Forced motion simulations of an overexpanded subscale rocket nozzle were performed to investigate the transient mechanisms that lead to self-exciting fluid–structure interaction as observed in preceded studies. The pressure response to the deformation could be separated into two regions upstream and downstream the flow separation position. Within these regions the transient part of the pressure was analyzed using fast Fourier transform based on the method of generalized aerodynamic forces. The amplitude spectrum and phase shift distribution of the pressure response could be explained by superposition of three independently acting mechanisms: the inclination effect, the existence of a moving axial pressure wave, and intrinsic oscillations caused by the turbulence created by the strong shock system. Simplified simulation setups using a bent flat plate and a detailed unsteady simulation of the flow in the undeformed nozzle were analyzed to validate these assumptions.
{"title":"Mechanisms Contributing to the Dynamic Stability of a Flexible Subscale Rocket Nozzle","authors":"S. Jack, Michael Oschwald, Thino Eggers","doi":"10.2514/1.b39178","DOIUrl":"https://doi.org/10.2514/1.b39178","url":null,"abstract":"Forced motion simulations of an overexpanded subscale rocket nozzle were performed to investigate the transient mechanisms that lead to self-exciting fluid–structure interaction as observed in preceded studies. The pressure response to the deformation could be separated into two regions upstream and downstream the flow separation position. Within these regions the transient part of the pressure was analyzed using fast Fourier transform based on the method of generalized aerodynamic forces. The amplitude spectrum and phase shift distribution of the pressure response could be explained by superposition of three independently acting mechanisms: the inclination effect, the existence of a moving axial pressure wave, and intrinsic oscillations caused by the turbulence created by the strong shock system. Simplified simulation setups using a bent flat plate and a detailed unsteady simulation of the flow in the undeformed nozzle were analyzed to validate these assumptions.","PeriodicalId":16903,"journal":{"name":"Journal of Propulsion and Power","volume":null,"pages":null},"PeriodicalIF":1.9,"publicationDate":"2023-09-05","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"45530467","PeriodicalName":null,"FirstCategoryId":null,"ListUrlMain":null,"RegionNum":4,"RegionCategory":"工程技术","ArticlePicture":[],"TitleCN":null,"AbstractTextCN":null,"PMCID":"","EPubDate":null,"PubModel":null,"JCR":null,"JCRName":null,"Score":null,"Total":0}
A critical decay time (CDT) model is developed to predict the critical energy of direct detonation initiation in gaseous mixtures. It is based on the global initiation criterion that the energy deposit should allow the decaying shock speed to stay in a specific range below the Chapman–Jouguet (CJ) speed at least for a critical decay time. The speed range is estimated with the sub-CJ Zel’dovich–von Neumann–Döring (ZND) simulations. The critical decay time is calculated as the minimum time to reach unity Mach number in the sub-CJ ZND simulations. The lower-speed bound is taken as a characteristic extinction speed below (which means the lower-speed bound) which the direct initiation should fail. This speed is calibrated using one-dimensional simulations for [Formula: see text] mixtures. The calibrated CDT model is then applied to estimate the critical initiation energy with the point-blast theory. The model yields better agreement with experimental data for hydrogen-fueled mixtures such as [Formula: see text] and [Formula: see text] mixtures than the well-known critical decay rate model. For small hydrocarbon-fueled mixtures such as [Formula: see text], [Formula: see text], [Formula: see text], and [Formula: see text] mixtures, the predicted critical energies also agree well with experimental results. The CDT model provides an efficient tool to evaluate the detonability of fuel–oxidizer mixtures, which could be beneficial for ignition initiation in propulsion and power devices such as rotating detonation engines.
{"title":"Critical Decay Time Model for Direct Detonation Initiation Energy in Gaseous Mixtures","authors":"Yuen Liu, Qing Xie, Yuxuan Chen, Rémy Mével, Zhuyin Ren","doi":"10.2514/1.b39263","DOIUrl":"https://doi.org/10.2514/1.b39263","url":null,"abstract":"A critical decay time (CDT) model is developed to predict the critical energy of direct detonation initiation in gaseous mixtures. It is based on the global initiation criterion that the energy deposit should allow the decaying shock speed to stay in a specific range below the Chapman–Jouguet (CJ) speed at least for a critical decay time. The speed range is estimated with the sub-CJ Zel’dovich–von Neumann–Döring (ZND) simulations. The critical decay time is calculated as the minimum time to reach unity Mach number in the sub-CJ ZND simulations. The lower-speed bound is taken as a characteristic extinction speed below (which means the lower-speed bound) which the direct initiation should fail. This speed is calibrated using one-dimensional simulations for [Formula: see text] mixtures. The calibrated CDT model is then applied to estimate the critical initiation energy with the point-blast theory. The model yields better agreement with experimental data for hydrogen-fueled mixtures such as [Formula: see text] and [Formula: see text] mixtures than the well-known critical decay rate model. For small hydrocarbon-fueled mixtures such as [Formula: see text], [Formula: see text], [Formula: see text], and [Formula: see text] mixtures, the predicted critical energies also agree well with experimental results. The CDT model provides an efficient tool to evaluate the detonability of fuel–oxidizer mixtures, which could be beneficial for ignition initiation in propulsion and power devices such as rotating detonation engines.","PeriodicalId":16903,"journal":{"name":"Journal of Propulsion and Power","volume":null,"pages":null},"PeriodicalIF":1.9,"publicationDate":"2023-09-03","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"47210475","PeriodicalName":null,"FirstCategoryId":null,"ListUrlMain":null,"RegionNum":4,"RegionCategory":"工程技术","ArticlePicture":[],"TitleCN":null,"AbstractTextCN":null,"PMCID":"","EPubDate":null,"PubModel":null,"JCR":null,"JCRName":null,"Score":null,"Total":0}
Simone D’Alessandro, Maria Luisa Frezzotti, Bernardo Favini, Francesco Nasuti
Several test cases in the literature have shown that both transverse and longitudinal high-frequency combustion instability can be driven by the injector dynamics. In these cases, pressure oscillations result in fluctuations in propellant mass flow rate, which yields pulsing heat release. This fundamental mechanism is the focus of the present work, with the aim of including this effect in a quasi-1D nonlinear model of Euler equations suited to studies of longitudinal combustion instability. In particular, the injection dynamics is represented through a simplified formulation, which is the core of the proposed response function. The analysis also addresses the influence of combustion efficiency on the main characteristics of the resulting limit cycle (frequency and amplitude). The obtained model is tested comparing the quasi-1D simulations against the experimental data of the continuously variable resonance combustor available in the literature, considering three different geometrical configurations, with different lengths of the oxidizer post. The proposed formulation is capable of reasonably reproducing the unstable behavior, as well as providing a simple model that explains the mechanism that leads to a low average combustion efficiency during unstable operation.
{"title":"Driving Mechanisms in Low-Order Modeling of Longitudinal Combustion Instability","authors":"Simone D’Alessandro, Maria Luisa Frezzotti, Bernardo Favini, Francesco Nasuti","doi":"10.2514/1.b39048","DOIUrl":"https://doi.org/10.2514/1.b39048","url":null,"abstract":"Several test cases in the literature have shown that both transverse and longitudinal high-frequency combustion instability can be driven by the injector dynamics. In these cases, pressure oscillations result in fluctuations in propellant mass flow rate, which yields pulsing heat release. This fundamental mechanism is the focus of the present work, with the aim of including this effect in a quasi-1D nonlinear model of Euler equations suited to studies of longitudinal combustion instability. In particular, the injection dynamics is represented through a simplified formulation, which is the core of the proposed response function. The analysis also addresses the influence of combustion efficiency on the main characteristics of the resulting limit cycle (frequency and amplitude). The obtained model is tested comparing the quasi-1D simulations against the experimental data of the continuously variable resonance combustor available in the literature, considering three different geometrical configurations, with different lengths of the oxidizer post. The proposed formulation is capable of reasonably reproducing the unstable behavior, as well as providing a simple model that explains the mechanism that leads to a low average combustion efficiency during unstable operation.","PeriodicalId":16903,"journal":{"name":"Journal of Propulsion and Power","volume":null,"pages":null},"PeriodicalIF":0.0,"publicationDate":"2023-09-01","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"136034796","PeriodicalName":null,"FirstCategoryId":null,"ListUrlMain":null,"RegionNum":4,"RegionCategory":"工程技术","ArticlePicture":[],"TitleCN":null,"AbstractTextCN":null,"PMCID":"","EPubDate":null,"PubModel":null,"JCR":null,"JCRName":null,"Score":null,"Total":0}
Electrical propulsion has been identified as one of the key fields of future research within the aerospace sector. This paper aims to contribute to the ongoing development of small electric-powered ducted fan engines with a thrust in the range of a few hundred newtons. A special emphasis is placed on introducing contra-rotating fans to such engines. Recently, such an innovative contra-rotating engine has been developed and put into operation on a novel test bench. This study intends to take a closer look at the performance of the engine at design speed ratio by means of experimental and numerical investigations. Global performance parameters are analyzed to validate parallel numerical simulations. Results indicate that both experiment and simulation are in good agreement. Moreover, detailed flow studies of the spanwise distribution of relevant parameters are conducted, and results are discussed. From these results, the stall behavior of the stage is identified. Finally, a comprehensive performance map is established with the help of a number of numerical simulations, which reveal a good off-design performance behavior of the engine.
{"title":"Detailed Performance Studies on a Small Electric-Powered Contra-Rotating Ducted Fan Engine","authors":"Tobias Ebus, M. Dietz, A. Hupfer","doi":"10.2514/1.b39046","DOIUrl":"https://doi.org/10.2514/1.b39046","url":null,"abstract":"Electrical propulsion has been identified as one of the key fields of future research within the aerospace sector. This paper aims to contribute to the ongoing development of small electric-powered ducted fan engines with a thrust in the range of a few hundred newtons. A special emphasis is placed on introducing contra-rotating fans to such engines. Recently, such an innovative contra-rotating engine has been developed and put into operation on a novel test bench. This study intends to take a closer look at the performance of the engine at design speed ratio by means of experimental and numerical investigations. Global performance parameters are analyzed to validate parallel numerical simulations. Results indicate that both experiment and simulation are in good agreement. Moreover, detailed flow studies of the spanwise distribution of relevant parameters are conducted, and results are discussed. From these results, the stall behavior of the stage is identified. Finally, a comprehensive performance map is established with the help of a number of numerical simulations, which reveal a good off-design performance behavior of the engine.","PeriodicalId":16903,"journal":{"name":"Journal of Propulsion and Power","volume":null,"pages":null},"PeriodicalIF":1.9,"publicationDate":"2023-08-25","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"44788649","PeriodicalName":null,"FirstCategoryId":null,"ListUrlMain":null,"RegionNum":4,"RegionCategory":"工程技术","ArticlePicture":[],"TitleCN":null,"AbstractTextCN":null,"PMCID":"","EPubDate":null,"PubModel":null,"JCR":null,"JCRName":null,"Score":null,"Total":0}
In this paper, we complete a full-thrust audit of an iodine-based gridded ion thruster. Prior results have demonstrated excellent agreement between indirect and direct laboratory thrust estimates. Here, thrust estimates from numerical modeling, indirect laboratory testing from diagnostic probes and propulsion system telemetry, indirect in-space testing from onboard propulsion system telemetry, and direct in-space testing by analyzing orbital maneuvers are compared to demonstrate consistency between the four methods and complete the thrust audit. Results from recent in-space testing of the iodine-based thruster demonstrate that thrust estimates from all four methods agree to within three standard deviations of uncertainty for the 11 maneuvers studied. This thrust audit represents a critical step toward improving the understanding and technological maturity of iodine-based gridded ion thrusters for future mission applications, and it demonstrates the utility of recently developed in-space thrust inference techniques for analyzing low-thrust maneuvers.
{"title":"Iodine Electric Propulsion System Thrust Validation: From Numerical Modeling to In-Space Testing","authors":"Oliver Jia-Richards, T. Lafleur","doi":"10.2514/1.b39198","DOIUrl":"https://doi.org/10.2514/1.b39198","url":null,"abstract":"In this paper, we complete a full-thrust audit of an iodine-based gridded ion thruster. Prior results have demonstrated excellent agreement between indirect and direct laboratory thrust estimates. Here, thrust estimates from numerical modeling, indirect laboratory testing from diagnostic probes and propulsion system telemetry, indirect in-space testing from onboard propulsion system telemetry, and direct in-space testing by analyzing orbital maneuvers are compared to demonstrate consistency between the four methods and complete the thrust audit. Results from recent in-space testing of the iodine-based thruster demonstrate that thrust estimates from all four methods agree to within three standard deviations of uncertainty for the 11 maneuvers studied. This thrust audit represents a critical step toward improving the understanding and technological maturity of iodine-based gridded ion thrusters for future mission applications, and it demonstrates the utility of recently developed in-space thrust inference techniques for analyzing low-thrust maneuvers.","PeriodicalId":16903,"journal":{"name":"Journal of Propulsion and Power","volume":null,"pages":null},"PeriodicalIF":1.9,"publicationDate":"2023-08-24","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"49336905","PeriodicalName":null,"FirstCategoryId":null,"ListUrlMain":null,"RegionNum":4,"RegionCategory":"工程技术","ArticlePicture":[],"TitleCN":null,"AbstractTextCN":null,"PMCID":"","EPubDate":null,"PubModel":null,"JCR":null,"JCRName":null,"Score":null,"Total":0}
Yihao Tang, Joo-Heon Kim, Brandon A. Sforzo, R. Scarcelli, V. Raman
The spark discharge of an aircraft plasma jet igniter is studied using high-fidelity numerical simulations and X-ray radiography measurements. The target problem here features the thermal expansion of hot gas introduced by the electric spark within a confined igniter cavity, which eventually evolves into a pulsed jet of a high-temperature kernel. A comprehensive set of models adapted from existing strategies for internal combustion engine spark plug discharge is extended to the target problem, including the modeling of energy deposition, plasma reactions, thermodynamic properties, and heat losses. A series of validation and parameter studies are performed and presented. The kernel size is found to be sensitive to heat losses arising from radiation and hot gas remained within the discharge cavity, rather than heat conduction to the wall in the discharge cavity. Depending on the enforced shape of the post-breakdown electric arc, the spark kernel can be off-centered, tilted, and considerably asymmetric. These features have been previously not considered when studying such igniter configurations and may have a first-order impact on the ignition process. Provided a proper setup of the heat loss models and electric arc shape, the numerical results are quantitatively comparable to the experimental results in terms of the kernel size, shape, and velocity throughout different stages after the spark discharge.
{"title":"Numerical and Experimental Study of an Aircraft Igniter Plasma Jet Discharge","authors":"Yihao Tang, Joo-Heon Kim, Brandon A. Sforzo, R. Scarcelli, V. Raman","doi":"10.2514/1.b38815","DOIUrl":"https://doi.org/10.2514/1.b38815","url":null,"abstract":"The spark discharge of an aircraft plasma jet igniter is studied using high-fidelity numerical simulations and X-ray radiography measurements. The target problem here features the thermal expansion of hot gas introduced by the electric spark within a confined igniter cavity, which eventually evolves into a pulsed jet of a high-temperature kernel. A comprehensive set of models adapted from existing strategies for internal combustion engine spark plug discharge is extended to the target problem, including the modeling of energy deposition, plasma reactions, thermodynamic properties, and heat losses. A series of validation and parameter studies are performed and presented. The kernel size is found to be sensitive to heat losses arising from radiation and hot gas remained within the discharge cavity, rather than heat conduction to the wall in the discharge cavity. Depending on the enforced shape of the post-breakdown electric arc, the spark kernel can be off-centered, tilted, and considerably asymmetric. These features have been previously not considered when studying such igniter configurations and may have a first-order impact on the ignition process. Provided a proper setup of the heat loss models and electric arc shape, the numerical results are quantitatively comparable to the experimental results in terms of the kernel size, shape, and velocity throughout different stages after the spark discharge.","PeriodicalId":16903,"journal":{"name":"Journal of Propulsion and Power","volume":null,"pages":null},"PeriodicalIF":1.9,"publicationDate":"2023-08-24","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"43605557","PeriodicalName":null,"FirstCategoryId":null,"ListUrlMain":null,"RegionNum":4,"RegionCategory":"工程技术","ArticlePicture":[],"TitleCN":null,"AbstractTextCN":null,"PMCID":"","EPubDate":null,"PubModel":null,"JCR":null,"JCRName":null,"Score":null,"Total":0}
An experimental study was conducted to visualize the dynamics and spectra of single-element shear coaxial liquid oxygen (LOX)/gaseous methane flames at high but still subcritical pressure for the LOX, and at supercritical temperature for the methane. High-speed shadowgraphs were employed to image and track the flame and jet dynamics close to the liquid oxygen’s post region. High-speed [Formula: see text] and [Formula: see text] chemiluminescence were also simultaneously collected along the same line of sight as the shadowgraphs. The results were compared to previous LOX/hydrogen results using the same injector at the same momentum flux ratios. The initial temperature of the methane was varied between 200 and 300 K. The LOX/methane flames were found to share many features in common with the LOX/hydrogen flames previously studied, but there were notable differences. The spreading rate of the LOX/hydrogen flames was larger than that of LOX/methane flames. The amplitude of the spreading rate fluctuations for hydrogen was also larger, although the relative spreading rate fluctuations normalized by the average spreading rates were about the same. Despite the differences, a wave amplification mechanism found previously to be active for shear coaxial LOX/hydrogen flames was found to also be active for LOX/methane flames. Overall, both the LOX/methane flames and the LOX/hydrogen were both found to be spectrally fairly quiet.
{"title":"Liquid Oxygen/Gaseous Methane Single-Element Shear Coaxial Flame Dynamics","authors":"M. Roa, D. Talley, R. Munipalli","doi":"10.2514/1.b39163","DOIUrl":"https://doi.org/10.2514/1.b39163","url":null,"abstract":"An experimental study was conducted to visualize the dynamics and spectra of single-element shear coaxial liquid oxygen (LOX)/gaseous methane flames at high but still subcritical pressure for the LOX, and at supercritical temperature for the methane. High-speed shadowgraphs were employed to image and track the flame and jet dynamics close to the liquid oxygen’s post region. High-speed [Formula: see text] and [Formula: see text] chemiluminescence were also simultaneously collected along the same line of sight as the shadowgraphs. The results were compared to previous LOX/hydrogen results using the same injector at the same momentum flux ratios. The initial temperature of the methane was varied between 200 and 300 K. The LOX/methane flames were found to share many features in common with the LOX/hydrogen flames previously studied, but there were notable differences. The spreading rate of the LOX/hydrogen flames was larger than that of LOX/methane flames. The amplitude of the spreading rate fluctuations for hydrogen was also larger, although the relative spreading rate fluctuations normalized by the average spreading rates were about the same. Despite the differences, a wave amplification mechanism found previously to be active for shear coaxial LOX/hydrogen flames was found to also be active for LOX/methane flames. Overall, both the LOX/methane flames and the LOX/hydrogen were both found to be spectrally fairly quiet.","PeriodicalId":16903,"journal":{"name":"Journal of Propulsion and Power","volume":null,"pages":null},"PeriodicalIF":1.9,"publicationDate":"2023-08-11","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"43251059","PeriodicalName":null,"FirstCategoryId":null,"ListUrlMain":null,"RegionNum":4,"RegionCategory":"工程技术","ArticlePicture":[],"TitleCN":null,"AbstractTextCN":null,"PMCID":"","EPubDate":null,"PubModel":null,"JCR":null,"JCRName":null,"Score":null,"Total":0}
Boundary-layer ingested engines have the potential to offer significantly reduced fuel burn, but the fan stage must be designed to run efficiently with a distorted inflow. It must also be able to withstand unsteady aerodynamic loads resulting from a complex nonuniform flowfield. This paper applies different numerical methods for an improved understanding of the aerodynamic interaction between a transonic fan and inlet distortion. A single-stage transonic tail cone thruster fan was designed using both in-house and commercial tools operating in an inlet distortion flowfield. This paper demonstrates that the relevant metrics required to compute the aerodynamic performance of a fan stage in distorted conditions can be reasonably modeled with a few harmonics using the nonlinear harmonic method in a fraction of time in comparison to a full annulus time marching solution. The nonlinear harmonic method also reduces the computational domain, and hence reduces the solution runtime by an order of magnitude. However, it fails to accurately resolve the wake and potential field transfer across the blade rows due to a limited number of harmonics being applied. A detailed aerodynamic description of the unsteady inflow distortion, the interacting blade-row mechanisms, the flow redistribution upstream of the rotor, the distortion transfer across the different blade rows, and the corresponding aerodynamic losses can be analyzed accurately using only a full annulus time-marching method.
{"title":"Numerical Modeling to Investigate the Aerodynamics of a Boundary-Layer Ingested Transonic Fan","authors":"Ritangshu Giri, M. Turner","doi":"10.2514/1.b38963","DOIUrl":"https://doi.org/10.2514/1.b38963","url":null,"abstract":"Boundary-layer ingested engines have the potential to offer significantly reduced fuel burn, but the fan stage must be designed to run efficiently with a distorted inflow. It must also be able to withstand unsteady aerodynamic loads resulting from a complex nonuniform flowfield. This paper applies different numerical methods for an improved understanding of the aerodynamic interaction between a transonic fan and inlet distortion. A single-stage transonic tail cone thruster fan was designed using both in-house and commercial tools operating in an inlet distortion flowfield. This paper demonstrates that the relevant metrics required to compute the aerodynamic performance of a fan stage in distorted conditions can be reasonably modeled with a few harmonics using the nonlinear harmonic method in a fraction of time in comparison to a full annulus time marching solution. The nonlinear harmonic method also reduces the computational domain, and hence reduces the solution runtime by an order of magnitude. However, it fails to accurately resolve the wake and potential field transfer across the blade rows due to a limited number of harmonics being applied. A detailed aerodynamic description of the unsteady inflow distortion, the interacting blade-row mechanisms, the flow redistribution upstream of the rotor, the distortion transfer across the different blade rows, and the corresponding aerodynamic losses can be analyzed accurately using only a full annulus time-marching method.","PeriodicalId":16903,"journal":{"name":"Journal of Propulsion and Power","volume":null,"pages":null},"PeriodicalIF":1.9,"publicationDate":"2023-08-08","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"45723674","PeriodicalName":null,"FirstCategoryId":null,"ListUrlMain":null,"RegionNum":4,"RegionCategory":"工程技术","ArticlePicture":[],"TitleCN":null,"AbstractTextCN":null,"PMCID":"","EPubDate":null,"PubModel":null,"JCR":null,"JCRName":null,"Score":null,"Total":0}
R. Murugesan, S. Chakravarthy, J. Kandasamy, R. Sarathi
Major challenges in developing and realizing a novel aluminum–water reaction-based water ramjet propulsion system for high-speed underwater vehicles and demonstration of a water-breathing jet propulsion test facility are investigated. Two stages of combustion, propellant grain combustion and subsequent water combustion, with primary combustion products are adopted. High-pressure-molded propellant grains up to 45% of micro–nano ([Formula: see text]) aluminum were prepared and combusted in the primary chamber, which exhibits mild ignition delay, and a residue of 4–6% was retained. Once water is injected into the secondary chamber, the net thrust generation is increased more than twice from the exhaust jet and improves the specific impulse by 40%. The lean fuel conditions in the secondary chamber lead to reduction in combustion propensity, which causes drop in [Formula: see text] efficiency. The ultrafine iron-oxide-catalyzed micro–nano blended propellants marginally improved the propulsive performance than the uncatalyzed compositions. The [Formula: see text] efficiency of the catalyzed propellants was enhanced up to 38.6%. Aluminum agglomeration in primary combustion considerably occurred; apparently, only a fraction of aluminum particles or agglomerates are completely burnt within the secondary chamber, and the remaining aluminum particles are either partially burnt or go unreacted.
{"title":"Experimental Investigation on Aluminum-Based Water Ramjet for Propelling High-Speed Underwater Vehicles","authors":"R. Murugesan, S. Chakravarthy, J. Kandasamy, R. Sarathi","doi":"10.2514/1.b39133","DOIUrl":"https://doi.org/10.2514/1.b39133","url":null,"abstract":"Major challenges in developing and realizing a novel aluminum–water reaction-based water ramjet propulsion system for high-speed underwater vehicles and demonstration of a water-breathing jet propulsion test facility are investigated. Two stages of combustion, propellant grain combustion and subsequent water combustion, with primary combustion products are adopted. High-pressure-molded propellant grains up to 45% of micro–nano ([Formula: see text]) aluminum were prepared and combusted in the primary chamber, which exhibits mild ignition delay, and a residue of 4–6% was retained. Once water is injected into the secondary chamber, the net thrust generation is increased more than twice from the exhaust jet and improves the specific impulse by 40%. The lean fuel conditions in the secondary chamber lead to reduction in combustion propensity, which causes drop in [Formula: see text] efficiency. The ultrafine iron-oxide-catalyzed micro–nano blended propellants marginally improved the propulsive performance than the uncatalyzed compositions. The [Formula: see text] efficiency of the catalyzed propellants was enhanced up to 38.6%. Aluminum agglomeration in primary combustion considerably occurred; apparently, only a fraction of aluminum particles or agglomerates are completely burnt within the secondary chamber, and the remaining aluminum particles are either partially burnt or go unreacted.","PeriodicalId":16903,"journal":{"name":"Journal of Propulsion and Power","volume":null,"pages":null},"PeriodicalIF":1.9,"publicationDate":"2023-08-08","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"45551978","PeriodicalName":null,"FirstCategoryId":null,"ListUrlMain":null,"RegionNum":4,"RegionCategory":"工程技术","ArticlePicture":[],"TitleCN":null,"AbstractTextCN":null,"PMCID":"","EPubDate":null,"PubModel":null,"JCR":null,"JCRName":null,"Score":null,"Total":0}
Brian Legros, L. Léger, A. Kourta, A. Chpoun, M. Sellam
The dual-bell nozzle (DBN) is a rocket nozzle concept that could provide a substantial payload gain. The present paper focuses on the impact of radial secondary injection on DBN behavior during its ascent and descent in the atmosphere. The influence of the secondary injection settling chamber (cavity) volume is discussed. The positive influence of the cavity volume on the DBN behavior during transition phases has been exposed, and its effects during retransition were shown to be negligible. The use of secondary injection allowed the transition nozzle pressure ratio to be increased by nearly 24%, and the lateral forces were reduced to less than 1% of the nozzle thrust. A first approach regarding secondary injection optimization is finally proposed.
{"title":"Parametrical Investigation of Transverse Injection in a Dual-Bell Nozzle During Altitude-Varying Conditions","authors":"Brian Legros, L. Léger, A. Kourta, A. Chpoun, M. Sellam","doi":"10.2514/1.b39077","DOIUrl":"https://doi.org/10.2514/1.b39077","url":null,"abstract":"The dual-bell nozzle (DBN) is a rocket nozzle concept that could provide a substantial payload gain. The present paper focuses on the impact of radial secondary injection on DBN behavior during its ascent and descent in the atmosphere. The influence of the secondary injection settling chamber (cavity) volume is discussed. The positive influence of the cavity volume on the DBN behavior during transition phases has been exposed, and its effects during retransition were shown to be negligible. The use of secondary injection allowed the transition nozzle pressure ratio to be increased by nearly 24%, and the lateral forces were reduced to less than 1% of the nozzle thrust. A first approach regarding secondary injection optimization is finally proposed.","PeriodicalId":16903,"journal":{"name":"Journal of Propulsion and Power","volume":null,"pages":null},"PeriodicalIF":1.9,"publicationDate":"2023-08-02","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"42061567","PeriodicalName":null,"FirstCategoryId":null,"ListUrlMain":null,"RegionNum":4,"RegionCategory":"工程技术","ArticlePicture":[],"TitleCN":null,"AbstractTextCN":null,"PMCID":"","EPubDate":null,"PubModel":null,"JCR":null,"JCRName":null,"Score":null,"Total":0}
京公网安备 11010802042870号
京ICP备2023020795号-1
扫码关注我们
求助内容：
应助结果提醒方式：