Research

26 pages, 5993 KiB

Open AccessArticle

Thermal Behaviour of the Cooling Jacket Belonging to a Liquid Oxygen/Liquid Methane Rocket Engine Demonstrator in the Operation Box

by Daniele Ricci, Francesco Battista, Manrico Fragiacomo and Ainslie Duncan French

Aerospace 2023, 10(7), 607; https://doi.org/10.3390/aerospace10070607 - 30 Jun 2023

Viewed by 1571

Abstract

The cooling jackets of liquid rocket engines are composed of narrow passages surrounding the thrust chambers and ensure the reliable operation of the engine. Critical conditions may also be encountered, since the cooling jackets of cryogenic engines, such as those using LO_X [...] Read more.

The cooling jackets of liquid rocket engines are composed of narrow passages surrounding the thrust chambers and ensure the reliable operation of the engine. Critical conditions may also be encountered, since the cooling jackets of cryogenic engines, such as those using LO_X/LCH₄ propellants, are based on a regenerative strategy, where the fuel is used as a refrigerant. Consequently, deterioration modes near where pseudocritical conditions are reached or low heat transfer coefficients where the fuel becomes a vapour and must therefore be managed. The verification of the cooling jacket behaviour to consolidate the design solutions in all the extreme points of the operating box represents a very important phase. The present paper discusses the full characterization of the HYPROB (HYdrocarbon PROpulsion test Bench Program) first unit of the final demonstrator, (DEMO-0A), by considering the working points within the limits of the operating box and comparisons with the nominal conditions are given. In this way, a full understanding of the cooling system behaviour, affecting the working of the entire thrust chamber, is accomplished. Moreover, the design strategy and choices have been confirmed, since the verifications also include potentially even more extreme conditions with respect to the nominal ones. The investigation has been numerically performed and supported the thermo-structural analyses accomplished before the final firing campaign, completed in December 2022. Since little information is available in the literature on LO_X/LCH₄ engines, suggestions are given as to the organization of the numerical simulations, which support the design of such rocket engine cooling systems. Full article

(This article belongs to the Special Issue Liquid Rocket Engines)

► Show Figures

Figure 1

17 pages, 2562 KiB

Open AccessArticle

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

by Jihyoung Cha, Erik Andersson and Alexis Bohlin

Aerospace 2023, 10(7), 582; https://doi.org/10.3390/aerospace10070582 - 23 Jun 2023

Cited by 2 | Viewed by 3739

Abstract

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 [...] Read more.

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. Full article

(This article belongs to the Special Issue Liquid Rocket Engines)

► Show Figures

Figure 1

14 pages, 1883 KiB

Open AccessArticle

Experimental Investigation of Flame Anchoring Behavior in a LOX/LNG Rocket Combustor

by Jan Martin, Michael Börner, Justin Hardi, Dmitry Suslov and Michael Oschwald

Aerospace 2023, 10(6), 542; https://doi.org/10.3390/aerospace10060542 - 06 Jun 2023

Viewed by 1174

Abstract

Hot fire tests of a multi-injector research combustor were performed with liquid-oxygen and liquefied-natural-gas (LOX/LNG) propellants at chamber pressures from 30 up to 67 bar, hence at conditions similar to an upper stage rocket engine. Within these tests shear coaxial injectors were tested [...] Read more.

Hot fire tests of a multi-injector research combustor were performed with liquid-oxygen and liquefied-natural-gas (LOX/LNG) propellants at chamber pressures from 30 up to 67 bar, hence at conditions similar to an upper stage rocket engine. Within these tests shear coaxial injectors were tested with and without a recessed LOX post. In both configurations, operating conditions with flames anchored at the LOX post tip and thus, if available, pre-combustion in the recess volume as well as lifted flames were observed. Flame anchoring was indirectly detected via acoustic measurements, using mean speed of sound to indicate the presence of flame in the head end of the combustion chamber. While the injector without recess showed only stable combustion irrespective of the flame anchoring behavior, the recessed injector featured short-lived bursts of oscillatory combustion and sustained combustion instabilities. Analysis of the test data showed that stable flame anchoring could not be ensured at momentum flux ratios below 20 for a non-recessed and below 45 for a recessed injector. Full article

(This article belongs to the Special Issue Liquid Rocket Engines)

► Show Figures

Figure 1

16 pages, 8124 KiB

Open AccessArticle

Orifice Flow Dynamics in a Rocket Injector as an Excitation Source of Injector-Driven Combustion Instabilities

by Min Son, Michael Börner, Wolfgang Armbruster and Justin S. Hardi

Aerospace 2023, 10(5), 452; https://doi.org/10.3390/aerospace10050452 - 15 May 2023

Cited by 2 | Viewed by 1897

Abstract

To investigate a hypothesis of the orifice flow-induced instability in rocket engine thrust chambers, a single liquid oxygen (LOX) injector with an optically accessible orifice module was used for experiments, with water as a simulant for LOX. The unsteady pressure downstream of the [...] Read more.

To investigate a hypothesis of the orifice flow-induced instability in rocket engine thrust chambers, a single liquid oxygen (LOX) injector with an optically accessible orifice module was used for experiments, with water as a simulant for LOX. The unsteady pressure downstream of the orifice was measured using high-speed piezoelectric sensors under cavitating and non-cavitating intra-injector flow conditions. The cavitating orifice flows were directly visualized via backlight imaging with a high-speed camera through the optically accessible orifice module. Cavitation initiated at the cavitation number of 2.05, and the downstream bubble cloud formation started below 1.91. The unsteady pressure spectrum arising from cavitation comprises multiple peaks over a broad frequency range, which can cause low- and high-frequency instabilities. The dominant frequencies from cavitation decrease with increasing pressure drop, while the frequencies during non-cavitating flow increase. The non-cavitating orifice flow excites the second longitudinal acoustic mode of the injector tube. The acoustic mode excited by the non-cavitating flow becomes stronger when the pressure peak in the range of whistling phenomenon is close to the first longitudinal acoustic mode. In conclusion, the excitation mechanisms of the orifice-induced instability for the cavitating and non-cavitating flows were well identified, despite the limitations of water as a simulant for LOX. Full article

(This article belongs to the Special Issue Liquid Rocket Engines)

► Show Figures

Graphical abstract

15 pages, 8919 KiB

Open AccessArticle

Improved Wall Temperature Prediction for the LUMEN Rocket Combustion Chamber with Neural Networks

by Kai Dresia, Eldin Kurudzija, Jan Deeken and Günther Waxenegger-Wilfing

Aerospace 2023, 10(5), 450; https://doi.org/10.3390/aerospace10050450 - 12 May 2023

Cited by 3 | Viewed by 1262

Abstract

Accurate calculations of the heat transfer and the resulting maximum wall temperature are essential for the optimal design of reliable and efficient regenerative cooling systems. However, predicting the heat transfer of supercritical methane flowing in cooling channels of a regeneratively cooled rocket combustor [...] Read more.

Accurate calculations of the heat transfer and the resulting maximum wall temperature are essential for the optimal design of reliable and efficient regenerative cooling systems. However, predicting the heat transfer of supercritical methane flowing in cooling channels of a regeneratively cooled rocket combustor presents a significant challenge. High-fidelity CFD calculations provide sufficient accuracy but are computationally too expensive to be used within elaborate design optimization routines. In a previous work it has been shown that a surrogate model based on neural networks is able to predict the maximum wall temperature along straight cooling channels with convincing precision when trained with data from CFD simulations for simple cooling channel segments. In this paper, the methodology is extended to cooling channels with curvature. The predictions of the extended model are tested against CFD simulations with different boundary conditions for the representative LUMEN combustor contour with varying geometries and heat flux densities. The high accuracy of the extended model’s predictions, suggests that it will be a valuable tool for designing and analyzing regenerative cooling systems with greater efficiency and effectiveness. Full article

(This article belongs to the Special Issue Liquid Rocket Engines)

► Show Figures

Figure 1

15 pages, 5259 KiB

Open AccessArticle

Analyzing Combustion Efficiency According to Spray Characteristics of Gas-Centered Swirl-Coaxial Injector

by Seongphil Woo, Jungho Lee, Ingyu Lee, Seunghan Kim, Yeoungmin Han and Youngbin Yoon

Aerospace 2023, 10(3), 274; https://doi.org/10.3390/aerospace10030274 - 10 Mar 2023

Cited by 2 | Viewed by 1714

Abstract

The momentum flux ratio (MFR) significantly affects the mixing characteristics and combustion efficiency of propellants in rocket engine injectors. The spray characteristics of three gas-centered swirl-coaxial injectors used in a full-scale combustion test were investigated according to the change in the momentum flux [...] Read more.

The momentum flux ratio (MFR) significantly affects the mixing characteristics and combustion efficiency of propellants in rocket engine injectors. The spray characteristics of three gas-centered swirl-coaxial injectors used in a full-scale combustion test were investigated according to the change in the momentum flux ratio. The difference in combustion efficiency was analyzed through the comparison with combustion test results using spray visualization and quantification. The spray cross-sectional shape and droplet distribution were measured using a structured laser illumination planar imaging technique. As the swirl effect was more apparent at a low MFR, the flow rate of the liquid that was sprayed outside was high. The flow rate of the liquid sprayed around the gas injection increased with the MFR. The Sauter mean diameter (SMD) of each injector liquid spray was obtained using the laser shadow imaging method. The SMD decreased as the MFR of all injector types increased, and the injector with a high liquid flow rate and small SMD injected towards the gas center exhibited higher combustion efficiency than the injector with a dominant liquid spray and the large SMD at a large injection angle. The outcomes of the study could help contribute to the increase in the combustion efficiency of the full-scale staged combustion cycle engine combustor. Full article

(This article belongs to the Special Issue Liquid Rocket Engines)

► Show Figures

Figure 1

22 pages, 7565 KiB

Open AccessArticle

HCF and LCF Analysis of a Generic Full Admission Turbine Blade

by Jörg R. Riccius and Evgeny B. Zametaev

Aerospace 2023, 10(2), 154; https://doi.org/10.3390/aerospace10020154 - 08 Feb 2023

Viewed by 1804

Abstract

A numerical turbine-blade fatigue-life analysis method is suggested. This method comprises a stationary thermal 3D finite element (FE) analysis of the hot run for the combined high-cycle fatigue (HCF) and creep analysis, and a follow-on (one-way coupled) quasi-stationary structural 3D FE analysis (including [...] Read more.

A numerical turbine-blade fatigue-life analysis method is suggested. This method comprises a stationary thermal 3D finite element (FE) analysis of the hot run for the combined high-cycle fatigue (HCF) and creep analysis, and a follow-on (one-way coupled) quasi-stationary structural 3D FE analysis (including six load steps) of a single and two half turbine blades and the related disk and rotor section and a (modified Goodman equation based) post-processing fatigue life analysis for the highest HCF-loaded point of the turbine blade. For the low-cycle fatigue (LCF) analysis, this includes a transient thermal 3D FE analysis of two full loading cycles, a follow-on (one-way coupled) quasi-stationary structural 3D finite element analysis of a single and two half turbine blades and the related disk and rotor section and a (modified-Langer-equation-based) post-processing fatigue life analysis approach for the highest LCF-loaded point of the turbine blade. Finally, this approach is demonstrated by the numerical HCF, LCF and creep analysis of a generic turbine blade of the first rotor row of a full admission hydrogen turbo pump of a 1 MN thrust class gas generator LOX-LH₂ liquid rocket engine (LRE). For this numerical example, the LCF loading turned out to be dominant. Creep turned out to be negligible. Full article

(This article belongs to the Special Issue Liquid Rocket Engines)

► Show Figures

Figure 1

18 pages, 708 KiB

Open AccessArticle

Supercritical Injection Modeling by an Incompressible but Variable Density Approach

by Leandro B. Magalhães, André R. R. Silva and Jorge M. M. Barata

Aerospace 2023, 10(2), 114; https://doi.org/10.3390/aerospace10020114 - 25 Jan 2023

Viewed by 1101

Abstract

Supercritical nitrogen jet behavior is modeled using an incompressible but variable density approach developed for variable density jets. Following mechanical and thermal breakup concepts, several injection conditions relevant to liquid rocket propulsion are analyzed, considering heat transfer in the injector. Regarding axial density [...] Read more.

Supercritical nitrogen jet behavior is modeled using an incompressible but variable density approach developed for variable density jets. Following mechanical and thermal breakup concepts, several injection conditions relevant to liquid rocket propulsion are analyzed, considering heat transfer in the injector. Regarding axial density distributions, different levels of agreement with experimental data are encountered for potential core, subsided core, and plateau formations. Further comparisons with compressible formulations from the literature are a good indicator of the proposed methodology’s suitability for the simulation of supercritical injection behavior. Full article

(This article belongs to the Special Issue Liquid Rocket Engines)

► Show Figures

Figure 1

18 pages, 4740 KiB

Open AccessArticle

Evaluation of Large-Eddy Simulation Coupled with an Homogeneous Equilibrium Model for the Prediction of Coaxial Cryogenic Flames under Subcritical Conditions

by Thomas Schmitt and Sébastien Ducruix

Aerospace 2023, 10(2), 98; https://doi.org/10.3390/aerospace10020098 - 18 Jan 2023

Cited by 1 | Viewed by 1356

Abstract

Large Eddy Simulations of liquid O

_{2}

/gaseous H

_{2}

coaxial flames at subcritical pressure conditions are reported in this paper. These simulations reproduce the experimental Mascotte cases A1, A10 and A30, operating at 1, 10 and 30 bar, respectively, and for which [...] Read more.

Large Eddy Simulations of liquid O

_{2}

/gaseous H

_{2}

coaxial flames at subcritical pressure conditions are reported in this paper. These simulations reproduce the experimental Mascotte cases A1, A10 and A30, operating at 1, 10 and 30 bar, respectively, and for which temperature measurements and experimental visualisations are available. The main objective of this work is to assess the accuracy of the multi-fluid Homogeneous Equilibrium Model (HEM) described in Pelletier et al. (Computers & Fluids, 2020) for rocket engine applications. Of particular interest is the comparison with the experimental temperature measurements from Grisch et al. (Aerospace science and technology, 2003). To that purpose, numerical simulations are conducted with care, in order to ensure a proper statistical convergence and estimate the influence of the grid resolution for each case. Despite the crude assumptions—no surface tension and no atomisation model, for instance—that are made with the HEM used in this work, results are found to be in reasonable agreements with the measurements for case A10, even with the coarser grid. For case A30, a fine mesh resolution is required to capture the low intensity recirculation zone downstream of the inner jet necessary to reproduce the shape of the experimental profile. Finally, case A1 simulations, with the lowest Weber number, show large departures with the experimental measurements. This is expected to be due to a deficiency of the model to properly reproduce the two-phase dispersed flow. Full article

(This article belongs to the Special Issue Liquid Rocket Engines)

► Show Figures

Figure 1

17 pages, 2568 KiB

Open AccessArticle

Thrust Control Method and Technology of Variable-Thrust Liquid Engine for Reusable Launch Rocket

by Zhaohui Yao, Yiwen Qi, Wen Bao and Tianhong Zhang

Aerospace 2023, 10(1), 32; https://doi.org/10.3390/aerospace10010032 - 30 Dec 2022

Cited by 3 | Viewed by 3484

Abstract

A high-precision variable-thrust control method based on real-time measurement of pintle displacement and closed-loop feedback control is proposed to solve the technical problems of deep throttling variable-thrust regulation and control of pintle liquid rocket engines (LRE). By optimizing the system structure and control [...] Read more.

A high-precision variable-thrust control method based on real-time measurement of pintle displacement and closed-loop feedback control is proposed to solve the technical problems of deep throttling variable-thrust regulation and control of pintle liquid rocket engines (LRE). By optimizing the system structure and control parameters, the closed-loop control of displacement with high precision and a fast response under a wide range of variable thrust can be realized, and thus the large-range, fast-response, and high-precision control of the chamber pressure, equivalent to thrust, can be indirectly realized. The chamber pressure response time is not more than 0.3 s, the overshoot is not more than ±3%, and the pulsation amplitude is not more than ±5%, which can meet the technical requirements of the large-range thrust adjustment and control of variable-thrust LRE of reusable launch rockets. The proposed variable-thrust LRE thrust control system is simple, reliable, and easy to use and maintain, which solves the problem of the large range, high precision, and fast response of thrust adjustment and control. The proposed system can provide important technical support for carrier rocket recycling and launch cost reduction. This is the first time a closed-loop control method of displacement of an integrated gas generator/flow regulator to achieve a 5:1 large-range continuous-variable-thrust control for the LRE of a reusable launch rocket has been proposed. Full article

(This article belongs to the Special Issue Liquid Rocket Engines)

► Show Figures

Figure 1

15 pages, 3583 KiB

Open AccessArticle

Life Analysis of Reusable Liquid Rocket Engine Thrust Chamber

by Yuanjie Qi, Yuqiang Cheng and Yan Zhang

Aerospace 2022, 9(12), 788; https://doi.org/10.3390/aerospace9120788 - 02 Dec 2022

Viewed by 2556

Abstract

The thrust chamber’s inner wall suffers high temperature and pressure differences from the coolant channel, which limits the life of the rocket engine. Life prediction of the thrust chamber really plays an important role in reusable launch vehicle propulsion systems. The Porowski beam [...] Read more.

The thrust chamber’s inner wall suffers high temperature and pressure differences from the coolant channel, which limits the life of the rocket engine. Life prediction of the thrust chamber really plays an important role in reusable launch vehicle propulsion systems. The Porowski beam model is widely used in the life prediction of reusable liquid rocket engine thrust chambers, which calculates the life caused by fatigue, creep, and thinning after each firing cycle. In order to analyze the life of the thrust chamber, a LOX/Kerosene rocket engine is investigated in this paper. The life analysis consists of pressure and temperature differences and structural parameters. Two kinds of inner wall materials were chosen for comparison in this research: OFHC copper and Narloy-Z alloy. The results are presented to offer a reference for the design and manufacture of reusable rocket engine thrust chambers in the future. Full article

(This article belongs to the Special Issue Liquid Rocket Engines)

► Show Figures

Figure 1

20 pages, 3223 KiB

Open AccessArticle

Effects of Compounds in Liquefied Methane on Rocket Engine Operation

by Jan van Schyndel, Elke Goos, Clemens Naumann, Justin S. Hardi and Michael Oschwald

Aerospace 2022, 9(11), 698; https://doi.org/10.3390/aerospace9110698 - 09 Nov 2022

Cited by 4 | Viewed by 2687

Abstract

Methane (CH₄) is a promising rocket fuel for various future space mission scenarios. It has advantages in terms of cost, performance, and environmental friendliness. Currently, there is no clear definition on standards and specifications for liquefied methane or similar liquids such [...] Read more.

Methane (CH₄) is a promising rocket fuel for various future space mission scenarios. It has advantages in terms of cost, performance, and environmental friendliness. Currently, there is no clear definition on standards and specifications for liquefied methane or similar liquids such as liquefied natural gas (LNG) for their use as rocket fuel. However, those regulations are necessary for the commercial, safe, and proper operation of methane rocket engines. Composition and impurities of liquefied methane gas mixtures obtained from natural gas or biogenic sources depend on location of the natural gas source (Europe, Asia, or America), its extraction method and treatment, used cleaning methods or conditions of the gasification process, and biomass sources. In the present work, effects of impurities (N₂, CO₂, C₂H₆) within liquid natural gas/liquid methane on the methalox rocket engine operation behavior are analyzed. Regarding the cold cryogenic side, phase diagrams are discussed and critical temperatures for the fuel side are outlined. Carbon dioxide is identified as a rather problematic pollutant. The combustion processes are investigated with several numerical simulations (1D and 2D CFD). The results indicate a minor influence on the overall combustion temperature and a minor but potentially relevant influence on the pressure within the combustion chamber. Additionally, the results indicate that with respect to temperature and pressure, no complex NOx nitrogen chemistry is required. Full article

(This article belongs to the Special Issue Liquid Rocket Engines)

► Show Figures

Figure 1

27 pages, 33029 KiB

Open AccessArticle

Investigations on an All-Oxide Ceramic Composites Based on Al₂O₃ Fibres and Alumina–Zirconia Matrix for Application in Liquid Rocket Engines

by Christian Bach, Frank Wehner and Jan Sieder-Katzmann

Aerospace 2022, 9(11), 684; https://doi.org/10.3390/aerospace9110684 - 03 Nov 2022

Cited by 3 | Viewed by 2100

Abstract

High performance ceramics, particularly Ceramic Matrix Composite (CMC) materials found their way into liquid rocket engines. Yet, so far, mainly carbide or nonoxide CMCs have been of interest. This paper explores the potential and challenges of oxide–oxide ceramic matrix composites (OCMCs) for application [...] Read more.

High performance ceramics, particularly Ceramic Matrix Composite (CMC) materials found their way into liquid rocket engines. Yet, so far, mainly carbide or nonoxide CMCs have been of interest. This paper explores the potential and challenges of oxide–oxide ceramic matrix composites (OCMCs) for application in rocket thrust chambers. Therefore, strength, leakage and hot gas tests are conducted with material samples. A particular focus lies on the application of coatings to seal the permeability inherent to the material. Furthermore, prototypes in the form of flame tubes, ceramic chambers with nozzles and ceramic chambers with graphite inlays are developed and investigated experimentally in test firings. The results show that a recrystallised glass of a Y-Al-Si-O compound can successfully create an impermeable coating of the OCMC without affecting its damag-tolerant behaviour. However, the prototype developments show that it is still very challenging to manufacture even slightly complex structures without critical failures. Nevertheless, OCMC structures of relatively simple geometries showed promising results in hot firings and could be used as a lightweight housing, while the inner contour of the chamber and nozzle are realised, e.g., by a graphite inlay of appropriate quality. Full article

(This article belongs to the Special Issue Liquid Rocket Engines)

► Show Figures

Graphical abstract

18 pages, 25684 KiB

Open AccessArticle

Selection Rules for Resonant Longitudinal Injector-Coupling in Experimental Rocket Combustors

by Tim Horchler

Aerospace 2022, 9(11), 669; https://doi.org/10.3390/aerospace9110669 - 29 Oct 2022

Cited by 2 | Viewed by 1428

Abstract

This paper investigates different types of longitudinal mode coupling in subscale rocket combustion chambers using experimental data and numerical simulations. Based on a one-dimensional planar wave acoustic model of coupled cavity resonators with two acoustic inlet boundary conditions, mode selection rules are derived, [...] Read more.

This paper investigates different types of longitudinal mode coupling in subscale rocket combustion chambers using experimental data and numerical simulations. Based on a one-dimensional planar wave acoustic model of coupled cavity resonators with two acoustic inlet boundary conditions, mode selection rules are derived, providing a simple way of predicting which injector and combustion chamber modes have matching frequencies. Longitudinal mode coupling of an injector with an acoustically open inlet boundary condition has been reported in the literature for the start-up transient of a research combustor experiment. In this experiment, every third injector mode couples to a corresponding chamber longitudinal mode, which is explained in terms of the selection rules derived in this paper. Numerical simulation results for a different combustor experiment show an unexpected mode coupling behavior when an acoustically closed injector inlet is used. Theoretical analysis by using the one-dimensional wave model and applying the derived selection rules shows that in this setup, the injector acoustic mode can accommodate two different acoustic boundary conditions at the injector-chamber interface simultaneously. This results in different acoustic mode shapes in the injector, explaining the unexpected behavior for the resonant coupling with an acoustically closed injector inlet. Full article

(This article belongs to the Special Issue Liquid Rocket Engines)

► Show Figures

Figure 1

21 pages, 1312 KiB

Open AccessArticle

Limits of Fluid Modeling for High Pressure Flow Simulations

by Nelson P. Longmire and Daniel T. Banuti

Aerospace 2022, 9(11), 643; https://doi.org/10.3390/aerospace9110643 - 24 Oct 2022

Cited by 3 | Viewed by 1583

Abstract

Flows in liquid propellant rocket engines (LRE) are characterized by high pressures and extreme temperature ranges, resulting in complex fluid behavior that requires elaborate thermo-physical models. In particular, cubic equations of state and dedicated models for transport properties are firmly established for LRE [...] Read more.

Flows in liquid propellant rocket engines (LRE) are characterized by high pressures and extreme temperature ranges, resulting in complex fluid behavior that requires elaborate thermo-physical models. In particular, cubic equations of state and dedicated models for transport properties are firmly established for LRE simulations as a way to account for the non-idealities of the high-pressure fluids. In this paper, we review some shortcomings of the current modeling paradigm. We build on the common study of property errors, as a direct measure of the density or heat capacity accuracy, to evaluate the quality of cubic equations of state with respect to pseudo boiling of rocket-relevant fluids. More importantly, we introduce the sampling error as a new category, measuring how likely a numerical scheme is to capture real fluid properties during a simulation, and show how even reference quality property models may lead to errors in simulations because of the failure of our numerical schemes to capture them. Ultimately, a further evolution of our non-ideal fluid models is needed, based on the gained insight over the last two decades. Full article

(This article belongs to the Special Issue Liquid Rocket Engines)

► Show Figures

Figure 1

30 pages, 10220 KiB

Open AccessArticle

Data Driven Models for the Design of Rocket Injector Elements

by José Felix Zapata Usandivaras, Annafederica Urbano, Michael Bauerheim and Bénédicte Cuenot

Aerospace 2022, 9(10), 594; https://doi.org/10.3390/aerospace9100594 - 12 Oct 2022

Cited by 2 | Viewed by 2221

Abstract

Improving the predictive capabilities of reduced-order models for the design of injector and chamber elements of rocket engines could greatly improve the quality of early rocket chamber designs. In the present work, we propose an innovative methodology that uses high-fidelity numerical simulations of [...] Read more.

Improving the predictive capabilities of reduced-order models for the design of injector and chamber elements of rocket engines could greatly improve the quality of early rocket chamber designs. In the present work, we propose an innovative methodology that uses high-fidelity numerical simulations of turbulent reactive flows and artificial intelligence for the generation of surrogate models. The surrogate models that were generated and analyzed are deep learning networks trained on a dataset of 100 large eddy simulations of a single-shear coaxial injector chamber. The design of experiments was created considering three design parameters: chamber diameter, recess length, and oxidizer–fuel ratio. The paper presents the methodology developed for training and optimizing the data-driven models. Fully connected neural networks (FCNNs) and U-Nets were utilized as surrogate-modeling technology. Eventually, the surrogate models for the global quantity, average, and root mean square fields were used in order to analyze the impact of the length of the post’s recess on the performances obtained and the behavior of the flow. Full article

(This article belongs to the Special Issue Liquid Rocket Engines)

► Show Figures

Figure 1

16 pages, 5078 KiB

Open AccessArticle

Comparative Analysis of Real-Time Fault Detection Methods Based on Certain Artificial Intelligent Algorithms for a Hydrogen–Oxygen Rocket Engine

by Peihao Huang, Tao Wang, Lin Ding, Huhuang Yu, Yong Tang and Dianle Zhou

Aerospace 2022, 9(10), 582; https://doi.org/10.3390/aerospace9100582 - 07 Oct 2022

Cited by 2 | Viewed by 1683

Abstract

The real-time fault detection and diagnosis algorithm of a liquid rocket engine is the basis of online reconfiguration of guidance and the control system of a launch vehicle, which is directly related to the success or failure of space mission. Based on previous [...] Read more.

The real-time fault detection and diagnosis algorithm of a liquid rocket engine is the basis of online reconfiguration of guidance and the control system of a launch vehicle, which is directly related to the success or failure of space mission. Based on previous related works, this paper carries out comparative experimental studies of relevant intelligent algorithm models for real-time fault detection engineering application requirements of a liquid hydrogen–oxygen rocket engine. Firstly, the working state and detection parameters’ selection of a hydrogen–oxygen engine are analyzed, and the proposed three real-time intelligent fault detection algorithm model design methods are elaborated again. Fault detection calculation and analysis are carried out through normal test data and fault test data. The comparative analysis results of real-time intelligent fault detection algorithm models is presented from three dimensions: detection time, fault detection, and stability and consistency. Finally, based on a correlation analysis, a comprehensive intelligent fault diagnosis model design framework is given to further solve the requirements of real-time fault detection and diagnosis engineering development of a liquid rocket engine, a complex piece of equipment. Full article

(This article belongs to the Special Issue Liquid Rocket Engines)

► Show Figures

Figure 1

18 pages, 3376 KiB

Open AccessArticle

Rotating Detonation Combustion for Advanced Liquid Propellant Space Engines

by Stephen D. Heister, John Smallwood, Alexis Harroun, Kevin Dille, Ariana Martinez and Nathan Ballintyn

Aerospace 2022, 9(10), 581; https://doi.org/10.3390/aerospace9100581 - 07 Oct 2022

Cited by 3 | Viewed by 3282

Abstract

Rotating (also termed continuous spin) detonation technology is gaining interest in the global research and development community due to the potential for increased performance. Potential performance benefits, thrust chamber design, and thrust chamber cooling loads are analyzed for propellant applications using liquid oxygen [...] Read more.

Rotating (also termed continuous spin) detonation technology is gaining interest in the global research and development community due to the potential for increased performance. Potential performance benefits, thrust chamber design, and thrust chamber cooling loads are analyzed for propellant applications using liquid oxygen or high-concentration hydrogen peroxide oxidizers with kerosene, hydrogen, and methane fuels. Performance results based on a lumped parameter treatment show that theoretical specific impulse gains of 3–14% are achievable with the highest benefit coming from hydrogen-fueled systems. Assessment of thrust chamber designs for notional space missions shows that both thrust chamber length and diameter benefits are achievable given the tiny annular chamber volume associated with the rotating detonation combustion. While the passing detonation front drastically increases local heat fluxes, global energy balances can be achieved if operating pressures are limited to be comparable to existing or prior space engines. Full article

(This article belongs to the Special Issue Liquid Rocket Engines)

► Show Figures

Figure 1

12 pages, 5177 KiB

Open AccessArticle

Effect of Local Momentum Ratio on Spray Windward Distribution of a Gas–Liquid Pintle Injector Element

by Xuan Jin, Yang Yang, Xiaomei Cao and Jinshui Wu

Aerospace 2022, 9(9), 494; https://doi.org/10.3390/aerospace9090494 - 03 Sep 2022

Cited by 1 | Viewed by 1672

Abstract

The variable-area pintle injector has unique geometry and spray characteristics compared to traditional coaxial injectors, and is advantageous for weight lightening and deep throttling of liquid rocket engines. To obtain an accurate prediction of the spray windward distribution of a gas–liquid pintle injector [...] Read more.

The variable-area pintle injector has unique geometry and spray characteristics compared to traditional coaxial injectors, and is advantageous for weight lightening and deep throttling of liquid rocket engines. To obtain an accurate prediction of the spray windward distribution of a gas–liquid pintle injector with discrete radial orifices, a pintle injector element using air and water as simulants was designed for spray experiments in the atmospheric environment. The air-film injection pressure drop and water-jet injection orifice diameter were both adjusted for a wide variance range from 0.19 to 2.85 for the local momentum ratio. Backlight imaging was adopted for shooting the frozen spray pattern from one side, and a new dimensionless parameter, i.e., the spray fraction, was defined to quantitatively analyze the time-averaged windward boundary band. The dimensionless spray windward boundary band model for a circular-orifice jet and the corresponding derivative formula of the spray half angle were summarized through parameter study. The predicted results of empirical models were in good agreement with the experimental results. It was found that when the local momentum ratio was about 1, the spray distribution range basically overlapped with the coverage scope of gas film with uniform liquid mist. Full article

(This article belongs to the Special Issue Liquid Rocket Engines)

► Show Figures

Figure 1

14 pages, 409 KiB

Open AccessArticle

Research and Development of Fault Diagnosis Methods for Liquid Rocket Engines

by Tao Wang, Lin Ding and Huahuang Yu

Aerospace 2022, 9(9), 481; https://doi.org/10.3390/aerospace9090481 - 29 Aug 2022

Cited by 6 | Viewed by 2421

Abstract

Currently, considerable efforts are being focused on the development of reusable rockets and smart rockets due to the heavy requirements of future next-generation aerospace transportation. Safety, low-launching cost, and repeatability are expected from liquid rocket for fulfilling the big dreams of space transportation, [...] Read more.

Currently, considerable efforts are being focused on the development of reusable rockets and smart rockets due to the heavy requirements of future next-generation aerospace transportation. Safety, low-launching cost, and repeatability are expected from liquid rocket for fulfilling the big dreams of space transportation, exploration, and travelling. Therefore, research on fault detection of the liquid rocket engines (LRE) is critical for satisfying the above claims. Therefore, a comprehensive survey on the research and development of fault diagnosis systems and methods for the liquid rocket engines is presented. First, development history of liquid rocket engine diagnostic systems is reviewed thoroughly. Then three broad headings of the fault detection approaches of liquid rocket engines are divided through the summary and analysis of the existing methods, including approaches using signal processing, model-driven approach, and approach using artificial intelligence (AI). Then the paper discusses the concrete algorithms according to the classification features of the algorithms. In the end, the future developments of the fault detection approaches are presented, which will mainly pay attention to the reusability and intelligence of the rockets. Full article

(This article belongs to the Special Issue Liquid Rocket Engines)

► Show Figures

Figure 1

19 pages, 2294 KiB

Open AccessArticle

Development of Ultra-Low Specific Speed Centrifugal Pumps Design Method for Small Liquid Rocket Engines

by Hye In Kim, Tae-Seong Roh, Hwanil Huh and Hyoung Jin Lee

Aerospace 2022, 9(9), 477; https://doi.org/10.3390/aerospace9090477 - 28 Aug 2022

Cited by 8 | Viewed by 3567

Abstract

With the growth of the satellite industry, the demand for a propulsion system for small launch vehicles and spacecraft has increased. Small liquid rocket engines may require Ultra-Low specific speed centrifugal pumps due to the low required thrust and volumetric flow rate and [...] Read more.

With the growth of the satellite industry, the demand for a propulsion system for small launch vehicles and spacecraft has increased. Small liquid rocket engines may require Ultra-Low specific speed centrifugal pumps due to the low required thrust and volumetric flow rate and high combustion chamber pressure. Therefore, in this study, a design method of Ultra-Low specific speed centrifugal pumps for several hundred Newton class small liquid rocket engines was developed by combining various empirical formulas. In addition, centrifugal pump impellers were designed using the Stepanoff method, which is typically used in pump design, and the circular arc method. The most appropriate method for designing Ultra-Low specific speed centrifugal pumps was determined through a comparative analysis with other methods and validated through CFD. As a result, the pump designed using the proposed method exhibited a performance of pumping and suction superior to the Stepanoff method. Although the number of arcs did not considerably influence the pump performance, the single arc method was confirmed to be the most appropriate design approach in terms of the design productivity and simplicity. Full article

(This article belongs to the Special Issue Liquid Rocket Engines)

► Show Figures

Figure 1

17 pages, 8183 KiB

Open AccessArticle

An Eddy Dissipation Concept Performance Study for Space Propulsion Applications

by Daniel Martinez-Sanchis, Andrej Sternin, Jaroslaw Shvab, Oskar Haidn and Xiangyu Hu

Aerospace 2022, 9(9), 476; https://doi.org/10.3390/aerospace9090476 - 27 Aug 2022

Cited by 3 | Viewed by 1388

Abstract

In this study, Direct Numerical Simulations (DNS) of a turbulent diffusion flame are conducted to investigate the performance of the Eddy Dissipation Concept in turbulent combustion for space propulsion applications. A 20-bar methane-oxygen diffusion flame is simulated to resemble the conditions encountered in [...] Read more.

In this study, Direct Numerical Simulations (DNS) of a turbulent diffusion flame are conducted to investigate the performance of the Eddy Dissipation Concept in turbulent combustion for space propulsion applications. A 20-bar methane-oxygen diffusion flame is simulated to resemble the conditions encountered in modern rocket combustors. The numerical simulations were conducted using the software EBI-DNS within the OpenFOAM framework. An approach for analysis and validation of the combustion model with DNS is developed. The EDC model presents a good agreement with DNS observations in the most prevalent species. Nevertheless, the EDC struggles to predict the mean chemical production rate of intermediate species. It is found that local adaption of the model constants is essential for maximizing the prediction capabilities. The relationship of these parameters with the Reynolds number and the Damköhler number are mostly in good agreement with the trends proposed in recent research . Full article

(This article belongs to the Special Issue Liquid Rocket Engines)

► Show Figures

Figure 1

16 pages, 1894 KiB

Open AccessArticle

Flame Characteristics and Response of a High-Pressure LOX/CNG Rocket Combustor with Large Optical Access

by Jan Martin, Wolfgang Armbruster, Dmitry Suslov, Robert Stützer, Justin S. Hardi and Michael Oschwald

Aerospace 2022, 9(8), 410; https://doi.org/10.3390/aerospace9080410 - 29 Jul 2022

Cited by 7 | Viewed by 1657

Abstract

Hot-fire tests were performed with a single-injector research combustor featuring a large optical access (255 × 38 mm) for flame imaging. These tests were conducted with the propellant combination of liquid oxygen and compressed natural gas (LOX/CNG) at conditions relevant for main- and [...] Read more.

Hot-fire tests were performed with a single-injector research combustor featuring a large optical access (255 × 38 mm) for flame imaging. These tests were conducted with the propellant combination of liquid oxygen and compressed natural gas (LOX/CNG) at conditions relevant for main- and upper-stage engines. The large optical access enabled synchronized flame imaging using OH* and CH* radiation wavelengths covering an area of the combustion chamber from the injection plane to shortly before the contraction section of the nozzle for two sets of operating conditions. Combined with temperature, pressure and unsteady pressure measurements, these data provide a high-quality basis for validation of numerical modeling. Flame width and opening angle were extracted from the imaging in order to determine the flame topology. A two dimensional Rayleigh Index was calculated for an acoustically unexcited and excited interval. These Rayleigh Indices are in good agreement with the thermoacoustic state of the chamber. Full article

(This article belongs to the Special Issue Liquid Rocket Engines)

► Show Figures

Figure 1

16 pages, 4057 KiB

Open AccessArticle

Fault Detection and Diagnosis for Liquid Rocket Engines Based on Long Short-Term Memory and Generative Adversarial Networks

by Lingzhi Deng, Yuqiang Cheng and Yehui Shi

Aerospace 2022, 9(8), 399; https://doi.org/10.3390/aerospace9080399 - 26 Jul 2022

Cited by 9 | Viewed by 2185

Abstract

The development of health monitoring technology for liquid rocket engines (LREs) can effectively improve the safety and reliability of launch vehicles, which has important theoretical and engineering significance. Therefore, we propose a fault detection and diagnosis (FDD) method for a large LOX/kerosene rocket [...] Read more.

The development of health monitoring technology for liquid rocket engines (LREs) can effectively improve the safety and reliability of launch vehicles, which has important theoretical and engineering significance. Therefore, we propose a fault detection and diagnosis (FDD) method for a large LOX/kerosene rocket engine based on long short-term memory (LSTM) and generative adversarial networks (GANs). Specifically, we first modeled a large LOX/kerosene rocket engine using MATLAB/Simulink and simulated the engine’s normal and fault operation states involving various startup and steady-state stages utilizing fault injection. Second, we created an LSTM-GAN model trained with normal operating data using LSTM as the generator and a multilayer perceptron (MLP) as the discriminator. Third, the test data were input into the discriminator to obtain the discrimination results and realize fault detection. Finally, the test data were input into the generator to obtain the predicted samples and calculate the absolute error between the predicted and the real value of each parameter. Then the fault diagnosis index, standardized absolute error (SAE), was constructed. SAE was analyzed to realize fault diagnosis. The simulated results highlight that the proposed method effectively detects faults in the startup and steady-state processes, and diagnoses the faults in the steady-state process without missing an alarm or being affected by false alarms. Compared with the conventional redline cut-off system (RCS), adaptive threshold algorithm (ATA), and support vector machine (SVM), the fault detection process of LSTM-GAN is more concise and more timely. Full article

(This article belongs to the Special Issue Liquid Rocket Engines)

► Show Figures

Figure 1

16 pages, 6009 KiB

Open AccessArticle

The Confirmation of Thermal Boundary Parameters in an Oxygen Kerosene Fuel-Rich Rocket Engine

by Xianggeng Wei, Zhongxu Yang, Shaohua Zhu, Zhixin Zhao, Jinying Ye and Oskar J. Haidn

Aerospace 2022, 9(7), 343; https://doi.org/10.3390/aerospace9070343 - 26 Jun 2022

Cited by 3 | Viewed by 1594

Abstract

The thermal environment is an important factor in the design of liquid rockets. In this paper, theoretical analysis, numerical simulation and experimental testing are conducted to study the boundary thermal characteristics of a GOX/kerosene liquid rocket motor with a total flow rate of [...] Read more.

The thermal environment is an important factor in the design of liquid rockets. In this paper, theoretical analysis, numerical simulation and experimental testing are conducted to study the boundary thermal characteristics of a GOX/kerosene liquid rocket motor with a total flow rate of 120 g/s and an oxygen-fuel ratio of 1:1. We measured the axial temperature in different positions in the combustor using thermocouples and the heat flux using a flux meter. We found that the heat flux at 182 mm increases by 6.8% when a carbon deposit exists. For the theoretical results, after correcting the thermal conductivity by the volume fraction of carbon deposition, the theoretical heat flux (1.11 MW/m², using the corrected thermal conductivity) and the numerical result (0.89 MW/m², considering the injectors) are similar to the experimental value (0.937 MW/m²). This study validates the accuracy of theoretical and simulation calculation in this case, and provides verification data for future numerical calculation, as well as data for setting gas temperature at the wall in the simulation of the gas phase. Full article

(This article belongs to the Special Issue Liquid Rocket Engines)

► Show Figures

Figure 1

17 pages, 31015 KiB

Open AccessArticle

Influence of Mass Flow Rate on the Atomization Characteristics of Screw Conveyor Swirl Injectors

by Xianggeng Wei, Yiming Feng, Jinying Ye, Na Li and Oskar J. Haidn

Aerospace 2022, 9(6), 293; https://doi.org/10.3390/aerospace9060293 - 27 May 2022

Cited by 4 | Viewed by 2175

Abstract

This study conducted cold flow experimental research on the influence of mass flow rate on the atomization characteristics of screw conveyer swirl injectors in an opening environment. The Phase Doppler Particle Analyzer (PDPA) and high-speed photography were utilized to obtain experimental data. The [...] Read more.

This study conducted cold flow experimental research on the influence of mass flow rate on the atomization characteristics of screw conveyer swirl injectors in an opening environment. The Phase Doppler Particle Analyzer (PDPA) and high-speed photography were utilized to obtain experimental data. The results showed that the mass flow rate greatly influenced the atomization establishment and working characteristics of the injectors. The design point selection of the injectors exerted significant influence on the flow range and the performances of the injectors in a steady-state operation. The Sauter mean diameter of the atomization field continued to decrease with the increase in the mass flow rate. As the distance to the injector exit increased, the Sauter mean diameter continued to decrease, and finally tended to be stable. The average particle diameter obtained by the current image-processing method was greater than that by PDPA; therefore, the image-processing method needs improvement. Full article

(This article belongs to the Special Issue Liquid Rocket Engines)

► Show Figures

Figure 1

Journal Menu

Journal Browser

Liquid Rocket Engines

Share This Special Issue

Special Issue Editor

Special Issue Information

Keywords

Published Papers (26 papers)

Research

Further Information

Guidelines

MDPI Initiatives

Follow MDPI