Liquid Propellant Rocket Engines Research Articles

Model-Based Sequential Design of Experiments with Machine Learning for Aerospace Systems

Traditional experimental design methods often face challenges in handling complex aerospace systems due to the high dimensionality and nonlinear behavior of such systems, resulting in nonoptimal experimental designs. To address these challenges, machine learning techniques can be used to further increase the application areas of modern Bayesian Optimal Experimental Design (BOED) approaches, enhancing their efficiency and accuracy. The proposed method leverages neural networks as surrogate models to approximate the underlying physical processes, thereby reducing computational costs and allowing for full differentiability. Additionally, the use of reinforcement learning enables the optimization of sequential designs and essential real-time capability. Our framework is validated by optimizing experimental designs that are used for the efficient characterization of turbopumps for liquid propellant rocket engines. The reinforcement learning approach yields superior results in terms of the expected information gain related to a sequence of 15 experiments, exhibiting mean performance increases of 9.07% compared to random designs and 6.47% compared to state-of-the-art approaches. Therefore, the results demonstrate significant improvements in experimental efficiency and accuracy compared to conventional methods. This work provides a robust framework for the application of advanced BOED methods in aerospace testing, with implications for broader engineering applications.

Eкспериментально-розрахункове визначення коефіцієнтів гідродинамічної моделі кавітуючих насосів рідинних ракетних двигунів

Cavitation phenomena in liquid-propellant rocket engine (LPRE) pumps not only affect the power performance characteristics of the pumps, but they also affect the pump dynamics and pogo vibrations. The theoretical characterization of cavitation phenomena in LPRE pumps is not a widely used practice because theoretical and experimental data are in unsatisfactory agreement. Because of this, use is made of approaches that employ experimental data. The goal of this work is to determine the coefficients of a hydrodynamic model of cavitating LPRE pumps throughout the cavity existence region based on the experimental frequencies of cavitation oscillations and cavitation self-oscillation boundaries. In determining the cavity elasticity and negative resistance, use was made of the experimental cavitation oscillation frequencies of 26 LPRE pumps differing in dimensions and capacity. In determining the cavitation resistance distribution coefficient and the cavity-due disturbance transfer time, the experimental cavitation self-oscillation boundaries of 14 more pumps were used. To extend the cavity elasticity determination region, the extrapolation dependence of the cavity elasticity in cavitation stall regimes was updated. To make the stratification of the cavity resistance dependence more uniform in the range of large discharge coefficients, incipient cavitation numbers were refined. Using he qualitative dependence of the cavitation resistance distribution coefficient obtained from theoretical transfer matrices of cavitating pumps and its lower estimate (at zero disturbance transfer time) and upper estimate (for a uniform stratification of pump transfer matrix determinants), its analytical dependence was found. Using it and the coefficients of a mathematical model of cavitation oscillations on the cavitation-self oscillation boundary, disturbance transfer times were found and approximated.

Additively-manufactured shear tri-coaxial rocket injector mixing and combustion characteristics

A monolithic tri-coaxial propellant injection scheme for enhanced mixing of methane-oxygen in liquid-propellant rocket systems is enabled by additive manufacturing. Mixing and combustion characteristics of the tri-coaxial design are assessed experimentally from 1–69 bar using laser absorption tomography and chemiluminescence imaging, and are compared to a traditionally-manufactured bi-coaxial design. Quantitative two-dimensional images of temperature and carbon monoxide mole fraction are generated from the laser absorption spectroscopy methods, while OH* chemiluminescence provides an approximate metric for combustion heat release defining flame length and injector standoff distance. At similar pressures and oxidizer-to-fuel ratios, the tri-coaxial injector design is shown to enhance mixing and combustion progress, reducing characteristic mixing length scales and achieving improved combustion performance relative to more conventional bi-coaxial designs. Despite enhanced mixing, the tri-coaxial design exhibits more limited reduction in flame standoff distance from the injector face, suggesting that increased heat flux to the injector face can be managed. The tri-coaxial injector highlights the potential to leverage additive manufacturing to enhance performance and simplify the fabrication of liquid-propellant rocket engines.

Heat transfer characteristics of a turbulent swirl jet issuing from a circular nozzle

Experimental and numerical studies on the flow field and cooling performance of a swirling jet, impinging on a flat surface is presented. A new nozzle that resembles the swirl injector of a liquid propellant rocket engine is used. This study considers different parameters including swirl number(S), Reynolds number(Re), normalised orifice to target spacing (Z/D) and confinement. The performance of various RANS and LES turbulence models is assessed using the data obtained from present experimental studies and that reported in literature. RNG k-ϵ turbulent model is successful in predicting heat transfer in non swirling flow. In the case of swirling flow, LES WALEM(Wall Adapting Local Eddy viscosity Model) predicts the heat transfer better than the other models. The performance of Swirling Impinging Jets(SIJ) and Cylindrical Impinging Jets (CIJ) is compared. At higher Z/D, introduction of swirl results in reduction in heat transfer and this is because of the increased spreading of the jet and reduction in velocity of the jet impinging on the target. Better performance is achieved at lower Z/D when a cylindrical jet is introduced with a small amount of swirl. For Z/D = 2 and S = 0.2, the peak Nu increases by 10 % and average Nu in the stagnation region increases by 6 % in comparison to CIJ. The position of the peak Nu moves away from the axis of the jet and there is better uniformity in Nu in the stagnation region. For Z/D = 2 swirling flow creates secondary peak(which usually occurs at high Re)even at low Re. At very low Z/D, top wall confinement causes increase in both peak heat transfer and average heat transfer in the stagnation region.

Mass minimization approach for the optimal preliminary design of CMC inner liners in rocket thrust chambers

In the past decade, the world has witnessed a new space race, driven by a growing commitment to reducing the environmental impact of space missions. This has led to the widespread adoption of liquid-propellant rocket engines, which offer several advantages over their solid-propellant counterparts. One key advantage is their reusability, which not only helps to reduce the generation of space debris but also makes space exploration cheaper. To further enhance the performance of liquid rocket engines, researchers have been exploring innovative cooling techniques and advanced materials. Among these materials, Ceramic Matrix Composites (CMCs) have shown great potential in reducing the overall engine weight when used instead of high-tech metal alloys, resulting in lower fuel consumption and emissions during launches. This paper focuses on the mass minimization of inner liners made of CMCs in rocket thrust chambers. At this aim, a computationally efficient preliminary design approach, based on an analytical one-dimensional thermo-mechanical model, is proposed. A case study of mass minimization of an inner liner of rocket thrust chamber is also presented and discussed, by considering five different CMC materials.

Adaptive physics-encoded graph neural network for health stage assessment of liquid-propellant rocket engines

The improvement of reliable health monitoring system for liquid-propellant rocket engines (LREs) is a crucial part for reusable launch vehicle, which contributes to providing competitive and cost-effective propulsion systems. Thus, it accentuates the need for reliable and quick health stage assessment of system and follow-up damage-mitigating control. In this paper, we propose a novel adaptive physics-encoded graph neural network for health stage assessment of LREs. Our approach embeds the relations of different sensors obtained through expert experience, which contributes to constructing a physical graph layer. To better capture the information contained in all the sensor data, a novel convolutional layer of adaptive auto-regressive moving average filters is designed, which considers the personalized information propagation needs of each neural network layer. The performance of the proposed method is quantified with data obtained from physics simulations and real-world engineering systems. The results show that our model has potential applicability for the health stage assessment of LREs with high accuracy.

Particle-Filter-Based Fault Diagnosis for the Startup Process of an Open-Cycle Liquid-Propellant Rocket Engine.

This study introduces a fault diagnosis algorithm based on particle filtering for open-cycle liquid-propellant rocket engines (LPREs). The algorithm serves as a model-based method for the startup process, accounting for more than 30% of engine failures. Similar to the previous fault detection and diagnosis (FDD) algorithm for the startup process, the algorithm in this study is composed of a nonlinear filter to generate residuals, a residual analysis, and a multiple-model (MM) approach to detect and diagnose faults from the residuals. In contrast to the previous study, this study makes use of the modified cumulative sum (CUSUM) algorithm, widely used in change-detection monitoring, and a particle filter (PF), which is theoretically the most accurate nonlinear filter. The algorithm is confirmed numerically using the CUSUM and MM methods. Subsequently, the FDD algorithm is compared with an algorithm from a previous study using a Monte Carlo simulation. Through a comparative analysis of algorithmic performance, this study demonstrates that the current PF-based FDD algorithm outperforms the algorithm based on other nonlinear filters.

Визначення параметрів вібрацій космічного апарата на активній ділянці польоту ракети-носія на основі вогневих випробувань рiдинного ракетного двигуна

In orbital injection, the launch vehicle (LV) structure and the spacecraft are subjected to extreme dynamic loads, in particular to vibroacoustic loads (from rocket engine thrust oscillations and aerodynamic loads), which may cause spacecraft instrumentation malfunction and damage spacecraft light-weight thin-walled structures. This paper is dedicated to the development of an approach to predicting dynamic loads on spacecraft in orbital injection by LVs of various layouts under propulsion system thrust oscillations in active flight. The paper presents an approach to predicting dynamic loads on spacecraft in orbital injection by LVs of various layouts. The approach makes it possible to evaluate dynamic loads (spectral densities of vibration accelerations) on spacecraft under propulsion system thrust oscillations acting on the liquid-propellant LV structure in active flight. The approach includes a mathematical simulation of the spatial oscillations of the LV structure according to its structural layout scheme and the experimental pre-determination of the spectral density of the rocket engine power. The workability of the proposed approach in predicting the spacecraft dynamic loads is demonstrated by the example of a computational analysis of the spectral densities of spacecraft oscillations in orbital injection by LVs of various structural layouts. It is shown that the approach allows one to predict, as early as at the initial LV design stage, the spacecraft vibratory load parameters at different times of the LV first-stage liquid-propellant rocket engine operation accounting for the rocket layout (with the spacecraft) and design features and using the vibroacoustic characteristics of the liquid-propellant rocket engine (known from the results of its fire tests).

Визначення коефіцієнтов гідродинамічної моделі кавітуючих насосів рідин-них ракетних двигунів за їх теоретичними передатними матрицями

The characterization of cavitating pumps of liquid-propellant rocket engines (LPRE) is an important problem because of the need to provide the pogo stability of liquid-propellant launch vehicles and the stability of liquid-propellant propulsion systems for cavitation oscillations. The development of a reliable mathematical model of LPRE cavitating pumps allows this problem to be resolved. The goal of this work is to determine the cavitation number and operating parameter dependences of the coefficients of a lumped-parameter hydrodynamic model of LPRE cavitating pumps from their theoretical transfer matrices obtained by a distributed-parameter model. The following coefficients are found as a function of operating parameters: the cavitation elasticity, the cavitation resistance, the cavity-caused disturbance transfer delay time, and the cavitation resistance distribution coefficient. The last two coefficients are new in the hydrodynamic model of cavitating pumps, and they were introduced when verifying the model using experimental and theoretical pump transfer matrices. Analyzing the cavitation resistance distribution coefficient as a function of operating parameters shows that it markedly decreases with increasing cavitation number. This testifies to that the location of the lumped cavity compliance is shifted from the mid position towards the pump inlet. Therefore, the assumption that the lumped cavity compliance is located in the middle of the attached cavity regardless of the cavitation number is not justified. The fact that the distribution coefficient as a function of cavitation number intersects the abscissa axis near a cavitation number of 0.25 may indicate the boundary of existence of attached cavities and thus the applicability boundary of the theoretical model. The disturbance transfer delay time as a function of cavitation number sharply increases at cavitation numbers of about 0.05. At cavitation numbers of about 0.25, it is close to a constant.

Design optimization of bellow joints used in liquid propellant rocket engines

Abstract Bellow joints are frequently used in hydraulic lines, constructions, and various areas such as nuclear stations to absorb the energy caused by flow and external forces, provide flexibility to the lines, and prevent damages such as cracking and deterioration in the flow lines. There exist various types of bellow joints (e.g., axial type, gimbal type, and hinge type) that allow axial, lateral, and angular movements. Bellow joints that assist thrust vector control in liquid propellant rocket engines prevent the hydraulic lines from being damaged during the orientation movements of the missile. While providing this flexibility to the lines in rocket engines, they create additional force against the linear actuators that move the liquid motor nozzle. This additional force causes the need for larger actuators, resulting in more weight and volume. In this study, design optimization of the bellow joint used in liquid propellant rocket engines is conducted to minimize the force transferred to the actuators by minimizing the bending moment developed in the bellow joint. It is found that the bending moment developed in the bellow joint could be reduced by a significant rate of 75 % without compromising the structural integrity of the bellow joint.

IMPROVEMENT OF DESIGN PARAMETERS OF INERTIAL TYPE FUEL CONTINUITY SYSTEMS

Multiple restart of liquid-propellant rocket engines of spacecraft in zero gravity is one of the most difficult tasks that arise in the creation of new products of rocket technology. When performing a flight task, due to the movement of the spacecraft along the passive section of the trajectory, the components of the liquid fuel are mixed with the pressurized gas in its tanks in an arbitrary manner. There is a possibility that the gas phase will be located close to the drain hole. At the time of re-turning on the propulsion engines of the spacecraft, along with fuel, gas bubbles will enter the consumable line from the tank, which can, in turn, cause a failure of their launch.
 In order to avoid this emergency, special fuel continuity systems are used. Despite the significant diversity of these systems, all of them have significant shortcomings and limited working conditions. The main purpose of using fuel continuity systems of any type is to prevent the penetration of boost gas from the tank cavity into the flow line until the tank is completely emptied.
 The paper is devoted to the analysis of the possibilities of improving the design parameters of the two most common systems for ensuring fuel continuity by creating a combined "hybrid" system on their basis. Inertial and mesh systems for ensuring fuel continuity are considered. Theoretically, an analysis of the possibility of reducing the effect of pre-start acceleration for an inertial system for ensuring fuel continuity due to the use of a mesh phase separator in the fuel supply system is carried out. The paper considers the main design parameters of this combined system, which directly affect the level of its technical perfection, as well as the possibilities of their optimization.
 The results of the work can be useful in engineering practice in the creation of new promising products of rocket and space technology.

The violent collapse of vapor bubbles in cryogenic liquids

Vapor bubbles in cryogenic fluids may collapse violently under subcooled and pressurized conditions. Despite important implications for engineering applications such as cavitation erosion in liquid propellant rocket engines, these intense phenomena are still largely unexplored. In this paper, we systematically investigate the ambient conditions leading to the occurrence of violent collapses in liquid nitrogen and analyze their thermodynamic characteristics. Using Brenner’s time ratio χ, the regime of violent collapse is identified in the ambient pressure–temperature parameter space. Complete numerical simulations further refine the prediction and illustrate two classes of collapses. At 1 < χ < 10, the collapse is impacted by significant thermal effects and attains only moderate wall velocity. Only when χ > 10 does the collapse show more inertial features. A mechanism analysis pinpoints a critical time when the surrounding liquid enters supercritical state. The ultimate collapse intensity is shown to be closely associated with the dynamics at this moment. Our study provides a fresh perspective to the treatment of cavitation in cryogenic fluids. The findings can be instrumental in engineering design to mitigate adverse effects arising from intense cavitational activities.

Prognostics Aware Control Design for Extended Remaining Useful Life

As most of the safety critical industrial systems remain sensitive to functional degradation and operate under closed loop, it becomes imperative to take into account the state of health within the control design process. To that end, an effective assessment and extension of the Remaining Useful Life (RUL) of complex systems is a standing challenge that seeks novel solutions at the cross-over of Prognostics and Health Management (PHM) domain as well as automatic control. This paper considers a dynamical system subjected to functional degradation presents a novel control design strategy. Wherein the assessment of state of health of the system is taken into account leading to effective prediction of the RUL as well as its extension. To that end, the degradation model is considered unknown but input-dependent. The control design is formulated as an optimization problem wherein a suitable comprise is reached between the performance and desired RUL of the system. The main contribution of the paper remains in proposal of set-point modulation based approach wherein the control input at a given present time stage is modulated in such way that futuristic health of the system over a long time horizon is extended whilst assuring acceptable performance. The effectiveness of the proposed strategy is assessed in simulation using a numerical example as well as liquid propellant rocket engine case.

Simulation of a liquid propellant rocket engine manufacturing assembly line utilizing the PFS/PN methodology

The system addressed in this paper pertains to an assembly line for liquid propellant rocket engines equipped with turbopump. The challenge involves automating the process of metal 3D printing for the turbopump and combustion chamber, followed by heat treatment for stress relief, X-ray computed tomography inspection, and the subsequent movement of components throughout the factory. To achieve this, the PFS/PN methodology was employed to facilitate a detailed analysis of the system, ensuring control over its inputs and outputs. The system was simulated using the ProModel Discrete Event Simulator, and the results were scrutinized, indicating the robustness of the simulation. Following the Ladder methodology, programming for Programmable Logic Controllers was also developed.

Regenerative Cooling Comparison of LOX/LCH4 and LOX/LC3H8 Rocket Engines Using the One-Dimensional Regenerative Cooling Modelling Tool ODREC

Due to the extreme temperatures inside the combustion chambers of liquid propellant rocket engines, the walls of the combustion chamber and the nozzle are cooled by either the fuel or the oxidizer in what is known as regenerative cooling. This study presents an indigenous computational tool developed for the analysis of heat transfer in regenerative cooling of such rocket engines. The developed tool incorporates a one-dimensional (1-D) combustion analysis to calculate the thermophysical properties of the combustion gas. Basic engine properties were calculated and used to generate a thrust chamber profile based on a bell-shaped nozzle. The hot gas side was analyzed using 1-D isentropic flow assumptions, along with heat transfer correlations. The coolant side was evaluated using the hydraulic analysis in the axial direction and the heat transfer analysis in the radial direction. Thermophysical properties and the phase of the coolant were determined using the given property tables and the instantaneous state of the coolant. This flexible and computationally less demanding tool was used to analyze two small-scale engines utilizing liquid hydrocarbon fuels, which are used in modern rocket propulsion. The wall cooling analyses of a liquid oxygen (LOX)/liquid methane (LCH4) engine and a liquid oxygen (LOX)/liquid propane (LC3H8) engine are presented. Fuel and oxidizer were used separately as coolants for both engines, and both of them experienced phase change. Results reveal the advantage of the high mass flow rate of the oxidizer in cooling performance. In addition, the results of this study show that the cooling of the LOX/LC3H8 engine is somewhat more challenging compared to the LOX/LCH4 engine.

Fault Diagnosis in Partially Observable Petri Nets with Quantum Bayesian Learning

This paper investigates the quantum Bayesian probability estimation of fault diagnosis based on Partially Observable Petri Nets (POPN) for a liquid-propellant rocket engine system. To solve the problem of a poor environment, a complex structure, and limited observable information in the liquid-propellant rocket engine system, a method of fault diagnosis based on POPN and quantum Bayesian probability estimation is proposed. According to the operating state and key actions of the system model, the places and transitions are set, and the unobservable key actions will become unobservable transitions. Combined with the trigger relationship of the transitions, a POPN model is established. All path estimation system states that satisfy the observable transition sequence information are traversed. If the diagnosis result is a possible failure, we establish a quantum Bayesian Petri net model corresponding to the failure transition, manually adjust the quantum parameters to calculate the quantum probability of the failure transition, and determine the system failure state. Finally, the model of the start-up process of the engine system based on the POPN is built to verify the effectiveness of the algorithm with the data in the simulation experiment.

Devising a method to design supersonic nozzles of rocket engines by using numerical analysis methods

The object of research is supersonic nozzles of liquid-propellant rocket engines. The work considers the problem of the lack of an effective method for profiling the supersonic contour of the nozzle, which will generate maximum thrust. For its solution, a method is proposed, the essence of which is approximating the nozzle contour with a power-law polynomial and determining the values of its coefficients by solving a multidimensional minimization problem using numerical modeling methods. The expression for the axial component of the thrust with the opposite sign at the specified values of atmospheric pressure and radius at the nozzle section was chosen as the objective function in this paper. Using the proposed method, contours of optimal nozzles were obtained based on polynomials of powers 2, 3, and 4, which were compared with nozzles obtained by the generally accepted Rao method. The maximum value of the relative deviation modulus calculated during the comparison did not exceed 3 %, which allows us to assert the correctness of the obtained results. The existence of such a discrepancy is explained by the difference in the numerical modeling method used. In contrast to the method of characteristics common in similar problems, the method of finite volumes of the Godunov type was used in the work. This has made it possible to reduce the sensitivity of the calculation to initial and boundary conditions and make decisions regardless of the flow regime of combustion products. In addition, the use of the extended cells method for the integration of finite volumes at the boundary of the calculation domain significantly reduced the total time of solving the problem of profiling the contour of the optimal nozzle

Результати проектування надзвукового сопла з подвійним розширенням для рідинного ракетного двигуна першого ступеня методами обчислювального аналізу

The subject of this study is dual-bell supersonic nozzles of liquid rocket engines. The goal is to design a dual-bell supersonic nozzle that will provide maximum thrust for a liquid rocket engine in the first stage of a launch vehicle by solving an optimization problem using numerical simulation methods. The goal of the study is to choose the exit pressures of each part of the dual-bell nozzle based on the known flight trajectory of the launch vehicle, set and solve the problem of optimizing the dual-bell nozzle contour, and assess the impact of the use of a dual-bell nozzle on the efficiency of a liquid-propellant rocket engine. The methods used are: numerical methods for solving the hyperbolic system of unsteady equations of gas dynamics and multidimensional minimization problems. The following results were obtained. The formulation and solution of the optimization problem of the contour of a dual-bell supersonic nozzle, considering the design limitations in the form of a fixed length of the nozzle, is carried out. By analyzing the flight path of the first stage of the launch vehicle, a first approximation of the nozzle contour was obtained. For further calculations, both sections were approximated by parabolas, the coefficients of which, together with the lengths of the sections, formed a vector of optimized parameters. An expression for the axial component of thrust with the opposite sign was used as the objective function to solve the minimization problem. Because of solving the optimization problem, a dual-bell nozzle contour was designed to provide maximum thrust. An assessment of the effectiveness of his work was also conducted. The calculated value of the average along the trajectory of the specific impulse when using a dual-bell nozzle was higher by 1.6% than when using a standard nozzle. Conclusions. The scientific novelty of the obtained results is as follows: the design of the dual-bell nozzle contour was performed by solving the problem of multidimensional optimization, taking into account the design constraints, and using computational analysis methods

FEATURES OF THE DEVELOPMENT OF ADDITIVE MANUFACTURING METHODS IN APPLICATION TO LIQUID PROPELLANT ROCKET ENGINES

Abstract. Modern development of additive manufacturing methods makes it necessary to seriously consider an alternative type of production for the manufacture of liquid rocket engine structures. On the one hand, additive manufacturing has a number of undeniable advantages compared to classical methods traditionally used in the rocket and space industry. On the other hand, products obtained by this method have characteristic features that must be taken into account at the early stages of design: relatively high surface roughness, the need to adapt the geometry of existing technical solutions, material properties, etc. Moreover, a full-fledged quality control system for the resulting products is currently the moment is missing, and the standardization of the production class is just emerging. It is also important to understand the direction of development of additive manufacturing, which in the future can lead to a significant increase in the characteristics of the resulting products. This paper discusses options for classifying additive manufacturing methods, provides definitions and classification of metal 3D printing methods, according to the existing ISO/ASTM 52900 standard, as well as the main features of the technology. The main advantages and disadvantages of additive manufacturing methods are considered, and a structural decomposition of the typical design of a liquid propellant rocket engine chamber is carried out. The work aimed at manufacturing structural elements for further integration into the production cycle of liquid-propellant rocket engines is considered. The work is considered, the results of which were a combination of additive technology methods, which made it possible to apply the concept of bimetallic additive production. Such products were produced without the use of complex technological equipment and belong to the product of “evolving” additive manufacturing. Based on the results of the work, it was concluded that it is possible to use additive technologies to improve the performance of new liquid-propellant rocket engine designs.

Maximum Likelihood Identification of Backflow Vortex Instability in Rocket Engine Inducers

Abstract Bayesian estimation is applied to the analysis of backflow vortex instabilities in typical three- and four-bladed liquid propellant rocket engine inducers. The flow in the impeller eye is modeled as a set of equally intense and evenly spaced 2D axial vortices, located at the same radial distance from the axis and rotating at a fraction of the impeller speed. The circle theorem is used to predict the flow pressure in terms of the vortex number, intensity, rotational speed, and radial position. The theoretical spectra so obtained are frequency broadened to mimic the dispersion of the experimental results and parametrically fitted to the measured data by maximum likelihood estimation with equal and independent Gaussian errors. The method is applied to three inducers, tested in water at room temperature and different operating conditions. It successfully characterizes backflow instabilities using the signals of a single pressure transducer flush-mounted in the impeller eye, effectively bypassing the aliasing limitations and the data acquisition/reduction complexities of traditional multiple-sensor cross-correlation methods. The identification returns the estimates of the model parameters and their standard deviations, providing the information necessary for assessing the accuracy and statistical significance of the results. The flowrate is found to be the major factor affecting the backflow vortex instability, which, on the other hand, is rather insensitive to the occurrence of cavitation. The results are consistent with the data reported in the literature, as well as with those generated by the auxiliary models specifically developed for initializing the maximum likelihood searches and supporting the identification procedure.