Although supersonic combustion ramjets—scramjets—provide a fuel-efficient method for propulsion at hypersonic speeds, current challenges with the engine prohibit the robustness necessary for space accessibility and trans-atmospheric flight. One such challenge the engine faces is the vehicle and inlet’s compliance under harsh thermal and mechanical loads at hypersonic speeds. The deformation of the inlet has ramifications on the downstream components and the engine as a whole, creating conditions outside of the original design envelope. Additionally, the deformations impact the vehicle’s aerodynamic performance due to the integrated airframe/inlet design. One mitigation technique that works in tandem with thermal management is active cooling. It is important to understand the impacts of active cooling on the inlet and engine performance; in order to do so, a multiphysics modeling approach is used to capture the coupled aerothermostructural response of the inlet, and a multifidelity approach is used to model the remaining components of the scramjet. The system is found to be extremely sensitive to the changes in deformation, leading to increased flow separation and heating and to deviations of the engine performance and efficiency from the original design point.
{"title":"Fully Coupled Analysis of Aerothermoelastic Deformation of a Scramjet Inlet","authors":"Jennifer A. Horing, Iain D. Boyd, Kurt K. Maute","doi":"10.2514/1.b39345","DOIUrl":"https://doi.org/10.2514/1.b39345","url":null,"abstract":"Although supersonic combustion ramjets—scramjets—provide a fuel-efficient method for propulsion at hypersonic speeds, current challenges with the engine prohibit the robustness necessary for space accessibility and trans-atmospheric flight. One such challenge the engine faces is the vehicle and inlet’s compliance under harsh thermal and mechanical loads at hypersonic speeds. The deformation of the inlet has ramifications on the downstream components and the engine as a whole, creating conditions outside of the original design envelope. Additionally, the deformations impact the vehicle’s aerodynamic performance due to the integrated airframe/inlet design. One mitigation technique that works in tandem with thermal management is active cooling. It is important to understand the impacts of active cooling on the inlet and engine performance; in order to do so, a multiphysics modeling approach is used to capture the coupled aerothermostructural response of the inlet, and a multifidelity approach is used to model the remaining components of the scramjet. The system is found to be extremely sensitive to the changes in deformation, leading to increased flow separation and heating and to deviations of the engine performance and efficiency from the original design point.","PeriodicalId":16903,"journal":{"name":"Journal of Propulsion and Power","volume":"84 1 1","pages":"0"},"PeriodicalIF":0.0,"publicationDate":"2023-10-18","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"135883103","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}
Rafid Bendimerad, Abu Taqui Md Tahsin, Adam Yonas, Caleb Colucci, Elaine M. Petro
Electrospray thrusters fulfill the main propulsion requirements for long-term small-satellite missions. However, the molecules present in the plume are susceptible to collisions, chemical reactions, and fragmentation, which may introduce different new species with various mass-to-charge ratios inside the plume. Prediction of the byproducts that appear upon collisions is of prime importance to predicting the evolution of the plume and estimating the performance and the lifetime expectancy of the thruster. In this work, we use molecular dynamics simulations to investigate monomer–neutral collisions at different impact configurations, impact energies, and impact parameters, and we provide the mass spectra of the resulting species. We predict that 1) collisions within a center-of-mass distance of 6 Å can result in momentum exchange and molecular fragmentation, 2) higher-energy impacts produce more byproducts, and 3) heavy molecules (e.g., 1-ethyl-3-methylimidazolium [EMI] and [Formula: see text]) are more likely to result from weak collisions ([Formula: see text]), whereas light molecules (e.g., H, F, and [Formula: see text]) are more likely to result from strong collisions. Collisional fragmentation is shown to negatively affect key performance indicators, including reductions in thrust, specific impulse, and propulsive efficiency. This phenomenon potentially accounts for the observed discrepancies in experimental measurements of current and mass loss rates.
{"title":"Investigating the Chemical Stability of Electrospray Plumes During Particle Collisions","authors":"Rafid Bendimerad, Abu Taqui Md Tahsin, Adam Yonas, Caleb Colucci, Elaine M. Petro","doi":"10.2514/1.b39118","DOIUrl":"https://doi.org/10.2514/1.b39118","url":null,"abstract":"Electrospray thrusters fulfill the main propulsion requirements for long-term small-satellite missions. However, the molecules present in the plume are susceptible to collisions, chemical reactions, and fragmentation, which may introduce different new species with various mass-to-charge ratios inside the plume. Prediction of the byproducts that appear upon collisions is of prime importance to predicting the evolution of the plume and estimating the performance and the lifetime expectancy of the thruster. In this work, we use molecular dynamics simulations to investigate monomer–neutral collisions at different impact configurations, impact energies, and impact parameters, and we provide the mass spectra of the resulting species. We predict that 1) collisions within a center-of-mass distance of 6 Å can result in momentum exchange and molecular fragmentation, 2) higher-energy impacts produce more byproducts, and 3) heavy molecules (e.g., 1-ethyl-3-methylimidazolium [EMI] and [Formula: see text]) are more likely to result from weak collisions ([Formula: see text]), whereas light molecules (e.g., H, F, and [Formula: see text]) are more likely to result from strong collisions. Collisional fragmentation is shown to negatively affect key performance indicators, including reductions in thrust, specific impulse, and propulsive efficiency. This phenomenon potentially accounts for the observed discrepancies in experimental measurements of current and mass loss rates.","PeriodicalId":16903,"journal":{"name":"Journal of Propulsion and Power","volume":"17 1","pages":"0"},"PeriodicalIF":0.0,"publicationDate":"2023-10-11","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"136098156","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}
The objective of this work was to assess the unstart reliability of the Hypersonic International Flight Research Experimentation Flight 2 system. To do this, a quantification of margins and uncertainties framework was used for comparing the predicted combustion-induced shock location to the predicted last stable shock location within the isolator. Uncertainty sources included parametric uncertainty in the flight conditions, the heat release model, and turbulence modeling, as well as model verification errors. Additionally, an estimate of the model-form uncertainty was established by comparing the model to measured ground-test data. A computationally efficient nonintrusive polynomial chaos approach was used to propagate parametric uncertainty through the computational fluids dynamics models of both the ground-test configuration and the flight vehicle. Compared to direct-connect ground-test data, computational fluid dynamics predictions yielded about two duct heights of model-form uncertainty. This was applied to a prediction of the flight vehicle unstart margin at the Mach 6.5 flight condition. Building up all of the computational model uncertainty (including parametric uncertainty, verification errors, and the determined model-form uncertainty), the 95%-probability-level-based confidence ratio, which is a ratio of a statistical margin measure to the total uncertainty, was found to be 0.31 for the flight system.
{"title":"Hypersonic International Flight Research Experimentation Flight 2 Unstart Reliability Analysis","authors":"Thomas K. West, Michael D. Bynum","doi":"10.2514/1.b39108","DOIUrl":"https://doi.org/10.2514/1.b39108","url":null,"abstract":"The objective of this work was to assess the unstart reliability of the Hypersonic International Flight Research Experimentation Flight 2 system. To do this, a quantification of margins and uncertainties framework was used for comparing the predicted combustion-induced shock location to the predicted last stable shock location within the isolator. Uncertainty sources included parametric uncertainty in the flight conditions, the heat release model, and turbulence modeling, as well as model verification errors. Additionally, an estimate of the model-form uncertainty was established by comparing the model to measured ground-test data. A computationally efficient nonintrusive polynomial chaos approach was used to propagate parametric uncertainty through the computational fluids dynamics models of both the ground-test configuration and the flight vehicle. Compared to direct-connect ground-test data, computational fluid dynamics predictions yielded about two duct heights of model-form uncertainty. This was applied to a prediction of the flight vehicle unstart margin at the Mach 6.5 flight condition. Building up all of the computational model uncertainty (including parametric uncertainty, verification errors, and the determined model-form uncertainty), the 95%-probability-level-based confidence ratio, which is a ratio of a statistical margin measure to the total uncertainty, was found to be 0.31 for the flight system.","PeriodicalId":16903,"journal":{"name":"Journal of Propulsion and Power","volume":"44 1","pages":"0"},"PeriodicalIF":0.0,"publicationDate":"2023-10-11","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"136211401","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}
Gridded ion thrusters are tested in ground vacuum chambers to verify their performance when deployed in space. However, the presence of high background pressure and conductive walls in the chamber leads to facility effects that increase uncertainty in the performance of the thruster in space. To address this issue, this study utilizes a fully kinetic simulation to investigate the facility effects on the thruster plume. The in-chamber condition shows a downstream neutral particle density 100 times larger than the in-space case due to ion neutralization at the wall and limited vacuum pump capability, resulting in a significant difference in the density and distribution of charge-exchange ions. The flux, energy, and angle of charge-exchange ions incident on the chamber wall are found to be altered by the electron sheath, which can only be simulated by the fully kinetic approach, as opposed to the conventionally used quasi-neutral Boltzmann approach. We also examine the effect of backsputtering, another important facility effect, and find that it does not necessarily require a fully kinetic simulation as the incident flux and energy of the sampled charge-exchange ion are negligibly small. Finally, we demonstrate that the carbon deposition rate on the thruster is significantly influenced by the angular dependence of the sputtered carbon, with a nearly 50% effect.
{"title":"Three-Dimensional Kinetic Simulations of Carbon Backsputtering in Vacuum Chambers from Ion Thruster Plumes","authors":"Keita Nishii, Deborah A. Levin","doi":"10.2514/1.b39194","DOIUrl":"https://doi.org/10.2514/1.b39194","url":null,"abstract":"Gridded ion thrusters are tested in ground vacuum chambers to verify their performance when deployed in space. However, the presence of high background pressure and conductive walls in the chamber leads to facility effects that increase uncertainty in the performance of the thruster in space. To address this issue, this study utilizes a fully kinetic simulation to investigate the facility effects on the thruster plume. The in-chamber condition shows a downstream neutral particle density 100 times larger than the in-space case due to ion neutralization at the wall and limited vacuum pump capability, resulting in a significant difference in the density and distribution of charge-exchange ions. The flux, energy, and angle of charge-exchange ions incident on the chamber wall are found to be altered by the electron sheath, which can only be simulated by the fully kinetic approach, as opposed to the conventionally used quasi-neutral Boltzmann approach. We also examine the effect of backsputtering, another important facility effect, and find that it does not necessarily require a fully kinetic simulation as the incident flux and energy of the sampled charge-exchange ion are negligibly small. Finally, we demonstrate that the carbon deposition rate on the thruster is significantly influenced by the angular dependence of the sputtered carbon, with a nearly 50% effect.","PeriodicalId":16903,"journal":{"name":"Journal of Propulsion and Power","volume":"20 1","pages":"0"},"PeriodicalIF":0.0,"publicationDate":"2023-10-04","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"135592300","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}
Porous bleed systems are a common technique to control shock-/boundary-layer interactions and/or supersonic boundary layers. However, the influence of various design parameters is still unknown. Even though porous bleed models are required to minimize the costs of the design process, they often do not include parameter effects. In the present study, the effect of the plate length, the hole diameter, the porosity level, the thickness-to-diameter ratio, and the stagger angle are investigated by means of three-dimensional Reynolds-averaged Navier–Stokes simulations. The bleed efficiency and the effectiveness in thinning a Mach [Formula: see text] turbulent boundary layer are determined. The findings show a crucial influence of the hole diameter on both the efficiency and effectiveness of the porous bleed. Similar findings are made for the porosity and stagger angle but with a smaller significance. The thickness-to-diameter ratio and plate length are shown to mainly affect the bleed efficiency.
{"title":"Parameter Influence on Porous Bleed Performance for Supersonic Turbulent Flows","authors":"Julian Giehler, Pierre Grenson, Reynald Bur","doi":"10.2514/1.b39236","DOIUrl":"https://doi.org/10.2514/1.b39236","url":null,"abstract":"Porous bleed systems are a common technique to control shock-/boundary-layer interactions and/or supersonic boundary layers. However, the influence of various design parameters is still unknown. Even though porous bleed models are required to minimize the costs of the design process, they often do not include parameter effects. In the present study, the effect of the plate length, the hole diameter, the porosity level, the thickness-to-diameter ratio, and the stagger angle are investigated by means of three-dimensional Reynolds-averaged Navier–Stokes simulations. The bleed efficiency and the effectiveness in thinning a Mach [Formula: see text] turbulent boundary layer are determined. The findings show a crucial influence of the hole diameter on both the efficiency and effectiveness of the porous bleed. Similar findings are made for the porosity and stagger angle but with a smaller significance. The thickness-to-diameter ratio and plate length are shown to mainly affect the bleed efficiency.","PeriodicalId":16903,"journal":{"name":"Journal of Propulsion and Power","volume":"67 1","pages":"0"},"PeriodicalIF":0.0,"publicationDate":"2023-09-25","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"135816460","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}
S. She-Ming Lau-Chapdelaine, Matei I. Radulescu, Zekai Hong
A numerical simulation of an annular rotating detonation engine with stoichiometric hydrogen–oxygen is performed. A generic, well-posed, and easily implemented approach using a quasi-two-dimensional method to model the area variations through the rotating detonation engine’s injector and combustor is presented. The detonation–injector interaction is studied for the case with a ratio of four between the combustor and injector’s throat areas. A shock wave is formed in the divergent portion of the injector due to the high backpressure created by the detonation in the combustor. A Favre-averaged steady-state analysis of stream lines and particle paths reveals that the shock causes an irrecoverable loss of stagnation pressure. Stagnation pressure gain in the combustor is insufficient to make up for the loss, and the flow leaves the engine with lower stagnation pressure than in the plenum.
{"title":"Quasi-Two-Dimensional Simulation of a Rotating Detonation Engine Combustor and Injector","authors":"S. She-Ming Lau-Chapdelaine, Matei I. Radulescu, Zekai Hong","doi":"10.2514/1.b39214","DOIUrl":"https://doi.org/10.2514/1.b39214","url":null,"abstract":"A numerical simulation of an annular rotating detonation engine with stoichiometric hydrogen–oxygen is performed. A generic, well-posed, and easily implemented approach using a quasi-two-dimensional method to model the area variations through the rotating detonation engine’s injector and combustor is presented. The detonation–injector interaction is studied for the case with a ratio of four between the combustor and injector’s throat areas. A shock wave is formed in the divergent portion of the injector due to the high backpressure created by the detonation in the combustor. A Favre-averaged steady-state analysis of stream lines and particle paths reveals that the shock causes an irrecoverable loss of stagnation pressure. Stagnation pressure gain in the combustor is insufficient to make up for the loss, and the flow leaves the engine with lower stagnation pressure than in the plenum.","PeriodicalId":16903,"journal":{"name":"Journal of Propulsion and Power","volume":"44 1","pages":"0"},"PeriodicalIF":0.0,"publicationDate":"2023-09-25","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"135768441","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}
Derek A. Nichols, Bojan Vukasinovic, Ari Glezer, Bradley Rafferty
The flow within the inlet of an engine nacelle model in the absence of a fan and the presence of crosswind is investigated in wind-tunnel experiments, with specific emphasis on the effects of separation over the inlet’s inner windward surface on the flow distortion and pressure recovery. The inlet’s entrance plane is tilted forward, and its cross section is asymmetric about the horizontal centerline. The flow topology within the inlet is characterized over a range of Mach numbers and crosswind speeds up to [Formula: see text] and [Formula: see text], respectively. It is shown that in the presence of sufficiently high crosswind to the inlet speed ratio, a three-dimensional horseshoe-like separation domain is formed over the inlet’s inner windward surface. Owing to the cross-sectional asymmetry of the entrance plane, the separation domain migrates azimuthally downward and expands azimuthally with increased crosswind to the inlet speed ratio. The present investigations demonstrate the utility of flow control for mitigating the adverse effects of the separation. The actuation is based on controllable distributed aerodynamic air bleed that is driven by the pressure differences across the nacelle’s inner and outer surfaces and reattaches the separated base flow up to crosswind speeds of [Formula: see text], resulting in a gain of up to 38% in total pressure recovery and a decrease of up to 55% in total pressure distortion. The efficacy of the bleed actuation can be further improved by tailoring the bleed distribution to the topology of the separated flow domain.
{"title":"Aerodynamic Control of an Inlet Flow in Crosswind Using Peripheral Bleed Actuation","authors":"Derek A. Nichols, Bojan Vukasinovic, Ari Glezer, Bradley Rafferty","doi":"10.2514/1.b38944","DOIUrl":"https://doi.org/10.2514/1.b38944","url":null,"abstract":"The flow within the inlet of an engine nacelle model in the absence of a fan and the presence of crosswind is investigated in wind-tunnel experiments, with specific emphasis on the effects of separation over the inlet’s inner windward surface on the flow distortion and pressure recovery. The inlet’s entrance plane is tilted forward, and its cross section is asymmetric about the horizontal centerline. The flow topology within the inlet is characterized over a range of Mach numbers and crosswind speeds up to [Formula: see text] and [Formula: see text], respectively. It is shown that in the presence of sufficiently high crosswind to the inlet speed ratio, a three-dimensional horseshoe-like separation domain is formed over the inlet’s inner windward surface. Owing to the cross-sectional asymmetry of the entrance plane, the separation domain migrates azimuthally downward and expands azimuthally with increased crosswind to the inlet speed ratio. The present investigations demonstrate the utility of flow control for mitigating the adverse effects of the separation. The actuation is based on controllable distributed aerodynamic air bleed that is driven by the pressure differences across the nacelle’s inner and outer surfaces and reattaches the separated base flow up to crosswind speeds of [Formula: see text], resulting in a gain of up to 38% in total pressure recovery and a decrease of up to 55% in total pressure distortion. The efficacy of the bleed actuation can be further improved by tailoring the bleed distribution to the topology of the separated flow domain.","PeriodicalId":16903,"journal":{"name":"Journal of Propulsion and Power","volume":"73 1","pages":"0"},"PeriodicalIF":0.0,"publicationDate":"2023-09-18","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"135149614","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}
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":" ","pages":""},"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":" ","pages":""},"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":"23 1","pages":"0"},"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}
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