Rocket Propulsion Systems Research Articles

On the Issue of Calculating the Reverse Arc Breakdown Voltage of a High-Voltage Thermionic Diode and Its Maximum Specific Electrical Power

The issues of creating powerful power propulsion systems (PPS) based on a thermionic reactor-converter (TRC) and electric rocket propulsion systems (ERPS) are related to the development of compact high-temperature radiation–resistant current conversion systems (CCS) using plasma electric power devices - grid key elements (GKE) and high-voltage plasma thermionic diodes (HVPTD). The main problem of the development of GCE and HVPTD is the determination of their operating voltage (or reverse arc breakdown voltage in the locked state), which together with the current density in the conductive state characterizes the electric power density of the valve and is its main parameter. Within the framework of this work, the authors, based on the study of the locked state, numerically and analytically obtained new dependences for calculating the reverse arc breakdown voltage and determined the limiting values of the electrical power density of the HVPTD.

RESULTS OF DESIGN AND THEORETICAL STUDIES OF LV PRESSURIZATION SPRAYERS PARAMETERS

The pressurization system of the fuel tanks of launch vehicles is one of the important and critical subsystems of the pneumatic-hydro supply system. The pressurization system is designed to create the necessary pressures in the fuel tanks in order to ensure their required values at the engine inlets (both during launch and during operation in specified modes), as well as to ensure the strength and stability of the tanks during the flight of launch vehicles. These systems determine the configuration and appearance of launch vehicles, their energy characteristics and operating conditions. At the same time, the pressurization system is functionally interconnected with the operation of many other subsystems of the pneumatic-hydraulic supply system (refueling of fuel components, control systems, fuel consumption control systems, units of a liquid-propellant rocket propulsion system, etc.).
 One of the main functional elements of the pressurization system is the atomizer. The atomizer is designed to introduce pressurization gas into the tank in such a way that its speed and direction provide the required pressure and gas temperature distribution in the free volume of the tank, without causing unacceptable heating of the tank structure and fuel component, as well as possible disturbances (vibrations, etc.) .
 On the example of specific launch vehicles on high- and low-boiling fuel components, a generalizing review of the pressurization system sprayers developed in the Yuzhnoye State Design Office (YSDO) was carried out, as a result of which the authors determined and described in detail: their design features, requirements, classification, and the procedure for their development. The possibilities and tools for modeling the parameters of atomizers are shown. The experimental data of some of the developed atomizers are presented. Promising directions for the development of pressurization systems are outlined, which in the future will allow the development of highly efficient systems with optimal parameters.

High entropy alloy as metal additive for hybrid rocket propellant

A typical solid propellant used in rocket propulsion systems is made from polymeric fuel, crystalline liquid binder and additives. In comparison to conventional propulsion systems, hybrid propulsion is more safe, reliable and environmentally friendly. However, the hybrid systems have some drawbacks, including low fuel regression rates and low combustion efficiency. As a result, several approaches have been developed to increase the rate of solid fuel regression rate including the introduction of metal additives to the system. Nowadays, high entropy alloys (HEA) are among the most advanced materials for high-performance applications. This alloy has a potential to replace the existing metal additives in hybrid propulsion systems. The present work proposes a new composition based on Paraffin wax fuel, loaded with high entropy alloy as a metal additive for hybrid propulsion systems using the following ratio of 80:20. Therefore, this research explores the efficacy of high entropy alloys as additives for hybrid propulsion systems. Characterizations and thermal analysis of solid fuel using X-ray diffraction (XRD), differential scanning calorimetry (DSC) and thermogravimetric analysis are carried out to determine the structure, thermal decomposition and energy characteristics of the solid fuel with the HEA as additive. XRD result shows that HEA in a form of solid solutions produces BCC and FCC structures. Thermal analysis of paraffin wax with the addition of HEA gives the maximum value of 136.34 J/g for the total latent heat of the mixture. It is concluded that HEA increases the energy released from the fuel by 86.92 % compared to pure paraffin wax fuel. The mixture of paraffin and HEA fuel has a prospect to be used as future rocket fuel.

The Mechanism Research of Low-Frequency Pressure Oscillation in the Feeding Pipe of Cryogenic Rocket Propulsion System

In the propulsion system of cryogenic liquid rockets, low-frequency pressure oscillation is a long-standing issue occurring in its feeding pipe, and is not conducive to the normal operation of the rocket. Its mechanism and excitation process are not very clear due to the limitation of the existing numerical method and the difficulty of the real dynamic experiment. Inspired by the periodic suck-back flow phenomenon of steam condensation, the fluctuation of the two-phase interface might be the crucial factor to initiate the low-frequency pressure oscillation. To simulate this interfacial characteristic of cryogenic propellant, a novel numerical model is proposed to predict the mass transfer rate weighted by the interfacial curvature. Aiming at the oxygen jet condensation simulation, the low-frequency pressure oscillation phenomenon is obtained successfully with the excitation frequency of 10.6 Hz, consistent with the natural frequency of the engine test run. It is conducted so the low-frequency pressure oscillation is caused by the periodic condensation of the continuous oxygen vapour plume, along with an oxygen suck-back flow phenomenon. In addition, the results indicate that both the oxygen and liquid oxygen mass flux promote the rise in the frequency of pressure oscillation. These conclusions provide theoretical instructions for the design and operation of the propulsion system of a cryogenic liquid rocket.

3D Printing Techniques for Paraffin-Based Fuel Grains

Hybrid rocket propulsion systems have proved to be a suitable option for some specific applications in the space transportation domain such as in launch vehicle upper stages, orbit transfer spacecrafts, decelerator engines for re-entry capsules, and small satellites launchers. Part of the renewed interest in hybrid rocket propulsion is due mainly to the safety aspects, cost reduction, and the use of paraffin-based fuel that impacts positively in terms of the solid fuel regression rate. However, paraffin solid fuel grains have poor structural characteristics and sometimes low performance due to the fuel internal ballistics behaviour. More recently, various studies have been carried out to overcome these drawbacks of paraffin-based fuels, such as the addition of energetic nano-sized metallic powder and 3D printing techniques. This study presents a review of the principal concepts of 3D printing processes and extrusion techniques that can be suitable for paraffin grains manufacturing and the conceptual design of a prototype for a 3D printer system under development at the Aero-Thermo-Mechanics Department of Université Libre de Bruxelles.

Reactivity of hypergolic hybrid solid fuel with industrial grade hydrogen peroxide

Hypergolic solid fuels and hydrogen peroxide oxidizers have attracted considerable attention for the development of safe hybrid rocket propulsion systems. In this study, a new approach was explored to develop a safe hypergolic hybrid propulsion system using industrial-grade hydrogen peroxide (∼70 wt%). Initially, the propulsion performance of different hypergolic solid fuels with an industrial-grade hydrogen peroxide was calculated using the NASA CEA code to find an appropriate propellant combination. The theoretical specific impulse of the hypergolic solid fuel was 27.8 % over hydrazine and 13.4 % over LMP-103S. Based on theoretical investigations, hypergolic solid fuels, named such as HSF-1 to HSF-10, were fabricated using molding and pressing methods. The ignition delay time was measured using the drop-test method. Interestingly, HSF-3 and HSF-4 exhibited very short ignition delay times of 4.92 and 8.75 ms, respectively, with industrial-grade hydrogen peroxide. In addition, the drop test results of hypergolic solid fuel with varying compression pressure in the pressing method confirmed that a pressure above the compressive strength of the binder shortens the ignition delay times deviation. Further, HSF was also ignited even at low concentration of IGHP (>40 wt%); opens the new area of research in combustion science. Overall, low-toxicity hypergolic solid fuels and industrial-grade hydrogen peroxide are promising alternatives to conventional toxic hypergolic propellants.

2.5 kN pump-pressure pintle engine ignition experiment in different loading cases

Advanced technology of motors and batteries has stimulated the application of the emerging electric pump in the liquid propulsion systems. However, the coupling characteristics between the electric pump and the pintle injector with high momentum ratio are not clear. In this study, suitable approaches were applied to conduct the ignition experiment that the hydrogen peroxide was catalyzed in full flow with kerosene in different cases. In the experiment, the electric pumps were used to pressurize the hydrogen peroxide and kerosene respectively. The experiment results showed that, the catalytic process would be realized between 1.0 s and 2.0 s in different cases while the ignition process could be reached between 0.1 s and 0.2 s, which was mainly caused by the overshoot between 15% and 30% in the catalytic process. When the injection valve remained closed after the expected speed was reached, the outlet pressure of electric pump would be subject to vibration. However, the low-frequency vibration predominated from 12 Hz to 31 Hz. After the injection valve opened, the pressure vibration of the kerosene electric pump was between 0.15 MPa and 0.8 MPa. As the load dropped, the pressure vibration would obviously escalate. Each time when the speed increased 3000 r/min, the propellant temperature at the outlet of the hydrogen peroxide pump would increase about 1°C and that of the kerosene pump 5°C. The experiment can promote the application of the electric pump with hydrogen peroxide and kerosene in rocket propulsion systems with pintle injectors to realize a large regulating ratio.

A Review on Vibro-Acoustic Analysis of a Launch Vehicle Structure

Abstract: Space vehicles are subjected to significant dynamic pressure loads when their rocket propulsion systems are in use during flying missions. During the aerodynamic and launch phases, launch vehicles, payloads, and their parts are subjected to extremely high random acoustic loads. The noise from the engine exhaust gas, aerodynamic boundary layer noise, transonic buffering, structure-borne vibration, engine thrust fluctuation, etc. is the source of these loads, which also result in a secondary acoustic load. When the vehicle is lifting off and traveling at a speed greater than Mach number, acoustic stresses to the spacecraft and payload are very harsh and significant. This loading causes the structure to vibrate randomly, which could be dangerous for some vehicle parts, avionics, propulsion systems, and payloads like satellites. This paper discusses the vibration of the structure subjected to acoustic excitation on a diffuse acoustic field and the software used for the analysis.

Modelling of Hybrid Rocket Flow-Fields with Computational Fluid Dynamics

A low regression rate is a major limitation in the hybrid rocket propulsion system. This paper is to study the regression rate characteristics of a cylindrical solid grain with cryogenic propellants. Paraffin (P) fuel is coupled with two types of oxidizers, namely gaseous oxygen (GOX) and nitrous oxide (N2O). ANSYS software is used as the CFD platform to observe the hybrid rocket flow-fields. The modelling values obtained from the pressure, temperature, velocity, and wall heat flux contours are used to calculate the hybrid rocket performance in terms of regression rate, thrust, specific impulse, characteristic velocity, and exit Mach number. The numerical results show that the mass-flow-inlet boundary conditions, initial design feature, and type of propellant play an important role in the enhancement of hybrid rocket performance. Result shows that enhanced of 68 % of the regression rate, 59 % of thrust, 6 % of specific impulse and exit Mach number, and % of characteristic velocity by improved 50% of mass flow rate. Due to the high flame temperature, the GOX/P propellants produce the best hybrid rocket motor compared with the others.

A method for verifying the model of tank operational pressurization process during development of propellant supply system in liquid rocket engines of launch vehicle propulsion systems

During development of propulsion systems for launch vehicles, in order to verify the design results, standalone developmental tests of propellant supply system components were run in specially designed test setups outfitted with the required instrumentation. As an example of such work, this paper describes the experience of standalone developmental testing of operational pressurization of propellant tanks. During tests in the constructed test setups, heat and mass exchange processes were tried out with simulation of design modes of operational pressurization of propellant tanks filled with oxygen and hydrogen. As a result of these tests, a rationale was provided for selection of working mediums as the pressurization gas, design and characteristics of gas injection into propellant tanks, a method for evaluating the efficiency of the pressurization process. Empirical dependences were obtained that were used as the boundary conditions in the models for practical calculations for refining the modes and evaluating the efficiency of the propellant supply subsystem of propulsion systems. The paper calls attention to the need to use for the purpose of development of the propellant supply systems that involves computer simulations the results of such earlier work when designing advanced rocket propulsion systems. Key words: propulsion system, propellant supply system, development of pressurization process, test setup, pressurization gas, propellant components, heat and mass transfer intensity, process efficiency, liquid surface sloshing, computer simulations.

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.

Comparison between the dynamic characteristics of electric pump fed engine and expander cycle engine

A successful space launch mainly depends on the start-up process and variable operating conditions of rocket propulsion system. With the development of motor and battery technology, new electric pump is gradually used, in place of the traditional turbine pump, in rocket propulsion system. However, it is still unclear about the dynamic characteristics of the two processes in the two schemes (electric pump scheme and turbine pump scheme). In this research, simulation models of the basic components and overall system of an expander cycle hydrogen-oxygen engine were established on the basis of turbine pump and electric pump respectively. In the simulation models, the Centralized Parameter Method and the Stepwise Parameter Method are used for different components. The research results show that the power of hydrogen pump in the electric pump scheme is 25.5% lower than that in the turbine pump scheme, with speed and torque respectively 10.5% and 16.6% lower than those in the latter scheme. In the start-up process of the electric pump scheme, the flow of hydrogen pump underwent an overshoot of 3.7%, and the settling times of hydrogen and oxygen flows were both more than 40% shorter than those in the turbine pump scheme. Under variable operating conditions of turbine pump scheme, the speed of pump experienced an overshoot of about 20%, and the settling times of hydrogen and oxygen flows were 37% and 21% longer than those in the electric pump scheme. These findings are helpful to analyze the said two processes of expander cycle hydrogen-oxygen engine and further understand the difference in dynamic characteristics between the electric pump scheme and the turbine pump scheme.

Observation of aluminum interaction with the binder melt layer using high-speed synchrotron-based phase contrast imaging

Particle interactions with the binder melt layer are a major factor in the combustion efficiency and stability of the propellant in a solid rocket motor. Metal particles tend to agglomerate on the surface of burning solid propellants, inhibiting the combustion process. Therefore, reduction of molten aluminum agglomeration in a solid rocket motor is vital to the improvement of solid rocket propulsion system performance. In other work, quenched samples have been used to study the impact of how metal particles alter the flow properties of the molten propellant surface. In-situ optical measurements have also been attempted for these particle-condensed phase interactions, but with little success due to the opacity and strong gradients within the flame. This study expanded on previous work using synchrotron–based x-rays to directly image these aluminum agglomerates as they interacted with the binder. X-ray absorption and phase contrast in the images allowed for the direct measurement of the particle size in combusting propellants in-situ at typical rocket pressures. Particle size distributions were collected in an optically accessible combustion vessel over a pressure range of 1.4- to 6.9-MPa (200- to 1000-psig). This investigation measured the aggregate sizes ranging from 300-to 650-µm at a 6.9 MPa chamber pressure. Interestingly, the lowest binder melting temperature produced the longest aluminum chaining at 1.1 mm and the highest binder melting temperature produced the second longest chain at 524 µm long. The study observed aluminum interactions with the binder melt layer on the propellant surface that may be a contributing factor in a plateau propellant. Additionally, data pointed to possible relationships between binder viscosity, burning rate, aluminum clusters, and chain formation in a solid propellant. The study concluded with collecting data on the molten aluminum volumetric changes in a propellant environment, thus defining a starting particle diameter for particle regression rates. Collecting the regression rates also provided information on the evaporation rate of aluminum, which burned in the propellant flame. All of these data are critical for understanding the combustion of aluminum-based solid rocket propellants and outlines knowledge gaps that need additional data.

Hydrogen peroxide – A promising oxidizer for rocket propulsion and its application in solid rocket propellants

This paper discusses the potential use of hydrogen peroxide as oxidizer for solid rocket propulsion. While hydrogen peroxide is a liquid in normal conditions, it may be used in solid rocket motor grains. Use of hydrogen peroxide of HTP class (High Test Peroxide) is proposed. It has been known for many decades but its utilization has been historically limited to liquid propellants. Until recently is application even in liquid state in rocket propulsion for space and defence has been limited due to storability and safety issues. With the availability of new grades of HTP with higher purity and concentrations, enhanced performance, safety and storability are possible. This results in HTP being considered for a wide range of rocket propulsion systems: as oxidizer in bipropellant and hybrid propulsion systems, as well a monopropellant. To ensure proper analysis of the potential of using HTP in next-generation solid rocket propellants, this paper reviews existing rocket propulsion applications of HTP. Modern use requires high performance and common current composite propellant compositions utilize ammonium perchlorate (AP) as oxidizer, what has several disadvantages, which are discussed. Alternative oxidizing compounds include ammonium dinitiramide (ADN), ammonium nitrate (AN), hydrazinium nitroformate (HNF), hexanitrohexaazaisowurtzitane (HNIW) and several other secondary explosives. Key properties of solid propellant oxidizers are listed and discussed. The need for further alternatives, despite numerous recent advances in solid rocket oxidizer technology is justified. The global push forward high performance green propulsion is one of the main motivations behind considering HTP for solid rocket propulsion. Theoretical performance of solid rocket motors using solid grains containing a high mass fraction of HTP is presented. This includes performance considering several fuels and additives and different oxidizer loadings. Up to date concepts of using hydrogen peroxide in solid propellants are reviewed. Solid cryogenic propellants using hydrogen peroxide are mentioned, but focus is given to solid propellants which could ensure flexible operations, thus use in state-of-the-art solid rocket motors. Challenges of HTP application as an oxidizer for solid propulsion are listed. This includes a discussion concerning its reactivity, thus limited compatibility with organic materials. Recommendations for further experimental work are proposed. Potential technology applications are listed with explanations why these particular niches may benefit from utilization of the new solid propellant technology.

Research on the thermal shock simulation of the super high speed aircraft

Ultra high temperature ceramics (UHTC) with good mechanical properties and thermal physical properties, including high melting point (>30000C), high conductivity, high thermal conductivity, high thermal shock resistance rate, provides a sharp leading edge and smooth surface of a new generation of aerospace vehicle design to the aerospace industry; Due to ultra-high temperature ceramic materials with high strength, oxidation resistance and thermal shock resistance, it can be applied to extreme environments such as hypersonic flight, atmospheric reentry and rocket propulsion system, and have been widely used in key components of aircraft front cone, engine, wing and engine hot end. Based on the application of ultra-high temperature superconducting materials in practical application, the solid modeling of ultra-high temperature superconducting front-end components of ultra-high speed aircraft wing is carried out by using ABAQUS software, the ultra-high temperature thermal field of loaded aircraft flying at ultra-high speed is simulated, the stress-strain field at the flange is obtained, and the thermal mechanical coupling analysis is carried out.

Optical Diagnostics for Solid Rocket Plumes Characterization: A Review

In recent decades, solid fuel combustion propulsion of spacecraft has become one of the most popular choices for rocket propulsion systems. The reasons for this success are a wide range of applications, lower production costs, simplicity, and safety. The rocket’s plumes leave the nozzle at high temperatures; hence, the knowledge of produced infrared (IR) emissions is a crucial aspect during the design and tests of the rocket motors. Furthermore, rocket plume composition is given by N2, H2, H2O, CO and CO2, while solid rocket motors (SRM) additionally inject some solid particles, given by metal fuel additives in the propellant grain, i.e., aluminum oxide (Al2O3) particles. The main issue is the detection of the particles remaining in the atmosphere due to the exhaust gas of the solid rocket propulsion system that could have effects on ozone depletion. The experimental characterization of SRM plumes in the presence of alumina particles can be conducted using different optical techniques. The present study aims to review the most promising ones with a description of the optics system and their potential applications for SRM plume measurements. The most common measurement techniques are infrared spectroscopy imaging, IR imaging. UV–VIS measurements, shadowgraph, and Schlieren optical methods. The choice of these techniques among many others is due to the ability to study the plume without influencing the physical conditions existing in and around the study object. This paper presents technical results concerning the study of rocket engines plumes with the above-mentioned methods and reveals the feasibility of the measurement techniques applied.

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

Use of the Low-Potential Heat for Heating Helium in Rocket-Carrier Tank Pressurisation Systems

The energy efficiency of new technical developments is a critical issue. It should be noted that today the focus in this issue has seen a major shift to the maximum use of renewable energy sources. The purpose of this research is to reduce the weight of helium heat exchangers of the fuel tank pressurisation systems in modern rocket propulsion systems that use fuel components like liquid oxygen and kerosene-type fuel. This is the first time that the question has been raised about the possibility and advisability of increasing the temperature of helium at the heat exchanger inlet without the use of additional resources. The paper addresses the use of the waste (“low-potential”) heat and ”industrial wastes” present in propulsion systems. Basic laws of complex heat exchange and the retrospective review of applicable heat exchanger structures are applied as a research methodology. Two sources of low-potential heat are identified that have been previously used in the rocket engine building in an inconsistent and piecemeal manner to obtain and heat the pressurisation working fluid. These are the rammedair pressurisation during the motion of the rocket carrier in the atmosphere, and the tank pressurisation as a result of boiling of the top layer of oxidiser which is on the saturation line. This is the first time that the advisability has been substantiated of increasing the temperature of the working fluid at the heat exchanger inlet, first of all due to the use of the low-potential heat. This is also the first time that unemployed sources of low-potential heat and “industrial wastes” are found in modern deep throttling propulsion systems. These are the high-boiling-point fuel in the tank, behind the highpressure pump, at the exit of the combustion chamber cooling duct, and also the fuel tank structures, and the engine plume. A possibility is proved, and an advisability demonstrated of their implementation to increase the efficiency of pressurisation system heat exchangers. This is the first time that the methodology of combustion chamber cooling analysis has been proposed to be adopted for the heating of heat exchanger by the engine plume. This is the first time that a classification of waste heat sources has been developed which can be used to increase the pressurisation working fluid temperature. The identified reserves help to increase the efficiency of the helium heat exchangers of the tank pressurisation systems in the propulsion systems

МОЖЛИВОСТІ ЗАСТОСУВАННЯ СОНЯЧНОГО ВІТРИЛЬНОГО ПРИСТРОЮ ДЛЯ ВІДВЕДЕННЯ КОСМІЧНИХ ОБ’ЄКТІВ ІЗ НАПІВСИНХРОННИХ ОРБІТ

The current state of the near-Earth space environment poses a threat to spacecraft moving not only in low and geostationary but also in semi-synchronous orbits. The main methods that can be used in these orbits for space debris removal are the active method using a rocket propulsion system and the passive method using a solar sail device. The last method is of particular interest because of the low overall mass and mass characteristics of the removal system, as well as the availability of unlimited solar light radiation, which can be converted into a driving force. Solar sails are currently undergoing experimental testing in the near-Earth space, but they have not been used for medium and high orbits. Two constructive variants of solar sails construction are considered: frame and rotor. The article is devoted to the evaluation of the possibility of removing space objects from semi-synchronous orbits using a solar sail. Two variants of the removal scheme are considered: into a circular disposal orbit 200 km below the target orbit and into the dense layers of the Earth's atmosphere. A computer simulation of the removal is performed using the example of the Galileo global positioning system spacecraft with a frame-type and rotor-type solar sail. The dependence of the orbit altitude at the ending of removal on the solar sail area was obtained. The values of the solar sail area required to provide for the maneuvering and the weight of the structure were determined. It was demonstrated that the total mass of the frame-type solar sail is 9 kg (1.3% of the spacecraft mass) and that of the rotary-type solar sail is 11 kg (1.6%). As applied to disposal into dense layers of Earth's atmosphere for frame-type solar sail - 52 kg (7.4%), for rotary-type solar sail - 25.5 kg (3.6%). It follows from the obtained results that for moving to the disposal orbit 200 km below the target orbit it is more advantageous to use a frame-type solar sail, and in the dense layers of the atmosphere - the rotor type.

Hydraulic similitude assessment for cryogenic cavitation in propellant lines: The case of thick orifice

Cryogenic cavitation can be easily encountered in rocket liquid propulsion systems at the very initial phase of the launch due to the propellant passing through orifices and valves. This leads to flow instability, performance degradation, and even damages to the structures. Hence, the design of the propulsion systems cannot overlook cavitation phenomena. Despite the relevance of the topic, their accurate prediction is still hampered by the difficulty in modeling the high non-equilibrium phenomena involved. From an engineering point of view, the main interest is to design such restrictions present in propellant systems and, hence, to predict the maximum flow rate they can reach. This work questions the possibility to design orifices for cryogenic applications based on purely isothermal testing. Therefore, we performed cavitation experiments with both water and liquid nitrogen by considering two orifices hydraulically similar. This implies a geometric similitude in terms of the orifice characteristic ratio (β) and its dimensionless thickness (th). The results show that the hydraulic similitude is not sufficient to predict the orifice behavior for a cryogenic flow correctly. In particular, the level of liquid subcooling (ΔTsub) upstream of the orifice plays an important role since it impacts the minimal value of pressure of the flow downstream the orifice and thus its thermodynamic state.