Knowledge Mapping in the Field of Liquid Rocket Propulsion System

mapping is a process of surveying, assessing and linking the information, knowledge, competencies and proficiencies held by individuals and groups with in an organization and in the subject Components that can be found on a knowledge map are primarily the organization‟s knowledge assets, also it will reflect, where the subject experts are lagging and its importance. The map will also include information as to their source, their format, the meta-data as well as stakeholders. Liquid Propulsion Rocket Systems (LPRS) is a subject domain, which deals with the Liquid Propulsion systems for Space Launch Vehicles and Satellites. By analyzing the literature published all over the world and also the internal reports published in the organization, extracting the concepts, it is proposed to build the knowledge map, by linking various concepts to model the LPRS domain. By mapping the above domain, one can visualize how the concepts are evolved to form new concepts. This paper explains the design methodology to develop the Conceptual model of Knowledge Mapping in the LPRS domain.

Research of Pipeline Fault Diagnosis for Liquid Rocket Propulsion System

The applications of healthy monitoring and fault diagnosis's system can enhance the reliability and safety of the whole vehicle, which has significance to detect and isolate fault as early as possible and that tragedy can be avoided. In this paper, pipeline fault simulation analysis and fault diagnosis for LRE is studied as direct-inverse problem. Firstly, the failure model was constituted for pipeline to analysis the necessary condition for fault isolation. Then the conclusion that strategy of fault diagnosis was build by analysis the simulation result based on AMESim. Finally, the fault diagnosis algorithm was validated by test data. The results indicate that the fault diagnosis algorithm can detect fault exactly and effectively. There are no false alarm to normal test data and no false alarm to other components fault. Consequently the fault isolation result could be reached.

Transient thermal analysis of a cryogenic oxidizer tank in the liquid rocket propulsion system during the prelaunch helium gas pressurization

The transient thermal behavior in the cryogenic oxidizer tank for the liquid propulsion system of the KSLV-I satellite launching space vehicle is theoretically investigated for the pressurization process by gaseous helium injection followed by the readiness check stage. The numerical model is established using the transient mass and energy conservation of oxygen/helium mixtures both in ullage and liquid regions. At the liquid-gas interface, various modes of heat and mass transfer are considered, including the liquid evaporation and the helium absorption. The present study focuses on the effects of the increasing pressurization level on some of important properties that should be considered in the propulsion system design, such as tank pressure, ullage gas temperature, and evaporation of liquid oxygen. Particularly, the tank pressure drop after the pressurization is investigated in the proposed design of propulsion system. Also, the effect of the initial loading of liquid oxygen in the oxidizer tank is studied and discussed. As for the helium absorption into the liquid region, its mass is negligibly small and it should not be the concern in design and operation of the oxidizer tank system.