Liquid Propellant Rocket Engines Research Articles

Simple Learning-Based Robust Nonlinear Control of an Electric Pump for Liquid-Propellant Rocket Engines

This paper presents a robust nonlinear control strategy for an electric pump for liquid-propellant rocket engines. In order to compensate for model uncertainties and disturbances, a gradient-descent-based simple learning control strategy is employed that minimizes the cost function defined on the error dynamics of the nonlinear system. Detailed stability analysis for the nonlinear system is provided. Computer simulation results are included to demonstrate the effectiveness of the nonlinear control method using an electric pump model consisting of a brushless permanent-magnet direct current (DC) motor and a centrifugal pump. In particular, it is shown that by employing the developed nonlinear controller, the mass flow rate can be successfully kept at a certain level, can be changed instantly from one level to another (immediate decrease or increase), or can be changed linearly/nonlinearly, gradually, and continually for a certain period.

AN OVERVIEW ON THE DESIGN VARIANTS FOR ORGANIZATION OF THE LIQUID FILM COOLING IN LPRE COMBUSTION CHAMBERS

Abstract. New technologies always create a trend to modernize and improve existing designs with new opportunities. At present, additive technologies have one of the most significant development rates among other manufacturing technologies. For modern liquid-propellant rocket engines, there is also a general trend towards increasing the efficiency of engines, finding new technological solutions that make it possible to simplify manufacturing technology, reduce technological costs and increase reliability. In this paper, known design solutions which employ film cooling for thermal protection of the walls of the engine chamber are considered. The main ways of organizing internal cooling are analyzed, the main mechanisms that define the thermal protection of the walls of the LRE chamber are considered. An overview of the existing design variants of devices which organize internal film cooling is given. The design solutions for film cooling of both the list of the real engines RD-105, RD-106, RD-0110, RD-115, RD-119 and known patent solutions are considered. The main features of such devices are shown and the influence of some design differences on the efficiency of the film cooling is analyzed. The main characteristics defining the efficiency of the device for film cooling are given. A review and analysis of the implemented design solutions for film cooling showed that the most efficient designs use the following elements, such as a gas-dynamic redan, a profiled wall of the predetermined space, and various types of inlet and outlet slots. Each design decision is based on the results of experimental testing of the engine. All these elements of film cooling design are very dependent on the accuracy of the manufacturing technology. Capabilities of additive manufacturing allow make it possible to create new designs with stable geometric characteristics in the form of a single part. The paper considers the options for applying the SLM additive manufacturing technology to create the design of the film cooling rings for liquid-propellant rocket engines.

TEORETICAL-EXPERIMENTAL COMPARISON OF THE MODELS OF ADDITIVELY MANUFACTURED SWIRL INJECTORS BASED ON THE HYDRAULIC TESTING RESULTS

Annotation. Recent success in development of the technology stipulated the appearance of advanced additive manufacturing aimed to shift conventional subtractive manufacturing from its position in the engineering field. Thus, for LPRE manufacturing there is an obvious advantage of additive manufacturing in terms of ability to produce single-piece parts with complex geometry must be seriously considered. In the modern world there is a constant competition between complex engineering decisions and cost efficiency of the production on the market of liquid propellant rocket engines. On the one hand, mixing head is one of the most demanding parts of LPRE in terms of designing. On the other hand, the earliest stage of its development starts with the choice of a type of mixing elements to ensure combustion stability, efficiency, and overall production cost [1]. Typically this step requires advanced knowledge in the real characteristics of produced injectors, or extensive experimental study should be carried out otherwise. In this work a problem of determining the characteristics of additively manufactured swirl injectors is considered. Laser Powder Bed Fusion [10] was chosen as a production method using Haynes 230 as a main material. A range of the swirl injectors with varied geometrical parameters was manufactured. Hydraulic testing was carried out in a wide range of Reynolds numbers for determining flow characteristics. Typically, discharge coefficient and spray angle are the desired values to know. Sauter mean diameter and film thickness were not studied due to lack of equipment. Considering a problem of prediction of the main characteristics of the produced injectors, a review on known equations for calculation of the main parameters was performed. Comparative analysis between calculated and experimental data was carried out. Results show that there is severe discrepancy between calculated values of characteristics of injectors and experimentally obtained data for additively manufactured swirl injectors. Even though the conducted study gives a base to form consistent conclusions of the behavior of the obtained characteristics, additional investigation is required.

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.

Analytical and Numerical Study of Mechanical Behaviour of Bolted Joints –A Review

Any component that is subjected to excessive heat or cold experiences behavioural changes. The nut bolt assembly at the flange joint may not function as planned or may fail under such circumstances. Temperature fluctuations tend to cause the inner and outer threads to react differently, resulting in the pair becoming looser. Flange materials have a propensity to flex, applying unneeded additional stress on the bolt head and nut. Notably, liquid propellant rocket engines display this type of temperature variation. To assess mechanical performance while accounting for different nut factors and flange materials, analytical and numerical approaches are applied.

Sensor Data Recovery of Faulty Rocket Engine Based on Multistage Graph Convolutional Network

Engine, the indispensable core of a rocket, has a significant impact on space exploration, especially the high-thrust liquid-propellant rocket engine. Most new-generation manned rockets for space stations or lunar exploration prefer the engine for its high performance and environmental friendliness. However, the engine is susceptible to failure under extreme conditions, which could cause catastrophic consequences without timely warning. Real-time state detection and fault location can prevent some catastrophic outcomes, but they require reliable sensor data. Nevertheless, some sensor data could be lost due to signal interruptions or equipment shutdown caused by system faults. Therefore, recovering the lost data based on the remaining measurements is a critical challenge that involves dealing with the distribution gap between normal and faulty data. To tackle the data drift and achieve real-time and high-precision sensor data recovery of the faulty engine, a multistage model based on graph convolutional networks is proposed in this paper. Trained by a multiloss function, the model primarily recognizes the status of the engine and passes the status to the next stage. Then the second stage recovers the lost data by two graph convolutional networks specific to the normal or faulty state. Evaluated on the practical experimental data from Xi’an Aerospace Propulsion Institute, our method successfully identifies the state of the system with accuracy above 99.99% and recovers the incomplete sensor data with a mean absolute error under 0.0065. Moreover, some ablation studies demonstrate that the blocks of two-stage and graph convolution could achieve a 26% improvement over the vanilla neural network.

A Numerical Approach to Optimize the Design of a Pintle Injector for LOX/GCH4 Liquid-Propellant Rocket Engine

This study presents an optimal design approach of a pintle injector for a deep throttlable liquid-propellant rocket engine (LPRE). Even though the pintle injector is used in rocket engines, it has become more important since reusable launch vehicles (RLVs) recently became a trend due to their economic and environmental benefits. However, since many variables must be determined to design a pintle injector, optimizing the pintle injector design is complicated. For this, we design a pintle injector to optimize the performance parameters; the spray angle, vaporization distance, and Sauter mean diameter (SMD). To confirm the approach, we design a pintle injector using an optimization method based on convex quadratic programming (CQP) for a 1000 N thrust and a throttle ability of 5 to 1 LPRE with liquid oxygen and gaseous methane. Then, we verify the performance using a numerical simulation. Through this work, we check the effectiveness of the optimization method for a pintle injector design.

Investigation of turbulent heat transfer performance of aviation turbine fuel multi-wall carbon nanotube nanofluid

Nanofluid is a novel heat transfer fluid prepared by suspending high thermal conductivity nano-sized particles in conventional fluids (water, engine oil and ethylene glycol). Thermo-physical properties (Thermal conductivity, dynamic viscosity and specific heat) and turbulent heat transfer performanceof Aviation Turbine Fuel (ATF) based Multiwall Carbon Nanotube (MWCNT) nanofluid are investigated experimentally for particle volume concentrations of 0–1% and at mean fluid temperatures of 30οC and 50οC for a potential regenerative heat transfer application in semi-cryogenic liquid propellant rocket engine. The experimental results show that the heat transfer coefficient of the nanofluid increases with particle volume concentration, with a maximum enhancement at 1% particle volume concentration of approximately 23% and 50% observed at 30οC and 50οC respectively. Two different numerical modelling approaches (a single phase fluid model with enhanced thermo-physical properties and an Eulerian-Lagrangian model called the “discrete phase model”) are employed to simulate the experimental conditions. The predictions from both numerical modelling approaches are found to compare reasonably well with the experimental data. The enhanced heat transfer performance is expressed on an equal power penalty basis to clearely show the advantage of the nanofluid.

Combustion Instabilities and its Control Techniques for Liquid Propellant Rocket Engine

It is a great challenge for us to understand and to predict the concept of liquid propellant rocket engines combustion instability due to high complex turbulent combustion process. Traditional methods have been adopted to eliminate the combustion instabilities which will increase system damping. Such methods usually suffer deficiencies with lack of study on fundamental mechanism. Difficulties have been faced during the process of various phenomena such include spray formation process, turbulent multiphase flow conditions, hydrodynamics of injection and transport characteristics of individual droplets. The ultimate aim of every engine design is to prevent occurrence of this combustion instability and to maintain reliable operation during entire span of the combustion process. This article is focused on occurrence of various type of combustion instabilities in liquid propellant rocket engine and its preventing methods in detail.

Разработка технологии механической обработки для получения турбулизаторов в камерах сгорания жидкостных ракетных двигателей

The paper considers design of the liquid-propellant rocket engine combustion chamber. Artificial roughness application area in the chamber shells is shown, and the existing methods to obtain artificial roughness in the rocket propulsion engineering are analyzed. The paper presents advantages and disadvantages of the existing technologies used in mechanical engineering for surface treatment of products being a part of the liquid-propellant rocket engine combustion chamber. Promising methods for obtaining artificial roughness on the rocket engine parts surfaces are demonstrated. Progressive means of technological equipment are considered to obtain artificial roughness by the mechanical machining on the surfaces of those parts that ensure integrity of the surface layer of the critical structural elements in the liquid-propellant rocket engines. The results obtained make it possible to significantly expand the production technological capabilities, as well as to seriously improve technical characteristics of the special equipment products in the aerospace and space industries.

Evolution in Propellant of the Rocket Engine

Evolution of rocket engine had contributed global opportunity for human civilization on the base of expanding science and technology, artificial intelligence and economic expansion which have given us potential to explore the universe as well as solve the burning issue over population, scarcity of basics and industrial resources, global warming and most prominently safety of mother earth and mankind. Hence, collecting and studying the data of various engine’s propellants and structures made till date to overcome various efficiency challenges, analyzing the current potential and challenges; and future visions with possible challenges is important, conducted as an spontaneous effort by NASA, ISRO, CNSA, Space-X and other space agencies. This theoretical research will help us to find scope, challenge and method to solve those chaos related to space engine’s efficiency in multiple time frame. Historic data will give us a mind set about our past experience of solving restriction giving us general ideas to solve future probable problems in propellant or fuel and design of rocket engine. History of primitive engine initiated since 1794 in Thomas Mead’s gas engine, Robert Street’s combustion engine in 1794, Francois Isaac de Rivaz’s electric spark followed by industrial utilization of Samuel Brown’s engine in 1823. Subsequently, all had set a base for the first liquid propellant rocket engine of Robert Goddard’s facilitated by Felix Wankel’s piston-less eccentric rotary engine design triggering today’s biggest space travel vision ultimately necessary for modern world in order to understand universe, and space exploration coupled with mass inter-stellar emigration. Hence, we shall discuss about the historic problems of space engine with their solution made till date, instantaneous situation and complication in space engine coupled with future vision and obstacles with their suitable solutions. For instance; the the factors responsible for the improvement of propellant efficiency, Isp relations with ve, Fn produced giving us mass ratio and various issues variability with the propellant used to produce an exhaust impact. By the help of comparative study of various engine’s propellant and design, we can figure out future space engine challenges and solution. Here, we shall also discuss the changes brought in structure of physically powered rocket (cold gas thruster, water rocket), chemically powered rocket(solid rocket, hybrid rocket, mono-propellant rocket, bi-propellant rocket, dual mode propulsion rocket, tri-propellant rocket, air-augmented rocket, turbo rocket, precool-ed jet engine/LACE, electrically powered (resistojet rocket, arc jet rocket, variable specific impulse magneto-plasma rocket, pulsed plasma thruster, ion propulsion system), thermal powered, solar thermal powered, beamed thermal powered, nuclear thermal powered(radioscope rocket, nuclear fission energy), nuclear propulsion powered(gas core reactor rocket, fission fragment rocket,fission sail, nuclear salt-water rocket, nuclear pulse propulsion, antimatter catalyzed nuclear pulse propulsion fusion energy, and antimatter rocket) due to properties of propellant used in combustion chamber.

Highly efficient technologies for obtaining small diameter holes in liquid propellant rocket engine parts

The paper considers design and use of parts and assemblies of a liquid propellant rocket engine — a combustion chamber, nozzle head, nozzles. Traditional and advanced technologies for making holes of small diameter in the details of these engines are analyzed. The features of the process of drilling small holes are indicated. The advantages and disadvantages of mechanical processing of component parts and assemblies of a liquid propellant rocket engine, as well as promising methods of electron beam and laser processing are given. It has been found that it is most expedient to obtain small diameter holes in such parts by combining electroerosive and electrochemical methods. The advantages and disadvantages of the combined processing method are described. It is noted that the presented technological processes, already implemented in aerospace production, increase the productivity and accuracy of holes, as well as reduce material and labor costs for the manufacture of special equipment products.

Theoretical studies of parabolic and hyperbolic equations of motion to refine the design parameters of processes in a liquid rocket engine

The problems of operation of a liquid-propellant rocket engine (LRE) on real gases with viscosity and compressibility are considered. The equations for laminar flow, which take into account the viscoelastic term characterizing the real flow, are analyzed and compared with the Euler equation for an ideal gas. The possibility of profiling supersonic nozzles of liquidpropellant rocket engines and other rocket engines is shown, taking into account the viscoelastic properties of real gas. Keywords:second order equations, laminar flow, liquid rocket engine, parameters, numerical methods, supersonic flow. mger_97@mail.ru

Перспективы применения комбинированных пористо-сетчатых материалов в конструкции внутрибаковых устройств двигательных установок космических аппаратов и разгонных блоков

The paper presents theoretical and experimental substantiation of the prospects to use materials of the new type, i.e. the combined porous-mesh materials (CPMM), in the in-tank devices (ITD) structures of spacecraft propulsion systems and upper stages to ensure multiple liquid-propellant rocket engine start in practical zero gravity. It is shown that CPMM operation was accompanied by phenomena that were missing, when using mesh materials previously used for this purpose, the so-called mesh separators (MS). It is concluded that at present it should be considered expedient to widely introduce CPMM in the ITD structure elements, since in this capacity the CPMM properties are noticeably superior to those of conventional MS. As a result, the ITD efficiency based on CPMM increases significantly, which finally makes it possible to minimize unusable fuel residues in the tanks and, thereby, increase the energy-mass efficiency of the aerial vehicle as a whole.

Petri Nets Applied in Purge Algorithm Analysis for a Rocket Engine Test with Liquid Propellant

During the development stage of a space vehicle, instrumented tests are carried out on the ground to prove the operational capacity of each liquid-propellant rocket engine, which is installed in this type of vehicle. The task of elaborating a Test Bench project for a propulsion unit with this application is complex and involves several steps, one of these steps being related to the analysis of this bench capacity to meet the algorithms for the liquid-propellant rocket-engine full run of tests, which is considered fundamental for this project’s operational success. Due to the high costs involved in this project’s elaboration and execution, it is strategic to use computational resources to evaluate, by simulation, the main operational functionalities that are previously established for this bench to perform. In this context, this work presents a model proposal through Petri Nets to evaluate, by computer simulation, an architecture capacity that was designed for the Test Bench to meet an algorithm dedicated to the liquid-propellant pipelines purge during the run of hot tests with the liquid-propellant rocket engine. The method used in this work to carry out the simulation shows the operational response of each module of this architecture, in accordance with the steps contained in the purge algorithm, which allows for analyzing, for each event of the process, the Petri Nets properties, mainly those related to the conservativeness, liveliness, deadlock-type, and confusion-type conflicts. The simulation carried out with the proposed model allows for the portrayal of the physical architecture and the operational states of the purge system according to the steps foreseen in the algorithm, showing that the conservation property is met because the number of marks remains constant, the vivacity property is also met since all positions have been reached, and there is no mortal-type conflict, as the simulation is not stopped; only confusion-type conflict is identified, which was solved with the strategic insertion of resources in the model in order to fix crashes related to the competition for tokens in the transition-enabled entries. The satisfactory results obtained in these simulations suggest that the modules provided for this architecture are sufficient and appropriate for carrying out all the steps contained in the purge algorithm, which will minimize or even eliminate the disorders that may be caused by the presence of foreign elements in the propellant supply lines during the tests with the rocket engine.

ВИКОРИСТАННЯ ПІДШИПНИКІВ КОВЗАННЯ У АВІАЦІЙНІЙ ТА КОСМІЧНОЇ ТЕХНІЦІ

Examples and prospects of various types of sliding bearings applications in power elements of aircraft engines and in turbopump units (TPU) of liquid propellant rocket engines (LPRE) are discussed in this article.In the second half of the last century, piston aircraft engines were replaced by turboprop (TPE) and turbojet (TJE) engines. Almost all leading manufacturers of these motors use rolling bearings as rotor supports for the main power units. The use of sliding bearings is extremely rare and is the exception rather than the rule. These exceptions include the company Pratt & Whitney, which has been using sliding bearings since the 1950s (engine PT6A). The sliding bearings are also used in modern turbofan engines Pratt & Whitney PW1000G series. The use of sliding bearings as a support for planetary gearbox elements has a number of advantages over rolling bearings, since under normal operating conditions such supports have a significantly longer service life.In rocketry, sliding bearings are used as supports for a LPRE TPU. Examples of such supports are hydrodynamic sliding bearings of the first German missiles during the Second World War (one of them was the FAU-2) and Soviet missiles developed by S.P. Korolev and V.P. Glushko. As our analysis has shown, in the 50s the use of sliding bearings in LPRE was abandoned due to a number of advantages of rolling bearings over sliding bearings, especially taking into account the very short service life of these engines (hundreds of seconds). Recently, however, sliding bearings as supports for a TPU have been attracting more attention again due to the development of reusable missile systems with a sharp increase in required service life. Although the results achieved for TPU with rolling bearings still provide the required durability, studies of the sliding bearings for TPU design with increased service life are relevant.The data obtained as the result of the analysis describe existing practice of using sliding bearings in units of aircraft and rocket engines, as well as the prospects for their further use.

HYDRO-GAS-DYNAMIC MODEL OF MOTION OF GAS-SATURATED LIQUID IN A CENTRIFUGAL POROUS PUMP


 
 
 An approach has been developed to the formation of a theoretical model of the movement of a gas- saturated liquid in the flow path of a porous centrifugal pump, as an assembly of the system for feeding the reaction mass of liquid-propellant rocket engines. The developed approach is based on a physical model of the turbulent motion of a gas-liquid mixture, within which conservation equations are used. Two models are proposed that describe the movement of a gas-liquid mixture in the flow path of a centrifugal porous pump in a two-phase one- dimensional formulation and a quasi-homogeneous one, complementing each other. A porous impeller of a closed- type centrifugal pump of constant width is considered. The energy losses of the gas-liquid mixture flow at the inlet to the pump impeller, as well as the losses of disk friction associated with rotation in a fixed housing, are not taken into account. The parameters of the viscous and inertial hydraulic resistance coefficients of the porous body are used. In particular, the influence of the static pressure gradient in a gas-saturated liquid and the friction force on the accumulation of gas at the inlet to the impeller and the stability of the liquid supply by the pumping unit is noted. The use of the developed approach makes it possible, by calculation methods, to evaluate the values of parameters characterizing the movement of gas-liquid media in centrifugal porous pumps of the designed systems for feeding propellant components to liquid-propellant rocket engines, to reduce the amount of experimental testing of new and modernized propulsion systems for space stages and launch vehicles.
 
 

Investigation on heat transfer and flow of a sonic steam jet condensation in crossflow of water in pipes

Direct contact condensation is widely used in nuclear reactor thermal-hydraulic systems, underwater propulsion systems, liquid propellant rocket engines and other industrial fields due to its highly efficient heat and mass transfer. Understanding the distribution of the thermal-hydraulic parameters of steam jet condensation systems has become an urgent need for optimal design and safe operation of industrial processes. In this paper, the flow characteristics and heat transfer for the condensation of sonic steam jet in crossflow in a subcooled water pipe are numerically simulated based on experimental results. A two-resistance model under the framework of two-fluid model is employed to consider the mass and heat transfer across the steam-water interface. The steam plume transforms from contraction-cone into expansion-cone when steam mass flux rises from 500 to 900 kg/m2s, and the plume length gradually increases approximately linearly from 2.43de to 4.17de. When the steam mass flux is 900 kg/m2s, the expanding channel of the steam plume allows the jet to accelerate from sonic to supersonic with Ma = 1.43. At the transition of expansion to contraction, the steam dynamic pressure reaches a maximum and the steam total pressure decreases to a minimum. Because the water dynamic pressure is significantly higher than the steam dynamic pressure, the total pressure is maximum at the end of the two-phase region. Mass transfer is much higher in two-phase region than in other regions. The heat transfer coefficient calculated from the interface area decreases from 1.98 to 1.56 MW/m2K, and that calculated from the two-phase region volume increases from 3.37 × 103 to 4.05 × 103 MW/m2K.

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.

Расчет параметров газа наддува при заполнении тупиковых зон в магистралях пневмогидравлических систем жидкостных ракетных двигательных установок

The paper proposes a mathematical model to estimate boost gas parameters at filling the feeding lines dead-end zones of the pneumohydraulic systems (PHS) of the liquid-propellant rocket engines (LRE). These calculations are required in preliminary assessment of the design characteristics of the PHS feed lines designed to ensure operation of the LRE pneumoautomatic systems. Engineering methodology currently existing to solve this problem provides for solution of a system of ordinary differential equations and does not allow tracking alteration in pressure and temperature along the length of the dead-end zones at filling the PHS feed lines. The proposed model takes into account inhomogeneity of the flow parameters along the length of the dead-end zones, as well as the nonadiabaticity of the flow, i.e. it assumes presence of the heat exchange between the walls of the feeding line and the boost gas. The presented model was used to construct dependences of gas parameters alterations in the feeding line. Results obtained were compared with calculations according to the evaluation method, which is traditionally used in the engineering practice to solve this class of problems.