A new method for the identification of the peak temperature of the heat shield based on remote spectral irradiance measurements is proposed and tested through the analysis of results from the Hayabusa sample return capsule. To identify the peak temperature of the heat shield, a parameterized empirical model for the surface temperature distribution on the geometry is developed, and it is then used to simulate spectra for optimized fitting with a measured spectral radiance. This new peak temperature identification method is shown to be insensitive to the wavelengths resolved, the view of the geometry, and measurement noise. The peak temperature of the Hayabusa capsule geometry at the representative condition considered (3330 K) was resolved to within 20 K for the three instruments investigated, while the effective temperature varied by around 600 K depending on which instrument was used. The error in the peak temperature model is also shown to be an order of magnitude less sensitive to measurement noise than the effective temperature approach. The new peak temperature method facilitates the direct comparison of results from instruments with different bandwidths and/or different view angles of the capsule, which was previously not possible.
{"title":"Improved Surface Temperature Identification Method for Remote Observations of Sample Return Capsules","authors":"Byrenn Birch, David Buttsworth, Fabian Zander","doi":"10.2514/1.a35692","DOIUrl":"https://doi.org/10.2514/1.a35692","url":null,"abstract":"A new method for the identification of the peak temperature of the heat shield based on remote spectral irradiance measurements is proposed and tested through the analysis of results from the Hayabusa sample return capsule. To identify the peak temperature of the heat shield, a parameterized empirical model for the surface temperature distribution on the geometry is developed, and it is then used to simulate spectra for optimized fitting with a measured spectral radiance. This new peak temperature identification method is shown to be insensitive to the wavelengths resolved, the view of the geometry, and measurement noise. The peak temperature of the Hayabusa capsule geometry at the representative condition considered (3330 K) was resolved to within 20 K for the three instruments investigated, while the effective temperature varied by around 600 K depending on which instrument was used. The error in the peak temperature model is also shown to be an order of magnitude less sensitive to measurement noise than the effective temperature approach. The new peak temperature method facilitates the direct comparison of results from instruments with different bandwidths and/or different view angles of the capsule, which was previously not possible.","PeriodicalId":50048,"journal":{"name":"Journal of Spacecraft and Rockets","volume":"57 6","pages":"0"},"PeriodicalIF":0.0,"publicationDate":"2023-11-05","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"135725969","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}
Phillip Post, Kaitlyn Fleming, Catharine Canavan, Stephanie Cho, Gati Aher, Whitney Lohmeyer
The Federal Communications Commission (FCC) acts as the de facto gatekeeper to space for all commercial satellites, so filings to the commission can detail the direction of the industry. Between January 1, 2000, and December 31, 2021, the Commission’s International Bureau (IB) received 3289 geostationary (GEO) satellite filings from 180 distinct companies. GEO satellite filings are in decline, with their 262 applications peak reached in 2003 versus only 69 in 2021. One contributing factor is the steady decline of satellite TV subscribers. Ku band was the most requested frequency band with its ideal combination of bandwidth and resistance to rain fade, but 10 GHz of potential E-band bandwidth remains unused due to technological barriers. Approximately 69.8% of applications were granted, followed by 9.6% withdrawn, and only 7.7% were denied or dismissed, but the top 12 filers constituted only 3.8% of all applications dismissed, meaning that smaller filers’ applications were far less likely to be granted. An application took 172 days on average to be acted on, but Special Temporary Authority applications, which make up 43.7% of all applications, took only 43 days on average. The most requested legal waivers were for frequency allocations, telemetry signals, orbital debris mitigation, and file formatting.
{"title":"Analysis of Geostationary Federal Communications Commission Satellite Applications from 2000 to 2022","authors":"Phillip Post, Kaitlyn Fleming, Catharine Canavan, Stephanie Cho, Gati Aher, Whitney Lohmeyer","doi":"10.2514/1.a35660","DOIUrl":"https://doi.org/10.2514/1.a35660","url":null,"abstract":"The Federal Communications Commission (FCC) acts as the de facto gatekeeper to space for all commercial satellites, so filings to the commission can detail the direction of the industry. Between January 1, 2000, and December 31, 2021, the Commission’s International Bureau (IB) received 3289 geostationary (GEO) satellite filings from 180 distinct companies. GEO satellite filings are in decline, with their 262 applications peak reached in 2003 versus only 69 in 2021. One contributing factor is the steady decline of satellite TV subscribers. Ku band was the most requested frequency band with its ideal combination of bandwidth and resistance to rain fade, but 10 GHz of potential E-band bandwidth remains unused due to technological barriers. Approximately 69.8% of applications were granted, followed by 9.6% withdrawn, and only 7.7% were denied or dismissed, but the top 12 filers constituted only 3.8% of all applications dismissed, meaning that smaller filers’ applications were far less likely to be granted. An application took 172 days on average to be acted on, but Special Temporary Authority applications, which make up 43.7% of all applications, took only 43 days on average. The most requested legal waivers were for frequency allocations, telemetry signals, orbital debris mitigation, and file formatting.","PeriodicalId":50048,"journal":{"name":"Journal of Spacecraft and Rockets","volume":"18 5","pages":"0"},"PeriodicalIF":0.0,"publicationDate":"2023-11-03","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"135821268","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}
Ruth A. Miller, Chun Y. Tang, Jose A. B. Santos, Todd R. White, Brett A. Cruden
The Mars Entry, Descent, and Landing Instrumentation 2 (MEDLI2) sensor suite on the Mars 2020 mission included two total heat flux sensors and one radiometer on the backshell to directly measure the aftbody aerothermal environments during entry into the Martian atmosphere. All three sensors successfully returned aftbody entry heating measurements. Comparisons between the total heat flux sensor measurements and predictions by NASA simulation tools (DPLR/NEQAIR) show excellent agreement and provide confidence in the models. The radiometer measured significant radiative heating, but compared to the model predictions the signal was attenuated by 48% at the end of the entry heat pulse. The loss of signal is attributed to blockage by thermal protection system (TPS) ablation product deposits on the radiometer window. Ground-based testing in the NASA Ames arcjet facilities was conducted to understand the impact of ablation product deposits on the measured radiometer signal. A discussion of the test results, how flight-like the test conditions were, and future work to further characterize the effect of TPS ablation product deposits on the radiometer performance are presented. In addition to measuring the entry heat pulse, all three sensors were sensitive enough to measure solar radiation during cruise, the radiometer measured solar flux during the entry heat pulse, and the leeside total heat flux sensor picked up the descent reaction control system firings.
{"title":"Aftbody Heat Flux Measurements During Mars 2020 Entry","authors":"Ruth A. Miller, Chun Y. Tang, Jose A. B. Santos, Todd R. White, Brett A. Cruden","doi":"10.2514/1.a35783","DOIUrl":"https://doi.org/10.2514/1.a35783","url":null,"abstract":"The Mars Entry, Descent, and Landing Instrumentation 2 (MEDLI2) sensor suite on the Mars 2020 mission included two total heat flux sensors and one radiometer on the backshell to directly measure the aftbody aerothermal environments during entry into the Martian atmosphere. All three sensors successfully returned aftbody entry heating measurements. Comparisons between the total heat flux sensor measurements and predictions by NASA simulation tools (DPLR/NEQAIR) show excellent agreement and provide confidence in the models. The radiometer measured significant radiative heating, but compared to the model predictions the signal was attenuated by 48% at the end of the entry heat pulse. The loss of signal is attributed to blockage by thermal protection system (TPS) ablation product deposits on the radiometer window. Ground-based testing in the NASA Ames arcjet facilities was conducted to understand the impact of ablation product deposits on the measured radiometer signal. A discussion of the test results, how flight-like the test conditions were, and future work to further characterize the effect of TPS ablation product deposits on the radiometer performance are presented. In addition to measuring the entry heat pulse, all three sensors were sensitive enough to measure solar radiation during cruise, the radiometer measured solar flux during the entry heat pulse, and the leeside total heat flux sensor picked up the descent reaction control system firings.","PeriodicalId":50048,"journal":{"name":"Journal of Spacecraft and Rockets","volume":"14 S1","pages":"0"},"PeriodicalIF":0.0,"publicationDate":"2023-11-02","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"135974699","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}
Carlo Girardello, Carsten Scharlemann, Wolfgang Treberer-Treberspurg, Markus Trenker, Christoph Obertscheider, Anand Nair Radhakrishnan, Martin Tajmar
CubeSats have a 65% success rate. Failures derive from design mistake or components malfunctions. To improve the success rate, Technical University of Dresden and FHWN (University of Applied Sciences Wiener Neustadt) developed GREATCUBE+, a software tool for the conceptual design of CubeSats. Its layered structure comprises three levels: empirical, where successfully flown missions are used for an initial tradeoff; analytical, where a design refinement is performed; and numerical, for the final assessment of the proposed architecture. The tool provides teams with information related to commercial off-the-shelf products which will satisfy the mission requirements. To turn this software into a universally applicable tool, it is possible to perform the design of CubeSat mission with many payload’s typologies such as attitude determination and control subsystem, telemetry telecommunication and command, onboard computer, propulsion unit and technology demonstration or scientific payloads. GREATCUBE+ has been validated using the information of existing CubeSats as baseline for its simulation. The achievable accuracy when comparing the simulated outcomes and the real design is of almost 100% for volumes, 90% for masses, and 80% for power generation. By implementing this tool during the conceptual development phase, it is hoped that teams could benefit in reliability thanks to the usage of flight proven equipment recommended via GREATCUBE+ together with a quicker development time.
{"title":"Conceptual Development of CubeSat Missions with GREATCUBE+: Methodology and Possible Applications","authors":"Carlo Girardello, Carsten Scharlemann, Wolfgang Treberer-Treberspurg, Markus Trenker, Christoph Obertscheider, Anand Nair Radhakrishnan, Martin Tajmar","doi":"10.2514/1.a35741","DOIUrl":"https://doi.org/10.2514/1.a35741","url":null,"abstract":"CubeSats have a 65% success rate. Failures derive from design mistake or components malfunctions. To improve the success rate, Technical University of Dresden and FHWN (University of Applied Sciences Wiener Neustadt) developed GREATCUBE+, a software tool for the conceptual design of CubeSats. Its layered structure comprises three levels: empirical, where successfully flown missions are used for an initial tradeoff; analytical, where a design refinement is performed; and numerical, for the final assessment of the proposed architecture. The tool provides teams with information related to commercial off-the-shelf products which will satisfy the mission requirements. To turn this software into a universally applicable tool, it is possible to perform the design of CubeSat mission with many payload’s typologies such as attitude determination and control subsystem, telemetry telecommunication and command, onboard computer, propulsion unit and technology demonstration or scientific payloads. GREATCUBE+ has been validated using the information of existing CubeSats as baseline for its simulation. The achievable accuracy when comparing the simulated outcomes and the real design is of almost 100% for volumes, 90% for masses, and 80% for power generation. By implementing this tool during the conceptual development phase, it is hoped that teams could benefit in reliability thanks to the usage of flight proven equipment recommended via GREATCUBE+ together with a quicker development time.","PeriodicalId":50048,"journal":{"name":"Journal of Spacecraft and Rockets","volume":"5 5","pages":"0"},"PeriodicalIF":0.0,"publicationDate":"2023-11-02","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"135933682","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}
This investigation demonstrates hypersonic trajectory optimization using planar rigid body dynamics within the indirect trajectory optimization framework. Employing rigid body dynamics captures the coupling between the optimal trajectory, the vehicle geometry, mass distribution, and control configuration. This provides trajectories that implicitly account for the maneuverability of the vehicle, wherein the vehicle is guaranteed to follow the angle-of-attack profile. This is unlike point-mass dynamics, wherein the angle-of-attack or angle-of-attack rate is directly used as the control variable, and maneuverability must be accounted for using bounds on these quantities. This is not straightforward because these bounds are dependent on flight conditions and are not constant for the entire trajectory. As a result, point-mass dynamics can produce infeasible solutions if these bounds are not properly handled. When using rigid body dynamics, these bounds are a consequence of the flight dynamics and are not required to be explicitly enforced, thereby circumventing the challenge with dynamic bounds on angle-of-attack and angle-of-attack rate altogether. Additionally, optimal trajectories calculated using rigid body dynamics more accurately reflect the drag penalties incurred when maneuvering the vehicle, such as when deflecting the aerodynamic control surfaces. The incurred drag penalties become critical in high-performance applications, wherein the terminal velocity is required to be maximized. Also, because the trajectory is coupled to the vehicle geometry, mass distribution, and control architecture, the optimal trajectory can be concurrently analyzed with the vehicle configuration, thereby enabling multidisciplinary design analysis. Despite these benefits offered by employing rigid body dynamics in trajectory optimization, there is limited literature in this regard, and none of them explored in this investigation employs indirect methods. This investigation fills this gap in the indirect trajectory optimization arena.
{"title":"Incorporation of Rigid Body Dynamics into Indirect Hypersonic Trajectory Optimization","authors":"Harish Saranathan, Michael J. Grant","doi":"10.2514/1.a35717","DOIUrl":"https://doi.org/10.2514/1.a35717","url":null,"abstract":"This investigation demonstrates hypersonic trajectory optimization using planar rigid body dynamics within the indirect trajectory optimization framework. Employing rigid body dynamics captures the coupling between the optimal trajectory, the vehicle geometry, mass distribution, and control configuration. This provides trajectories that implicitly account for the maneuverability of the vehicle, wherein the vehicle is guaranteed to follow the angle-of-attack profile. This is unlike point-mass dynamics, wherein the angle-of-attack or angle-of-attack rate is directly used as the control variable, and maneuverability must be accounted for using bounds on these quantities. This is not straightforward because these bounds are dependent on flight conditions and are not constant for the entire trajectory. As a result, point-mass dynamics can produce infeasible solutions if these bounds are not properly handled. When using rigid body dynamics, these bounds are a consequence of the flight dynamics and are not required to be explicitly enforced, thereby circumventing the challenge with dynamic bounds on angle-of-attack and angle-of-attack rate altogether. Additionally, optimal trajectories calculated using rigid body dynamics more accurately reflect the drag penalties incurred when maneuvering the vehicle, such as when deflecting the aerodynamic control surfaces. The incurred drag penalties become critical in high-performance applications, wherein the terminal velocity is required to be maximized. Also, because the trajectory is coupled to the vehicle geometry, mass distribution, and control architecture, the optimal trajectory can be concurrently analyzed with the vehicle configuration, thereby enabling multidisciplinary design analysis. Despite these benefits offered by employing rigid body dynamics in trajectory optimization, there is limited literature in this regard, and none of them explored in this investigation employs indirect methods. This investigation fills this gap in the indirect trajectory optimization arena.","PeriodicalId":50048,"journal":{"name":"Journal of Spacecraft and Rockets","volume":"39 7","pages":"0"},"PeriodicalIF":0.0,"publicationDate":"2023-11-02","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"135934804","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 demanding working environment of hypersonic missiles imposes high requirements on the reliability and stability of the reaction control system. This work aims to investigate the impact of uncertainties in freestream and jet parameters on the flowfield and aerodynamic performance of a missile with a lateral jet. Initially, the influence of angle of attack and Mach number on aerodynamic characteristics is evaluated. The findings indicate that the lateral jet control effect is enhanced as the Mach number increases. Moreover, the interference force/moment does not vary linearly with the angle of attack, and there are notable differences in the flow characteristics when the jet is on the windward and leeward sides. Subsequently, six parameters are selected as input variables. The uncertainties of the output results are analyzed using a stochastic expansion based on the nonintrusive polynomial chaos method. Sensitivity analysis is performed using Sobol indices to estimate the relative contributions of each parameter to the overall uncertainty. Within the given uncertainty interval, the surface pressure of the missile exhibits significant variation, with a maximum uncertainty of approximately 480%. The uncertainty of the recirculation zone range is 50.14%, while the pitch moment amplification factor has an uncertainty of 52.44%. The freestream Mach number, angle of attack, and jet pressure ratio are identified as key factors affecting the aerodynamic performance, surface pressure coefficient, and jet penetration height. Conversely, the contribution of other parameters can be considered negligible.
{"title":"Uncertainty and Sensitivity Study on Lateral Jet Interaction for Hypersonic Missile","authors":"Mu-liang Jia, Shu-sheng Chen, Jin-ping Li, Wu Yuan, Bo-xi Lin, Zheng-hong Gao","doi":"10.2514/1.a35841","DOIUrl":"https://doi.org/10.2514/1.a35841","url":null,"abstract":"The demanding working environment of hypersonic missiles imposes high requirements on the reliability and stability of the reaction control system. This work aims to investigate the impact of uncertainties in freestream and jet parameters on the flowfield and aerodynamic performance of a missile with a lateral jet. Initially, the influence of angle of attack and Mach number on aerodynamic characteristics is evaluated. The findings indicate that the lateral jet control effect is enhanced as the Mach number increases. Moreover, the interference force/moment does not vary linearly with the angle of attack, and there are notable differences in the flow characteristics when the jet is on the windward and leeward sides. Subsequently, six parameters are selected as input variables. The uncertainties of the output results are analyzed using a stochastic expansion based on the nonintrusive polynomial chaos method. Sensitivity analysis is performed using Sobol indices to estimate the relative contributions of each parameter to the overall uncertainty. Within the given uncertainty interval, the surface pressure of the missile exhibits significant variation, with a maximum uncertainty of approximately 480%. The uncertainty of the recirculation zone range is 50.14%, while the pitch moment amplification factor has an uncertainty of 52.44%. The freestream Mach number, angle of attack, and jet pressure ratio are identified as key factors affecting the aerodynamic performance, surface pressure coefficient, and jet penetration height. Conversely, the contribution of other parameters can be considered negligible.","PeriodicalId":50048,"journal":{"name":"Journal of Spacecraft and Rockets","volume":"14 5","pages":"0"},"PeriodicalIF":0.0,"publicationDate":"2023-11-02","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"135974697","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}
Adam Herrmann, Mark A. Stephenson, Hanspeter Schaub
This work explores single-agent reinforcement learning for the multi-satellite agile Earth-observing scheduling problem. The objective of the problem is to maximize the weighted sum of imaging targets collected and downlinked while avoiding resource constraint violations on board the spacecraft. To avoid the computational complexity associated with multi-agent deep reinforcement learning while creating a robust and scalable solution, a policy is trained in a single satellite environment. This policy is then deployed on board each satellite in a Walker-delta constellation. A global set of targets is distributed to each satellite based on target access. The satellites communicate with one another to determine whether an imaging target is imaged or downlinked. Free communication, line-of-sight communication, and no communication are explored to determine how the communication assumptions and constellation design impact performance. Free communication is shown to produce the best performance, and no communication is shown to produce the worst performance. Line-of-sight communication performance is shown to depend heavily on the design of the constellation and how frequently the satellites can communicate with one another. To explore how higher-level coordination can impact performance, a centralized mixed-integer programming optimization approach to global target distribution is explored and compared to a decentralized approach. A genetic algorithm is also implemented for comparison purposes, and the proposed method is shown to achieve higher reward on average at a fraction of the computational cost.
{"title":"Single-Agent Reinforcement Learning for Scalable Earth-Observing Satellite Constellation Operations","authors":"Adam Herrmann, Mark A. Stephenson, Hanspeter Schaub","doi":"10.2514/1.a35736","DOIUrl":"https://doi.org/10.2514/1.a35736","url":null,"abstract":"This work explores single-agent reinforcement learning for the multi-satellite agile Earth-observing scheduling problem. The objective of the problem is to maximize the weighted sum of imaging targets collected and downlinked while avoiding resource constraint violations on board the spacecraft. To avoid the computational complexity associated with multi-agent deep reinforcement learning while creating a robust and scalable solution, a policy is trained in a single satellite environment. This policy is then deployed on board each satellite in a Walker-delta constellation. A global set of targets is distributed to each satellite based on target access. The satellites communicate with one another to determine whether an imaging target is imaged or downlinked. Free communication, line-of-sight communication, and no communication are explored to determine how the communication assumptions and constellation design impact performance. Free communication is shown to produce the best performance, and no communication is shown to produce the worst performance. Line-of-sight communication performance is shown to depend heavily on the design of the constellation and how frequently the satellites can communicate with one another. To explore how higher-level coordination can impact performance, a centralized mixed-integer programming optimization approach to global target distribution is explored and compared to a decentralized approach. A genetic algorithm is also implemented for comparison purposes, and the proposed method is shown to achieve higher reward on average at a fraction of the computational cost.","PeriodicalId":50048,"journal":{"name":"Journal of Spacecraft and Rockets","volume":"33 2","pages":"0"},"PeriodicalIF":0.0,"publicationDate":"2023-11-02","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"135934234","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}
Justin St. P. Walsh, Rao Hamza Ali, Alice C. Gorman, Amir Kanan Kashefi
This paper presents an analysis of data derived from thousands of publicly available photographs showing life on the International Space Station (ISS) between 2000 and 2020. Our analysis uses crew and locational information from the photographs’ metadata to identify the distribution of different population groups—by gender, nationality, and space agency affiliation—across modules of the ISS, for the first time. Given the significance of the ISS as the most intensively inhabited space habitat to date, an international cooperative initiative involving 26 countries and five space agencies, and one of the most expensive building projects ever undertaken by humans, developing an understanding of which people are using different parts of the space station is critical for future usage of this and other stations. This study also sheds light on problems faced by future space station designers who are concerned with optimal usage of their habitats. The data from this investigation have been permanently deposited with Open Context. It is freely available for use under a Creative Commons license (CC BY 4.0) at https://doi.org/10.6078/M7668B9H .
本文分析了2000年至2020年期间国际空间站(ISS)上数千张公开照片的数据。我们的分析首次使用了来自照片元数据的人员和位置信息，根据性别、国籍和隶属于航天机构的情况，确定了不同人口群体在国际空间站各个模块中的分布情况。考虑到国际空间站作为迄今为止居住最密集的空间栖息地的重要性，涉及26个国家和5个空间机构的国际合作倡议，以及人类有史以来承担的最昂贵的建筑项目之一，了解人们正在使用空间站的不同部分对于未来使用这个和其他空间站至关重要。这项研究还揭示了未来空间站设计师所面临的问题，他们关心的是如何最佳地利用他们的栖息地。本次调查的数据已永久保存在Open Context中。在知识共享许可(CC BY 4.0)下，可以在https://doi.org/10.6078/M7668B9H上免费使用。
{"title":"First Approximation of Population Distributions on the International Space Station","authors":"Justin St. P. Walsh, Rao Hamza Ali, Alice C. Gorman, Amir Kanan Kashefi","doi":"10.2514/1.a35686","DOIUrl":"https://doi.org/10.2514/1.a35686","url":null,"abstract":"This paper presents an analysis of data derived from thousands of publicly available photographs showing life on the International Space Station (ISS) between 2000 and 2020. Our analysis uses crew and locational information from the photographs’ metadata to identify the distribution of different population groups—by gender, nationality, and space agency affiliation—across modules of the ISS, for the first time. Given the significance of the ISS as the most intensively inhabited space habitat to date, an international cooperative initiative involving 26 countries and five space agencies, and one of the most expensive building projects ever undertaken by humans, developing an understanding of which people are using different parts of the space station is critical for future usage of this and other stations. This study also sheds light on problems faced by future space station designers who are concerned with optimal usage of their habitats. The data from this investigation have been permanently deposited with Open Context. It is freely available for use under a Creative Commons license (CC BY 4.0) at https://doi.org/10.6078/M7668B9H .","PeriodicalId":50048,"journal":{"name":"Journal of Spacecraft and Rockets","volume":"5 4","pages":"0"},"PeriodicalIF":0.0,"publicationDate":"2023-10-26","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"135013366","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}
Sarah A. Mecklem, Damian Curran, Will O. Landsberg, Ananthanarayanan Veeraragavan
A fundamental investigation of combustion behavior in an ethylene-fueled Mach 8 axisymmetric scramjet combustor was numerically conducted using US3D Reynolds-averaged Navier–Stokes to evaluate the impact of fueling strategies employed with a tandem cavity-equipped combustor. Five fueling configurations with a consistent total fuel-to-air equivalence ratio at 75% of stoichiometric were assessed using a combination of injector ports, located either upstream of, within, between, or downstream of the cavities. Cold-wall and hot-wall conditions were examined, representing shock tunnel equivalent test conditions with minimal wall heating and steady, in-flight wall temperatures, respectively. Two distributed fueling schemes resulted in higher combustion performance when operating in cold conditions and comparable performance when operating in hot conditions. Total pressure losses sustained under each distributed fueling method were equivalent to or lower than those sustained by cases with fuel injected only upstream of the cavities. All distributed schemes examined delayed the onset of flow separation, resulting in improved operational stability and higher resistance to engine unstart.
{"title":"Examining Distributed Fueling Schemes Using Tandem Cavities in a Mach 8 Scramjet","authors":"Sarah A. Mecklem, Damian Curran, Will O. Landsberg, Ananthanarayanan Veeraragavan","doi":"10.2514/1.a35848","DOIUrl":"https://doi.org/10.2514/1.a35848","url":null,"abstract":"A fundamental investigation of combustion behavior in an ethylene-fueled Mach 8 axisymmetric scramjet combustor was numerically conducted using US3D Reynolds-averaged Navier–Stokes to evaluate the impact of fueling strategies employed with a tandem cavity-equipped combustor. Five fueling configurations with a consistent total fuel-to-air equivalence ratio at 75% of stoichiometric were assessed using a combination of injector ports, located either upstream of, within, between, or downstream of the cavities. Cold-wall and hot-wall conditions were examined, representing shock tunnel equivalent test conditions with minimal wall heating and steady, in-flight wall temperatures, respectively. Two distributed fueling schemes resulted in higher combustion performance when operating in cold conditions and comparable performance when operating in hot conditions. Total pressure losses sustained under each distributed fueling method were equivalent to or lower than those sustained by cases with fuel injected only upstream of the cavities. All distributed schemes examined delayed the onset of flow separation, resulting in improved operational stability and higher resistance to engine unstart.","PeriodicalId":50048,"journal":{"name":"Journal of Spacecraft and Rockets","volume":"11 1","pages":"0"},"PeriodicalIF":0.0,"publicationDate":"2023-10-26","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"135018368","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}
Jason Hartwig, Narottama Esser, Shreykumar Jain, David Souders, Allen Prasad Varghese, Angelo Tafuni
Future cryogenic propulsion systems will require efficient methods for transferring cryogenic propellants from a depot storage tank to a customer receiver tank to minimize costs and maximize reusability. The Reduced Gravity Cryogenic Transfer Project is currently developing advanced cryogenic fluid management technology and developing and validating new numerical models for three phases of transfer: line chill down, tank chill down, and tank fill. Additionally, multiple liquid nitrogen ([Formula: see text]) parabolic flight transfer rigs are being designed by universities and NASA to investigate the gravitational sensitivities that exist in these three technologies. To maximize the collection of low-g data during flights, it is required to extract as much [Formula: see text] as possible from the supply tank, despite variable gravity levels. The purpose of this study is to present computational fluid dynamics volume of fluid simulations of [Formula: see text] behavior in the supply tank onboard parabolic flights to validate the optimal design of a bidirectional propellant management device (PMD) using the commercial software FLOW-3D. A parametric study was conducted on the effects of gravity level, fill level, pore size, open area, thickness, and type of baffle on PMD performance. Based on the results, the designed PMD exceeded the targeted expulsion efficiency.
{"title":"CFD Modeling of Bidirectional PMDs Inside Cryogenic Propellant Tanks Onboard Parabolic Flights","authors":"Jason Hartwig, Narottama Esser, Shreykumar Jain, David Souders, Allen Prasad Varghese, Angelo Tafuni","doi":"10.2514/1.a35808","DOIUrl":"https://doi.org/10.2514/1.a35808","url":null,"abstract":"Future cryogenic propulsion systems will require efficient methods for transferring cryogenic propellants from a depot storage tank to a customer receiver tank to minimize costs and maximize reusability. The Reduced Gravity Cryogenic Transfer Project is currently developing advanced cryogenic fluid management technology and developing and validating new numerical models for three phases of transfer: line chill down, tank chill down, and tank fill. Additionally, multiple liquid nitrogen ([Formula: see text]) parabolic flight transfer rigs are being designed by universities and NASA to investigate the gravitational sensitivities that exist in these three technologies. To maximize the collection of low-g data during flights, it is required to extract as much [Formula: see text] as possible from the supply tank, despite variable gravity levels. The purpose of this study is to present computational fluid dynamics volume of fluid simulations of [Formula: see text] behavior in the supply tank onboard parabolic flights to validate the optimal design of a bidirectional propellant management device (PMD) using the commercial software FLOW-3D. A parametric study was conducted on the effects of gravity level, fill level, pore size, open area, thickness, and type of baffle on PMD performance. Based on the results, the designed PMD exceeded the targeted expulsion efficiency.","PeriodicalId":50048,"journal":{"name":"Journal of Spacecraft and Rockets","volume":"9 1","pages":"0"},"PeriodicalIF":0.0,"publicationDate":"2023-10-25","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"135167045","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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