Jay V. Evans, Brandon T. Reid, Rohan M. Gejji, Carson D. Slabaugh
The instantaneous fuel regression rate within a solid-fuel ramjet combustor was characterized using X-ray radiography and ultrasonic transducer measurements. Experiments were performed with cylindrical center-perforated hydroxyl-terminated polybutadiene fuel grains at three mass fluxes ([Formula: see text]) with consistent inlet total temperatures and chamber pressures. Ultrasonic transducer measurements demonstrated changes of web thickness ranging from 7.50 to 9.85 mm and regression rate measurements ranging from 1.35 to [Formula: see text]. The local maxima of change in the web thickness due to flow reattachment and erosive burning were consistently measured with the ultrasonic transducers. Changes in the port radius on the order of 8–9 mm and regression rates of approximately [Formula: see text] were deduced from the X-ray radiography images. The structure of the flow reattachment region was evident in measurements from the X-ray radiography images captured near the combustor entrance, whereas images captured at the midlength of the combustor exhibited more uniform fuel regression profiles. Ultrasonic measurements of change in the web thickness were consistently greater in magnitude relative to X-ray radiography measurements. X-ray radiography imaging allowed for the more accurate measurement of fuel regression with the greatest axial spatial resolution, whereas ultrasonic transducer measurements yielded the greatest radial spatial resolution. The change in web thickness calculated with weight-based techniques yielded smaller-magnitude measurements of change in the web thickness relative to X-ray radiography. The regression rate was largely invariant with the mass flux within the investigated operating regime.
{"title":"Solid-Fuel Ramjet Regression Rate Measurements Using X-Ray Radiography and Ultrasonic Transducers","authors":"Jay V. Evans, Brandon T. Reid, Rohan M. Gejji, Carson D. Slabaugh","doi":"10.2514/1.b39210","DOIUrl":"https://doi.org/10.2514/1.b39210","url":null,"abstract":"The instantaneous fuel regression rate within a solid-fuel ramjet combustor was characterized using X-ray radiography and ultrasonic transducer measurements. Experiments were performed with cylindrical center-perforated hydroxyl-terminated polybutadiene fuel grains at three mass fluxes ([Formula: see text]) with consistent inlet total temperatures and chamber pressures. Ultrasonic transducer measurements demonstrated changes of web thickness ranging from 7.50 to 9.85 mm and regression rate measurements ranging from 1.35 to [Formula: see text]. The local maxima of change in the web thickness due to flow reattachment and erosive burning were consistently measured with the ultrasonic transducers. Changes in the port radius on the order of 8–9 mm and regression rates of approximately [Formula: see text] were deduced from the X-ray radiography images. The structure of the flow reattachment region was evident in measurements from the X-ray radiography images captured near the combustor entrance, whereas images captured at the midlength of the combustor exhibited more uniform fuel regression profiles. Ultrasonic measurements of change in the web thickness were consistently greater in magnitude relative to X-ray radiography measurements. X-ray radiography imaging allowed for the more accurate measurement of fuel regression with the greatest axial spatial resolution, whereas ultrasonic transducer measurements yielded the greatest radial spatial resolution. The change in web thickness calculated with weight-based techniques yielded smaller-magnitude measurements of change in the web thickness relative to X-ray radiography. The regression rate was largely invariant with the mass flux within the investigated operating regime.","PeriodicalId":16903,"journal":{"name":"Journal of Propulsion and Power","volume":null,"pages":null},"PeriodicalIF":0.0,"publicationDate":"2023-11-01","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"134996582","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}
Nuclear thermal propulsion can potentially reduce the time of flight and spacecraft system mass needed for human spaceflight beyond cislunar space. This nuclear propulsion system has comparable thrust capability to chemically impulsive systems, which at about twice the specific impulse, can double the delta-velocity ([Formula: see text]) for the same propellant mass. However, the canonical problem for nuclear propulsion has always been that its benefits are shadowed by low technology readiness of a complex system. This paper describes a combined cycle nuclear thermal rocket (CCNTR) system architecture for propulsion and electrical power that comprises a 42-MWt-capable nuclear reactor core to provide 9.4 kN thrust on demand at a specific impulse of 940 s. The liquid hydrogen propellant flow through the rocket chamber cools the reactor during burns, thereby producing thrust while concurrently rejecting waste heat to space. The reactor also produces up to 100 kWe power for the spacecraft, eliminating the need for solar power generation and averting challenges associated with restarting a cold reactor for propulsive burns. Radiators reject the waste heat from electrical power production. Earth-to-Mars orbital transfers less than 100 days appear feasible assuming 680,000 kg of liquid hydrogen propellant and a vehicle dry mass of 83,000 kg that includes the 13,000 kg CCNTR system. Together, these results suggest that a CCNTR could be most promising to enable crewed missions to Mars.
{"title":"Combined Cycle Nuclear System Architecture for Crewed Mars Spacecraft Propulsion and Power","authors":"Jack V. Maydan, James A. Nabity","doi":"10.2514/1.b39149","DOIUrl":"https://doi.org/10.2514/1.b39149","url":null,"abstract":"Nuclear thermal propulsion can potentially reduce the time of flight and spacecraft system mass needed for human spaceflight beyond cislunar space. This nuclear propulsion system has comparable thrust capability to chemically impulsive systems, which at about twice the specific impulse, can double the delta-velocity ([Formula: see text]) for the same propellant mass. However, the canonical problem for nuclear propulsion has always been that its benefits are shadowed by low technology readiness of a complex system. This paper describes a combined cycle nuclear thermal rocket (CCNTR) system architecture for propulsion and electrical power that comprises a 42-MWt-capable nuclear reactor core to provide 9.4 kN thrust on demand at a specific impulse of 940 s. The liquid hydrogen propellant flow through the rocket chamber cools the reactor during burns, thereby producing thrust while concurrently rejecting waste heat to space. The reactor also produces up to 100 kWe power for the spacecraft, eliminating the need for solar power generation and averting challenges associated with restarting a cold reactor for propulsive burns. Radiators reject the waste heat from electrical power production. Earth-to-Mars orbital transfers less than 100 days appear feasible assuming 680,000 kg of liquid hydrogen propellant and a vehicle dry mass of 83,000 kg that includes the 13,000 kg CCNTR system. Together, these results suggest that a CCNTR could be most promising to enable crewed missions to Mars.","PeriodicalId":16903,"journal":{"name":"Journal of Propulsion and Power","volume":null,"pages":null},"PeriodicalIF":0.0,"publicationDate":"2023-10-24","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"135316516","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}
Bradley Gobin, Paul Reiter, Sean Whalen, Gregory Young
An experimental study was conducted on electrically controllable solid propellants (ECSPs) created using a polyethylene oxide polymer binder, lithium perchlorate, and multiwalled carbon nanotubes. The propellants decompose and ignite shortly after the application of a voltage potential and extinguish when the voltage is removed under atmospheric conditions. The ignition delay as a function of the applied voltage magnitude was determined for a range of ECSP compositions. Pressurized experiments were conducted in an optically accessible strand burner to characterize the burning properties of the ECSPs as a function of pressure and electrical power. Additional experiments were conducted at elevated pressures where the voltage potential was removed and reapplied to extinguish and reignite the propellant and determine the self-extinction limits of the ECSPs. The results demonstrate that small compositional changes can drastically impact the ability to extinguish the ECSPs at elevated pressures.
{"title":"Extinguishing and Combustion Characteristics of Electrically Controllable Solid Propellants Under Elevated Pressures","authors":"Bradley Gobin, Paul Reiter, Sean Whalen, Gregory Young","doi":"10.2514/1.b39189","DOIUrl":"https://doi.org/10.2514/1.b39189","url":null,"abstract":"An experimental study was conducted on electrically controllable solid propellants (ECSPs) created using a polyethylene oxide polymer binder, lithium perchlorate, and multiwalled carbon nanotubes. The propellants decompose and ignite shortly after the application of a voltage potential and extinguish when the voltage is removed under atmospheric conditions. The ignition delay as a function of the applied voltage magnitude was determined for a range of ECSP compositions. Pressurized experiments were conducted in an optically accessible strand burner to characterize the burning properties of the ECSPs as a function of pressure and electrical power. Additional experiments were conducted at elevated pressures where the voltage potential was removed and reapplied to extinguish and reignite the propellant and determine the self-extinction limits of the ECSPs. The results demonstrate that small compositional changes can drastically impact the ability to extinguish the ECSPs at elevated pressures.","PeriodicalId":16903,"journal":{"name":"Journal of Propulsion and Power","volume":null,"pages":null},"PeriodicalIF":0.0,"publicationDate":"2023-10-24","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"135316229","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}
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":null,"pages":null},"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":null,"pages":null},"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":null,"pages":null},"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":null,"pages":null},"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":null,"pages":null},"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":null,"pages":null},"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":null,"pages":null},"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}
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