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Acta Astronautica
Journal Prestige (SJR): 0.758  Citation Impact (citeScore): 2 Number of Followers: 419  Hybrid journal (It can contain Open Access articles)ISSN (Print) 0094-5765 Published by Elsevier  [3161 journals]
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- Design of distant retrograde orbits based on a higher order analytical
solutiont- Abstract: Publication date: Available online 14 February 2019Source: Acta AstronauticaAuthor(s): Martin Lara Because the distant retrograde orbits dynamics inherently depends on special functions, approximate analytical solutions in the literature are commonly constrained to providing rough approximations of the qualitative behavior. We rely on perturbation methods and succeed in computing a higher order analytical solution that extends the range of applicability of previous solutions to the problem of relative, quasi-satellite orbits with large librations. Besides, the analytical solution provides two design parameters that are effectively used in the computation of periodic, distant retrograde orbits without constraining to the typical 1:1 resonance.
- Abstract: Publication date: Available online 14 February 2019Source: Acta AstronauticaAuthor(s): Martin Lara Because the distant retrograde orbits dynamics inherently depends on special functions, approximate analytical solutions in the literature are commonly constrained to providing rough approximations of the qualitative behavior. We rely on perturbation methods and succeed in computing a higher order analytical solution that extends the range of applicability of previous solutions to the problem of relative, quasi-satellite orbits with large librations. Besides, the analytical solution provides two design parameters that are effectively used in the computation of periodic, distant retrograde orbits without constraining to the typical 1:1 resonance.
- Folding analysis for thin-walled deployable composite boom
- Abstract: Publication date: Available online 13 February 2019Source: Acta AstronauticaAuthor(s): Jiang-Bo Bai, Di Chen, Jun-Jiang Xiong, R. Ajit Shenoi Thin-walled deployable composite structures with high strain ability are of considerable interest and increasingly used in aerospace field due to their superior mechanical behaviour and folding function. This paper seeks to investigate large folding deformation behaviour of a thin-walled deployable composite boom (DCB) made of high strain carbon-fibre-reinforced-plastics fabricated by vacuum-bag and co-bonding technologies. Large deformation function tests are conducted to determine tensile, compression and folding behaviours of the DCB. Geometrically nonlinear, explicit and standard, finite element models and an analytical model are implemented to characterize the tensile and compression behaviours of the DCB, and the deviations of predictions from experiments are in an acceptable scatter. Furthermore, a geometrically nonlinear and explicit finite element model is also generated for folding analysis of the DCB, and a good correlation between predictions and experiments has been achieved.
- Abstract: Publication date: Available online 13 February 2019Source: Acta AstronauticaAuthor(s): Jiang-Bo Bai, Di Chen, Jun-Jiang Xiong, R. Ajit Shenoi Thin-walled deployable composite structures with high strain ability are of considerable interest and increasingly used in aerospace field due to their superior mechanical behaviour and folding function. This paper seeks to investigate large folding deformation behaviour of a thin-walled deployable composite boom (DCB) made of high strain carbon-fibre-reinforced-plastics fabricated by vacuum-bag and co-bonding technologies. Large deformation function tests are conducted to determine tensile, compression and folding behaviours of the DCB. Geometrically nonlinear, explicit and standard, finite element models and an analytical model are implemented to characterize the tensile and compression behaviours of the DCB, and the deviations of predictions from experiments are in an acceptable scatter. Furthermore, a geometrically nonlinear and explicit finite element model is also generated for folding analysis of the DCB, and a good correlation between predictions and experiments has been achieved.
- Mixing and combustion characteristics in a cavity-based supersonic
combustor with different injection schemes- Abstract: Publication date: Available online 13 February 2019Source: Acta AstronauticaAuthor(s): Xiliang Song, Hongbo Wang, Mingbo Sun, Zun Cai, Chaoyang Liu, Jiangfei Yu Mixing and combustion characteristics in an ethylene fueled cavity-based scramjet combustor with different injection schemes are investigated experimentally and numerically in low equivalence ratio conditions, with inflow conditions of Ma = 2.52, stagnation pressure P0 = 1.65 MPa and stagnation temperature T0 = 1600 K. The wall-pressure distributions obtained by numerical simulations are in fairly good agreement with the experimental data. The numerical results also indicate that multiple-orifice cases have higher mixing efficiencies and combustion efficiencies than single-orifice cases. In addition, it is observed in the experiments that flamebases are not stabilized in a fixed position. In single-orifice cases, the flamebases oscillate dramatically between the cavity leading edge and trailing edge, and the oscillations are similar for different injection pressures. In multiple-orifice cases, the flamebases oscillate slightly in the cavity shear layer just downstream of the leading edge and the amplitudes with lower injection pressure are larger than those with higher injection pressure.
- Abstract: Publication date: Available online 13 February 2019Source: Acta AstronauticaAuthor(s): Xiliang Song, Hongbo Wang, Mingbo Sun, Zun Cai, Chaoyang Liu, Jiangfei Yu Mixing and combustion characteristics in an ethylene fueled cavity-based scramjet combustor with different injection schemes are investigated experimentally and numerically in low equivalence ratio conditions, with inflow conditions of Ma = 2.52, stagnation pressure P0 = 1.65 MPa and stagnation temperature T0 = 1600 K. The wall-pressure distributions obtained by numerical simulations are in fairly good agreement with the experimental data. The numerical results also indicate that multiple-orifice cases have higher mixing efficiencies and combustion efficiencies than single-orifice cases. In addition, it is observed in the experiments that flamebases are not stabilized in a fixed position. In single-orifice cases, the flamebases oscillate dramatically between the cavity leading edge and trailing edge, and the oscillations are similar for different injection pressures. In multiple-orifice cases, the flamebases oscillate slightly in the cavity shear layer just downstream of the leading edge and the amplitudes with lower injection pressure are larger than those with higher injection pressure.
- Combustion possibility assessment for separating launch-vehicle components
during atmospheric phase of descent trajectory- Abstract: Publication date: Available online 12 February 2019Source: Acta AstronauticaAuthor(s): V. Trushlyakov, K. Zharikov, D. Davydovich The investigation is made of the processes associated with the combustion of separating launch vehicle components (LVCs) on the atmospheric phase of the descent trajectory. This allows us to provide a radical solution to the problem of reducing the areas of the LVC impact zones. Ballistic parameters values for the LVC motion are determined. Such areas are proposed for realizing LVC heating to the ignition temperature through the energy mixture burning. Further combustion of the LVC occurs when there is a sufficient amount of oxygen in the oncoming air stream.
- Abstract: Publication date: Available online 12 February 2019Source: Acta AstronauticaAuthor(s): V. Trushlyakov, K. Zharikov, D. Davydovich The investigation is made of the processes associated with the combustion of separating launch vehicle components (LVCs) on the atmospheric phase of the descent trajectory. This allows us to provide a radical solution to the problem of reducing the areas of the LVC impact zones. Ballistic parameters values for the LVC motion are determined. Such areas are proposed for realizing LVC heating to the ignition temperature through the energy mixture burning. Further combustion of the LVC occurs when there is a sufficient amount of oxygen in the oncoming air stream.
- Is it worth the risk' – An astronaut's approach to risk
awareness- Abstract: Publication date: Available online 12 February 2019Source: Acta AstronauticaAuthor(s): Reinhold Ewald The need to make human spaceflight as safe as technically possible is a characteristic of this special branch of space missions and drives the cost and feasibility of human space exploration. Between the time prospective astronauts first apply for a chance to fly into space and the actual time when they climb into the spacecraft on top of a rocket for the first flight, risk awareness, mitigation, and assessment are present as a constant background reflection. What drives human explorers to accept the remaining non-mitigatable risks and when is the individual “go” decision made' In light of future long-term missions; leading humans again from LEO into deep space a sound understanding of this decision process may lead to an improved selection and composition of capable space expedition crews.
- Abstract: Publication date: Available online 12 February 2019Source: Acta AstronauticaAuthor(s): Reinhold Ewald The need to make human spaceflight as safe as technically possible is a characteristic of this special branch of space missions and drives the cost and feasibility of human space exploration. Between the time prospective astronauts first apply for a chance to fly into space and the actual time when they climb into the spacecraft on top of a rocket for the first flight, risk awareness, mitigation, and assessment are present as a constant background reflection. What drives human explorers to accept the remaining non-mitigatable risks and when is the individual “go” decision made' In light of future long-term missions; leading humans again from LEO into deep space a sound understanding of this decision process may lead to an improved selection and composition of capable space expedition crews.
- Inverse design methodology of cone-derived waverider based on pre-defined
shock wave under strong geometric constraints- Abstract: Publication date: Available online 12 February 2019Source: Acta AstronauticaAuthor(s): Kaikai Yu, Jinglei Xu, Hao Gong, Rui Li, Shun Liu An inverse design methodology on a waverider is proposed under strong geometric constraints on the basis of the generation method of an arbitrarily-shaped shock wave to obtain an improved performance. First, the generation method of the body contour that can form an arbitrarily-shaped shock wave is introduced in detail based on the method of characteristics. The effectiveness and accuracy of the generation method are verified through numerical simulation. Second, with the aid of the generation method, an inverse design method on waveriders under strong geometric constraints is presented. The waverider designed by the proposed method can satisfy the geometric constraints completely. Then, the influences of the shock wave shapes on the performance of the waveriders are studied through numerical simulation. In comparison with a traditional cone-derived waverider, the waverider based on the shock wave with convex quality can considerably increase the aerodynamic performance under the current geometric constraints. Conversely, the shock wave with concave quality deteriorates the aerodynamic performance of waveriders. The volumetric efficiency can also be improved with the adoption of the shock wave with concave quality. In conclusion, the numerical simulation result demonstrates that the proposed waverider design method can provide an increased flexibility and meet the different requirements of an actual design of hypersonic vehicles in geometric constraints and aerodynamic performances.
- Abstract: Publication date: Available online 12 February 2019Source: Acta AstronauticaAuthor(s): Kaikai Yu, Jinglei Xu, Hao Gong, Rui Li, Shun Liu An inverse design methodology on a waverider is proposed under strong geometric constraints on the basis of the generation method of an arbitrarily-shaped shock wave to obtain an improved performance. First, the generation method of the body contour that can form an arbitrarily-shaped shock wave is introduced in detail based on the method of characteristics. The effectiveness and accuracy of the generation method are verified through numerical simulation. Second, with the aid of the generation method, an inverse design method on waveriders under strong geometric constraints is presented. The waverider designed by the proposed method can satisfy the geometric constraints completely. Then, the influences of the shock wave shapes on the performance of the waveriders are studied through numerical simulation. In comparison with a traditional cone-derived waverider, the waverider based on the shock wave with convex quality can considerably increase the aerodynamic performance under the current geometric constraints. Conversely, the shock wave with concave quality deteriorates the aerodynamic performance of waveriders. The volumetric efficiency can also be improved with the adoption of the shock wave with concave quality. In conclusion, the numerical simulation result demonstrates that the proposed waverider design method can provide an increased flexibility and meet the different requirements of an actual design of hypersonic vehicles in geometric constraints and aerodynamic performances.
- A novel non-ablative thermal protection system with combined spike and
opposing jet concept- Abstract: Publication date: Available online 11 February 2019Source: Acta AstronauticaAuthor(s): Jie Huang, Wei-Xing Yao In order to reduce the aerodynamic heating of the hypersonic vehicles, a combined non-ablative thermal protection system (CNA-TPS) is proposed in this paper, which consists of the blunt body, the spike, the front and rear jets. The heat reduction efficiency is studied by the CFD numerical method. The results show that the heat reduction efficiency of CNA-TPS is better than the other configurations, which verifies the advantage of the CNA-TPS. In addition, the effect of length of the spike, total pressure ratios and sizes of the front and rear jets are studied. The results show that the heat reduction efficiency of CNA-TPS increases with the increase of length of the spike, total pressure ratios and sizes of the front and rear jets. The improving rate of the heat reduction efficiency gradually decreases with the increase of length of the spike and total pressure ratios of the front and rear jets. The investigations in this paper show the feasibility and advantage of CNA-TPS in the future engineering applications.
- Abstract: Publication date: Available online 11 February 2019Source: Acta AstronauticaAuthor(s): Jie Huang, Wei-Xing Yao In order to reduce the aerodynamic heating of the hypersonic vehicles, a combined non-ablative thermal protection system (CNA-TPS) is proposed in this paper, which consists of the blunt body, the spike, the front and rear jets. The heat reduction efficiency is studied by the CFD numerical method. The results show that the heat reduction efficiency of CNA-TPS is better than the other configurations, which verifies the advantage of the CNA-TPS. In addition, the effect of length of the spike, total pressure ratios and sizes of the front and rear jets are studied. The results show that the heat reduction efficiency of CNA-TPS increases with the increase of length of the spike, total pressure ratios and sizes of the front and rear jets. The improving rate of the heat reduction efficiency gradually decreases with the increase of length of the spike and total pressure ratios of the front and rear jets. The investigations in this paper show the feasibility and advantage of CNA-TPS in the future engineering applications.
- New thermal protection system using high-temperature carbon
fibre-reinforced plastic sandwich panel- Abstract: Publication date: Available online 7 February 2019Source: Acta AstronauticaAuthor(s): Yuuki Kubota, Ousei Miyamoto, Takuya Aoki, Yuichi Ishida, Toshio Ogasawara, Shinjiro Umezu Herein, a new lightweight thermal protection system was proposed to reduce the weight of a re-entry capsule. The proposed system employs an ablator as a thermal protection material, and a high-temperature polyimide carbon fibre-reinforced plastic (CFRP) sandwich panel as a structure member. The sandwich panel was designed to exhibit high thermal insulation and good mechanical properties at high temperature, and simultaneously ensure that the ablator has low thickness. Three-point bending was carried out at a high temperature up to 300 °C to evaluate shear strength and flexural rigidity of the sandwich panel. Further, the proposed thermal protection system was examined using an arc-heated plasma wind tunnel to evaluate its thermal protection performance and thermal response. The shear strength and the flexural rigidity of this high-temperature structure were 76.8 and 86.9% of that at room temperature, respectively, when evaluated at 300 °C. The proposed system achieved more than 40% weight reduction while maintaining high thermal insulation and recession resistance.
- Abstract: Publication date: Available online 7 February 2019Source: Acta AstronauticaAuthor(s): Yuuki Kubota, Ousei Miyamoto, Takuya Aoki, Yuichi Ishida, Toshio Ogasawara, Shinjiro Umezu Herein, a new lightweight thermal protection system was proposed to reduce the weight of a re-entry capsule. The proposed system employs an ablator as a thermal protection material, and a high-temperature polyimide carbon fibre-reinforced plastic (CFRP) sandwich panel as a structure member. The sandwich panel was designed to exhibit high thermal insulation and good mechanical properties at high temperature, and simultaneously ensure that the ablator has low thickness. Three-point bending was carried out at a high temperature up to 300 °C to evaluate shear strength and flexural rigidity of the sandwich panel. Further, the proposed thermal protection system was examined using an arc-heated plasma wind tunnel to evaluate its thermal protection performance and thermal response. The shear strength and the flexural rigidity of this high-temperature structure were 76.8 and 86.9% of that at room temperature, respectively, when evaluated at 300 °C. The proposed system achieved more than 40% weight reduction while maintaining high thermal insulation and recession resistance.
- The orbit deployment strategy of OOS system for refueling near-earth orbit
satellites- Abstract: Publication date: Available online 6 February 2019Source: Acta AstronauticaAuthor(s): Bo Meng, Jianbin Huang, Zhi Li, Longfei Huang, Yujia Pang, Xu Han, Zhimin Zhang Extend lifetime of near-earth orbit satellites by on-orbit refueling has significant economic benefits, while the orbit deployment strategy of the on-orbit servicing (OOS) system which implements the refueling task directly affects the efficiency of on-orbit refueling mission. Three orbit deployment modes were proposed, including ground deployment, low earth orbit (LEO) deployment and the same orbit deployment. For the client satellites to be refueled on near-earth orbit including LEO, medium earth orbit (MEO) and geosynchronous orbit (GEO), the three deployment modes were compared quantitatively based on two evaluation indicators: the refueling response time and the economic benefits. The comparing results showed that for MEO and GEO client satellites, it is appropriate to adopt the same orbit deployment mode, while for LEO client satellites, ground deployment mode is more suitable. As the maximum economic benefits can be earned to refuel GEO client satellites, the orbit deployment scheme of OOS system for refueling GEO client satellites was further studied. A new architecture of OOS system called “1 + N” consist of 1 fuel storage station and N refueling vehicles was proposed. The fuel storage station carries a great deal of fuel, running on orbit steadily. The refueling vehicles maneuver on GEO and implement the refueling operations for GEO client satellites. If the vehicles run out of fuel, they arrive at the station to be refueled. The weight and orbital altitude of fuel storage station, the number, weight, orbital altitude and orbital phase of refueling vehicles are the key parameters of the orbit deployment scheme. The mathematical model for optimizing the scheme was constructed, which takes shorten the refueling response time and lower the cost of OOS system as multi-objective. The calculation process of the optimization was explained. By analyzing the optimization results, the optimal orbit deployment scheme of OOS system containing 1 fuel storage station and 4 to 6 refueling vehicles was proposed.
- Abstract: Publication date: Available online 6 February 2019Source: Acta AstronauticaAuthor(s): Bo Meng, Jianbin Huang, Zhi Li, Longfei Huang, Yujia Pang, Xu Han, Zhimin Zhang Extend lifetime of near-earth orbit satellites by on-orbit refueling has significant economic benefits, while the orbit deployment strategy of the on-orbit servicing (OOS) system which implements the refueling task directly affects the efficiency of on-orbit refueling mission. Three orbit deployment modes were proposed, including ground deployment, low earth orbit (LEO) deployment and the same orbit deployment. For the client satellites to be refueled on near-earth orbit including LEO, medium earth orbit (MEO) and geosynchronous orbit (GEO), the three deployment modes were compared quantitatively based on two evaluation indicators: the refueling response time and the economic benefits. The comparing results showed that for MEO and GEO client satellites, it is appropriate to adopt the same orbit deployment mode, while for LEO client satellites, ground deployment mode is more suitable. As the maximum economic benefits can be earned to refuel GEO client satellites, the orbit deployment scheme of OOS system for refueling GEO client satellites was further studied. A new architecture of OOS system called “1 + N” consist of 1 fuel storage station and N refueling vehicles was proposed. The fuel storage station carries a great deal of fuel, running on orbit steadily. The refueling vehicles maneuver on GEO and implement the refueling operations for GEO client satellites. If the vehicles run out of fuel, they arrive at the station to be refueled. The weight and orbital altitude of fuel storage station, the number, weight, orbital altitude and orbital phase of refueling vehicles are the key parameters of the orbit deployment scheme. The mathematical model for optimizing the scheme was constructed, which takes shorten the refueling response time and lower the cost of OOS system as multi-objective. The calculation process of the optimization was explained. By analyzing the optimization results, the optimal orbit deployment scheme of OOS system containing 1 fuel storage station and 4 to 6 refueling vehicles was proposed.
- Modal Propellant Gauging: High-resolution and non-invasive gauging of both
settled and unsettled liquids in reduced gravity- Abstract: Publication date: Available online 6 February 2019Source: Acta AstronauticaAuthor(s): Kevin M. Crosby, Nehemiah J. Williams, Rudolph J. Werlink, Eric A. Hurlbert The modal response of a liquid-filled tank to external acoustic excitation can be used to infer with high resolution the mass of contained liquid, the mass flow rate of liquids into and out of the tank, and changes in tank pressure. Both contained liquid mass and internal ullage pressure affect the modal response of the tank walls through fluid mass-loading of the tank walls and pressure-induced wall stiffening, respectively. Modal Propellant Gauging refers to the technology that exploits these shifts in modal frequencies to infer the mass of propellant in a tank. MPG is a non-invasive gauging technology that has demonstrated gauging resolutions of 1% for settled propellants and 2–3% for unsettled, sloshing propellants. Extensive parabolic flight testing of the MPG system on model tanks has been conducted to validate the technology in reduced gravity. MPG testing on a qualification tank for the Orion Program's European Service Module has also been conducted and is reported here. Finite element modeling of the Orion ESM ″upper” tank is discussed and compared with measurement data. Three computational approaches to mass determination, Peak Tracking, Point Sensor, and Spectral Density methods, are described here. Use cases are defined and analyzed in the context of the Orion ESM Qualification tank data, and an implementation scheme for continuous mass gauging on the Orion ESM is discussed.
- Abstract: Publication date: Available online 6 February 2019Source: Acta AstronauticaAuthor(s): Kevin M. Crosby, Nehemiah J. Williams, Rudolph J. Werlink, Eric A. Hurlbert The modal response of a liquid-filled tank to external acoustic excitation can be used to infer with high resolution the mass of contained liquid, the mass flow rate of liquids into and out of the tank, and changes in tank pressure. Both contained liquid mass and internal ullage pressure affect the modal response of the tank walls through fluid mass-loading of the tank walls and pressure-induced wall stiffening, respectively. Modal Propellant Gauging refers to the technology that exploits these shifts in modal frequencies to infer the mass of propellant in a tank. MPG is a non-invasive gauging technology that has demonstrated gauging resolutions of 1% for settled propellants and 2–3% for unsettled, sloshing propellants. Extensive parabolic flight testing of the MPG system on model tanks has been conducted to validate the technology in reduced gravity. MPG testing on a qualification tank for the Orion Program's European Service Module has also been conducted and is reported here. Finite element modeling of the Orion ESM ″upper” tank is discussed and compared with measurement data. Three computational approaches to mass determination, Peak Tracking, Point Sensor, and Spectral Density methods, are described here. Use cases are defined and analyzed in the context of the Orion ESM Qualification tank data, and an implementation scheme for continuous mass gauging on the Orion ESM is discussed.
- Stationary configurations of space tether anchored on smaller primary in
three-body problem- Abstract: Publication date: Available online 6 February 2019Source: Acta AstronauticaAuthor(s): A.F.S. Ferreira, A.A. Burov, A.D. Guerman, A.F.B.A. Prado, V.I. Nikonov Spatial dynamics of a system composed by a planet, its moon and a spacecraft tethered to the moon surface is studied in the framework of circular restricted three-body problem. The moon is assumed to be in synchronous condition (1:1 mean motion resonances) so as to keep its orientation with respect to the planet; the size of the moon is non-negligible. The tether is considered to be light and inextensible. Equilibrium configurations of the tether are identified; their stability is analyzed. The bifurcation points, where the number of equilibria changes and new branches arise, are determined. The theoretical results are applied to particular cases of the Earth-Moon, Mars-Phobos and Pluto-Charon systems.
- Abstract: Publication date: Available online 6 February 2019Source: Acta AstronauticaAuthor(s): A.F.S. Ferreira, A.A. Burov, A.D. Guerman, A.F.B.A. Prado, V.I. Nikonov Spatial dynamics of a system composed by a planet, its moon and a spacecraft tethered to the moon surface is studied in the framework of circular restricted three-body problem. The moon is assumed to be in synchronous condition (1:1 mean motion resonances) so as to keep its orientation with respect to the planet; the size of the moon is non-negligible. The tether is considered to be light and inextensible. Equilibrium configurations of the tether are identified; their stability is analyzed. The bifurcation points, where the number of equilibria changes and new branches arise, are determined. The theoretical results are applied to particular cases of the Earth-Moon, Mars-Phobos and Pluto-Charon systems.
- Pre-flight qualification test procedure for nanosatellites using sounding
rockets- Abstract: Publication date: Available online 6 February 2019Source: Acta AstronauticaAuthor(s): Leonardo Kessler Slongo, João Gabriel Reis, Daniel Gaiki, Pedro Von Hohendorff Seger, Sara Vega Martínez, Bruno Vale Barbosa Eiterer, Tulio Gomes Pereira, Mario Baldini Neto, Matheus dos Santos Frata, Henrique Daniel Hamisch, André Martins Pio de Mattos, Juan Pablo Florez Mera, Kleber Vieira de Paiva, Marcia Barbosa Henriques Mantelli, Luigi Dilillo, Eduardo Augusto Bezerra This work presents an innovative procedure to test nanosatellites subsystems on board sounding rockets. The procedure allows the subsystems to save their telemetry data during the whole rocket flight phases (including at the rocket lift-off). The subsystems of the FloripaSat (1U CubeSat) engineering model have been tested on board a VSB-30 rocket in order to validate their design and integration process. A dedicated embedded system has been proposed to operate as an electronic interface between the nanosatellite subsystems and the rocket electronics. Also, a system to process telemetry data was implemented on the ground station. The test procedure validates data frame definition, communication protocol specification, hardware integration specification, among other functionalities. Three FloripaSat subsystems have been tested: Electrical Power Subsystem (EPS); Telemetry, Tracking and Command (TT&C); and On Board Data Handling (OBDH). Several nanosatellites’ functions have been evaluated, including: battery monitoring; inertial measurement unit; temperature measurements; radio transceiver; and beacon. The flight results were considerably different from the results obtained at the laboratory. This has allowed important modifications to be made on the design of the CubeSat flight model as the revision on the radio circuit on TT&C and the implementation of an external battery charger circuit for the EPS.
- Abstract: Publication date: Available online 6 February 2019Source: Acta AstronauticaAuthor(s): Leonardo Kessler Slongo, João Gabriel Reis, Daniel Gaiki, Pedro Von Hohendorff Seger, Sara Vega Martínez, Bruno Vale Barbosa Eiterer, Tulio Gomes Pereira, Mario Baldini Neto, Matheus dos Santos Frata, Henrique Daniel Hamisch, André Martins Pio de Mattos, Juan Pablo Florez Mera, Kleber Vieira de Paiva, Marcia Barbosa Henriques Mantelli, Luigi Dilillo, Eduardo Augusto Bezerra This work presents an innovative procedure to test nanosatellites subsystems on board sounding rockets. The procedure allows the subsystems to save their telemetry data during the whole rocket flight phases (including at the rocket lift-off). The subsystems of the FloripaSat (1U CubeSat) engineering model have been tested on board a VSB-30 rocket in order to validate their design and integration process. A dedicated embedded system has been proposed to operate as an electronic interface between the nanosatellite subsystems and the rocket electronics. Also, a system to process telemetry data was implemented on the ground station. The test procedure validates data frame definition, communication protocol specification, hardware integration specification, among other functionalities. Three FloripaSat subsystems have been tested: Electrical Power Subsystem (EPS); Telemetry, Tracking and Command (TT&C); and On Board Data Handling (OBDH). Several nanosatellites’ functions have been evaluated, including: battery monitoring; inertial measurement unit; temperature measurements; radio transceiver; and beacon. The flight results were considerably different from the results obtained at the laboratory. This has allowed important modifications to be made on the design of the CubeSat flight model as the revision on the radio circuit on TT&C and the implementation of an external battery charger circuit for the EPS.
- A compact C-Band CP-SAR microsatellite antenna for Earth Observation
- Abstract: Publication date: Available online 5 February 2019Source: Acta AstronauticaAuthor(s): Katia N. Urata, Josaphat T. Sri Sumantyo, Cahya E. Santosa, Tor Viscor This work describes the development of a compact Circularly Polarized SAR C-band antenna system and the design considerations suitable to use on small spacecrafts. To reduce size and weight of the small spaceborne SAR, we utilize a lightweight deployable parabolic mesh reflector and operate at low Earth orbital altitudes. The antenna is a wrap-rib center-fed parabolic reflector with dedicated receiving and transmitting feeds. Antenna requirements are: center frequency of 5.3 GHz with bandwidth of 400 MHz and circular polarization with axial ratio better than 3 dB. Simulation of the parabolic reflector and effects of different structural elements to the main radiation pattern is analyzed, which include ribs, struts, feed blockage and mesh surface. A research model of the parabolic reflector was constructed. Surface verification was realized using two different approaches, one using a laser distance meter along ribs and the other using 3D scanning of the reflector surface, with respective surface accuracy of 1.92 mm and 3.86 mm RMS. Near-field antenna measurements of the deployable reflector mesh antenna was realized for final antenna validation, presenting good agreement with the simulation results. Future work comprises of prototyping and testing of the full polarimetric feed assembly.
- Abstract: Publication date: Available online 5 February 2019Source: Acta AstronauticaAuthor(s): Katia N. Urata, Josaphat T. Sri Sumantyo, Cahya E. Santosa, Tor Viscor This work describes the development of a compact Circularly Polarized SAR C-band antenna system and the design considerations suitable to use on small spacecrafts. To reduce size and weight of the small spaceborne SAR, we utilize a lightweight deployable parabolic mesh reflector and operate at low Earth orbital altitudes. The antenna is a wrap-rib center-fed parabolic reflector with dedicated receiving and transmitting feeds. Antenna requirements are: center frequency of 5.3 GHz with bandwidth of 400 MHz and circular polarization with axial ratio better than 3 dB. Simulation of the parabolic reflector and effects of different structural elements to the main radiation pattern is analyzed, which include ribs, struts, feed blockage and mesh surface. A research model of the parabolic reflector was constructed. Surface verification was realized using two different approaches, one using a laser distance meter along ribs and the other using 3D scanning of the reflector surface, with respective surface accuracy of 1.92 mm and 3.86 mm RMS. Near-field antenna measurements of the deployable reflector mesh antenna was realized for final antenna validation, presenting good agreement with the simulation results. Future work comprises of prototyping and testing of the full polarimetric feed assembly.
- Fixed-time pinpoint mars landing using two sliding-surface autonomous
guidance- Abstract: Publication date: Available online 5 February 2019Source: Acta AstronauticaAuthor(s): Yao Zhang, Yanning Guo, Guangfu Ma, Bong Wie Autonomous powered-descent guidance algorithm for the pinpoint Mars landing in the presence of various disturbances and uncertainties is necessary for next-generation Mars exploration rover mission. This paper proposes a novel two sliding-surfaces guidance scheme based on fixed-time stabilization technique, which is robust against unknown Martian atmospheric disturbances (with known upper bound). For a fixed-time pinpoint landing mission, the main advantage of the proposed guidance is that the landing mission reliability can be ensured that Martian surface collision would never encounter. The fuel efficiency can be guaranteed in the comparison with the offline fuel optimal solution. The capacity of avoiding collisions is guaranteed by the monotonic convergence design of the proposed sliding modes. Lyapunov stabilization theory is adopted to prove the global stability of the proposed guidance. Monte Carlo numerical simulations are implemented in a realistic scenario and the results confirm the collision avoidance capability, the fuel efficiency and the robustness of the proposed guidance.
- Abstract: Publication date: Available online 5 February 2019Source: Acta AstronauticaAuthor(s): Yao Zhang, Yanning Guo, Guangfu Ma, Bong Wie Autonomous powered-descent guidance algorithm for the pinpoint Mars landing in the presence of various disturbances and uncertainties is necessary for next-generation Mars exploration rover mission. This paper proposes a novel two sliding-surfaces guidance scheme based on fixed-time stabilization technique, which is robust against unknown Martian atmospheric disturbances (with known upper bound). For a fixed-time pinpoint landing mission, the main advantage of the proposed guidance is that the landing mission reliability can be ensured that Martian surface collision would never encounter. The fuel efficiency can be guaranteed in the comparison with the offline fuel optimal solution. The capacity of avoiding collisions is guaranteed by the monotonic convergence design of the proposed sliding modes. Lyapunov stabilization theory is adopted to prove the global stability of the proposed guidance. Monte Carlo numerical simulations are implemented in a realistic scenario and the results confirm the collision avoidance capability, the fuel efficiency and the robustness of the proposed guidance.
- Design and surface modification of a microfluidic chip for intercellular
interactions research during space flight- Abstract: Publication date: Available online 5 February 2019Source: Acta AstronauticaAuthor(s): Yu Chen, Sizhu Pei, Liben Yan, Xin Qiu, Rui Li, Shiyong Yu, Adil Lodhi, Bo Li, Robina Manzoor, Peng Zhang, Yulin Deng, Hong Ma Intercellular interactions widely exist in multicellular organisms. The exposure of astronaut's body to space environment results in a series of biological effects including intercellular interactions. However, these interactions have not been studied extensively in space because of the difficulties faced in performing such experiments in the space. To solve this problem, we have designed a co-culture microfluidic chip for studying intercellular interactions and provides an effectively dynamic co-culture method to both adherent cells and suspension cells. Its structure consists of two cell chambers which are divided by polycarbonate semipermeable membrane. The membrane is permeable to signal molecules which are secreted by cells but it is impermeable to the cells itself. Each cell chamber is divided by bolting silk. This results in a control of flow shear stress exerted on the cells and it also results in trapping the suspended cells. As surface property of the base of any microfluidic chip is important, therefore, we optimized a surface modification strategy using MTS assays and water contact angle test. The results show that optimum surface modification strategy is using air plasma treated polystyrene surface for 90s. Moreover, the contact angle recovery after plasma treatment indicated that the co-culture microfluidic chip should be seeded within 6 days after surface modification. Our results show that a successful dynamic cell co-culture was achieved using this design. We conclude that this co-culture microfluidic chip can be a valuable tool for investigating intercellular interactions in space as it can be operated automatically during a space flight.
- Abstract: Publication date: Available online 5 February 2019Source: Acta AstronauticaAuthor(s): Yu Chen, Sizhu Pei, Liben Yan, Xin Qiu, Rui Li, Shiyong Yu, Adil Lodhi, Bo Li, Robina Manzoor, Peng Zhang, Yulin Deng, Hong Ma Intercellular interactions widely exist in multicellular organisms. The exposure of astronaut's body to space environment results in a series of biological effects including intercellular interactions. However, these interactions have not been studied extensively in space because of the difficulties faced in performing such experiments in the space. To solve this problem, we have designed a co-culture microfluidic chip for studying intercellular interactions and provides an effectively dynamic co-culture method to both adherent cells and suspension cells. Its structure consists of two cell chambers which are divided by polycarbonate semipermeable membrane. The membrane is permeable to signal molecules which are secreted by cells but it is impermeable to the cells itself. Each cell chamber is divided by bolting silk. This results in a control of flow shear stress exerted on the cells and it also results in trapping the suspended cells. As surface property of the base of any microfluidic chip is important, therefore, we optimized a surface modification strategy using MTS assays and water contact angle test. The results show that optimum surface modification strategy is using air plasma treated polystyrene surface for 90s. Moreover, the contact angle recovery after plasma treatment indicated that the co-culture microfluidic chip should be seeded within 6 days after surface modification. Our results show that a successful dynamic cell co-culture was achieved using this design. We conclude that this co-culture microfluidic chip can be a valuable tool for investigating intercellular interactions in space as it can be operated automatically during a space flight.
- Comment on “the ultimate limits of the relativistic rocket equation. The
Planck photon rocket”- Abstract: Publication date: Available online 2 February 2019Source: Acta AstronauticaAuthor(s): Daniele Tommasini, Angel Paredes, Humberto Michinel We disprove the claim that was published in a recent paper [1] about an absolute maximum velocity related to the value of the Planck mass for a photon-propelled rocket in vacuum. This claim is based on incorrect hypotheses and is incompatible with the theory of relativity. We also review the relative bound that can actually be derived for the velocity of such an ideal rocket, as measured in the inertial frame in which it was at rest at a given initial time.
- Abstract: Publication date: Available online 2 February 2019Source: Acta AstronauticaAuthor(s): Daniele Tommasini, Angel Paredes, Humberto Michinel We disprove the claim that was published in a recent paper [1] about an absolute maximum velocity related to the value of the Planck mass for a photon-propelled rocket in vacuum. This claim is based on incorrect hypotheses and is incompatible with the theory of relativity. We also review the relative bound that can actually be derived for the velocity of such an ideal rocket, as measured in the inertial frame in which it was at rest at a given initial time.
- Two regimes of a single n-heptane droplet combustion
- Abstract: Publication date: Available online 2 February 2019Source: Acta AstronauticaAuthor(s): Veronika Tyurenkova This paper is devoted to the burning of an isolated n-heptane droplet in microgravity and analysis of the experimental data presented by FLEX experiment. Developing the classical analytical research methodology, proposed by Forman A. Williams, and using our obtained analytical solutions for the problem of droplet equilibrium evaporation and combustion, the analytical solutions to solve the problem of two regimes of droplet burning were found. To explain the presence of the radiative flame extinction and further droplet burning in cool flame regime, we assumed that two flame surfaces exist, but the first flame surface can be fixed with observation devices, and the second flame surface is an invisible one. According to this hypothesis, we suggested the analytical formulas for droplet and flame histories. The analytical solutions demonstrate a good coincidence with the experimental and numerical results.
- Abstract: Publication date: Available online 2 February 2019Source: Acta AstronauticaAuthor(s): Veronika Tyurenkova This paper is devoted to the burning of an isolated n-heptane droplet in microgravity and analysis of the experimental data presented by FLEX experiment. Developing the classical analytical research methodology, proposed by Forman A. Williams, and using our obtained analytical solutions for the problem of droplet equilibrium evaporation and combustion, the analytical solutions to solve the problem of two regimes of droplet burning were found. To explain the presence of the radiative flame extinction and further droplet burning in cool flame regime, we assumed that two flame surfaces exist, but the first flame surface can be fixed with observation devices, and the second flame surface is an invisible one. According to this hypothesis, we suggested the analytical formulas for droplet and flame histories. The analytical solutions demonstrate a good coincidence with the experimental and numerical results.
- Effect of multi-walled carbon nanotubes on thermal stability and ablation
properties of EPDM insulation materials for solid rocket motors- Abstract: Publication date: Available online 1 February 2019Source: Acta AstronauticaAuthor(s): Mengfei Guo, Jiang Li, Kun Xi, Yang Liu, Jianbo Ji Multi-walled carbon nanotubes (MWCNTs) have excellent physicochemical properties and can be used as fillers to improve the performance of insulation materials. However, there is currently no proper understanding of the mechanism by which MWCNTs improve the ablation performance of insulation materials. Here, we prepare two formulations of ethylene propylene diene monomer (EPDM) insulation materials with and without MWCNTs and study their thermal stability and ablation characteristics via thermogravimetric analysis, high temperature char layer residue experiments, oxyacetylene ablation tests, and other analytical methods along with scanning electron microscopy and energy dispersive X-ray spectroscopy. The intrinsic mechanism by which MWCNTs improve the ablation performance of insulation materials is revealed from the perspectives of char layer formation, consumption, and micro-morphology. The addition of MWCNTs increased the carbon residue rate after thermal decomposition and can suppress the consumption reaction of the char layer under high-temperature conditions. Microscopic morphological analyses show that the insulation material containing MWCNTs has a denser char layer surface and a network-like structure inside, which improve the ability of the char layer to resist erosion by gas flow. The combination of these factors improves the ablation performance of EPDM. Our findings regarding the MWCNT reinforcement mechanism can guide the design of new insulation materials.
- Abstract: Publication date: Available online 1 February 2019Source: Acta AstronauticaAuthor(s): Mengfei Guo, Jiang Li, Kun Xi, Yang Liu, Jianbo Ji Multi-walled carbon nanotubes (MWCNTs) have excellent physicochemical properties and can be used as fillers to improve the performance of insulation materials. However, there is currently no proper understanding of the mechanism by which MWCNTs improve the ablation performance of insulation materials. Here, we prepare two formulations of ethylene propylene diene monomer (EPDM) insulation materials with and without MWCNTs and study their thermal stability and ablation characteristics via thermogravimetric analysis, high temperature char layer residue experiments, oxyacetylene ablation tests, and other analytical methods along with scanning electron microscopy and energy dispersive X-ray spectroscopy. The intrinsic mechanism by which MWCNTs improve the ablation performance of insulation materials is revealed from the perspectives of char layer formation, consumption, and micro-morphology. The addition of MWCNTs increased the carbon residue rate after thermal decomposition and can suppress the consumption reaction of the char layer under high-temperature conditions. Microscopic morphological analyses show that the insulation material containing MWCNTs has a denser char layer surface and a network-like structure inside, which improve the ability of the char layer to resist erosion by gas flow. The combination of these factors improves the ablation performance of EPDM. Our findings regarding the MWCNT reinforcement mechanism can guide the design of new insulation materials.
- Performance assessment of energy deposition based drag reduction technique
for Earth and Mars flight conditions- Abstract: Publication date: Available online 1 February 2019Source: Acta AstronauticaAuthor(s): Dipankar Das, Siddesh Desai, Vinayak Kulkarni, Hrishikesh Gadgil Energy deposition based drag reduction technique is numerically investigated for Earth and Mars atmospheric conditions using in-house non-equilibrium flow solvers. Golden section search algorithm is successfully integrated with these CFD solvers to evaluate the optimum amount of energy to be deposited for maximum power effectiveness. During these studies for spherical configuration, it has been noticed that, the amount of energy deposited corresponding to maximum power effectiveness is more for Mars case than the Earth case. Higher temperature dependence of specific heat for carbon dioxide than air is accounted to be responsible for this observation. Further, maximum power effectiveness is marked to be lower for carbon dioxide flow as compared to that for air flow. It has also been noticed that the performance of this technique gets lowered with increase in stagnation enthalpy of the flow, in either cases, with remarkable drop in case of air.
- Abstract: Publication date: Available online 1 February 2019Source: Acta AstronauticaAuthor(s): Dipankar Das, Siddesh Desai, Vinayak Kulkarni, Hrishikesh Gadgil Energy deposition based drag reduction technique is numerically investigated for Earth and Mars atmospheric conditions using in-house non-equilibrium flow solvers. Golden section search algorithm is successfully integrated with these CFD solvers to evaluate the optimum amount of energy to be deposited for maximum power effectiveness. During these studies for spherical configuration, it has been noticed that, the amount of energy deposited corresponding to maximum power effectiveness is more for Mars case than the Earth case. Higher temperature dependence of specific heat for carbon dioxide than air is accounted to be responsible for this observation. Further, maximum power effectiveness is marked to be lower for carbon dioxide flow as compared to that for air flow. It has also been noticed that the performance of this technique gets lowered with increase in stagnation enthalpy of the flow, in either cases, with remarkable drop in case of air.
- Effects of an accompanied gas jet on transverse liquid injection in a
supersonic crossflow- Abstract: Publication date: Available online 31 January 2019Source: Acta AstronauticaAuthor(s): Runsheng Hu, Qinglian Li, Chun Li, Chenyang Li A liquid injection scheme in a supersonic crossflow, namely transverse liquid injection accompanied with a gas jet, is proposed and investigated experimentally and numerically. High-speed photography was adopted to capture the jet spray in the experimental work; while a discrete phase model was utilized to simulate the liquid jet penetration and gas-liquid interaction in the numerical study. The differences between transverse liquid injection with and without an accompanied gas jet were analyzed. The effects of an accompanied gas jet were determined. Both the experimental and computational results indicate that an accompanied gas jet can significantly enhance liquid jet penetration. Simulation results also indicate that gas injection would cause additional total pressure losses.
- Abstract: Publication date: Available online 31 January 2019Source: Acta AstronauticaAuthor(s): Runsheng Hu, Qinglian Li, Chun Li, Chenyang Li A liquid injection scheme in a supersonic crossflow, namely transverse liquid injection accompanied with a gas jet, is proposed and investigated experimentally and numerically. High-speed photography was adopted to capture the jet spray in the experimental work; while a discrete phase model was utilized to simulate the liquid jet penetration and gas-liquid interaction in the numerical study. The differences between transverse liquid injection with and without an accompanied gas jet were analyzed. The effects of an accompanied gas jet were determined. Both the experimental and computational results indicate that an accompanied gas jet can significantly enhance liquid jet penetration. Simulation results also indicate that gas injection would cause additional total pressure losses.
- Mars entry trajectory robust optimization based on evidence under
epistemic uncertainty- Abstract: Publication date: Available online 31 January 2019Source: Acta AstronauticaAuthor(s): Yuechen Huang, Haiyang Li, Xin Du, Xiangyue He The epistemic uncertainties caused by insufficient knowledge of the atmosphere, aerodynamic coefficient, and entry state render the entry process challenging, as they not only result in the deviation of the preplanned trajectory, but also may lead to the nonsatisfaction of path constraints. Herein, a robust epistemic uncertainty optimization (REUO) method based on evidence is proposed to solve the Mars entry trajectory optimization problem under epistemic uncertainty. A two-loop nested robust optimization (RO) model is formulated, in which the outer-loop optimization searches the optimal control while the inner-loop optimization calculates the extremal trajectory performances within each focal element (FE) to evaluate the evidence level. To solve the path constraint violation problem under uncertainties, the constraint design based on limitation bounds is considered in the RO model. The polynomial chaos expansion (PCE) is employed to obtain the approximate analytic function of the trajectory performance under uncertainties. Thereafter, the optimization based on ordinary stochastic entry dynamics in the inner loop is replaced by a simple parameter optimization of the analytic functions, which can be readily and rapidly solved. The REUO method is tested in a specific Mars entry mission. The simulation results show that the proposed method can identify the most robust solutions with the optimal trajectory performance under epistemic uncertainties.
- Abstract: Publication date: Available online 31 January 2019Source: Acta AstronauticaAuthor(s): Yuechen Huang, Haiyang Li, Xin Du, Xiangyue He The epistemic uncertainties caused by insufficient knowledge of the atmosphere, aerodynamic coefficient, and entry state render the entry process challenging, as they not only result in the deviation of the preplanned trajectory, but also may lead to the nonsatisfaction of path constraints. Herein, a robust epistemic uncertainty optimization (REUO) method based on evidence is proposed to solve the Mars entry trajectory optimization problem under epistemic uncertainty. A two-loop nested robust optimization (RO) model is formulated, in which the outer-loop optimization searches the optimal control while the inner-loop optimization calculates the extremal trajectory performances within each focal element (FE) to evaluate the evidence level. To solve the path constraint violation problem under uncertainties, the constraint design based on limitation bounds is considered in the RO model. The polynomial chaos expansion (PCE) is employed to obtain the approximate analytic function of the trajectory performance under uncertainties. Thereafter, the optimization based on ordinary stochastic entry dynamics in the inner loop is replaced by a simple parameter optimization of the analytic functions, which can be readily and rapidly solved. The REUO method is tested in a specific Mars entry mission. The simulation results show that the proposed method can identify the most robust solutions with the optimal trajectory performance under epistemic uncertainties.
- Non-invasive wearable ECG-patch system for astronauts and patients on
Earth- Abstract: Publication date: Available online 31 January 2019Source: Acta AstronauticaAuthor(s): Natalia Glazkova, Tatiana Podladchikova, Rupert Gerzer, Daria Stepanova Ambulatory long term cardiac diagnostics is still not optimally solved. We have developed a new system that can overcome this gap in diagnostics. The new system consists of an adhesive patch-like wearable miniaturized device for electrocardiogram (ECG) data acquisition with further processing of acquired data. The new patch-like wearable device is able to record data for up to 7 days and significantly reduces discomfort to the user by its cable-less design, small size and direct attachment to the skin surface. This three lead system also detects ECG waveform morphology. It is aimed to spot heart rate, irregular beat episodes and circadian variations. Such a device has high potential to be used for both aerospace and terrestrial applications.
- Abstract: Publication date: Available online 31 January 2019Source: Acta AstronauticaAuthor(s): Natalia Glazkova, Tatiana Podladchikova, Rupert Gerzer, Daria Stepanova Ambulatory long term cardiac diagnostics is still not optimally solved. We have developed a new system that can overcome this gap in diagnostics. The new system consists of an adhesive patch-like wearable miniaturized device for electrocardiogram (ECG) data acquisition with further processing of acquired data. The new patch-like wearable device is able to record data for up to 7 days and significantly reduces discomfort to the user by its cable-less design, small size and direct attachment to the skin surface. This three lead system also detects ECG waveform morphology. It is aimed to spot heart rate, irregular beat episodes and circadian variations. Such a device has high potential to be used for both aerospace and terrestrial applications.
- Abnormal enhancement of separated turbulent air flow and heat transfer in
inclined single-row oval-trench dimples at the narrow channel wall- Abstract: Publication date: Available online 31 January 2019Source: Acta AstronauticaAuthor(s): Sergey Isaev, Мichael Grintsevich, Аlexandr Leontiev, Igor Popov New shapes of surface vortex generators are proposed which allow improving the energy performance of channels of propulsion devices and contribute towards the safety of space flights. Abnormal enhancement of heat transfer and separated turbulent air flow in oval-trench dimples of depth 0.3 inclined at an angle of 45о to the incoming flow and located in-line at the heated wall of the narrow channel is accompanied by a four-fold growth of relative negative friction and an almost five-fold increase of relative heat transfer in the separation zone. At the same time, a maximum secondary flow velocity in the vicinity of the dimple reaches a maximum flow velocity in the plane-parallel channel. A reason for the onset of this phenomenon (ultra intensification of transport processes) caused by the formation of a high pressure drop between closely spaced stagnation zones on the windward side of the dimple and a low pressure on the entrance spherical portion of the dimple is disclosed.
- Abstract: Publication date: Available online 31 January 2019Source: Acta AstronauticaAuthor(s): Sergey Isaev, Мichael Grintsevich, Аlexandr Leontiev, Igor Popov New shapes of surface vortex generators are proposed which allow improving the energy performance of channels of propulsion devices and contribute towards the safety of space flights. Abnormal enhancement of heat transfer and separated turbulent air flow in oval-trench dimples of depth 0.3 inclined at an angle of 45о to the incoming flow and located in-line at the heated wall of the narrow channel is accompanied by a four-fold growth of relative negative friction and an almost five-fold increase of relative heat transfer in the separation zone. At the same time, a maximum secondary flow velocity in the vicinity of the dimple reaches a maximum flow velocity in the plane-parallel channel. A reason for the onset of this phenomenon (ultra intensification of transport processes) caused by the formation of a high pressure drop between closely spaced stagnation zones on the windward side of the dimple and a low pressure on the entrance spherical portion of the dimple is disclosed.
- Spray coating process of MWCNT/epoxy nanocomposite films for aerospace
applications: Effects of process parameters on surface electrical
properties- Abstract: Publication date: Available online 31 January 2019Source: Acta AstronauticaAuthor(s): Susanna Laurenzi, Marialaura Clausi, Federica Zaccardi, Ugo Curt, M. Gabriella Santonicola Nanocomposite coatings on flexible membranes possess many interesting properties useful in aerospace applications, such as thermal blankets and charging mitigation layers. However, their fabrication is not trivial and the overall performance is strongly dependent on the manufacturing process, since it influences the filler distribution and so the homogeneity of the coated film. In this work, carbon nanotube/epoxy nanocomposite films were deposited on Mylar substrates by spray coating process. The effects of the process parameters on the morphology and on the electrical properties of the nanocomposite films were investigated. In particular, the influence of the nozzle diameters and the distance between the nozzle and the target substrate, as well as the concentration of the multiwalled carbon nanotubes (MWCNT) were studied. The electrical properties of the MWCNT/epoxy nanocomposite surfaces were determined using impedance spectroscopy in the frequency range from 20 Hz to 2 MHz. The analysis was performed on samples extracted from different regions of the nanocomposite films, in order to assess their property homogeneity.
- Abstract: Publication date: Available online 31 January 2019Source: Acta AstronauticaAuthor(s): Susanna Laurenzi, Marialaura Clausi, Federica Zaccardi, Ugo Curt, M. Gabriella Santonicola Nanocomposite coatings on flexible membranes possess many interesting properties useful in aerospace applications, such as thermal blankets and charging mitigation layers. However, their fabrication is not trivial and the overall performance is strongly dependent on the manufacturing process, since it influences the filler distribution and so the homogeneity of the coated film. In this work, carbon nanotube/epoxy nanocomposite films were deposited on Mylar substrates by spray coating process. The effects of the process parameters on the morphology and on the electrical properties of the nanocomposite films were investigated. In particular, the influence of the nozzle diameters and the distance between the nozzle and the target substrate, as well as the concentration of the multiwalled carbon nanotubes (MWCNT) were studied. The electrical properties of the MWCNT/epoxy nanocomposite surfaces were determined using impedance spectroscopy in the frequency range from 20 Hz to 2 MHz. The analysis was performed on samples extracted from different regions of the nanocomposite films, in order to assess their property homogeneity.
- Performance of polyimide film under hypervelocity impact of micro flyer:
Experiments and simulations- Abstract: Publication date: Available online 31 January 2019Source: Acta AstronauticaAuthor(s): Tao Liu, Zhixin Zeng, Xinghua Zhang, Xinming Qiu, ZhengAi Cheng, Li Wang, Shaoxia Jia, Jian Cai The characteristics of polyimide film hypervelocity impacted (HVI) by micro flyer are investigated though experimental and numerical methods. Using the laser driven flyer (LDF) method combined with the technique of scanning electron microscopy (SEM) and explicit dynamics numerical algorithm, the fracture morphology and detailed mechanical responses of polyimide film, both in the impact and non–impact zones, are captured; and a dynamic fracture strain criteria, combing temperature and strain rate, is presented. A parametric study on the performance of polyimide film is implemented and the influence of the factors, including the impact velocity, film thickness, strain rate constant and pre–stress, are assessed. The research reveals that polyimide film exits two failure modes, i.e. ductile punching and brittle cracks. Polyimide materials in the impact zone have experienced a high temperature (>Tg = 685 K), ultrahigh strain rate (108∼109) and plastic state; while in the non–impact zone, it is normal temperature (293–295 K), high strain rate (105∼106) and elastic state.
- Abstract: Publication date: Available online 31 January 2019Source: Acta AstronauticaAuthor(s): Tao Liu, Zhixin Zeng, Xinghua Zhang, Xinming Qiu, ZhengAi Cheng, Li Wang, Shaoxia Jia, Jian Cai The characteristics of polyimide film hypervelocity impacted (HVI) by micro flyer are investigated though experimental and numerical methods. Using the laser driven flyer (LDF) method combined with the technique of scanning electron microscopy (SEM) and explicit dynamics numerical algorithm, the fracture morphology and detailed mechanical responses of polyimide film, both in the impact and non–impact zones, are captured; and a dynamic fracture strain criteria, combing temperature and strain rate, is presented. A parametric study on the performance of polyimide film is implemented and the influence of the factors, including the impact velocity, film thickness, strain rate constant and pre–stress, are assessed. The research reveals that polyimide film exits two failure modes, i.e. ductile punching and brittle cracks. Polyimide materials in the impact zone have experienced a high temperature (>Tg = 685 K), ultrahigh strain rate (108∼109) and plastic state; while in the non–impact zone, it is normal temperature (293–295 K), high strain rate (105∼106) and elastic state.
- Development of effective and safe compound disinfectant for space cabins
- Abstract: Publication date: Available online 30 January 2019Source: Acta AstronauticaAuthor(s): Lei Li, Yuming Fu, Hong Liu Space-enclosed cabins such as space stations provide a suitable living environment for astronauts, meanwhile they also facilitate various microorganisms to breed and corrode the materials’ surfaces. To effectively prevent and control the microbial contamination in space-closed cabins and ensure the health of the astronauts, it is necessary to develop effective and safe disinfectants. In this study, organosilicon quaternary ammonium salt(OQAS), sodium dehydroacetate (SD) and polyhexamethylene biguanidine hydrochloride(PHMB) were selected as the target disinfectants and tested on four kinds of corrosive fungi, including Cladosporium cladosporioides, Aspergillus niger, Penicillium aurantiogriseum and Cladosporium herbarum. Firstly, the mycelial growth inhibition effects of these three disinfectants on the four fungi were tested separately, and the effective concentration range of the three disinfectants were determined. Then we combined these three disinfectants to get a new highly effective and safe compound bactericides, consisting of OQAS (0.37 mg/ml), SD (0.7 mg/ml) and PHMB (0.5 mg/ml), in which the total concentrations of bactericides were 1.57 mg/ml. This study provided a new perspective for the development of safe microbial bacteriostatic agents in spacecraft.Graphical abstract
- Abstract: Publication date: Available online 30 January 2019Source: Acta AstronauticaAuthor(s): Lei Li, Yuming Fu, Hong Liu Space-enclosed cabins such as space stations provide a suitable living environment for astronauts, meanwhile they also facilitate various microorganisms to breed and corrode the materials’ surfaces. To effectively prevent and control the microbial contamination in space-closed cabins and ensure the health of the astronauts, it is necessary to develop effective and safe disinfectants. In this study, organosilicon quaternary ammonium salt(OQAS), sodium dehydroacetate (SD) and polyhexamethylene biguanidine hydrochloride(PHMB) were selected as the target disinfectants and tested on four kinds of corrosive fungi, including Cladosporium cladosporioides, Aspergillus niger, Penicillium aurantiogriseum and Cladosporium herbarum. Firstly, the mycelial growth inhibition effects of these three disinfectants on the four fungi were tested separately, and the effective concentration range of the three disinfectants were determined. Then we combined these three disinfectants to get a new highly effective and safe compound bactericides, consisting of OQAS (0.37 mg/ml), SD (0.7 mg/ml) and PHMB (0.5 mg/ml), in which the total concentrations of bactericides were 1.57 mg/ml. This study provided a new perspective for the development of safe microbial bacteriostatic agents in spacecraft.Graphical abstract
- Knockdown of CD44 inhibits the alteration of osteoclast function induced
by simulated microgravity- Abstract: Publication date: Available online 25 January 2019Source: Acta AstronauticaAuthor(s): Yuheng Li, Xingcheng Gao, Shukuan Ling, Guohui Zhong, Weijia Sun, Caizhi Liu, Jianwei Li, Jinping Song, Dingsheng Zhao, Xiaoyan Jin, Yingxian Li, Xi-qing Sun Mechanical forces are essential to maintain skeletal homeostasis, and microgravity exposure leads to bone loss. During space flight, bone mineral density is decreased because of the inactivation of osteoblast and activation of osteoclast. However, the underlying molecular mechanism is still not clear. CD44 acts as a cellular surface adhesion molecule, which plays an important role in signal transduction between cells. Numerous studies have shown that CD44 plays diverse roles in promoting pre-osteoclast fusion. However, the role of CD44 in the alteration of osteoclast function induced by simulated microgravity remains to be fully elucidated. In this work, we utilized a ground-based, microgravity-simulation system, the Rotating Wall Vessel Bioreactor (RWVB). Using the RWVB, we demonstrated that simulated microgravity enhanced osteoclast function, meantime CD44 mRNA and protein levels were significantly increased. To verify the effects of CD44 on alteration of osteoclast function induced by simulated microgravity, specific siRNAs targeting CD44 were transfected. The results showed knockdown of CD44 inhibited the alteration of osteoclast function induced by simulated microgravity. Moreover, we found that clinorotation activated the NF-κB/NFATc1-mediated signaling pathway, which was downregulated after knockdown of CD44. Our study provided evidence that CD44 positively regulated osteoclast function and therapeutic suppression of CD44 may counteract bone loss induced by simulated microgravity.
- Abstract: Publication date: Available online 25 January 2019Source: Acta AstronauticaAuthor(s): Yuheng Li, Xingcheng Gao, Shukuan Ling, Guohui Zhong, Weijia Sun, Caizhi Liu, Jianwei Li, Jinping Song, Dingsheng Zhao, Xiaoyan Jin, Yingxian Li, Xi-qing Sun Mechanical forces are essential to maintain skeletal homeostasis, and microgravity exposure leads to bone loss. During space flight, bone mineral density is decreased because of the inactivation of osteoblast and activation of osteoclast. However, the underlying molecular mechanism is still not clear. CD44 acts as a cellular surface adhesion molecule, which plays an important role in signal transduction between cells. Numerous studies have shown that CD44 plays diverse roles in promoting pre-osteoclast fusion. However, the role of CD44 in the alteration of osteoclast function induced by simulated microgravity remains to be fully elucidated. In this work, we utilized a ground-based, microgravity-simulation system, the Rotating Wall Vessel Bioreactor (RWVB). Using the RWVB, we demonstrated that simulated microgravity enhanced osteoclast function, meantime CD44 mRNA and protein levels were significantly increased. To verify the effects of CD44 on alteration of osteoclast function induced by simulated microgravity, specific siRNAs targeting CD44 were transfected. The results showed knockdown of CD44 inhibited the alteration of osteoclast function induced by simulated microgravity. Moreover, we found that clinorotation activated the NF-κB/NFATc1-mediated signaling pathway, which was downregulated after knockdown of CD44. Our study provided evidence that CD44 positively regulated osteoclast function and therapeutic suppression of CD44 may counteract bone loss induced by simulated microgravity.
- Global fixed-time attitude tracking control for the rigid spacecraft with
actuator saturation and faults- Abstract: Publication date: February 2019Source: Acta Astronautica, Volume 155Author(s): Xiao-Ning Shi, Yong-An Zhang, Di Zhou, Zhi-Gang Zhou This paper investigates the global finite-time attitude tracking problem for the rigid spacecraft subject to inertial uncertainties, external disturbances, actuator faults, and input saturation constraints. The exponential coordinates vector in conjunction with a hysteretic-based jump condition is introduced to overcome the topological obstacles of global stability on the special orthogonal group. A novel nonsingular fixed-time-based sliding mode is designed, which not only avoids the singularity but also guarantees that the convergence time of tracking errors along the sliding surface is independent of the state value. Then, an adaptive fault-tolerant control law is constructed to enforce the system state to reach a neighborhood of the sliding surface in the sense of the fixed-time concept, which can accommodate actuator failures under limited control torque. The total convergence time is independent of the initial conditions information. A rigorous mathematical stability Proof is given. Numerical simulations are finally performed to demonstrate the effectiveness of the proposed finite-time controller.
- Abstract: Publication date: February 2019Source: Acta Astronautica, Volume 155Author(s): Xiao-Ning Shi, Yong-An Zhang, Di Zhou, Zhi-Gang Zhou This paper investigates the global finite-time attitude tracking problem for the rigid spacecraft subject to inertial uncertainties, external disturbances, actuator faults, and input saturation constraints. The exponential coordinates vector in conjunction with a hysteretic-based jump condition is introduced to overcome the topological obstacles of global stability on the special orthogonal group. A novel nonsingular fixed-time-based sliding mode is designed, which not only avoids the singularity but also guarantees that the convergence time of tracking errors along the sliding surface is independent of the state value. Then, an adaptive fault-tolerant control law is constructed to enforce the system state to reach a neighborhood of the sliding surface in the sense of the fixed-time concept, which can accommodate actuator failures under limited control torque. The total convergence time is independent of the initial conditions information. A rigorous mathematical stability Proof is given. Numerical simulations are finally performed to demonstrate the effectiveness of the proposed finite-time controller.
- Orbit information of predetermined accuracy and its sharing in the space
situational awareness context- Abstract: Publication date: Available online 25 January 2019Source: Acta AstronauticaAuthor(s): Vitali Braun, Heiner Klinkrad, Enrico Stoll A method to deduce orbit solutions of predetermined accuracy with respect to a reference orbit is presented. The idea is to guarantee the applicability to all orbits and provide both the orbit and the covariance matrix with a constrained method error. Such a solution may be useful, if different users expect different levels of accuracy or have limited privileges to access high-precision data. The method proposed in this paper performs a non-linear least squares fit to the catalogued reference orbit with an modified force model. Subsequently, that solution is subjected to an ephemeris compression method and the results thereof can be provided to the users in terms of Chebyshev polynomial coefficients, after being put in a standardised data message. One of the main advantages is that there is no need for a propagation on the user's side anymore. The orbit information is obtained directly from the series representation in the data message for any time in the in the provided interval. Moreover, the orbit theory used in the orbit determination and catalog maintenance is not required to be shared or synchronized with the community any longer.
- Abstract: Publication date: Available online 25 January 2019Source: Acta AstronauticaAuthor(s): Vitali Braun, Heiner Klinkrad, Enrico Stoll A method to deduce orbit solutions of predetermined accuracy with respect to a reference orbit is presented. The idea is to guarantee the applicability to all orbits and provide both the orbit and the covariance matrix with a constrained method error. Such a solution may be useful, if different users expect different levels of accuracy or have limited privileges to access high-precision data. The method proposed in this paper performs a non-linear least squares fit to the catalogued reference orbit with an modified force model. Subsequently, that solution is subjected to an ephemeris compression method and the results thereof can be provided to the users in terms of Chebyshev polynomial coefficients, after being put in a standardised data message. One of the main advantages is that there is no need for a propagation on the user's side anymore. The orbit information is obtained directly from the series representation in the data message for any time in the in the provided interval. Moreover, the orbit theory used in the orbit determination and catalog maintenance is not required to be shared or synchronized with the community any longer.
- A dimension-reduction solution of free-time differential games for
spacecraft pursuit-evasion- Abstract: Publication date: Available online 23 January 2019Source: Acta AstronauticaAuthor(s): Zhen-yu Li, Hai Zhu, Zhen Yang, Ya-zhong Luo This paper presents a dimension-reduction method for solving the free-time pursuit-evasion game between two spacecraft near circular orbits. Theoretically, finding the saddle-point solution of the game results in solving a high-dimensional (twenty-four in our case) two-point boundary value problem (TPBVP), which is quite difficult and computationally intensive. By using the circular-orbit variational equations to model the relative states of two spacecraft near circular orbits, the 24-dimensional TPBVP is firstly transformed into solving a set of four-dimensional nonlinear equations. Then, a hybrid numerical algorithm is proposed to solve the equations, where the differential evolution algorithm is used to obtain an initial guess and the Newton's iteration method is used to find the accurate solution. Numerical results show that the proposed method is more efficient than previous methods, and that the proposed method are suitable for large-phasing-distance orbital pursuit-evasion problems.
- Abstract: Publication date: Available online 23 January 2019Source: Acta AstronauticaAuthor(s): Zhen-yu Li, Hai Zhu, Zhen Yang, Ya-zhong Luo This paper presents a dimension-reduction method for solving the free-time pursuit-evasion game between two spacecraft near circular orbits. Theoretically, finding the saddle-point solution of the game results in solving a high-dimensional (twenty-four in our case) two-point boundary value problem (TPBVP), which is quite difficult and computationally intensive. By using the circular-orbit variational equations to model the relative states of two spacecraft near circular orbits, the 24-dimensional TPBVP is firstly transformed into solving a set of four-dimensional nonlinear equations. Then, a hybrid numerical algorithm is proposed to solve the equations, where the differential evolution algorithm is used to obtain an initial guess and the Newton's iteration method is used to find the accurate solution. Numerical results show that the proposed method is more efficient than previous methods, and that the proposed method are suitable for large-phasing-distance orbital pursuit-evasion problems.
- Ignition and combustion of air/fuel mixture in a long tube induced by
microwave subcritical streamer discharge- Abstract: Publication date: Available online 23 January 2019Source: Acta AstronauticaAuthor(s): M.P. Bulat, P.V. Bulat, P.V. Denissenko, I.I. Esakov, L.P. Grachev, K.N. Volkov, I.A. Volobuev There have been consistent efforts in developing more efficient combustion for propulsion systems. Ignition and combustion control using cold and non-thermal plasma in microwave discharges have become a major topic of interest. In this study, a microwave subcritical streamer discharge is used to initiate ignition and combustion of premixed air/fuel mixture in a long cylindrical tube. The streamer discharge is arising on the internal surface of the dielectric tube using a passive vibrator in a single pulse regime at atmospheric pressure and temperature. The propagation speed of the combustion front in the quartz cylindrical tube filled by the air/propane mixture is analyzed experimentally and numerically. The streamer discharge creating a multitude of ignition points provides practically instantaneous ignition of the mixture in the entire volume. The speed of streamer induced combustion front has been shown to be higher compared to that initiated by a spark. Increasing the length of streamer discharge leads to increasing the flame propagation speed. The combustion efficiency has also been shown to be higher when using the microwave streamer ignition.
- Abstract: Publication date: Available online 23 January 2019Source: Acta AstronauticaAuthor(s): M.P. Bulat, P.V. Bulat, P.V. Denissenko, I.I. Esakov, L.P. Grachev, K.N. Volkov, I.A. Volobuev There have been consistent efforts in developing more efficient combustion for propulsion systems. Ignition and combustion control using cold and non-thermal plasma in microwave discharges have become a major topic of interest. In this study, a microwave subcritical streamer discharge is used to initiate ignition and combustion of premixed air/fuel mixture in a long cylindrical tube. The streamer discharge is arising on the internal surface of the dielectric tube using a passive vibrator in a single pulse regime at atmospheric pressure and temperature. The propagation speed of the combustion front in the quartz cylindrical tube filled by the air/propane mixture is analyzed experimentally and numerically. The streamer discharge creating a multitude of ignition points provides practically instantaneous ignition of the mixture in the entire volume. The speed of streamer induced combustion front has been shown to be higher compared to that initiated by a spark. Increasing the length of streamer discharge leads to increasing the flame propagation speed. The combustion efficiency has also been shown to be higher when using the microwave streamer ignition.
- Observability-based visual navigation using landmarks measuring angle for
pinpoint landing- Abstract: Publication date: February 2019Source: Acta Astronautica, Volume 155Author(s): Shengying Zhu, Dongchen Liu, Yang Liu, Pingyuan Cui Visual navigation by landmarks on the surface of the planet and small body is a potential main navigation method for pinpoint landing, which is a very challenging and necessary task in future Mars and asteroid exploration missions. In this paper, a method to reduce the complexity caused by nonlinear characteristic of six-degree-of-freedom (6-DOF) state estimation using pixel values of navigation landmark is proposed. Based on the invariance of angles in optical image, the position and attitude in pixel observation equations are decoupled by choosing angles between the line-of-sight (LoS) vectors of landmarks as observations, hence the solution can be obtained with high accuracy and low complexity. Then, the influence of landmark distribution on navigation accuracy can be analyzed, through evaluating the observable degree of landmark by LoS angle observation matrix. Based on the analysis, an optimal navigation landmarks selection method and the corresponding navigation algorithm are given. Finally, Monte Carlo simulations are used to verify the effectiveness of the proposed navigation algorithm, and evaluate the influence of related factors on navigation accuracy.
- Abstract: Publication date: February 2019Source: Acta Astronautica, Volume 155Author(s): Shengying Zhu, Dongchen Liu, Yang Liu, Pingyuan Cui Visual navigation by landmarks on the surface of the planet and small body is a potential main navigation method for pinpoint landing, which is a very challenging and necessary task in future Mars and asteroid exploration missions. In this paper, a method to reduce the complexity caused by nonlinear characteristic of six-degree-of-freedom (6-DOF) state estimation using pixel values of navigation landmark is proposed. Based on the invariance of angles in optical image, the position and attitude in pixel observation equations are decoupled by choosing angles between the line-of-sight (LoS) vectors of landmarks as observations, hence the solution can be obtained with high accuracy and low complexity. Then, the influence of landmark distribution on navigation accuracy can be analyzed, through evaluating the observable degree of landmark by LoS angle observation matrix. Based on the analysis, an optimal navigation landmarks selection method and the corresponding navigation algorithm are given. Finally, Monte Carlo simulations are used to verify the effectiveness of the proposed navigation algorithm, and evaluate the influence of related factors on navigation accuracy.
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S O 3
- Abstract: Publication date: Available online 21 January 2019Source: Acta AstronauticaAuthor(s): Yulin Wang, Haichao Hong, Shengjing Tang This paper proposes a geometric attitude control approach with the model predictive static programming (MPSP) for a rigid body modeled on the Lie group SO(3). The modeling scheme can avoid the defects of the singularity associated with Euler angles and the unwinding phenomenon associated with quaternion, and the MPSP technique features a high computational efficiency and is capable of achieving hard terminal constraints. In this paper, the MPSP is extended to the geometric control on SO(3), using the variational principle, the left-invariance principle of SO(3) and the topology structure of the Lie algebra space so(3). The desired attitude is considered as the hard terminal constraint and the energy consumption is optimized in the optimal control formulation of the MPSP. Moreover, the system disturbance and the measurement error of the states are taken into account in the online application of the MPSP. Simulation results of a spacecraft attitude control illustrate the effectiveness of the proposed method.
- Abstract: Publication date: Available online 21 January 2019Source: Acta AstronauticaAuthor(s): Yulin Wang, Haichao Hong, Shengjing Tang This paper proposes a geometric attitude control approach with the model predictive static programming (MPSP) for a rigid body modeled on the Lie group SO(3). The modeling scheme can avoid the defects of the singularity associated with Euler angles and the unwinding phenomenon associated with quaternion, and the MPSP technique features a high computational efficiency and is capable of achieving hard terminal constraints. In this paper, the MPSP is extended to the geometric control on SO(3), using the variational principle, the left-invariance principle of SO(3) and the topology structure of the Lie algebra space so(3). The desired attitude is considered as the hard terminal constraint and the energy consumption is optimized in the optimal control formulation of the MPSP. Moreover, the system disturbance and the measurement error of the states are taken into account in the online application of the MPSP. Simulation results of a spacecraft attitude control illustrate the effectiveness of the proposed method.
- Optimal scheduling for location geosynchronous satellites refueling
problem- Abstract: Publication date: Available online 19 January 2019Source: Acta AstronauticaAuthor(s): Tian-Jiao Zhang, Hong-Xin Shen, Bao-Hua Wang, Zhao Li, Heng-Nian Li, Yi-Kang Yang This paper addresses the scheduling problem arising from refueling multiple geosynchronous (GEO) satellites with multiple servicing vehicles [many-to-many (M2M) refueling]. The problem is defined by a set of potential fuel tanker locations, a homogeneous fleet of servicing vehicles with limited capacities, and a set of fuel-deficient GEO satellites with known fuel demands. The objective is to open a subset of fuel tankers, to assign GEO satellites to these tankers and to design vehicle servicing sequences, in order to minimize the total mission costs. To achieve such an economical refueling strategy, the fuel tanker location and routing decision are required to be determined simultaneously. It is shown that this problem can be formulated as a location-routing problem (LRP) which is obviously NP-hard. In order to solve it, an ant colony optimization (ACO) metaheuristic that features a special solution representation scheme is proposed. The proposed algorithm is then tested on two types of instances depending on whether the cost for opening a fuel tanker is considered or not. Numerical experiments show that both the consideration of the fuel tanker opening costs and their locations do affect the refueling schedule. Furthermore, the proposed algorithm outperforms the existing approaches in the M2M refueling mission design, and it is able to detect the optimal refueling strategy on all instances containing 85 GEO targets with 15 potential fuel tankers while previous works stopped at 15 satellites and 2 fuel tankers.
- Abstract: Publication date: Available online 19 January 2019Source: Acta AstronauticaAuthor(s): Tian-Jiao Zhang, Hong-Xin Shen, Bao-Hua Wang, Zhao Li, Heng-Nian Li, Yi-Kang Yang This paper addresses the scheduling problem arising from refueling multiple geosynchronous (GEO) satellites with multiple servicing vehicles [many-to-many (M2M) refueling]. The problem is defined by a set of potential fuel tanker locations, a homogeneous fleet of servicing vehicles with limited capacities, and a set of fuel-deficient GEO satellites with known fuel demands. The objective is to open a subset of fuel tankers, to assign GEO satellites to these tankers and to design vehicle servicing sequences, in order to minimize the total mission costs. To achieve such an economical refueling strategy, the fuel tanker location and routing decision are required to be determined simultaneously. It is shown that this problem can be formulated as a location-routing problem (LRP) which is obviously NP-hard. In order to solve it, an ant colony optimization (ACO) metaheuristic that features a special solution representation scheme is proposed. The proposed algorithm is then tested on two types of instances depending on whether the cost for opening a fuel tanker is considered or not. Numerical experiments show that both the consideration of the fuel tanker opening costs and their locations do affect the refueling schedule. Furthermore, the proposed algorithm outperforms the existing approaches in the M2M refueling mission design, and it is able to detect the optimal refueling strategy on all instances containing 85 GEO targets with 15 potential fuel tankers while previous works stopped at 15 satellites and 2 fuel tankers.
- Evacuation method and outgassing rate of a lunar regolith simulant for
deep drilling tests- Abstract: Publication date: Available online 19 January 2019Source: Acta AstronauticaAuthor(s): Tao Zhang, Xilun Ding, Kun Xu, Shuting Liu, Li He, Haifei Zhu, Yisheng Guan A novel method to evacuate large bins of lunar regolith simulant for deep drilling tests was proposed in the current work. This method can be used to simulate a vacuous lunar regolith environment to a maximum penetration depth of 2 m. An experimental apparatus was built and is composed of a vacuum chamber, a specially designed regolith container and a vacuum pumping system. A pressure on the order of 10 Pa could be reached with the 4.3 m3 vacuum chamber when compacted lunar regolith simulant with a volume of 0.4 m3 was loaded. A theoretical model to predict vacuum degree was proposed basing on the viscous flow theory. Evacuation experiments with or without lunar regolith simulant inside the chamber were performed and the outgassing properties of lunar regolith simulant was experimentally studied. The results show that the outgassing rate of the lunar regolith simulant was about 107 times to that of the electro-polished stainless-steel.
- Abstract: Publication date: Available online 19 January 2019Source: Acta AstronauticaAuthor(s): Tao Zhang, Xilun Ding, Kun Xu, Shuting Liu, Li He, Haifei Zhu, Yisheng Guan A novel method to evacuate large bins of lunar regolith simulant for deep drilling tests was proposed in the current work. This method can be used to simulate a vacuous lunar regolith environment to a maximum penetration depth of 2 m. An experimental apparatus was built and is composed of a vacuum chamber, a specially designed regolith container and a vacuum pumping system. A pressure on the order of 10 Pa could be reached with the 4.3 m3 vacuum chamber when compacted lunar regolith simulant with a volume of 0.4 m3 was loaded. A theoretical model to predict vacuum degree was proposed basing on the viscous flow theory. Evacuation experiments with or without lunar regolith simulant inside the chamber were performed and the outgassing properties of lunar regolith simulant was experimentally studied. The results show that the outgassing rate of the lunar regolith simulant was about 107 times to that of the electro-polished stainless-steel.
- Optical-aided, autonomous and optimal space rendezvous with a
non-cooperative target- Abstract: Publication date: Available online 19 January 2019Source: Acta AstronauticaAuthor(s): Renato Volpe, Christian Circi Performing rendezvous in close-proximity of a non-cooperative target is a challenging task, especially if certain performance requirements have to be met. When maneuvering in hazardous scenarios in order to accomplish either refuelling, inspection, repair or dismissal tasks, optimality and safety are relevant aspects to be sought. Additionally, autonomy is essential in unmanned missions. All these aspects can be achieved through purposely intended Guidance Navigation and Control (GNC) architectures. Optimality is obtained by means of the guidance system, that is by a proper optimal trajectory calculation. Safety and accuracy are guaranteed by evaluation of target's relative pose and shape, which is exploited by the navigation system. Among all the possible choices, optical hardware is becoming widely studied for its high accuracy and capability to reconstruct three-dimensional properties of the observed scene. The present work investigates the reciprocal influences between the guidance and navigation subsystems, putting higher focus on how the performance of an optimal rendezvous maneuver with a non-cooperative target can be influenced by the accuracy reached in the pose and shape estimation process. To this purpose, the chaser, or maneuvering spacecraft is considered to be equipped with a single camera and a distance sensor, as this architecture both provides high accuracy and meets mass, power and volume requirements to be implemented on-board a small platform. An algorithm including the optimization process and a purposely built filter, capable of receiving images as inputs and providing evaluation of target's relative pose and shape as output, is implemented and tested in several different scenarios to validate the feasibility of the maneuver.
- Abstract: Publication date: Available online 19 January 2019Source: Acta AstronauticaAuthor(s): Renato Volpe, Christian Circi Performing rendezvous in close-proximity of a non-cooperative target is a challenging task, especially if certain performance requirements have to be met. When maneuvering in hazardous scenarios in order to accomplish either refuelling, inspection, repair or dismissal tasks, optimality and safety are relevant aspects to be sought. Additionally, autonomy is essential in unmanned missions. All these aspects can be achieved through purposely intended Guidance Navigation and Control (GNC) architectures. Optimality is obtained by means of the guidance system, that is by a proper optimal trajectory calculation. Safety and accuracy are guaranteed by evaluation of target's relative pose and shape, which is exploited by the navigation system. Among all the possible choices, optical hardware is becoming widely studied for its high accuracy and capability to reconstruct three-dimensional properties of the observed scene. The present work investigates the reciprocal influences between the guidance and navigation subsystems, putting higher focus on how the performance of an optimal rendezvous maneuver with a non-cooperative target can be influenced by the accuracy reached in the pose and shape estimation process. To this purpose, the chaser, or maneuvering spacecraft is considered to be equipped with a single camera and a distance sensor, as this architecture both provides high accuracy and meets mass, power and volume requirements to be implemented on-board a small platform. An algorithm including the optimization process and a purposely built filter, capable of receiving images as inputs and providing evaluation of target's relative pose and shape as output, is implemented and tested in several different scenarios to validate the feasibility of the maneuver.
- Linear analysis and control of a tethered system with two rigid end bodies
- Abstract: Publication date: Available online 19 January 2019Source: Acta AstronauticaAuthor(s): Nickolas Sabey, Steven Tragesser A model is proposed for studying the dynamics of a tethered satellite system with identical rigid end-masses, offset tether attachment points, and a massless tether. The equations of motion are developed and linearized for use within a linear control system. A Linear Quadratic Regulator (LQR) paired with a reduced-order closed loop estimator is designed to control the attitude of each end-mass individually using Reaction Wheel Assemblies (RWAs), assuming measured angular rates from gyroscopes on the end-masses and tether. Modal analysis shows the existence of a slow tether libration mode and two fast modes of end-mass oscillation. The oscillation is shown to be approximated by a rigid pendulum with an accelerating pivot point. Two control methodologies are investigated: one which aligns each end-mass with the nadir vector and the other which aligns each end-mass with the tether. Results suggest that both modes of operation are controllable by a system architecture which employs LQR control and reduced-order closed loop estimation with a wide range of required control effort.
- Abstract: Publication date: Available online 19 January 2019Source: Acta AstronauticaAuthor(s): Nickolas Sabey, Steven Tragesser A model is proposed for studying the dynamics of a tethered satellite system with identical rigid end-masses, offset tether attachment points, and a massless tether. The equations of motion are developed and linearized for use within a linear control system. A Linear Quadratic Regulator (LQR) paired with a reduced-order closed loop estimator is designed to control the attitude of each end-mass individually using Reaction Wheel Assemblies (RWAs), assuming measured angular rates from gyroscopes on the end-masses and tether. Modal analysis shows the existence of a slow tether libration mode and two fast modes of end-mass oscillation. The oscillation is shown to be approximated by a rigid pendulum with an accelerating pivot point. Two control methodologies are investigated: one which aligns each end-mass with the nadir vector and the other which aligns each end-mass with the tether. Results suggest that both modes of operation are controllable by a system architecture which employs LQR control and reduced-order closed loop estimation with a wide range of required control effort.
- Instability of the solution of the problem on determining the reentry time
of satellites on elliptic orbits- Abstract: Publication date: Available online 17 January 2019Source: Acta AstronauticaAuthor(s): А.I. Nazarenko, I.V. Usovik The article considers specific examples, which show that the solution of the problem on determining the reentry time of elliptic-orbit satellites can be unstable. An analysis is executed on the basis of orbital data about reentry of two satellites: Molnija-1-44 and Zenit-2SB second stage. The presence of photos during their reentry and application of the improved methodology for updating of orbital parameters on measurements allowed to get more reliable results.
- Abstract: Publication date: Available online 17 January 2019Source: Acta AstronauticaAuthor(s): А.I. Nazarenko, I.V. Usovik The article considers specific examples, which show that the solution of the problem on determining the reentry time of elliptic-orbit satellites can be unstable. An analysis is executed on the basis of orbital data about reentry of two satellites: Molnija-1-44 and Zenit-2SB second stage. The presence of photos during their reentry and application of the improved methodology for updating of orbital parameters on measurements allowed to get more reliable results.
- Bounded lunar relative orbits
- Abstract: Publication date: Available online 17 January 2019Source: Acta AstronauticaAuthor(s): Tao Nie, Pini Gurfil, Shijie Zhang Finding relative orbits for satellite formations flying around the Earth, which are long-term bounded under various perturbations, has been a vibrant field of study. However, much less attention has been given to detecting such orbits for circumlunar formation flying missions. As opposed to low-Earth orbits, in circumlunar missions the third-body effect is large, and the magnitude of the C22 sectorial harmonic has the same order as the J2 zonal harmonic. This renders the analysis of bounded relative orbits more challenging. In this paper, we detect a new family of long-term bounded lunar relative orbits, which can be used for lunar formation flying missions. The main idea is to find an analytical solution for the mean relative distance among the formation satellites, and use it to derive formation geometries that are resilient to the gravitational and third-body perturbations. Two invariant mean-distance conditions and one bounded mean-distance conditions are derived. Numerical simulations indicate that the newly-derived invariance conditions yield long-term bounded relative motion.
- Abstract: Publication date: Available online 17 January 2019Source: Acta AstronauticaAuthor(s): Tao Nie, Pini Gurfil, Shijie Zhang Finding relative orbits for satellite formations flying around the Earth, which are long-term bounded under various perturbations, has been a vibrant field of study. However, much less attention has been given to detecting such orbits for circumlunar formation flying missions. As opposed to low-Earth orbits, in circumlunar missions the third-body effect is large, and the magnitude of the C22 sectorial harmonic has the same order as the J2 zonal harmonic. This renders the analysis of bounded relative orbits more challenging. In this paper, we detect a new family of long-term bounded lunar relative orbits, which can be used for lunar formation flying missions. The main idea is to find an analytical solution for the mean relative distance among the formation satellites, and use it to derive formation geometries that are resilient to the gravitational and third-body perturbations. Two invariant mean-distance conditions and one bounded mean-distance conditions are derived. Numerical simulations indicate that the newly-derived invariance conditions yield long-term bounded relative motion.
- Dynamic processes in the tether of a space tethered system
- Abstract: Publication date: Available online 16 January 2019Source: Acta AstronauticaAuthor(s): P.A. Diakov, A.A. Malashin, N.N. Smirnov sIn this work dynamical processes in the tether of а space tethered system are considered. The whole system is designed to “clean” orbits of space debris. The system consists of a «big» satellite moving along a circular or slightly elliptical orbit (orbital station or some satellite for orbital missions where fragments of space debris can be captured), two endmasses (counterweights, space debris hubs) and tethers between them. Containers (load) with space debris can be transported along the tethers. One of the main problems of stable working of the whole system is longitudinal and transverse wave propagation and vibrations of the tethers. The conditions on the system parameters are obtained for stable movement of the containers and the tethers. At these parameters the influence of dynamical processes is considered in this work.
- Abstract: Publication date: Available online 16 January 2019Source: Acta AstronauticaAuthor(s): P.A. Diakov, A.A. Malashin, N.N. Smirnov sIn this work dynamical processes in the tether of а space tethered system are considered. The whole system is designed to “clean” orbits of space debris. The system consists of a «big» satellite moving along a circular or slightly elliptical orbit (orbital station or some satellite for orbital missions where fragments of space debris can be captured), two endmasses (counterweights, space debris hubs) and tethers between them. Containers (load) with space debris can be transported along the tethers. One of the main problems of stable working of the whole system is longitudinal and transverse wave propagation and vibrations of the tethers. The conditions on the system parameters are obtained for stable movement of the containers and the tethers. At these parameters the influence of dynamical processes is considered in this work.
- Active vibration control of underactuated free-floating spacecraft via a
performance enhanced way- Abstract: Publication date: Available online 15 January 2019Source: Acta AstronauticaAuthor(s): Caisheng Wei, Jianjun Luo, Ziteng Guo, Zeyang Yin, Jianping Yuan This paper investigates an enhanced prescribed performance control approach for the postcapture active vibration suppression of the underactuated free-floating spacecraft via only using the output information. First, an enhanced performance function is developed to quantitatively characterize the transient and steady-state vibration responses. Then, based on the performance function, an output-feedback prescribed performance controller along with an adaptive nonlinear filter is devised to eliminate the vibration within arbitrarily specified time. Compared with the existing works, the primary advantage is that the convergence time of the vibration system can be offline preassigned arbitrarily by the users. In the meanwhile, the proposed control scheme can be directly extended to a class of nonstrict feedback systems. Finally, applications to the underactuated free-floating spacecraft are employed to validate the effectiveness of the proposed enhanced prescribed performance control approach.
- Abstract: Publication date: Available online 15 January 2019Source: Acta AstronauticaAuthor(s): Caisheng Wei, Jianjun Luo, Ziteng Guo, Zeyang Yin, Jianping Yuan This paper investigates an enhanced prescribed performance control approach for the postcapture active vibration suppression of the underactuated free-floating spacecraft via only using the output information. First, an enhanced performance function is developed to quantitatively characterize the transient and steady-state vibration responses. Then, based on the performance function, an output-feedback prescribed performance controller along with an adaptive nonlinear filter is devised to eliminate the vibration within arbitrarily specified time. Compared with the existing works, the primary advantage is that the convergence time of the vibration system can be offline preassigned arbitrarily by the users. In the meanwhile, the proposed control scheme can be directly extended to a class of nonstrict feedback systems. Finally, applications to the underactuated free-floating spacecraft are employed to validate the effectiveness of the proposed enhanced prescribed performance control approach.
- Simulation of hypersonic flows with equilibrium chemical reactions on
graphics processor units- Abstract: Publication date: Available online 14 January 2019Source: Acta AstronauticaAuthor(s): V. Emelyanov, A. Karpenko, K. Volkov When hypersonic vehicle travels in the atmosphere with a high speed, the surrounding gas experiences complicated physical and chemical processes producing high-temperature gas effects. High-temperature gas effects are a key issue related to hypersonic aerodynamic design and optimization. The finite volume method is applied to solve unsteady three-dimensional compressible Navier–Stokes equations on unstructured meshes. High-temperature gas effects altering the aerodynamics of vehicle are taken into account. Possibilities of the use of graphics processor units (GPUs) for the simulation of hypersonic flows are demonstrated. Solutions of some benchmark test cases on GPUs are reported, and a comparison between computational results of chemically equilibrium reacting and perfect air flowfields is performed. Speedup of solution of the problems of interest on GPUs with respect to their solution on central processor units (CPUs) is compared. The results obtained provide promising perspective for designing a GPU-based software framework for applications in CFD.
- Abstract: Publication date: Available online 14 January 2019Source: Acta AstronauticaAuthor(s): V. Emelyanov, A. Karpenko, K. Volkov When hypersonic vehicle travels in the atmosphere with a high speed, the surrounding gas experiences complicated physical and chemical processes producing high-temperature gas effects. High-temperature gas effects are a key issue related to hypersonic aerodynamic design and optimization. The finite volume method is applied to solve unsteady three-dimensional compressible Navier–Stokes equations on unstructured meshes. High-temperature gas effects altering the aerodynamics of vehicle are taken into account. Possibilities of the use of graphics processor units (GPUs) for the simulation of hypersonic flows are demonstrated. Solutions of some benchmark test cases on GPUs are reported, and a comparison between computational results of chemically equilibrium reacting and perfect air flowfields is performed. Speedup of solution of the problems of interest on GPUs with respect to their solution on central processor units (CPUs) is compared. The results obtained provide promising perspective for designing a GPU-based software framework for applications in CFD.
- A linear model for relative motion in an elliptical orbit based on a
spherical coordinate system- Abstract: Publication date: Available online 14 January 2019Source: Acta AstronauticaAuthor(s): Chao Han, Huan Chen, Gustavo Alonso, Yinrui Rao, Javier Cubas, Jianfeng Yin, Xiaohui Wang When compared with Cartesian coordinates, curvilinear coordinates have shown significant superior accuracy when applied to a linear model for relative motion in a circular reference orbit. Experts assume the same superiority for an elliptical reference orbit. However, the curvilinear model in an elliptical reference orbit has not been established, and the superiority has not been proved. After a strictly theoretical derivation this paper establishes a linear model with spherical coordinates for relative motion in an elliptical reference orbit. The differential equation has the same form as the Lawden or Tschauner-Hempel equation and can be easily solved with the solutions from the literature. In addition both the theoretical derivation and the numerical examples prove the model's superior accuracy in leader-follower formation and formations with a large along-track distance. The proposed model reduces the approximate limits of close relative motion, and it is expected to be applied in the propagation, guidance, and control for longer term relative motion with large along-track distances.
- Abstract: Publication date: Available online 14 January 2019Source: Acta AstronauticaAuthor(s): Chao Han, Huan Chen, Gustavo Alonso, Yinrui Rao, Javier Cubas, Jianfeng Yin, Xiaohui Wang When compared with Cartesian coordinates, curvilinear coordinates have shown significant superior accuracy when applied to a linear model for relative motion in a circular reference orbit. Experts assume the same superiority for an elliptical reference orbit. However, the curvilinear model in an elliptical reference orbit has not been established, and the superiority has not been proved. After a strictly theoretical derivation this paper establishes a linear model with spherical coordinates for relative motion in an elliptical reference orbit. The differential equation has the same form as the Lawden or Tschauner-Hempel equation and can be easily solved with the solutions from the literature. In addition both the theoretical derivation and the numerical examples prove the model's superior accuracy in leader-follower formation and formations with a large along-track distance. The proposed model reduces the approximate limits of close relative motion, and it is expected to be applied in the propagation, guidance, and control for longer term relative motion with large along-track distances.
- Supersonic jet and nozzle flows in uniform-flow and free-vortex
aerodynamic windows of gas lasers- Abstract: Publication date: Available online 14 January 2019Source: Acta AstronauticaAuthor(s): V.N. Emelyanov, A.V. Pustovalov, K.N. Volkov The mathematical models and computational tools for design, analysis and predictions of supersonic jet and nozzle flows in the aerodynamic windows of high-power gas lasers are considered. The steady-state Euler equations describing strong shock waves, contact discontinuities, rarefaction waves and their interactions are solved with the finite-volume solver and space-marching method. The results of numerical simulation of steady-state supersonic flows of inviscid compressible gas in nozzles and under- and over-expanded jets are obtained and analyzed for different pressure ratios in the laser cavity and ambient atmosphere. The flowfields corresponding to the uniform velocity profile and free-vortex velocity profile in the outlet nozzle boundary are compared. Nozzle profiling tools are developed on the basis of numerical solution of a sequence of direct problems. The aerodynamic performance of the window is evaluated in terms of the simulated laser cavity pressure and plenum pressure of the free-vortex supply nozzle. The pressure support characteristic for the aerodynamic window is established by determining the ambient to cavity pressure ratio over a range of aerodynamic window supply pressures.
- Abstract: Publication date: Available online 14 January 2019Source: Acta AstronauticaAuthor(s): V.N. Emelyanov, A.V. Pustovalov, K.N. Volkov The mathematical models and computational tools for design, analysis and predictions of supersonic jet and nozzle flows in the aerodynamic windows of high-power gas lasers are considered. The steady-state Euler equations describing strong shock waves, contact discontinuities, rarefaction waves and their interactions are solved with the finite-volume solver and space-marching method. The results of numerical simulation of steady-state supersonic flows of inviscid compressible gas in nozzles and under- and over-expanded jets are obtained and analyzed for different pressure ratios in the laser cavity and ambient atmosphere. The flowfields corresponding to the uniform velocity profile and free-vortex velocity profile in the outlet nozzle boundary are compared. Nozzle profiling tools are developed on the basis of numerical solution of a sequence of direct problems. The aerodynamic performance of the window is evaluated in terms of the simulated laser cavity pressure and plenum pressure of the free-vortex supply nozzle. The pressure support characteristic for the aerodynamic window is established by determining the ambient to cavity pressure ratio over a range of aerodynamic window supply pressures.
- Deployment control of tethered space systems with explicit velocity
constraint and invariance principle- Abstract: Publication date: Available online 12 January 2019Source: Acta AstronauticaAuthor(s): Latheepan Murugathasan, Zheng H. Zhu This paper develops control laws for the deployment of tethered spacecraft systems with an explicit non-negative velocity constraint using invariance principle of control systems. It addresses a class of space tether deployment control problems where the tether is deployed out only. This implies the tether deployment velocity must be always positive and approaches to zero at the end of deployment. This non-negative velocity constraint is explicitly satisfied in the proposed tension control laws by transforming it into an invariance property of control system. Two control laws with nonlinear feedback are derived with two complementary manifolds. The Lyapunov stability of the control laws is proved and the controllability of the tethered spacecraft system with tension input is analyzed. Finally, the effectiveness of control laws is demonstrated by numerical simulation.
- Abstract: Publication date: Available online 12 January 2019Source: Acta AstronauticaAuthor(s): Latheepan Murugathasan, Zheng H. Zhu This paper develops control laws for the deployment of tethered spacecraft systems with an explicit non-negative velocity constraint using invariance principle of control systems. It addresses a class of space tether deployment control problems where the tether is deployed out only. This implies the tether deployment velocity must be always positive and approaches to zero at the end of deployment. This non-negative velocity constraint is explicitly satisfied in the proposed tension control laws by transforming it into an invariance property of control system. Two control laws with nonlinear feedback are derived with two complementary manifolds. The Lyapunov stability of the control laws is proved and the controllability of the tethered spacecraft system with tension input is analyzed. Finally, the effectiveness of control laws is demonstrated by numerical simulation.
- Decomposition analysis of spacecraft relative motion with different
inter-satellite ranges- Abstract: Publication date: Available online 11 January 2019Source: Acta AstronauticaAuthor(s): Jiang Chao, Wang Zhaokui, Zhang Yulin With the development of satellite cluster and giant constellations in recent years, the orbit design and long-term maintenance of these systems become more difficult with the restriction of fuel consumption and the influence of orbital perturbations. It's necessary to study the analytical model of relative motion which applicable to different inter-satellite ranges. Then, considering the perturbed evolutions of orbital elements, the decomposed relative motion model was proposed by separating the higher-fidelity relative motion model into two different parts, which reflect the circling motion around chief, and their overall motion relative to chief respectively. Eccentricity expansions had been carried out to provide a time-explicit analytical solution of the decomposed model. Based on this first order time-explicit analytical solution, the bound relative motions with different inter-satellite ranges were studied in Kepler orbit. The initial condition calculation formula for general chief centered circular relative motion was derived at the same time. The accuracy of first order analytical solution of the decomposed model was estimated both theoretically and numerically, to show the effect of inter-satellite range and chief's eccentricity. At last, the first order analytical solution using mean orbital elements was compared with high precision numerical solution under perturbations to validate it's effective in long-term stability analysis.
- Abstract: Publication date: Available online 11 January 2019Source: Acta AstronauticaAuthor(s): Jiang Chao, Wang Zhaokui, Zhang Yulin With the development of satellite cluster and giant constellations in recent years, the orbit design and long-term maintenance of these systems become more difficult with the restriction of fuel consumption and the influence of orbital perturbations. It's necessary to study the analytical model of relative motion which applicable to different inter-satellite ranges. Then, considering the perturbed evolutions of orbital elements, the decomposed relative motion model was proposed by separating the higher-fidelity relative motion model into two different parts, which reflect the circling motion around chief, and their overall motion relative to chief respectively. Eccentricity expansions had been carried out to provide a time-explicit analytical solution of the decomposed model. Based on this first order time-explicit analytical solution, the bound relative motions with different inter-satellite ranges were studied in Kepler orbit. The initial condition calculation formula for general chief centered circular relative motion was derived at the same time. The accuracy of first order analytical solution of the decomposed model was estimated both theoretically and numerically, to show the effect of inter-satellite range and chief's eccentricity. At last, the first order analytical solution using mean orbital elements was compared with high precision numerical solution under perturbations to validate it's effective in long-term stability analysis.
- The impact of methane oxidation kinetics on a rocket nozzle flow
- Abstract: Publication date: Available online 11 January 2019Source: Acta AstronauticaAuthor(s): Victor P. Zhukov The impact of methane combustion kinetics on a rocket nozzle flow is theoretically studied in the work. To evaluate the effect of the kinetics, simulations of a rocket nozzle flow are carried out using the computational fluid dynamics solver Fluent. At first a methane kinetic model is selected. The comparison of different kinetic models shows that the recent skeletal mechanism of Zhukov and Kong has a very good accuracy at a small size. For comparison, the flow simulations are performed both for hydrogen/oxygen and methane/oxygen propellant combinations, and for three different reaction models: non-reactive (“frozen”) flow, reactive (non-equilibrium) flow, and chemically equilibrium flow. Chemically non-equilibrium flow is modelled using the Zhukov–Kong model. Simulations results show that the recombination reactions of combustion products should be taken into account for modelling rocket nozzle flow. For hydrogen the difference in results between chemically non-equilibrium and equilibrium flows is negligible while in methane the difference is small but noticeable.
- Abstract: Publication date: Available online 11 January 2019Source: Acta AstronauticaAuthor(s): Victor P. Zhukov The impact of methane combustion kinetics on a rocket nozzle flow is theoretically studied in the work. To evaluate the effect of the kinetics, simulations of a rocket nozzle flow are carried out using the computational fluid dynamics solver Fluent. At first a methane kinetic model is selected. The comparison of different kinetic models shows that the recent skeletal mechanism of Zhukov and Kong has a very good accuracy at a small size. For comparison, the flow simulations are performed both for hydrogen/oxygen and methane/oxygen propellant combinations, and for three different reaction models: non-reactive (“frozen”) flow, reactive (non-equilibrium) flow, and chemically equilibrium flow. Chemically non-equilibrium flow is modelled using the Zhukov–Kong model. Simulations results show that the recombination reactions of combustion products should be taken into account for modelling rocket nozzle flow. For hydrogen the difference in results between chemically non-equilibrium and equilibrium flows is negligible while in methane the difference is small but noticeable.
- Optimal regular reflection of oblique shocks
- Abstract: Publication date: Available online 11 January 2019Source: Acta AstronauticaAuthor(s): M.V. Chernyshov, O.A. Tolpegin The regular reflection of oblique steady and unsteady shocks in a perfect gas is considered. Non-monotonic variation of the mechanical and thermal loads on the obstacle with respect to shock/blast wave incidence angle is shown; the obstacle slope angles that correspond to pressure and temperature minima downwards of both steady and unsteady shock reflection point are determined analytically. Achieved minimization of dynamical and thermal loads at shock/blast wave reflection can be of use in space flight safety problems as blast safety onboard a spaceship, protection of spaceport facilities from blast of rocket carrier, prevention of fire onboard, and stable functioning of rocket engines, including ones based on detonation effects.
- Abstract: Publication date: Available online 11 January 2019Source: Acta AstronauticaAuthor(s): M.V. Chernyshov, O.A. Tolpegin The regular reflection of oblique steady and unsteady shocks in a perfect gas is considered. Non-monotonic variation of the mechanical and thermal loads on the obstacle with respect to shock/blast wave incidence angle is shown; the obstacle slope angles that correspond to pressure and temperature minima downwards of both steady and unsteady shock reflection point are determined analytically. Achieved minimization of dynamical and thermal loads at shock/blast wave reflection can be of use in space flight safety problems as blast safety onboard a spaceship, protection of spaceport facilities from blast of rocket carrier, prevention of fire onboard, and stable functioning of rocket engines, including ones based on detonation effects.
- Experimental investigation on propagation characteristics of rotating
detonation wave with a hydrogen-ethylene-acetylene fuel- Abstract: Publication date: Available online 9 January 2019Source: Acta AstronauticaAuthor(s): Shengbing Zhou, Hu Ma, Sihe Chen, Yepan Zhong, Changsheng Zhou It is attracting much attention that the hydrocarbon mixtures are used as the fuel of rotating detonation engine when the oxidizer is air. In this study, a hydrogen-ethylene-acetylene mixture was used as the fuel of rotating detonation combustor to investigate the propagation characteristics of rotating detonation wave. An annular combustor, with an inner diameter of 124 mm and an outer diameter of 152 mm, was used in this research. The air and fuel were separately injected into the combustor through a slot-orifice injection structure. High-frequency pressure sensors and ionization probes were used to measure the detonation characteristics. Experiments could obtain a stable detonation wave whose velocity was about 65% of the C-J value, showing that there was a large velocity loss compared to hydrogen fuel, which reached a C-J velocity value of 94%. The detonation intensity for the hydrogen-ethylene-acetylene mixture is weaker than that of hydrogen. The equivalence-ratio range of forming a detonation wave with hydrogen-ethylene-acetylene mixture was narrower than that of hydrogen fuel with the same mass flow rate. The detonation-wave height of hydrogen-ethylene-acetylene mixture had similar results with hydrogen fuel.
- Abstract: Publication date: Available online 9 January 2019Source: Acta AstronauticaAuthor(s): Shengbing Zhou, Hu Ma, Sihe Chen, Yepan Zhong, Changsheng Zhou It is attracting much attention that the hydrocarbon mixtures are used as the fuel of rotating detonation engine when the oxidizer is air. In this study, a hydrogen-ethylene-acetylene mixture was used as the fuel of rotating detonation combustor to investigate the propagation characteristics of rotating detonation wave. An annular combustor, with an inner diameter of 124 mm and an outer diameter of 152 mm, was used in this research. The air and fuel were separately injected into the combustor through a slot-orifice injection structure. High-frequency pressure sensors and ionization probes were used to measure the detonation characteristics. Experiments could obtain a stable detonation wave whose velocity was about 65% of the C-J value, showing that there was a large velocity loss compared to hydrogen fuel, which reached a C-J velocity value of 94%. The detonation intensity for the hydrogen-ethylene-acetylene mixture is weaker than that of hydrogen. The equivalence-ratio range of forming a detonation wave with hydrogen-ethylene-acetylene mixture was narrower than that of hydrogen fuel with the same mass flow rate. The detonation-wave height of hydrogen-ethylene-acetylene mixture had similar results with hydrogen fuel.
- Automated near real-time validation and exploitation of optical sensor
data for improved orbital safety- Abstract: Publication date: Available online 8 January 2019Source: Acta AstronauticaAuthor(s): Thomas Kelecy, Emily Lambert, Benjamin Sunderland, Jason Stauch, Vishnuu Mallik, Moriba Jah The orbital safety of operational spacecraft, regardless of the mission, relies on timely and actionable observations to maintain so-called “custody” of all trackable Resident Space Objects (RSOs), including space debris, that might pose a hazard to safe, secure, and sustainable operations. For operations in and around the Geosynchronous Earth Orbit (GEO) regime, electro-optical (EO) observations are the most prevalent observation type available for tracking and determining RSO orbits. The quality (both accuracy and precision) of the data affects the inferable kinematic, physical, and other characteristics of RSOs and, in particular, measurement biases will result in inaccurate orbital trajectories and subsequent predictions. Physically meaningful conjunction assessments rely on not only accurate orbit state prediction, but also the “realistic” covariances associated with said predictions. In this paper we demonstrate an automated near real-time (NRT) assessment of measurement biases with an appropriately implemented Unscented Schmidt Kalman Filter (USKF) [14]. Hypothesized biases that are deterministic but statistically unobservable in the measurement data and cannot be estimated are accounted for as so-called “consider” parameters. The method presented herein is assessed and quantified using both simulated and actual measurement data. This method will enable the exploitation and mining of so-called “non-traditional” sensor data to maximize Space Situational Awareness (SSA) in a robust and timely fashion toward improvement of orbital safety. The ultimate goal is to provide decision-making evidence required solve problems preventing the space domain from being safe, secure, and sustainable.
- Abstract: Publication date: Available online 8 January 2019Source: Acta AstronauticaAuthor(s): Thomas Kelecy, Emily Lambert, Benjamin Sunderland, Jason Stauch, Vishnuu Mallik, Moriba Jah The orbital safety of operational spacecraft, regardless of the mission, relies on timely and actionable observations to maintain so-called “custody” of all trackable Resident Space Objects (RSOs), including space debris, that might pose a hazard to safe, secure, and sustainable operations. For operations in and around the Geosynchronous Earth Orbit (GEO) regime, electro-optical (EO) observations are the most prevalent observation type available for tracking and determining RSO orbits. The quality (both accuracy and precision) of the data affects the inferable kinematic, physical, and other characteristics of RSOs and, in particular, measurement biases will result in inaccurate orbital trajectories and subsequent predictions. Physically meaningful conjunction assessments rely on not only accurate orbit state prediction, but also the “realistic” covariances associated with said predictions. In this paper we demonstrate an automated near real-time (NRT) assessment of measurement biases with an appropriately implemented Unscented Schmidt Kalman Filter (USKF) [14]. Hypothesized biases that are deterministic but statistically unobservable in the measurement data and cannot be estimated are accounted for as so-called “consider” parameters. The method presented herein is assessed and quantified using both simulated and actual measurement data. This method will enable the exploitation and mining of so-called “non-traditional” sensor data to maximize Space Situational Awareness (SSA) in a robust and timely fashion toward improvement of orbital safety. The ultimate goal is to provide decision-making evidence required solve problems preventing the space domain from being safe, secure, and sustainable.
- A study of droplets emission in Ultrasonic Electric Propulsion system:
Experiment and simulation- Abstract: Publication date: Available online 8 January 2019Source: Acta AstronauticaAuthor(s): Xiaoming Kang, Weiguo He, Xinyu Liu, Guanrong Hang, Lin Li, Wansheng Zhao As an innovative propulsion technology, Ultrasonic Electric Propulsion (UEP) is mainly applied to nano-and micro-spacecrafts (
- Abstract: Publication date: Available online 8 January 2019Source: Acta AstronauticaAuthor(s): Xiaoming Kang, Weiguo He, Xinyu Liu, Guanrong Hang, Lin Li, Wansheng Zhao As an innovative propulsion technology, Ultrasonic Electric Propulsion (UEP) is mainly applied to nano-and micro-spacecrafts (
- A new nonlinear Type3-PLL with noise rejection and fast locking
performance in tracking telemetry and command systems- Abstract: Publication date: Available online 8 January 2019Source: Acta AstronauticaAuthor(s): Congying Zhu, Xiaoping Li, Lei Shi, Yanming Liu, Bo Yao Type3-PLLs are widely used for carrier synchronization in tracking telemetry and command (TT&C) systems. A major challenge with the traditional Type3-PLL is how to further improve their acquisition and tracking speed without compromising the noise suppression ability and stability. In this article, drawn lessons from the nonlinear control theory, a new structural Type3-PLL is presented. The idea of the proposed Type3-PLL, which can be denoted by N-Type3-PLL, is to couple with a designed nonlinear element in the traditional Type3-PLL. The loop noise bandwidth is adaptively adjusted by the designed nonlinear element on the basis of the loop filter output, which is determined by the frequency-phase error. The output noise of the nonlinear element is then suppressed by the consequent low-pass filter. The influence of the nonlinear element parameters on the N-Type3-PLL is quantitatively analyzed and discussed. Simulation results indicate that the N-Type3-PLLs under various parameters can achieve locking faster than the traditional Type3-PLL in response to a step change in frequency. The nonlinear element parameters constraint the comprehensive performance of the N-Type3-PLL. The optimal parameters of the nonlinear element are solved via a suggested iterative feedback method. Compared with the traditional Type3-PLL, the output frequency of the N-Type3-PLL under the optimal parameters jitters much less and the standard deviation of frequency-phase error is reduced from 0.6198 (rad/s) 2 to 0.2264 (rad/s) 2 when the input SNR equals −15 dB. The lock time after the step is extended from 0.22 s to 0.24 s, which is within an acceptable range. The loss threshold of the proposed N-Type3-PLL is reduced from −20 dB to −23 dB.
- Abstract: Publication date: Available online 8 January 2019Source: Acta AstronauticaAuthor(s): Congying Zhu, Xiaoping Li, Lei Shi, Yanming Liu, Bo Yao Type3-PLLs are widely used for carrier synchronization in tracking telemetry and command (TT&C) systems. A major challenge with the traditional Type3-PLL is how to further improve their acquisition and tracking speed without compromising the noise suppression ability and stability. In this article, drawn lessons from the nonlinear control theory, a new structural Type3-PLL is presented. The idea of the proposed Type3-PLL, which can be denoted by N-Type3-PLL, is to couple with a designed nonlinear element in the traditional Type3-PLL. The loop noise bandwidth is adaptively adjusted by the designed nonlinear element on the basis of the loop filter output, which is determined by the frequency-phase error. The output noise of the nonlinear element is then suppressed by the consequent low-pass filter. The influence of the nonlinear element parameters on the N-Type3-PLL is quantitatively analyzed and discussed. Simulation results indicate that the N-Type3-PLLs under various parameters can achieve locking faster than the traditional Type3-PLL in response to a step change in frequency. The nonlinear element parameters constraint the comprehensive performance of the N-Type3-PLL. The optimal parameters of the nonlinear element are solved via a suggested iterative feedback method. Compared with the traditional Type3-PLL, the output frequency of the N-Type3-PLL under the optimal parameters jitters much less and the standard deviation of frequency-phase error is reduced from 0.6198 (rad/s) 2 to 0.2264 (rad/s) 2 when the input SNR equals −15 dB. The lock time after the step is extended from 0.22 s to 0.24 s, which is within an acceptable range. The loss threshold of the proposed N-Type3-PLL is reduced from −20 dB to −23 dB.
- Why planetary and exoplanetary protection differ: The case of long
- Abstract: Publication date: Available online 8 January 2019Source: Acta AstronauticaAuthor(s): Claudius Gros Time is arguably the key limiting factor for interstellar exploration. At high speeds, flyby missions to nearby stars by laser propelled wafersats taking 50–100 years would be feasible. Directed energy launch systems could accelerate on the other side also crafts weighing several tons to cruising speeds of the order of 1000 km/s (c/300). At these speeds, superconducting magnetic sails would be able to decelerate the craft by transferring kinetic energy to the protons of the interstellar medium. A tantalizing perspective, which would allow interstellar probes to stop whenever time is not a limiting factor. Prime candidates are in this respect Genesis probes, that is missions aiming to offer terrestrial life new evolutionary pathways on potentially habitable but hitherto barren exoplanets.Genesis missions raise important ethical issues, in particular with regard to planetary protection. Here we argue that exoplanetary and planetary protection differ qualitatively as a result of the vastly different cruising times for payload delivering probes, which are of the order of millennia for interstellar probes, but only of years for solar system bodies. Furthermore we point out that our galaxy may harbor a large number of habitable exoplanets, M-dwarf planets, which could be sterile due to the presence of massive primordial oxygen atmospheres. We believe that the prospect terrestrial life has in our galaxy would shift on a fundamental level in case that the existence of this type of habitable but sterile oxygen planets will be corroborated by future research. It may also explain why our sun is not a M dwarf, the most common star type, but a medium-sized G-class star.
- Abstract: Publication date: Available online 8 January 2019Source: Acta AstronauticaAuthor(s): Claudius Gros Time is arguably the key limiting factor for interstellar exploration. At high speeds, flyby missions to nearby stars by laser propelled wafersats taking 50–100 years would be feasible. Directed energy launch systems could accelerate on the other side also crafts weighing several tons to cruising speeds of the order of 1000 km/s (c/300). At these speeds, superconducting magnetic sails would be able to decelerate the craft by transferring kinetic energy to the protons of the interstellar medium. A tantalizing perspective, which would allow interstellar probes to stop whenever time is not a limiting factor. Prime candidates are in this respect Genesis probes, that is missions aiming to offer terrestrial life new evolutionary pathways on potentially habitable but hitherto barren exoplanets.Genesis missions raise important ethical issues, in particular with regard to planetary protection. Here we argue that exoplanetary and planetary protection differ qualitatively as a result of the vastly different cruising times for payload delivering probes, which are of the order of millennia for interstellar probes, but only of years for solar system bodies. Furthermore we point out that our galaxy may harbor a large number of habitable exoplanets, M-dwarf planets, which could be sterile due to the presence of massive primordial oxygen atmospheres. We believe that the prospect terrestrial life has in our galaxy would shift on a fundamental level in case that the existence of this type of habitable but sterile oxygen planets will be corroborated by future research. It may also explain why our sun is not a M dwarf, the most common star type, but a medium-sized G-class star.
- Experimental study on erosion characteristics of char layer of EPDM
composite by cold flow- Abstract: Publication date: Available online 8 January 2019Source: Acta AstronauticaAuthor(s): Jiang Li, Kai Ma, Xiang Lv, Yang Liu, Shuxian Wang, Peijin Liu Char layer erosion due to gas flow is an important part of the ablation process for insulation materials used in solid rocket motors (SRMs) or rocket ramjets. However, the current understanding of the erosion pattern is very limited. This paper reports on an experimental study on the erosion characteristics of the char layer of ethylene propylene diene monomer (EPDM) composite. First, the charring samples were prepared using a special ablation motor. A cold flow erosion test device was then developed, and the erosion experiments on the charring samples under cold airflow conditions were conducted. The experimental results show that under a pressure of 0.5 MPa, when the airflow velocity is lower than or equal to 180 m/s, the char layer erosion is largely in the form of wear; when the airflow velocity reaches 200 m/s, the large pieces of the char layer are peeled off; when the airflow velocity reaches 230 m/s, the entire char layer is peeled off. The tensile and shear strength of the char layer under cold conditions were tested using a simple method. It was found that the shear stress of the airflow when the char layer is peeled off is significantly lower than the shear strength of the char layer. The failure of the char layer is not only due to the pure shear force of the airflow. The local cracking, warping, and delamination not only reduce the char layer strength, but also enhance the destructive force of the airflow. This was the main cause of the char layer erosion due to airflow.
- Abstract: Publication date: Available online 8 January 2019Source: Acta AstronauticaAuthor(s): Jiang Li, Kai Ma, Xiang Lv, Yang Liu, Shuxian Wang, Peijin Liu Char layer erosion due to gas flow is an important part of the ablation process for insulation materials used in solid rocket motors (SRMs) or rocket ramjets. However, the current understanding of the erosion pattern is very limited. This paper reports on an experimental study on the erosion characteristics of the char layer of ethylene propylene diene monomer (EPDM) composite. First, the charring samples were prepared using a special ablation motor. A cold flow erosion test device was then developed, and the erosion experiments on the charring samples under cold airflow conditions were conducted. The experimental results show that under a pressure of 0.5 MPa, when the airflow velocity is lower than or equal to 180 m/s, the char layer erosion is largely in the form of wear; when the airflow velocity reaches 200 m/s, the large pieces of the char layer are peeled off; when the airflow velocity reaches 230 m/s, the entire char layer is peeled off. The tensile and shear strength of the char layer under cold conditions were tested using a simple method. It was found that the shear stress of the airflow when the char layer is peeled off is significantly lower than the shear strength of the char layer. The failure of the char layer is not only due to the pure shear force of the airflow. The local cracking, warping, and delamination not only reduce the char layer strength, but also enhance the destructive force of the airflow. This was the main cause of the char layer erosion due to airflow.
- Planetary landing site detection and selection using multilevel
optimization strategy- Abstract: Publication date: Available online 7 January 2019Source: Acta AstronauticaAuthor(s): Xu Liu, Shuang Li, Xiuqiang Jiang, Xiangyu Huang Reliable landing site detection and selection method is the key to enable a safe planetary landing for both robotic and manned missions. In this paper, a multilevel optimization strategy is newly employed to address this issue. This strategy decomposes the landing site detection and selection problem with constraints into three successive optimization sub-problems, which are then solved as a predetermined sequence in the multilevel structure. In the optimization procedures, an innovation derivatives based hazard detection algorithm is also proposed to solve the first sub-problem. Meanwhile, three performance indexes based on different constraints are defined to locate the optimal landing site. A typical lunar landing site selection scenario is simulated simultaneously using multilevel optimization and existing safety index based optimization, and comparative simulations illustrate the feasibility and superiority of the proposed approach. In addition, robustness assessments of multilevel optimization on initial flight states, guidance laws and landing terrains are also conducted.
- Abstract: Publication date: Available online 7 January 2019Source: Acta AstronauticaAuthor(s): Xu Liu, Shuang Li, Xiuqiang Jiang, Xiangyu Huang Reliable landing site detection and selection method is the key to enable a safe planetary landing for both robotic and manned missions. In this paper, a multilevel optimization strategy is newly employed to address this issue. This strategy decomposes the landing site detection and selection problem with constraints into three successive optimization sub-problems, which are then solved as a predetermined sequence in the multilevel structure. In the optimization procedures, an innovation derivatives based hazard detection algorithm is also proposed to solve the first sub-problem. Meanwhile, three performance indexes based on different constraints are defined to locate the optimal landing site. A typical lunar landing site selection scenario is simulated simultaneously using multilevel optimization and existing safety index based optimization, and comparative simulations illustrate the feasibility and superiority of the proposed approach. In addition, robustness assessments of multilevel optimization on initial flight states, guidance laws and landing terrains are also conducted.
- A space surveillance satellite for cataloging high-altitude small debris
- Abstract: Publication date: Available online 4 January 2019Source: Acta AstronauticaAuthor(s): Jianli Du, Xiangxu Lei, Jizhang Sang This paper proposes a space surveillance satellite whose mission is to detect, catalog and maintain space debris as small as 10 cm in the Geosynchronous Earth Orbit (GEO) region. To fulfill the objective, the satellite is designed to be placed on a zero-inclination orbit of 4163 km in altitude. The payload is a large optical telescope with a 1.5 m Entrance Pupil Diameter (EPD) and a 20° × 20° Field of View (FoV). Based on these configurations, a 60-day simulation experiment is organized to evaluate the debris detecting and cataloging performances of the satellite. The simulation results show that the satellite is capable of detecting small debris higher than 4163 km. About 94.7% of 10 cm debris in the GEO region are detected more than 60 times over the 60-day simulation span by the satellite. In addition, the rules-based cataloging and maintaining capability of the satellite is close to 50.0% with regard to an entire 10 cm reference population of 3267 debris. For larger size debris, the results are more optimistic. This paper also assesses the performances of our Initial Orbit Determination (IOD) and Un-Correlated Track (UCT) association methods for 74 debris in the GEO region. The IOD success rate is 97.4% by applying a range-search based method. The True Positive (TP) association rate is 89.7% and the False Positive (FP) rate is 1.8%. Accordingly, 66 of the 74 space objects are correctly identified that each has at least two detection arcs, making accurate orbit determination possible.
- Abstract: Publication date: Available online 4 January 2019Source: Acta AstronauticaAuthor(s): Jianli Du, Xiangxu Lei, Jizhang Sang This paper proposes a space surveillance satellite whose mission is to detect, catalog and maintain space debris as small as 10 cm in the Geosynchronous Earth Orbit (GEO) region. To fulfill the objective, the satellite is designed to be placed on a zero-inclination orbit of 4163 km in altitude. The payload is a large optical telescope with a 1.5 m Entrance Pupil Diameter (EPD) and a 20° × 20° Field of View (FoV). Based on these configurations, a 60-day simulation experiment is organized to evaluate the debris detecting and cataloging performances of the satellite. The simulation results show that the satellite is capable of detecting small debris higher than 4163 km. About 94.7% of 10 cm debris in the GEO region are detected more than 60 times over the 60-day simulation span by the satellite. In addition, the rules-based cataloging and maintaining capability of the satellite is close to 50.0% with regard to an entire 10 cm reference population of 3267 debris. For larger size debris, the results are more optimistic. This paper also assesses the performances of our Initial Orbit Determination (IOD) and Un-Correlated Track (UCT) association methods for 74 debris in the GEO region. The IOD success rate is 97.4% by applying a range-search based method. The True Positive (TP) association rate is 89.7% and the False Positive (FP) rate is 1.8%. Accordingly, 66 of the 74 space objects are correctly identified that each has at least two detection arcs, making accurate orbit determination possible.
- Three-dimensional reconstruction of incident shock/boundary layer
interaction using background-oriented schlieren- Abstract: Publication date: Available online 4 January 2019Source: Acta AstronauticaAuthor(s): Chengpeng Wang, Pei Xu, Longsheng Xue, Yun Jiao To study the three-dimensional flow structure of an incident shock/boundary layer interaction, a M2.7 nozzle model with a ramp was tested based on the background-oriented schlieren (BOS) technique. After comparing the results of CFD, shear-sensitive liquid crystal (SSLC), and high-speed schlieren, we conclude that BOS can capture the internal structure of the flow field and that its results are consistent with those of other methods in which the relative error can be controlled to 7%. In this research, a new type of experiment with several light paths was used with BOS technology. Three-dimensional reconstruction of incident oblique shock (primarily the width of Mach stem of the shock wave) and quantitative results of the flow field can be obtained successfully, and BOS shows great potential for the measurement of flow fields.
- Abstract: Publication date: Available online 4 January 2019Source: Acta AstronauticaAuthor(s): Chengpeng Wang, Pei Xu, Longsheng Xue, Yun Jiao To study the three-dimensional flow structure of an incident shock/boundary layer interaction, a M2.7 nozzle model with a ramp was tested based on the background-oriented schlieren (BOS) technique. After comparing the results of CFD, shear-sensitive liquid crystal (SSLC), and high-speed schlieren, we conclude that BOS can capture the internal structure of the flow field and that its results are consistent with those of other methods in which the relative error can be controlled to 7%. In this research, a new type of experiment with several light paths was used with BOS technology. Three-dimensional reconstruction of incident oblique shock (primarily the width of Mach stem of the shock wave) and quantitative results of the flow field can be obtained successfully, and BOS shows great potential for the measurement of flow fields.
- Downrange manoeuvre and oscillation suppression of a self-regulating
centrifugally deployed flexible heat shield using a controlled reaction
wheel- Abstract: Publication date: Available online 4 January 2019Source: Acta AstronauticaAuthor(s): Rui Wu, Peter C.E. Roberts, Constantinos Soutis, Carl Diver A recent study has introduced a flexible deployable heat shield that passively deploys and stiffens due to centrifugal forces generated from a self-regulated autorotation. This paper demonstrates that the heat shield is similar to a PI controlled second order nonlinear system, which explains why the deployment is accompanied by a limit cycle structural oscillation that persists throughout a simulated re-entry. The heat shield design offers a unique capability to actively adjust the deployment using conventional attitude control devices. This operation is explored by simulating the re-entry of a CubeSat-sized vehicle equipped with an off-the-shelf reaction wheel controlled by a switching phase shift controller and gain-scheduled controllers. The effects of the control parameters are investigated, and successful oscillation suppression as well as an open-loop downrange manoeuvre of over 300 km is predicted for re-entry from low earth orbit.
- Abstract: Publication date: Available online 4 January 2019Source: Acta AstronauticaAuthor(s): Rui Wu, Peter C.E. Roberts, Constantinos Soutis, Carl Diver A recent study has introduced a flexible deployable heat shield that passively deploys and stiffens due to centrifugal forces generated from a self-regulated autorotation. This paper demonstrates that the heat shield is similar to a PI controlled second order nonlinear system, which explains why the deployment is accompanied by a limit cycle structural oscillation that persists throughout a simulated re-entry. The heat shield design offers a unique capability to actively adjust the deployment using conventional attitude control devices. This operation is explored by simulating the re-entry of a CubeSat-sized vehicle equipped with an off-the-shelf reaction wheel controlled by a switching phase shift controller and gain-scheduled controllers. The effects of the control parameters are investigated, and successful oscillation suppression as well as an open-loop downrange manoeuvre of over 300 km is predicted for re-entry from low earth orbit.
- Design and performance evaluation of a piezoelectric aeroelastic energy
harvester based on the limit cycle oscillation phenomenon- Abstract: Publication date: Available online 4 January 2019Source: Acta AstronauticaAuthor(s): Hassan Elahi, Marco Eugeni, Paolo Gaudenzi In the present era, the demand for self-powered electronic instruments is increasing and their energy consumption is decreasing. The ability to extract energy from the operating environment is of great importance in advanced industrial applications particularly in the field of aerospace. In this research, a nonlinear piezoelectric aeroelastic energy harvester (PAEH) is modeled based on fluid-structure interaction (FSI) that represents an important area of research for the development of innovative energy harvesting solution. This PAEH operates on post critical aeroelastic behavior i.e., Limit Cycle Oscillations (LCOs) that arise after the flutter velocity. Moreover, it is emphasized that the determination of aerodynamic model is necessary for correct prediction of PAEH performance. The design of an aeroelastic harvester based on the use of two different types of piezoelectric materials is presented. The designed model is suitable for energy harvesting and can be utilized to drive nano and microelectronics. Maximum output power obtained for the designed PAEH is found to be 0.55 mW.
- Abstract: Publication date: Available online 4 January 2019Source: Acta AstronauticaAuthor(s): Hassan Elahi, Marco Eugeni, Paolo Gaudenzi In the present era, the demand for self-powered electronic instruments is increasing and their energy consumption is decreasing. The ability to extract energy from the operating environment is of great importance in advanced industrial applications particularly in the field of aerospace. In this research, a nonlinear piezoelectric aeroelastic energy harvester (PAEH) is modeled based on fluid-structure interaction (FSI) that represents an important area of research for the development of innovative energy harvesting solution. This PAEH operates on post critical aeroelastic behavior i.e., Limit Cycle Oscillations (LCOs) that arise after the flutter velocity. Moreover, it is emphasized that the determination of aerodynamic model is necessary for correct prediction of PAEH performance. The design of an aeroelastic harvester based on the use of two different types of piezoelectric materials is presented. The designed model is suitable for energy harvesting and can be utilized to drive nano and microelectronics. Maximum output power obtained for the designed PAEH is found to be 0.55 mW.
- Optimal commands based multi-stage drag de-orbit design for a tethered
system during large space debris removal- Abstract: Publication date: Available online 3 January 2019Source: Acta AstronauticaAuthor(s): Zhongyi Chu, Tao Wei, Tao Shen, Jingnan Di, Jing Cui, Jun Sun There is a serious challenge to the safe operation of orbiting satellites as the number of space debris increases for its high risks and the possible crippling effects of collisions. Consequently, the active removal scenario of space debris has drawn wide attention in recent years, and a tethered system is considered to be a promising method for its large operating distance and low power consumption. However, the flexible tether of the system will bring the coupling of the orbit motion, the sway motion and the variation of large debris attitude, which brings about great danger in de-orbiting phase and a huge challenge in later control. Hence, aiming to de-orbit large space debris safely with a tethered system, a multi-stage horizontal drag de-orbit strategy that consists of two stages is designed. At the first stage, the orbit altitude is rose with a tug thrust whose direction is consistent with the tether, which does not provoke the in-plane oscillation of the tethered system. At the second stage, the orbit is circularized by changing the size and direction of the tug thrust. Especially, optimal commands based on the minimum of tangling risks are planned using Gauss pseudospectral method to avoid target tangling and achieve decoupling of the orbit motion, the sway motion and the variation of the target attitude. Then, the stable attitude control of the tethered system is achieved by designing a hybrid fuzzy adaptive proportion differentiation (PD) with hierarchical sliding-mode controller (HSMC) in the de-orbit process. Finally, numerical simulation is implemented to verify the effectiveness of the proposed de-orbit design.
- Abstract: Publication date: Available online 3 January 2019Source: Acta AstronauticaAuthor(s): Zhongyi Chu, Tao Wei, Tao Shen, Jingnan Di, Jing Cui, Jun Sun There is a serious challenge to the safe operation of orbiting satellites as the number of space debris increases for its high risks and the possible crippling effects of collisions. Consequently, the active removal scenario of space debris has drawn wide attention in recent years, and a tethered system is considered to be a promising method for its large operating distance and low power consumption. However, the flexible tether of the system will bring the coupling of the orbit motion, the sway motion and the variation of large debris attitude, which brings about great danger in de-orbiting phase and a huge challenge in later control. Hence, aiming to de-orbit large space debris safely with a tethered system, a multi-stage horizontal drag de-orbit strategy that consists of two stages is designed. At the first stage, the orbit altitude is rose with a tug thrust whose direction is consistent with the tether, which does not provoke the in-plane oscillation of the tethered system. At the second stage, the orbit is circularized by changing the size and direction of the tug thrust. Especially, optimal commands based on the minimum of tangling risks are planned using Gauss pseudospectral method to avoid target tangling and achieve decoupling of the orbit motion, the sway motion and the variation of the target attitude. Then, the stable attitude control of the tethered system is achieved by designing a hybrid fuzzy adaptive proportion differentiation (PD) with hierarchical sliding-mode controller (HSMC) in the de-orbit process. Finally, numerical simulation is implemented to verify the effectiveness of the proposed de-orbit design.
- Space-native construction materials for earth-independent and sustainable
infrastructure- Abstract: Publication date: Available online 10 December 2018Source: Acta AstronauticaAuthor(s): M.Z. Naser A successful space exploration requires establishing permanent and earth-independent infrastructure that are not only resilient to the extreme environment of space but also preferably made of sustainable and indigenous “space-native” construction materials. This review covers feasibility of exploiting in-situ lunar and Martian resources as well as harvesting of elements and compounds, from near Earth objects (NEOs), to produce extraterrestrial materials suitable for construction of space-based infrastructure. This review also details material features and characteristics required to withstand the unique, and harsh, effects of space environment. In essence, this paper reviews past and recent advancements in construction-based materials that could be used in the design and development of space human bases and highlights design consideration for sustainable human settlements. Towards the end of this review, practical and technological challenges associated with development of lunar and Martian indigenous construction materials are identified and examined.
- Abstract: Publication date: Available online 10 December 2018Source: Acta AstronauticaAuthor(s): M.Z. Naser A successful space exploration requires establishing permanent and earth-independent infrastructure that are not only resilient to the extreme environment of space but also preferably made of sustainable and indigenous “space-native” construction materials. This review covers feasibility of exploiting in-situ lunar and Martian resources as well as harvesting of elements and compounds, from near Earth objects (NEOs), to produce extraterrestrial materials suitable for construction of space-based infrastructure. This review also details material features and characteristics required to withstand the unique, and harsh, effects of space environment. In essence, this paper reviews past and recent advancements in construction-based materials that could be used in the design and development of space human bases and highlights design consideration for sustainable human settlements. Towards the end of this review, practical and technological challenges associated with development of lunar and Martian indigenous construction materials are identified and examined.
- Design exploration of combinational spike and opposing jet concept in
hypersonic flows based on CFD calculation and surrogate model- Abstract: Publication date: Available online 10 December 2018Source: Acta AstronauticaAuthor(s): Min Ou, Li Yan, Wei Huang, Tian-tian Zhang The combined thermal protection system has led the drag and heat reduction of hypersonic reentry vehicles to a new developing direction. In order to obtain better resistance to aerodynamic drag and heat while maintain the work stability of the aircraft at the same time, the optimization design of the combined thermal protection system is indispensable. In this paper, a CFD numerical simulation method is combined with a surrogate model based multi-objective optimization algorithm to optimize the configuration of a combinatorial spike and opposing jet thermal protection system in a hypersonic flow with the freestream Mach number of 5.75. The obtained results show that when the total drag coefficient (Cd) and the heat fluxes on the head of the blunt body (Q) are the objective functions, the diameter ratio of the aerodisk to the blunt body bottom (d/D) has no significant effect on the drag and heat flux reduction, and it can be neglected in the optimization process. While the objective functions Cd and Q are affected by the three variables the length-to-diameter ratio of the aerospike (L/D), the jet pressure ratio (PR), and the nozzle radius (r0) in similar ways. Therefore, in this case, the optimization problem can be transformed into a single-objective optimization problem. The quadratic response surface model established by sample points obtained by the orthogonal experimental design method is of high simulation accuracy, with the determination coefficients of Cd and Q are 0.964 and 0.965 respectively, and there is only a difference of −6.96% in the objective function Cd, −0.93% in Q between the optimization results and the CFD results. The combined thermal protection system has better performance in both drag and heat reduction than the single spike or opposing jet systems. Compared with the pure blunt body, the total drag coefficient and heat flux on the head of the blunt body with the combined thermal protection system have significantly decreased, with the Cd goes down 86.66%, and the wall heat flux Q drops off 96.37%.
- Abstract: Publication date: Available online 10 December 2018Source: Acta AstronauticaAuthor(s): Min Ou, Li Yan, Wei Huang, Tian-tian Zhang The combined thermal protection system has led the drag and heat reduction of hypersonic reentry vehicles to a new developing direction. In order to obtain better resistance to aerodynamic drag and heat while maintain the work stability of the aircraft at the same time, the optimization design of the combined thermal protection system is indispensable. In this paper, a CFD numerical simulation method is combined with a surrogate model based multi-objective optimization algorithm to optimize the configuration of a combinatorial spike and opposing jet thermal protection system in a hypersonic flow with the freestream Mach number of 5.75. The obtained results show that when the total drag coefficient (Cd) and the heat fluxes on the head of the blunt body (Q) are the objective functions, the diameter ratio of the aerodisk to the blunt body bottom (d/D) has no significant effect on the drag and heat flux reduction, and it can be neglected in the optimization process. While the objective functions Cd and Q are affected by the three variables the length-to-diameter ratio of the aerospike (L/D), the jet pressure ratio (PR), and the nozzle radius (r0) in similar ways. Therefore, in this case, the optimization problem can be transformed into a single-objective optimization problem. The quadratic response surface model established by sample points obtained by the orthogonal experimental design method is of high simulation accuracy, with the determination coefficients of Cd and Q are 0.964 and 0.965 respectively, and there is only a difference of −6.96% in the objective function Cd, −0.93% in Q between the optimization results and the CFD results. The combined thermal protection system has better performance in both drag and heat reduction than the single spike or opposing jet systems. Compared with the pure blunt body, the total drag coefficient and heat flux on the head of the blunt body with the combined thermal protection system have significantly decreased, with the Cd goes down 86.66%, and the wall heat flux Q drops off 96.37%.
- Experimental investigations of cavity parameters leading to combustion
oscillation in a supersonic crossflow- Abstract: Publication date: Available online 8 December 2018Source: Acta AstronauticaAuthor(s): Guo-Yan Zhao, Ming-Bo Sun, Xi-Liang Song, Xi-Peng Li, Hong-Bo Wang The effects of cavity parameters on combustion oscillation inside an ethylene-fueled scramjet combustor equipped with a cavity flameholder are experimentally investigated for Mach 5.5 flight conditions. Three certain cases, such as i) longer length-to-depth ratio, ii) sharper aft degree of cavity, iii) closer air throttling downstream of the cavity, exhibit quasi-periodic combustion oscillation, which can be separately attributed to i) larger recirculation volume in the cavity and more mass and heat exchange between cavity shear layer and the core flow, ii) the stronger impinging shock wave in the cavity acting on shear layer, iii) improved fuel/air mixing owing to the interaction between separated boundary layer and combustion. High-speed and schlieren images demonstrate that the cavity and downstream of it act as the most sensitive areas, at the same time, the factors mentioned above can form a thermal throat further triggering flame flashback, which is an indispensable key sub-process of combustion oscillation. The quantitative analysis results obtained from iso-luminosity contour have shown the different distribution trends of flame front and distinct differences of quasi-periodic oscillation frequencies, whereas similar flame propagation speed distributions. In addition, a simplified combustion opening system model has been established to analyze combustion oscillation mechanisms, which theoretically demonstrates that above factors can destroy the balance of heat release and dissipation, causing the system cannot self-stabilize once certain temperature fluctuation thresholds in sensitive areas are exceeded.
- Abstract: Publication date: Available online 8 December 2018Source: Acta AstronauticaAuthor(s): Guo-Yan Zhao, Ming-Bo Sun, Xi-Liang Song, Xi-Peng Li, Hong-Bo Wang The effects of cavity parameters on combustion oscillation inside an ethylene-fueled scramjet combustor equipped with a cavity flameholder are experimentally investigated for Mach 5.5 flight conditions. Three certain cases, such as i) longer length-to-depth ratio, ii) sharper aft degree of cavity, iii) closer air throttling downstream of the cavity, exhibit quasi-periodic combustion oscillation, which can be separately attributed to i) larger recirculation volume in the cavity and more mass and heat exchange between cavity shear layer and the core flow, ii) the stronger impinging shock wave in the cavity acting on shear layer, iii) improved fuel/air mixing owing to the interaction between separated boundary layer and combustion. High-speed and schlieren images demonstrate that the cavity and downstream of it act as the most sensitive areas, at the same time, the factors mentioned above can form a thermal throat further triggering flame flashback, which is an indispensable key sub-process of combustion oscillation. The quantitative analysis results obtained from iso-luminosity contour have shown the different distribution trends of flame front and distinct differences of quasi-periodic oscillation frequencies, whereas similar flame propagation speed distributions. In addition, a simplified combustion opening system model has been established to analyze combustion oscillation mechanisms, which theoretically demonstrates that above factors can destroy the balance of heat release and dissipation, causing the system cannot self-stabilize once certain temperature fluctuation thresholds in sensitive areas are exceeded.
- Simple solution to optimal cotangential transfer between coplanar elliptic
orbits- Abstract: Publication date: Available online 7 December 2018Source: Acta AstronauticaAuthor(s): Alessandro A. Quarta, Giovanni Mengali The aim of this paper is to propose a semi-analytical method for the analysis of a two-impulse transfer between two coplanar elliptic orbits, assuming each maneuver to change the magnitude of the spacecraft velocity only, without affecting its direction. Using a recent mathematical model that describes the spacecraft dynamics in a compact analytical form within a two-dimensional multiple-impulse scenario, this work proposes an algorithm to calculate the global minimum velocity variation required to complete the transfer. The characteristics of the optimal transfer trajectory, which is tangent to both the parking and the target orbit, are obtained as a function of a single variable, which defines the angular position of the first maneuver. This feature allows the designer to analyze the cotangential transfer in a parametric form, thus obtaining a trade-off solution between the total velocity variation and the desired characteristics of the transfer orbit.
- Abstract: Publication date: Available online 7 December 2018Source: Acta AstronauticaAuthor(s): Alessandro A. Quarta, Giovanni Mengali The aim of this paper is to propose a semi-analytical method for the analysis of a two-impulse transfer between two coplanar elliptic orbits, assuming each maneuver to change the magnitude of the spacecraft velocity only, without affecting its direction. Using a recent mathematical model that describes the spacecraft dynamics in a compact analytical form within a two-dimensional multiple-impulse scenario, this work proposes an algorithm to calculate the global minimum velocity variation required to complete the transfer. The characteristics of the optimal transfer trajectory, which is tangent to both the parking and the target orbit, are obtained as a function of a single variable, which defines the angular position of the first maneuver. This feature allows the designer to analyze the cotangential transfer in a parametric form, thus obtaining a trade-off solution between the total velocity variation and the desired characteristics of the transfer orbit.
- Reliability modeling and analysis of environmental control and life
support systems of space stations: A literature survey- Abstract: Publication date: Available online 7 December 2018Source: Acta AstronauticaAuthor(s): Garima Sharma, Rajiv Nandan Rai Ultra reliable life support is needed for environmental control and life support systems (ECLSS) of space stations (SS) because provisioning of spare parts or a speedy crew recovery may not be a feasible option. There is limited work on the reliability modeling and analysis (RMA) of ECLSS of a SS. In addition, presently the literature review on RMA of space stations which is the central theme of the paper is not organized in a formalized manner and the ideas presented are spread out and limited in developing a consolidated literature review in a sequential manner. The paper presents a detailed literature survey on the RMA of SS ECLSS. The paper also highlights the research gaps and an overview of latest RMA models that can be applied for the reliability analysis of SS ECLSS. The paper contributes by providing sufficient insights into the complete development of RMA of the SS ECLSS comprehensively for better understanding of the readers.
- Abstract: Publication date: Available online 7 December 2018Source: Acta AstronauticaAuthor(s): Garima Sharma, Rajiv Nandan Rai Ultra reliable life support is needed for environmental control and life support systems (ECLSS) of space stations (SS) because provisioning of spare parts or a speedy crew recovery may not be a feasible option. There is limited work on the reliability modeling and analysis (RMA) of ECLSS of a SS. In addition, presently the literature review on RMA of space stations which is the central theme of the paper is not organized in a formalized manner and the ideas presented are spread out and limited in developing a consolidated literature review in a sequential manner. The paper presents a detailed literature survey on the RMA of SS ECLSS. The paper also highlights the research gaps and an overview of latest RMA models that can be applied for the reliability analysis of SS ECLSS. The paper contributes by providing sufficient insights into the complete development of RMA of the SS ECLSS comprehensively for better understanding of the readers.
- Mars atmospheric entry guidance for optimal terminal altitude
- Abstract: Publication date: Available online 6 December 2018Source: Acta AstronauticaAuthor(s): Jiateng Long, Ai Gao, Pingyuan Cui, Yang Liu Maximizing the terminal altitude for Mars atmospheric entry has long been investigated on trajectory design to allow a sufficient timeline margin for subsequent operations and the scientific requirements of exploring Mars ancient highland. The purpose of this paper is to design an onboard Mars atmospheric entry guidance algorithm, which can achieve the optimal terminal altitude at the predetermined terminal flight range. Two important characters that the optimal bank angle profile are revealed in this paper, offering the gateway to the application of the optimal guidance by onboard parameter searching, which will be accomplished by the numerical predictor-corrector strategy. Moreover, suboptimal situations are also investigated considering the performance restriction of reactive control system (RCS). Effectiveness of the proposed guidance algorithm is demonstrated using scenarios of the Mars Science Laboratory (MSL) mission.
- Abstract: Publication date: Available online 6 December 2018Source: Acta AstronauticaAuthor(s): Jiateng Long, Ai Gao, Pingyuan Cui, Yang Liu Maximizing the terminal altitude for Mars atmospheric entry has long been investigated on trajectory design to allow a sufficient timeline margin for subsequent operations and the scientific requirements of exploring Mars ancient highland. The purpose of this paper is to design an onboard Mars atmospheric entry guidance algorithm, which can achieve the optimal terminal altitude at the predetermined terminal flight range. Two important characters that the optimal bank angle profile are revealed in this paper, offering the gateway to the application of the optimal guidance by onboard parameter searching, which will be accomplished by the numerical predictor-corrector strategy. Moreover, suboptimal situations are also investigated considering the performance restriction of reactive control system (RCS). Effectiveness of the proposed guidance algorithm is demonstrated using scenarios of the Mars Science Laboratory (MSL) mission.
- Power system analysis and optimization of a modular experiment Carrier
during an analog lunar demo mission on a volcanic environment- Abstract: Publication date: Available online 5 December 2018Source: Acta AstronauticaAuthor(s): Georgios Tsakyridis, Caroline Lange, Stephan Siegfried Jahnke, Lars Witte, Norbert Toth, Marco Scharringhausen, Nikolas I. Xiros The ROBEX (Robotic Exploration of Extreme Environments) alliance, as formed by the German Helmholtz association, aims to explore synergies and bring together technological challenges and scientific questions between two, up to now unrelated, fields: space and deep sea. The final goal of the alliance targets field tests for available and newly developed instrumentation for the deep sea and on a terrestrial lunar analogue. In this regard, two different test campaigns were conducted, one in the area of Svalbard, Norway and one on mount Etna in Sicily, Italy. The volcano environment served as a lunar analogue, enabling seismic scientific experiments and testing of robotic mobility algorithms. The complete field mission infrastructure consists of a stationary lander, a mobile element and instrument carriers. The modular instrument carrier, commonly referred as Remote Unit (RU), was developed accounting for two different mass requirements: 3 kg (RU3) and 10 kg (RU10). While developed in the frame of ROBEX resumes the idea of a lightweight instrument carrier as developed for the MASCOT (Mobile Asteroid surface scout) mission. The RU houses the instrument, shelters it and provides all essential support functions such as rudimentary thermal control (via foil covering), power provision, data acquisition and handling and data transmission to the control centre. This paper presents theoretical and experimental results of the RU3 power subsystem analysis during the mount Etna field campaign. Drawing upon this analysis, necessary adjustments and revisions to further develop the system towards a more power efficient structure for terrestrial and extraterrestrial usage can be concluded.
- Abstract: Publication date: Available online 5 December 2018Source: Acta AstronauticaAuthor(s): Georgios Tsakyridis, Caroline Lange, Stephan Siegfried Jahnke, Lars Witte, Norbert Toth, Marco Scharringhausen, Nikolas I. Xiros The ROBEX (Robotic Exploration of Extreme Environments) alliance, as formed by the German Helmholtz association, aims to explore synergies and bring together technological challenges and scientific questions between two, up to now unrelated, fields: space and deep sea. The final goal of the alliance targets field tests for available and newly developed instrumentation for the deep sea and on a terrestrial lunar analogue. In this regard, two different test campaigns were conducted, one in the area of Svalbard, Norway and one on mount Etna in Sicily, Italy. The volcano environment served as a lunar analogue, enabling seismic scientific experiments and testing of robotic mobility algorithms. The complete field mission infrastructure consists of a stationary lander, a mobile element and instrument carriers. The modular instrument carrier, commonly referred as Remote Unit (RU), was developed accounting for two different mass requirements: 3 kg (RU3) and 10 kg (RU10). While developed in the frame of ROBEX resumes the idea of a lightweight instrument carrier as developed for the MASCOT (Mobile Asteroid surface scout) mission. The RU houses the instrument, shelters it and provides all essential support functions such as rudimentary thermal control (via foil covering), power provision, data acquisition and handling and data transmission to the control centre. This paper presents theoretical and experimental results of the RU3 power subsystem analysis during the mount Etna field campaign. Drawing upon this analysis, necessary adjustments and revisions to further develop the system towards a more power efficient structure for terrestrial and extraterrestrial usage can be concluded.
- Cost estimating of commercial smallsat launch vehicles
- Abstract: Publication date: Available online 4 December 2018Source: Acta AstronauticaAuthor(s): N.T. Drenthe, B.T.C. Zandbergen, R. Curran, M.O. Van Pelt Commercial launch service providers’ low-priced offerings have been a hotly debated topic. However, the strategies with which these firms reduce costs have seen little incorporation into the hardware Cost Estimating Methods (CEMs) and tools prevalent in the aerospace industry. This research changes this, by providing adaptations to agency-focused CEMs that befit a new commercial paradigm, with an emphasis on smallsat launch vehicles.A parametric model for estimating costs in an early phase of development was synthesized, with which it is possible to approximate the full life-cycle costs of small commercial liquid and solid propellant rockets, as well as their cost-based price per flight. Key elements included were reductions in cost achieved by commercial launch operators, by modeling reduced subcontractor management effort and profit retention experienced at lower subcontracting rates.Prices per flight of small commercial launch vehicles were approximated by combining a parametric cost estimating methodology used frequently in the context of space agencies such as ESA and NASA, called the T1 Equivalents method, with another parametric three-part estimate developed by Koelle for development, manufacture and operations phase costs. The first two phases were estimated through the T1 method, while the operations costs were modeled with TRANSCOST.Along with the newly developed methodologies, novel insights such as required launch rates have shone a light on small commercial launch systems’ cost feasibility in the age of public-private spaceflight partnerships.The model developed was able to approximate costs of development, manufacture and price per flight of three commercial rockets to within 20% of actual reported costs or prices. However, it is recommended the model is refined as more reference cost data, especially on a subsystem level, as well as pricing for these smaller rockets becomes available in the coming years.
- Abstract: Publication date: Available online 4 December 2018Source: Acta AstronauticaAuthor(s): N.T. Drenthe, B.T.C. Zandbergen, R. Curran, M.O. Van Pelt Commercial launch service providers’ low-priced offerings have been a hotly debated topic. However, the strategies with which these firms reduce costs have seen little incorporation into the hardware Cost Estimating Methods (CEMs) and tools prevalent in the aerospace industry. This research changes this, by providing adaptations to agency-focused CEMs that befit a new commercial paradigm, with an emphasis on smallsat launch vehicles.A parametric model for estimating costs in an early phase of development was synthesized, with which it is possible to approximate the full life-cycle costs of small commercial liquid and solid propellant rockets, as well as their cost-based price per flight. Key elements included were reductions in cost achieved by commercial launch operators, by modeling reduced subcontractor management effort and profit retention experienced at lower subcontracting rates.Prices per flight of small commercial launch vehicles were approximated by combining a parametric cost estimating methodology used frequently in the context of space agencies such as ESA and NASA, called the T1 Equivalents method, with another parametric three-part estimate developed by Koelle for development, manufacture and operations phase costs. The first two phases were estimated through the T1 method, while the operations costs were modeled with TRANSCOST.Along with the newly developed methodologies, novel insights such as required launch rates have shone a light on small commercial launch systems’ cost feasibility in the age of public-private spaceflight partnerships.The model developed was able to approximate costs of development, manufacture and price per flight of three commercial rockets to within 20% of actual reported costs or prices. However, it is recommended the model is refined as more reference cost data, especially on a subsystem level, as well as pricing for these smaller rockets becomes available in the coming years.
- Application of a planar air-bearing microgravity simulator for
demonstration of operations required for an orbital capture with a
manipulator- Abstract: Publication date: Available online 3 December 2018Source: Acta AstronauticaAuthor(s): Tomasz Rybus, Karol Seweryn, Jakub Oleś, Fatina Liliana Basmadji, Kamil Tarenko, Radosław Moczydłowski, Tomasz Barciński, Jan Kindracki, Łukasz Mężyk, Przemysław Paszkiewicz, Piotr Wolański Microgravity is the aspect of space environment that is important for robotic technologies. One possible approach to simulate microgravity conditions on Earth is to use planar air-bearing microgravity simulators. In this study we have assessed the possible application of such simulator for demonstration of operations required for the orbital capture manoeuvre. In our experiments we have used mock-up of satellite equipped with a manipulator and a set of cold-gas thrusters. Results of two sets of experiments are presented. In both sets the manipulator end-effector followed a straight-line trajectory. In the first set the satellite mock-up was commanded to hold fixed position and orientation, while in the second set the satellite mock-up was commanded to follow a pre-defined trajectory. Performed experiments show that the accuracy of the planar air-bearing microgravity simulator is sufficient to use this facility to validate control algorithms and control strategies. Presented analysis also shows that changes of the satellite-manipulator system parameters caused by the consumption of air and nitrogen from gas canisters located on the satellite mock-up have negligible influence on the dynamic behaviour of the system.
- Abstract: Publication date: Available online 3 December 2018Source: Acta AstronauticaAuthor(s): Tomasz Rybus, Karol Seweryn, Jakub Oleś, Fatina Liliana Basmadji, Kamil Tarenko, Radosław Moczydłowski, Tomasz Barciński, Jan Kindracki, Łukasz Mężyk, Przemysław Paszkiewicz, Piotr Wolański Microgravity is the aspect of space environment that is important for robotic technologies. One possible approach to simulate microgravity conditions on Earth is to use planar air-bearing microgravity simulators. In this study we have assessed the possible application of such simulator for demonstration of operations required for the orbital capture manoeuvre. In our experiments we have used mock-up of satellite equipped with a manipulator and a set of cold-gas thrusters. Results of two sets of experiments are presented. In both sets the manipulator end-effector followed a straight-line trajectory. In the first set the satellite mock-up was commanded to hold fixed position and orientation, while in the second set the satellite mock-up was commanded to follow a pre-defined trajectory. Performed experiments show that the accuracy of the planar air-bearing microgravity simulator is sufficient to use this facility to validate control algorithms and control strategies. Presented analysis also shows that changes of the satellite-manipulator system parameters caused by the consumption of air and nitrogen from gas canisters located on the satellite mock-up have negligible influence on the dynamic behaviour of the system.
- A high-accuracy autonomous navigation scheme for the Mars rover
- Abstract: Publication date: January 2019Source: Acta Astronautica, Volume 154Author(s): Yunan Zhao, Xinlong Wang, Qunsheng Li, Dun Wang, Yuanwen Cai High-accuracy, autonomous and reliable navigation systems are important foundations for Mars rovers to achieve exploration missions successfully. Based on the motion characteristics and working environment of the rover, the paper shows a high-accuracy strapdown inertial navigation system/visual navigation system/celestial navigation system (SINS/VNS/CNS) integrated navigation scheme suitable for the rover with long-time and long-distance motion. According to the feature point positions in the camera frame at adjacent time obtained by the binocular visual odometry, its velocity in the camera frame and the attitude are calculated, then a subsystem model of SINS/VNS integrated navigation is established. In addition, using the star vectors measured by the large field-of-view star sensor, a high-accuracy attitude matrix of the rover in the inertial frame can be obtained, and a SINS/CNS subsystem model is established. Furthermore, in order to make full use of the complementary advantages of the two subsystems in attitude and position estimation, an interacting multiple model filter is developed. By updating the model probabilities of the subsystems separately in real time, the filter can output accurate navigation information of the rover. Simulation results show that the proposed navigation scheme can significantly improve the estimation accuracy of attitude, position and velocity simultaneously.
- Abstract: Publication date: January 2019Source: Acta Astronautica, Volume 154Author(s): Yunan Zhao, Xinlong Wang, Qunsheng Li, Dun Wang, Yuanwen Cai High-accuracy, autonomous and reliable navigation systems are important foundations for Mars rovers to achieve exploration missions successfully. Based on the motion characteristics and working environment of the rover, the paper shows a high-accuracy strapdown inertial navigation system/visual navigation system/celestial navigation system (SINS/VNS/CNS) integrated navigation scheme suitable for the rover with long-time and long-distance motion. According to the feature point positions in the camera frame at adjacent time obtained by the binocular visual odometry, its velocity in the camera frame and the attitude are calculated, then a subsystem model of SINS/VNS integrated navigation is established. In addition, using the star vectors measured by the large field-of-view star sensor, a high-accuracy attitude matrix of the rover in the inertial frame can be obtained, and a SINS/CNS subsystem model is established. Furthermore, in order to make full use of the complementary advantages of the two subsystems in attitude and position estimation, an interacting multiple model filter is developed. By updating the model probabilities of the subsystems separately in real time, the filter can output accurate navigation information of the rover. Simulation results show that the proposed navigation scheme can significantly improve the estimation accuracy of attitude, position and velocity simultaneously.
- Impact of pure favorable pressure gradient on a supersonic flat-plate
turbulent boundary layer- Abstract: Publication date: January 2019Source: Acta Astronautica, Volume 154Author(s): Qian-cheng Wang, Zhen-guo Wang, Yu-xin Zhao By employing the particle image velocimetry and Nanoparticle-based Planar Laser Scattering method, the impact of streamwise favorable pressure gradient on the mean and turbulent characteristics of a Mach 2.95 turbulent boundary layer is experimentally investigated. Through a careful arrangement of the experiment, the possible influence of streamline convex curvature is minimized to have a flat-plate streamwise favorable-pressure-gradient boundary layer. While the log law is found to be well preserved at all streamwise positions, the wake region is weakened by the pressure gradient. Different from the boundary layers formed over the convex wall and over the sudden expansion ramp, both principal strain rate and bulk dilatation in the near-wall region are found to be barely changed along the streamwise direction in the flat-plate favorable-pressure-gradient boundary layer. Because of this, the impact of the favorable pressure gradient on the near-wall turbulent fluctuation is also found to be insignificant, which differs from the supersonic convex boundary layer where the near-wall turbulence is found to be greatly suppressed.
- Abstract: Publication date: January 2019Source: Acta Astronautica, Volume 154Author(s): Qian-cheng Wang, Zhen-guo Wang, Yu-xin Zhao By employing the particle image velocimetry and Nanoparticle-based Planar Laser Scattering method, the impact of streamwise favorable pressure gradient on the mean and turbulent characteristics of a Mach 2.95 turbulent boundary layer is experimentally investigated. Through a careful arrangement of the experiment, the possible influence of streamline convex curvature is minimized to have a flat-plate streamwise favorable-pressure-gradient boundary layer. While the log law is found to be well preserved at all streamwise positions, the wake region is weakened by the pressure gradient. Different from the boundary layers formed over the convex wall and over the sudden expansion ramp, both principal strain rate and bulk dilatation in the near-wall region are found to be barely changed along the streamwise direction in the flat-plate favorable-pressure-gradient boundary layer. Because of this, the impact of the favorable pressure gradient on the near-wall turbulent fluctuation is also found to be insignificant, which differs from the supersonic convex boundary layer where the near-wall turbulence is found to be greatly suppressed.
- The effect of coolant injection from the tip of spike on aerodynamic
heating of nose cone at supersonic flow- Abstract: Publication date: January 2019Source: Acta Astronautica, Volume 154Author(s): F. Pish, Rasoul Moradi, Amirhossein Edalatpour, M. Barzegar Gerdroodbary Aerodynamic heating is the main challenge for increasing the speed of hypervelocity vehicles. In this work, three-dimensional numerical simulations have been performed to investigate the influence of the cooling injection from the tip of the spike mounted on the nose cone. This work comprehensively focused on the effect of different types of gas injections (Air, He and CO2) on the flow feature and mechanism of cooling through three-dimensional simulations. Comprehensive parametric studies are performed to reveal the main effective parameters on the cooling performance along the nose cone. In order to perform numerical simulations, Reynolds-averaged Navier–Stokes equations with Menter's Shear Stress Transport (SST) turbulence model are applied. Our findings show that the injection of the CO2 jet from the tip of a spike is more efficient on the cooling of the nose cone due to the formation of larger circulation in the vicinity of the spike. In addition, the mass distribution of these two jets confirms that helium jets penetrate more in the upstream. Our results clearly show that the effect of the helium in low jet pressure is more significant on the cooling performance of the nose cone.
- Abstract: Publication date: January 2019Source: Acta Astronautica, Volume 154Author(s): F. Pish, Rasoul Moradi, Amirhossein Edalatpour, M. Barzegar Gerdroodbary Aerodynamic heating is the main challenge for increasing the speed of hypervelocity vehicles. In this work, three-dimensional numerical simulations have been performed to investigate the influence of the cooling injection from the tip of the spike mounted on the nose cone. This work comprehensively focused on the effect of different types of gas injections (Air, He and CO2) on the flow feature and mechanism of cooling through three-dimensional simulations. Comprehensive parametric studies are performed to reveal the main effective parameters on the cooling performance along the nose cone. In order to perform numerical simulations, Reynolds-averaged Navier–Stokes equations with Menter's Shear Stress Transport (SST) turbulence model are applied. Our findings show that the injection of the CO2 jet from the tip of a spike is more efficient on the cooling of the nose cone due to the formation of larger circulation in the vicinity of the spike. In addition, the mass distribution of these two jets confirms that helium jets penetrate more in the upstream. Our results clearly show that the effect of the helium in low jet pressure is more significant on the cooling performance of the nose cone.
- Characterization of heat release rate by OH* and CH* chemiluminescence
- Abstract: Publication date: January 2019Source: Acta Astronautica, Volume 154Author(s): Yao Liu, Jianguo Tan, Hao Wang, Liang Lv As single component chemiluminescence cannot accurately measure the distribution of heat release rate, this paper put forward the idea of employing OH* and CH* chemiluminescence for the measurement. However, it is difficult to characterize the relationship between chemiluminescence and heat release rate, so the method of deep learning was applied to process numerical simulation results of methane-air steady premixed flames and a deep neural network model was developed to determine the distribution of heat release rate with OH* and CH* chemiluminescence. The evaluation indexes of the model performed satisfactorily: root mean square error of the normalized heat release rate is less than 0.13, mean relative error of peak and opening distance is below 3%, and mean error of peak position less than 0.01 mm. The validation results showed that OH* and CH* chemiluminescence could properly measure the distribution of heat release rate: relative error of peak ranges from −6% to 6%; on test dataset, correlation coefficient between predictive results and simulation results is above 0.99 in terms of heat release rate peak, peak positions and opening distance.
- Abstract: Publication date: January 2019Source: Acta Astronautica, Volume 154Author(s): Yao Liu, Jianguo Tan, Hao Wang, Liang Lv As single component chemiluminescence cannot accurately measure the distribution of heat release rate, this paper put forward the idea of employing OH* and CH* chemiluminescence for the measurement. However, it is difficult to characterize the relationship between chemiluminescence and heat release rate, so the method of deep learning was applied to process numerical simulation results of methane-air steady premixed flames and a deep neural network model was developed to determine the distribution of heat release rate with OH* and CH* chemiluminescence. The evaluation indexes of the model performed satisfactorily: root mean square error of the normalized heat release rate is less than 0.13, mean relative error of peak and opening distance is below 3%, and mean error of peak position less than 0.01 mm. The validation results showed that OH* and CH* chemiluminescence could properly measure the distribution of heat release rate: relative error of peak ranges from −6% to 6%; on test dataset, correlation coefficient between predictive results and simulation results is above 0.99 in terms of heat release rate peak, peak positions and opening distance.
- Thermal-structural analysis for flexible spacecraft with single or double
solar panels: A comparison study- Abstract: Publication date: January 2019Source: Acta Astronautica, Volume 154Author(s): Lun Liu, Shupeng Sun, Dengqing Cao, Xiyu Liu The thermal-structural analysis for a flexible spacecraft with double solar panels is carried out in this paper through a comparison study with spacecraft having a single panel. The solar panels are composed of honeycomb panel and subjected to time-varying thermal loading. Taking into account the coupling effect among attitude motion, structural deformation and thermal loading, the rigid-flexible-thermal coupling dynamic model of the spacecraft is established by using the Hamiltonian Principle. Based on the finite difference method, an explicit algorithm is developed to solve the transient heat conduction problem of the solar panel. The coupled thermal-structural analysis reveals significant differences between the dynamic characteristics of thermally induced vibration of spacecraft with single and double solar panels. The thermally induced dynamic response significantly affects the attitude of spacecraft with a single solar panel, while it hardly affects the attitude of spacecraft with double solar panels. As the maneuver attitude or the initial incident angle of heat flux increase, the thermally induced vibration of spacecraft with a single solar panel changes from stable to unstable and thermal flutter occurs, while that of spacecraft with double solar panels always keeps stable.
- Abstract: Publication date: January 2019Source: Acta Astronautica, Volume 154Author(s): Lun Liu, Shupeng Sun, Dengqing Cao, Xiyu Liu The thermal-structural analysis for a flexible spacecraft with double solar panels is carried out in this paper through a comparison study with spacecraft having a single panel. The solar panels are composed of honeycomb panel and subjected to time-varying thermal loading. Taking into account the coupling effect among attitude motion, structural deformation and thermal loading, the rigid-flexible-thermal coupling dynamic model of the spacecraft is established by using the Hamiltonian Principle. Based on the finite difference method, an explicit algorithm is developed to solve the transient heat conduction problem of the solar panel. The coupled thermal-structural analysis reveals significant differences between the dynamic characteristics of thermally induced vibration of spacecraft with single and double solar panels. The thermally induced dynamic response significantly affects the attitude of spacecraft with a single solar panel, while it hardly affects the attitude of spacecraft with double solar panels. As the maneuver attitude or the initial incident angle of heat flux increase, the thermally induced vibration of spacecraft with a single solar panel changes from stable to unstable and thermal flutter occurs, while that of spacecraft with double solar panels always keeps stable.
- Optimization design of launch locking protective device (LLPD) based on
carbon fiber bracket for magnetically suspended flywheel (MSFW)- Abstract: Publication date: January 2019Source: Acta Astronautica, Volume 154Author(s): Liu Qiang, Wang Kun, Ren Yuan, Chen Xiaocen, Ma Limei, Zhao Yong Because of the launch vibration and shock, magnetically suspended flywheels (MSFWs) are equipped with an additional launch locking protective device (LLPD), and the LLPD performance has great influence on the attitude control precision of the flywheel system. In this paper, a LLPD that takes the carbon fiber bracket as the key clamped and releasable mechanism was presented. And the configuration, operating principle and functional performance requirements were introduced. The locking/unlocking force, maximum stress and contact force of the carbon fiber bracket were analyzed. The dynamic analysis of the single carbon fiber bracket equivalent to the cantilever beam model was carried out. Subsequently, the sensitivity of the constraint variables vs the structural parameters was calculated. The lower and upper parts of the carbon fiber bracket were separately optimized. The result shows that the mass of the carbon fiber bracket can reach to the minimum of 60.5 g when the number of the upper carbon fiber bracket slices is 12. Finally, the LLPD prototype was manufactured and its locking protection for the flywheel system was verified by the swept-sine vibration and the random vibration.
- Abstract: Publication date: January 2019Source: Acta Astronautica, Volume 154Author(s): Liu Qiang, Wang Kun, Ren Yuan, Chen Xiaocen, Ma Limei, Zhao Yong Because of the launch vibration and shock, magnetically suspended flywheels (MSFWs) are equipped with an additional launch locking protective device (LLPD), and the LLPD performance has great influence on the attitude control precision of the flywheel system. In this paper, a LLPD that takes the carbon fiber bracket as the key clamped and releasable mechanism was presented. And the configuration, operating principle and functional performance requirements were introduced. The locking/unlocking force, maximum stress and contact force of the carbon fiber bracket were analyzed. The dynamic analysis of the single carbon fiber bracket equivalent to the cantilever beam model was carried out. Subsequently, the sensitivity of the constraint variables vs the structural parameters was calculated. The lower and upper parts of the carbon fiber bracket were separately optimized. The result shows that the mass of the carbon fiber bracket can reach to the minimum of 60.5 g when the number of the upper carbon fiber bracket slices is 12. Finally, the LLPD prototype was manufactured and its locking protection for the flywheel system was verified by the swept-sine vibration and the random vibration.
- Velocity-free sliding mode control for spacecraft with input saturation
- Abstract: Publication date: January 2019Source: Acta Astronautica, Volume 154Author(s): Yong Guo, Bing Huang, Jin-hua Guo, Ai-jun Li, Chang-qing Wang This paper investigates the robust control for spacecraft without angular velocity measurement subject to the external disturbances and input saturation constraints. A novel integral sliding mode surface that is suitable to solve the problem of input saturation constraints is established based on the hyperbolic tangent function. Both controllers can deal with the actuator saturation and external disturbances simultaneously by using the sliding mode surface. The first controller is full state feedback, while the second one just contains attitude information, where the angular velocity is unnecessary due to the existence of a finite-time observer. Finally, Lyapunov theory and simulation results are provided to illustrate the effectiveness of the controllers.
- Abstract: Publication date: January 2019Source: Acta Astronautica, Volume 154Author(s): Yong Guo, Bing Huang, Jin-hua Guo, Ai-jun Li, Chang-qing Wang This paper investigates the robust control for spacecraft without angular velocity measurement subject to the external disturbances and input saturation constraints. A novel integral sliding mode surface that is suitable to solve the problem of input saturation constraints is established based on the hyperbolic tangent function. Both controllers can deal with the actuator saturation and external disturbances simultaneously by using the sliding mode surface. The first controller is full state feedback, while the second one just contains attitude information, where the angular velocity is unnecessary due to the existence of a finite-time observer. Finally, Lyapunov theory and simulation results are provided to illustrate the effectiveness of the controllers.
- IFC - Publication Information
- Abstract: Publication date: January 2019Source: Acta Astronautica, Volume 154Author(s):
- Abstract: Publication date: January 2019Source: Acta Astronautica, Volume 154Author(s):
- Flow characteristics of a pintle injector element
- Abstract: Publication date: January 2019Source: Acta Astronautica, Volume 154Author(s): Peng Cheng, Qinglian Li, Huiyuan Chen The flow characteristics of a pintle injector element were studied by theory and experiments. A spray angle model was firstly proposed based on our previous work, and then a model on the discharge coefficient of the orifice was also derived. Then, experiments were conducted by a pintle injector element which was simplified from pintle injectors with discrete radial orifices. Snapshots of the sprays were captured by using backlit photography technique. Results show that the pintle injector forms an arched cloak-like spray. The spray pattern and breakup are heavily related to the local momentum ratio and Reynolds numbers of jet and film. The spray angles obtained from the snapshots coincide with theory well. Meanwhile, a semi-empirical model of discharge coefficient of the orifice was proposed based on the experiments. It is found that the spray angle and discharge coefficient are dominated by the local momentum ratio. The spray angle and discharge coefficient increase along with local momentum ratio and tend to be stable for large local momentum ratios. Besides, discharge coefficient of the orifice is also found to be slightly affected by the ratio of jet diameter to film thickness.
- Abstract: Publication date: January 2019Source: Acta Astronautica, Volume 154Author(s): Peng Cheng, Qinglian Li, Huiyuan Chen The flow characteristics of a pintle injector element were studied by theory and experiments. A spray angle model was firstly proposed based on our previous work, and then a model on the discharge coefficient of the orifice was also derived. Then, experiments were conducted by a pintle injector element which was simplified from pintle injectors with discrete radial orifices. Snapshots of the sprays were captured by using backlit photography technique. Results show that the pintle injector forms an arched cloak-like spray. The spray pattern and breakup are heavily related to the local momentum ratio and Reynolds numbers of jet and film. The spray angles obtained from the snapshots coincide with theory well. Meanwhile, a semi-empirical model of discharge coefficient of the orifice was proposed based on the experiments. It is found that the spray angle and discharge coefficient are dominated by the local momentum ratio. The spray angle and discharge coefficient increase along with local momentum ratio and tend to be stable for large local momentum ratios. Besides, discharge coefficient of the orifice is also found to be slightly affected by the ratio of jet diameter to film thickness.
- Thermal protection characteristics for a combinational opposing jet and
platelet transpiration cooling nose-tip- Abstract: Publication date: Available online 1 December 2018Source: Acta AstronauticaAuthor(s): BinXian Shen, Liang Yin, HongPeng Liu, Weiqiang Liu A combinational opposing jet and platelet transpiration cooling nose-tip is analyzed. Platelet transpiration is introduced to the opposing jet thermal protection system (TPS) in hypersonic vehicles to enhance the cooling efficiency in the reattachment region without increasing the overall cooling intensity, thereby saving the total coolant consumption. In this study, the combinational cooling nose-tip is simplified with a limited number of enlarged transpiration orifices instead of micro-pores. The two-equation shear stress transport k-w turbulence model is utilized to study the flow field and surface heat transfer in hypersonic flow, and the numerical method is validated against experimental data available in the literature. The peak heat flux reduces more than 6.6% of the combinational cooling nose-tip than that of the opposing jet TPS with the same total coolant consumption. Thus, platelet transpiration can help save the coolant consumption to reach a similar cooling efficiency. The transpiration gas can cover the surface to insulate heating. Results show that the cooling efficiency increases with increasing transpiration intensity. Finally, the peak heat flux of the four-orifice model reduces more than 3.7% than that of the two-orifice model. The cooling efficiency is improved as the number of orifices is increased. This result indicates that the cooling efficiency of the simplified model is underestimated than that of real structures with a large number of transpiration micro-pores.
- Abstract: Publication date: Available online 1 December 2018Source: Acta AstronauticaAuthor(s): BinXian Shen, Liang Yin, HongPeng Liu, Weiqiang Liu A combinational opposing jet and platelet transpiration cooling nose-tip is analyzed. Platelet transpiration is introduced to the opposing jet thermal protection system (TPS) in hypersonic vehicles to enhance the cooling efficiency in the reattachment region without increasing the overall cooling intensity, thereby saving the total coolant consumption. In this study, the combinational cooling nose-tip is simplified with a limited number of enlarged transpiration orifices instead of micro-pores. The two-equation shear stress transport k-w turbulence model is utilized to study the flow field and surface heat transfer in hypersonic flow, and the numerical method is validated against experimental data available in the literature. The peak heat flux reduces more than 6.6% of the combinational cooling nose-tip than that of the opposing jet TPS with the same total coolant consumption. Thus, platelet transpiration can help save the coolant consumption to reach a similar cooling efficiency. The transpiration gas can cover the surface to insulate heating. Results show that the cooling efficiency increases with increasing transpiration intensity. Finally, the peak heat flux of the four-orifice model reduces more than 3.7% than that of the two-orifice model. The cooling efficiency is improved as the number of orifices is increased. This result indicates that the cooling efficiency of the simplified model is underestimated than that of real structures with a large number of transpiration micro-pores.
- Optical fragment tracking in hypervelocity impact experiments
- Abstract: Publication date: Available online 1 December 2018Source: Acta AstronauticaAuthor(s): Erkai Watson, Max Gulde, Lukas Kortmann, Masumi Higashide, Frank Schaefer, Stefan Hiermaier In-orbit impacts between satellites and space debris lead to varying degrees of fragmentation, ranging anywhere from minor damage to complete breakups. In this paper, we describe an experimental measurement approach for studying fragmentation caused by hypervelocity impact in the laboratory. We investigate impacts on thin aluminum bumper plates with the goal of measuring individual fragment velocities and sizes generated by hypervelocity impact. The experimental setup, commonly used in fluid dynamics for Particle Tracking Velocimetry, consists of using a laser plane and high-speed video camera to track the motion of debris fragments. We describe the fragment tracking algorithm and demonstrate its ability to determine fragment velocity and sizes in specific hypervelocity impact experiments performed at Fraunhofer EMI. The measurement technique enables quantitative data, at an unprecedented level of detail, to be measured from hypervelocity impact fragmentation experiments in the lab, which can be applied to understanding the effects of satellite collisions and improving breakup models.
- Abstract: Publication date: Available online 1 December 2018Source: Acta AstronauticaAuthor(s): Erkai Watson, Max Gulde, Lukas Kortmann, Masumi Higashide, Frank Schaefer, Stefan Hiermaier In-orbit impacts between satellites and space debris lead to varying degrees of fragmentation, ranging anywhere from minor damage to complete breakups. In this paper, we describe an experimental measurement approach for studying fragmentation caused by hypervelocity impact in the laboratory. We investigate impacts on thin aluminum bumper plates with the goal of measuring individual fragment velocities and sizes generated by hypervelocity impact. The experimental setup, commonly used in fluid dynamics for Particle Tracking Velocimetry, consists of using a laser plane and high-speed video camera to track the motion of debris fragments. We describe the fragment tracking algorithm and demonstrate its ability to determine fragment velocity and sizes in specific hypervelocity impact experiments performed at Fraunhofer EMI. The measurement technique enables quantitative data, at an unprecedented level of detail, to be measured from hypervelocity impact fragmentation experiments in the lab, which can be applied to understanding the effects of satellite collisions and improving breakup models.
- Numerical analysis of convergence property of heat flux next to the wall
- Abstract: Publication date: Available online 30 November 2018Source: Acta AstronauticaAuthor(s): Yipu Zhao, Yumeng Hu, Haiming Huang The chemical nonequilibrium mathematical model is used to study the effect of the first grid spacing next to the wall on numerical calculation results of the surface heat flux and the stagnation heat flux. This model is composed of two-dimensional axisymmetric Navier–Stokes equations, the thermodynamic relations and the chemical reaction model. The solver is derived by implicit finite volume schemes and is validated by comparing the numerical calculation results with that in the literature. The surface heat flux and the stagnation heat flux are calculated under nine cases in which the cell Reynolds number next to the wall equals to 14.7, 7.3, 5.1, 2.9, 1.5, 1.0, 0.7, 0.5 and 0.25, respectively, and the results indicate that the surface heat flux and the stagnation heat flux gradually converge with the decrease of the cell Reynolds number next to the wall. To further verify the conclusion, the stagnation heat flux is also calculated by changing a single factor, which considers nine cases and the cell Reynolds numbers above are selected for each case, and the consistent result is obtained under nine cases.
- Abstract: Publication date: Available online 30 November 2018Source: Acta AstronauticaAuthor(s): Yipu Zhao, Yumeng Hu, Haiming Huang The chemical nonequilibrium mathematical model is used to study the effect of the first grid spacing next to the wall on numerical calculation results of the surface heat flux and the stagnation heat flux. This model is composed of two-dimensional axisymmetric Navier–Stokes equations, the thermodynamic relations and the chemical reaction model. The solver is derived by implicit finite volume schemes and is validated by comparing the numerical calculation results with that in the literature. The surface heat flux and the stagnation heat flux are calculated under nine cases in which the cell Reynolds number next to the wall equals to 14.7, 7.3, 5.1, 2.9, 1.5, 1.0, 0.7, 0.5 and 0.25, respectively, and the results indicate that the surface heat flux and the stagnation heat flux gradually converge with the decrease of the cell Reynolds number next to the wall. To further verify the conclusion, the stagnation heat flux is also calculated by changing a single factor, which considers nine cases and the cell Reynolds numbers above are selected for each case, and the consistent result is obtained under nine cases.
- `Using predictive Bayesian Monte Carlo- Markov Chain methods to provide a
probablistic solution for the Drake equation- Abstract: Publication date: Available online 29 November 2018Source: Acta AstronauticaAuthor(s): Frederick Bloetscher Are we alone in the universe' It is an age-old question that continues to encourage interest and controversy among the public as well as academics. Development of explanations for life elsewhere ranges widely, but few mathematical models have been developed to measure the likelihood of concurrent, intelligent life, and those that exist are widely speculative due to the lack of information. However, with the addition of information from Kepler explorations for new solar systems within our galaxy, and calculation of the potential number of stars in the expanse of the universe, data for a useful probabilistic model to determine the likelihood of life beyond Earth may be possible with the use of predictive Bayesian statistics. Predictive Bayesian statistical methods are designed to use limited, uncertain data, to develop results. The result provides a probability curve of the likelihood of life in the universe that includes both uncertainty and potential variability within the result to provide a means to define the probability of life in the galaxy as well as life within proximity to earth. That said, the results indicate that the probability we are alone (
- Abstract: Publication date: Available online 29 November 2018Source: Acta AstronauticaAuthor(s): Frederick Bloetscher Are we alone in the universe' It is an age-old question that continues to encourage interest and controversy among the public as well as academics. Development of explanations for life elsewhere ranges widely, but few mathematical models have been developed to measure the likelihood of concurrent, intelligent life, and those that exist are widely speculative due to the lack of information. However, with the addition of information from Kepler explorations for new solar systems within our galaxy, and calculation of the potential number of stars in the expanse of the universe, data for a useful probabilistic model to determine the likelihood of life beyond Earth may be possible with the use of predictive Bayesian statistics. Predictive Bayesian statistical methods are designed to use limited, uncertain data, to develop results. The result provides a probability curve of the likelihood of life in the universe that includes both uncertainty and potential variability within the result to provide a means to define the probability of life in the galaxy as well as life within proximity to earth. That said, the results indicate that the probability we are alone (
- Seismic investigation of icy crust covering subsurface oceans of Europa
and Ganymede: Preliminary assessment of hypothetical experiment using
impactor- Abstract: Publication date: Available online 28 November 2018Source: Acta AstronauticaAuthor(s): Aline Franqui, Spencer Seufert, Masakata Okutsu Jupiter's satellite Europa is believed to harbor a global ocean beneath its ice-covered surface. But the thickness of this ice, despite its significance to the habitability of this moon, is unknown: estimates range from as thin as hundreds of meters to as thick as tens of kilometers. In this paper, we investigate the feasibility of a hypothetical experiment in which the ice's thickness is measured via seismic analysis. The assumed scenario calls for a seismometer to be placed on the satellite's surface to detect the ice surface's seismic response induced by an artificial impact event. Our hypothetical experiment could be applied at Europa as well as at Ganymede. For both satellites two impact scenarios are considered: a low-energy-impact case, in which an orbiter probe impacts the ice at the end of the mission, and a high-energy-impact case, in which a spent upper rocket stage impacts the ice upon Jupiter arrival. We find that an impactor-induced seismic investigation is a promising add-on experiment in future missions to the icy moons of Jupiter.
- Abstract: Publication date: Available online 28 November 2018Source: Acta AstronauticaAuthor(s): Aline Franqui, Spencer Seufert, Masakata Okutsu Jupiter's satellite Europa is believed to harbor a global ocean beneath its ice-covered surface. But the thickness of this ice, despite its significance to the habitability of this moon, is unknown: estimates range from as thin as hundreds of meters to as thick as tens of kilometers. In this paper, we investigate the feasibility of a hypothetical experiment in which the ice's thickness is measured via seismic analysis. The assumed scenario calls for a seismometer to be placed on the satellite's surface to detect the ice surface's seismic response induced by an artificial impact event. Our hypothetical experiment could be applied at Europa as well as at Ganymede. For both satellites two impact scenarios are considered: a low-energy-impact case, in which an orbiter probe impacts the ice at the end of the mission, and a high-energy-impact case, in which a spent upper rocket stage impacts the ice upon Jupiter arrival. We find that an impactor-induced seismic investigation is a promising add-on experiment in future missions to the icy moons of Jupiter.
- An active control strategy to suppress nonlinear vibrations of large space
membranes- Abstract: Publication date: Available online 27 November 2018Source: Acta AstronauticaAuthor(s): Hang Shi, Chao Wang, Lilan Liu, Zenggui Gao, Yangmin Xie Large space membranes have been widely used in many space projects to develop extremely large and lightweight SAR antennas, solar arrays and solar sails, and the nonlinear structural vibrations induced by on-orbit disturbances should be suppressed to satisfy the mission requirements. In this paper, the nonlinear dynamic behaviors of a typical large space membrane, which can not be accurately described by natural frequencies and mode shapes, are investigated by the trustworthy iterative membrane property modeling approach. Based on the discussions of the controller design challenges from such nonlinearity, this paper proposes a distributed control strategy to actively suppress the nonlinear membrane vibration, and comparatively evaluates three different actuating and sensing placements to obtain superior vibration suppression performance. Very promising results numerically demonstrate that the proposed control strategy is theoretically feasible and has great potential to be used in the design of active vibration suppression systems for large space membrane applications.
- Abstract: Publication date: Available online 27 November 2018Source: Acta AstronauticaAuthor(s): Hang Shi, Chao Wang, Lilan Liu, Zenggui Gao, Yangmin Xie Large space membranes have been widely used in many space projects to develop extremely large and lightweight SAR antennas, solar arrays and solar sails, and the nonlinear structural vibrations induced by on-orbit disturbances should be suppressed to satisfy the mission requirements. In this paper, the nonlinear dynamic behaviors of a typical large space membrane, which can not be accurately described by natural frequencies and mode shapes, are investigated by the trustworthy iterative membrane property modeling approach. Based on the discussions of the controller design challenges from such nonlinearity, this paper proposes a distributed control strategy to actively suppress the nonlinear membrane vibration, and comparatively evaluates three different actuating and sensing placements to obtain superior vibration suppression performance. Very promising results numerically demonstrate that the proposed control strategy is theoretically feasible and has great potential to be used in the design of active vibration suppression systems for large space membrane applications.
- Effects of total pressure on mode transition in a dual-mode combustor
- Abstract: Publication date: Available online 26 November 2018Source: Acta AstronauticaAuthor(s): Jianping Li, Guiqian Jiao, Jinyuan Luo, Wenyan Song The mode transition experiments of the dual-mode combustor were carried out under the conditions that the total pressure was 600 kPa, 700 kPa, 800 kPa and 900 kPa, the total temperature was 810 K, and the Mach number was 2.0. According to the wall pressure of the combustor measured in the experiment, the distribution of other airflow parameters of the combustor were obtained through the one-dimensional performance calculation method. The processes of mode transition of the combustor under different total pressures were studied. The study show that the dimensionless peak pressure of the combustor increased with the increase of the entrance total pressure under the same fuel equivalence ratio; the larger the total pressure, the smaller the fuel equivalence ratio required to reach the same dimensionless peak pressure. The results show that there were pure scramjet mode, dual-mode scramjet mode, dual-mode ramjet mode and the combustion state that the pressure disturbance had spread to the entrance of the isolator with the increase of the fuel equivalence ratio. The dimensionless peak pressures of the combustor when mode transition occurred were 0.25, 0.41, and 0.5, and did not change with the change of the total pressure of the incoming flow.
- Abstract: Publication date: Available online 26 November 2018Source: Acta AstronauticaAuthor(s): Jianping Li, Guiqian Jiao, Jinyuan Luo, Wenyan Song The mode transition experiments of the dual-mode combustor were carried out under the conditions that the total pressure was 600 kPa, 700 kPa, 800 kPa and 900 kPa, the total temperature was 810 K, and the Mach number was 2.0. According to the wall pressure of the combustor measured in the experiment, the distribution of other airflow parameters of the combustor were obtained through the one-dimensional performance calculation method. The processes of mode transition of the combustor under different total pressures were studied. The study show that the dimensionless peak pressure of the combustor increased with the increase of the entrance total pressure under the same fuel equivalence ratio; the larger the total pressure, the smaller the fuel equivalence ratio required to reach the same dimensionless peak pressure. The results show that there were pure scramjet mode, dual-mode scramjet mode, dual-mode ramjet mode and the combustion state that the pressure disturbance had spread to the entrance of the isolator with the increase of the fuel equivalence ratio. The dimensionless peak pressures of the combustor when mode transition occurred were 0.25, 0.41, and 0.5, and did not change with the change of the total pressure of the incoming flow.
- Numerical investigation on drag and heat reduction mechanism of combined
spike and rear opposing jet configuration- Abstract: Publication date: Available online 26 November 2018Source: Acta AstronauticaAuthor(s): Jie Huang, Wei-Xing Yao, Xian-Yang Shan In order to reduce the hypersonic aerodynamic drag and heating, a combined spike and rear opposing jet configuration is proposed in this paper, and the CFD method is adopted to analyze the drag and heat reduction efficiency. The results show that the spike pushes the bow shock wave away from the blunt body, which translates the normal shock wave into the oblique shock wave and reduces the shock wave intensity. In addition, the low temperature jet gas is injected into the flow field, which reduces the temperature of the flow field after the shock wave. So the combined configuration reduces the aerodynamic drag and heating of the blunt body by the reconstruction of flow field, and the drag and heat reduction efficiency is better than the other configurations that already exist. The influences of the length of spike, total pressure of the opposing jet and jet gas on the drag and heat reduction efficiency are studied. The results show that increasing the length of the spike and the total pressure of the opposing jet can effectively improve the drag and heat reduction efficiency, and the decreasing rates of the aerodynamic drag and heating slow down gradually with the increase of above two parameters. In addition, the nitrogen has the best drag reduction efficiency and the carbon dioxide has the best heat reduction efficiency. The investigations in this paper verify the advantages and application in engineering of the combined configuration proposed in this paper.
- Abstract: Publication date: Available online 26 November 2018Source: Acta AstronauticaAuthor(s): Jie Huang, Wei-Xing Yao, Xian-Yang Shan In order to reduce the hypersonic aerodynamic drag and heating, a combined spike and rear opposing jet configuration is proposed in this paper, and the CFD method is adopted to analyze the drag and heat reduction efficiency. The results show that the spike pushes the bow shock wave away from the blunt body, which translates the normal shock wave into the oblique shock wave and reduces the shock wave intensity. In addition, the low temperature jet gas is injected into the flow field, which reduces the temperature of the flow field after the shock wave. So the combined configuration reduces the aerodynamic drag and heating of the blunt body by the reconstruction of flow field, and the drag and heat reduction efficiency is better than the other configurations that already exist. The influences of the length of spike, total pressure of the opposing jet and jet gas on the drag and heat reduction efficiency are studied. The results show that increasing the length of the spike and the total pressure of the opposing jet can effectively improve the drag and heat reduction efficiency, and the decreasing rates of the aerodynamic drag and heating slow down gradually with the increase of above two parameters. In addition, the nitrogen has the best drag reduction efficiency and the carbon dioxide has the best heat reduction efficiency. The investigations in this paper verify the advantages and application in engineering of the combined configuration proposed in this paper.
- Space collision probability computation based on on-board optical cues
- Abstract: Publication date: Available online 23 November 2018Source: Acta AstronauticaAuthor(s): Meng Yu, Shuang Li, Shu Leng This paper presents a novel on-board space collision analysis method for space situational awareness. The framework is developed under the following assumptions: 1) A satellite can be equipped with on-board sensors for space object recognition. 2) No a–priori knowledge of the space objects is provided. A space object size and relative state estimation method is firstly proposed, wherein optical cues acquired from onboard sensors are utilized to achieve the estimation. Then, the unscented transform approach is employed to calculate the probability density function (PDF) of collision probability based on the estimate information. Monte Carlo simulations and an experimental test demonstrate that the proposed approach can achieve high-precision on-board collision probability estimation with an error less than 3%.
- Abstract: Publication date: Available online 23 November 2018Source: Acta AstronauticaAuthor(s): Meng Yu, Shuang Li, Shu Leng This paper presents a novel on-board space collision analysis method for space situational awareness. The framework is developed under the following assumptions: 1) A satellite can be equipped with on-board sensors for space object recognition. 2) No a–priori knowledge of the space objects is provided. A space object size and relative state estimation method is firstly proposed, wherein optical cues acquired from onboard sensors are utilized to achieve the estimation. Then, the unscented transform approach is employed to calculate the probability density function (PDF) of collision probability based on the estimate information. Monte Carlo simulations and an experimental test demonstrate that the proposed approach can achieve high-precision on-board collision probability estimation with an error less than 3%.
- Detecting migrant vessels in the Mediterranean Sea: Using Sentinel-2
images to aid humanitarian actions- Abstract: Publication date: Available online 23 November 2018Source: Acta AstronauticaAuthor(s): Urška Kanjir The process of migration by sea is often accompanied by a great deal of risk for the migrants. The need for reliable and on time information on migrant movements is essential, especially as the available information is often limited or inconsistent. The aim of this paper is to show how freely available Sentinel-2 optical images over large areas can support humanitarian actions with timely and accurate geospatial information by providing the exact location of vessels at sea at the time of satellite acquisition. With the proposed detection method, we would like to distance ourselves from border surveillance and fight against clandestine migration as the first associations when dealing with movements of people using satellite technology. Instead, we would like to provide a better understanding of the situation for relieve authorities. Using Sentinel-2 data we have developed an automatic vessel detection and classification procedure. By first removing land from the images with the use of Modified Normalized Difference Water Index (MNDWI) we obtained a sea mask. We then applied a “vessel index”, which eliminates most of the atmospherics influences that can significantly weaken the detection results from optical images and binomial logistic regression on the sea mask to obtain segments of possible vessels. We calculated a group of geometrical and spectral attributes of detected segments and removed all the non-vessel segments based on the value of segments area. Later we implemented classification using decision tree classifier. Finally, we performed an accuracy assessment of vessel classification. The results demonstrate that the methodology gives a reliable outcome in a timely and consistent manner but can overlook smaller vessels (the length less than 20 m). Freely available satellite technology can, therefore, offer an efficient and effective solution for frequent monitoring and tracking of vessels in real time across large areas. This approach would give effective results when complementary with other methods for search and rescue of migrants to help reduce the intolerable death toll of refugees while crossing the sea.
- Abstract: Publication date: Available online 23 November 2018Source: Acta AstronauticaAuthor(s): Urška Kanjir The process of migration by sea is often accompanied by a great deal of risk for the migrants. The need for reliable and on time information on migrant movements is essential, especially as the available information is often limited or inconsistent. The aim of this paper is to show how freely available Sentinel-2 optical images over large areas can support humanitarian actions with timely and accurate geospatial information by providing the exact location of vessels at sea at the time of satellite acquisition. With the proposed detection method, we would like to distance ourselves from border surveillance and fight against clandestine migration as the first associations when dealing with movements of people using satellite technology. Instead, we would like to provide a better understanding of the situation for relieve authorities. Using Sentinel-2 data we have developed an automatic vessel detection and classification procedure. By first removing land from the images with the use of Modified Normalized Difference Water Index (MNDWI) we obtained a sea mask. We then applied a “vessel index”, which eliminates most of the atmospherics influences that can significantly weaken the detection results from optical images and binomial logistic regression on the sea mask to obtain segments of possible vessels. We calculated a group of geometrical and spectral attributes of detected segments and removed all the non-vessel segments based on the value of segments area. Later we implemented classification using decision tree classifier. Finally, we performed an accuracy assessment of vessel classification. The results demonstrate that the methodology gives a reliable outcome in a timely and consistent manner but can overlook smaller vessels (the length less than 20 m). Freely available satellite technology can, therefore, offer an efficient and effective solution for frequent monitoring and tracking of vessels in real time across large areas. This approach would give effective results when complementary with other methods for search and rescue of migrants to help reduce the intolerable death toll of refugees while crossing the sea.
- Research on the hydrolysis of human urine using biological activated
carbon and its application in bioregenerative life support system- Abstract: Publication date: Available online 22 November 2018Source: Acta AstronauticaAuthor(s): Guorong Zhu, Guanghui Liu, Dianlei Liu, Haoxiang Chen, Chenhao Fang, Yue Yi, Ming Li, Beizhen Xie, Hong Liu The water and nitrogen recovery from human urine are the crucial issues for the water recycling in the bioregenerative life support system (BLSS). Most of the water in the urine could be recovered through physical/chemical methods, however the efficiency of recovering nitrogen remained unsatisfactory. In our previous work, immobilized urease catalysis followed by reduced pressure distillation has been utilized to purify the urine, and high recovery efficiency of both water and nitrogen has been gained. However, the source of the urease and the immobilization preparation method would limit its practical application. In this study, biological activated carbon (BAC) immobilizing urease-producing microorganisms was explored as the continuous provider of urease to hydrolyze urea existing in human urine. The batch experiment results illustrated that it is feasible to use BAC technology to hydrolyze urea in urine and the highest urea hydrolysis efficiency of 79.33% was gained in synthetic urine with the powder activated carbon dosage (PACD) of 100 g/L and the hydraulic retention time (HRT) of 5 d. Image results of scanning electron microscope and confocal laser scanning microscope presented that BAC succcessfully immobilized living bacteria, and 16S rRNA high throughput sequencing illustrated that the main urease-producing bacteria were genera of Bacillus, Sporosarcina, Pseudomonas and Paracoccus, and genera of Pseudomonas and Paracoccus possessed heterotrophic nitrification ability as well. A pilot-scale membrane biological reactor inoculated with urea-hydrolyzing BAC cultivated in batch experiment was applied to treat the crew's urine continuously inside an experimental facility for BLSS, and the urea hydrolysis efficiency could maintain at 99.84% during 203 d operation.Graphical abstract
- Abstract: Publication date: Available online 22 November 2018Source: Acta AstronauticaAuthor(s): Guorong Zhu, Guanghui Liu, Dianlei Liu, Haoxiang Chen, Chenhao Fang, Yue Yi, Ming Li, Beizhen Xie, Hong Liu The water and nitrogen recovery from human urine are the crucial issues for the water recycling in the bioregenerative life support system (BLSS). Most of the water in the urine could be recovered through physical/chemical methods, however the efficiency of recovering nitrogen remained unsatisfactory. In our previous work, immobilized urease catalysis followed by reduced pressure distillation has been utilized to purify the urine, and high recovery efficiency of both water and nitrogen has been gained. However, the source of the urease and the immobilization preparation method would limit its practical application. In this study, biological activated carbon (BAC) immobilizing urease-producing microorganisms was explored as the continuous provider of urease to hydrolyze urea existing in human urine. The batch experiment results illustrated that it is feasible to use BAC technology to hydrolyze urea in urine and the highest urea hydrolysis efficiency of 79.33% was gained in synthetic urine with the powder activated carbon dosage (PACD) of 100 g/L and the hydraulic retention time (HRT) of 5 d. Image results of scanning electron microscope and confocal laser scanning microscope presented that BAC succcessfully immobilized living bacteria, and 16S rRNA high throughput sequencing illustrated that the main urease-producing bacteria were genera of Bacillus, Sporosarcina, Pseudomonas and Paracoccus, and genera of Pseudomonas and Paracoccus possessed heterotrophic nitrification ability as well. A pilot-scale membrane biological reactor inoculated with urea-hydrolyzing BAC cultivated in batch experiment was applied to treat the crew's urine continuously inside an experimental facility for BLSS, and the urea hydrolysis efficiency could maintain at 99.84% during 203 d operation.Graphical abstract
- Study on electrons conduction paths in Hall thruster ignition processes
with the cathode located inside and outside the magnetic separatrix- Abstract: Publication date: Available online 22 November 2018Source: Acta AstronauticaAuthor(s): Wen-Bo Li, Hong Li, Yong-Jie Ding, Li-Qiu Wei, Qian Gao, Shi-Lin Yan, Tian-Hang Meng, Xi-ming Zhu, Da-Ren Yu A high speed charged coupled device (CCD) camera was used to examine the plume features over time when the cathode was located inside and outside the magnetic separatrix. Due to the different positions of the cathode, there are obvious differences in the electron impact excitation process, the characteristics of the plasma bridge, and the transition process from the end of ignition to the steady-state discharge process. The main reason for these differences may be due to the different conduction paths followed by emitted electrons into the acceleration channel. The CCD images method can be used as a technique to characterize the ignition process of Hall thruster.
- Abstract: Publication date: Available online 22 November 2018Source: Acta AstronauticaAuthor(s): Wen-Bo Li, Hong Li, Yong-Jie Ding, Li-Qiu Wei, Qian Gao, Shi-Lin Yan, Tian-Hang Meng, Xi-ming Zhu, Da-Ren Yu A high speed charged coupled device (CCD) camera was used to examine the plume features over time when the cathode was located inside and outside the magnetic separatrix. Due to the different positions of the cathode, there are obvious differences in the electron impact excitation process, the characteristics of the plasma bridge, and the transition process from the end of ignition to the steady-state discharge process. The main reason for these differences may be due to the different conduction paths followed by emitted electrons into the acceleration channel. The CCD images method can be used as a technique to characterize the ignition process of Hall thruster.
- CubeSat particle aggregation Collision Experiment (Q-PACE): Design of a 3U
CubeSat mission to investigate planetesimal formation- Abstract: Publication date: Available online 22 November 2018Source: Acta AstronauticaAuthor(s): Stephanie Jarmak, Julie Brisset, Joshua Colwell, Adrienne Dove, Douglas Maukonen, Samir A. Rawashdeh, Jürgen Blum, Larry Roe Observations of the collisional evolution of particle ensembles in a microgravity environment are necessary to characterize the processes that lead to the formation of planetesimals, km-size and larger bodies, within the protoplanetary disk. The two current theories of planetesimal formation, namely growth through binary sticking collisions and gravitational instability within the protoplanetary disk, have difficulties in explaining how particles grow beyond a centimeter in size. In this paper we describe the CubeSat Particle Aggregation and Collision Experiment (Q-PACE), a Low Earth Orbit 3U CubeSat mission that will provide a high-quality, long duration microgravity environment in which we will observe collisions between particles under conditions relevant to planetesimal formation. We have designed a series of experiments involving a broad range of particle size, density, surface properties, and collision velocities to observe collisional outcomes from bouncing to sticking as well as aggregate disruption in tens of thousands of collisions.
- Abstract: Publication date: Available online 22 November 2018Source: Acta AstronauticaAuthor(s): Stephanie Jarmak, Julie Brisset, Joshua Colwell, Adrienne Dove, Douglas Maukonen, Samir A. Rawashdeh, Jürgen Blum, Larry Roe Observations of the collisional evolution of particle ensembles in a microgravity environment are necessary to characterize the processes that lead to the formation of planetesimals, km-size and larger bodies, within the protoplanetary disk. The two current theories of planetesimal formation, namely growth through binary sticking collisions and gravitational instability within the protoplanetary disk, have difficulties in explaining how particles grow beyond a centimeter in size. In this paper we describe the CubeSat Particle Aggregation and Collision Experiment (Q-PACE), a Low Earth Orbit 3U CubeSat mission that will provide a high-quality, long duration microgravity environment in which we will observe collisions between particles under conditions relevant to planetesimal formation. We have designed a series of experiments involving a broad range of particle size, density, surface properties, and collision velocities to observe collisional outcomes from bouncing to sticking as well as aggregate disruption in tens of thousands of collisions.
- Cost analysis of solar thermal propulsion systems for microsatellite
applications- Abstract: Publication date: Available online 20 November 2018Source: Acta AstronauticaAuthor(s): Fiona Leverone, Angelo Cervone, Eberhard Gill In recent years, satellite design has extended towards miniaturisation to reduce associated cost with launching and conducting space missions. Small satellites provide low-cost platforms for space missions. However, this lower cost comes at the expense of the removal of key sub-systems, such as the propulsion system, due to the small available onboard volume and mass restrictions. For this reason, small, lightweight, high-performing and affordable propulsion systems are necessary. However, there is limited research available on the comparison of propulsion technologies with regards to cost. Motivated by the above challenges the objective of this paper is to provide a comparison of propulsion technologies that are compatible with small satellites with respect to cost and application. The different propulsion systems are investigated for three mission scenarios, a small on-orbit manoeuvre, a station-keeping, and a lunar orbit transfer mission. Each system is evaluated in terms of a total figure of merit which incorporates nine variables such as propellant mass, safety, and hardware price, that affect the total cost of a propulsion system. This figure of merit is used to quantitatively compare the propulsion systems to identify cost-effective solutions as a function of various mission scenarios. Solar thermal propulsion has been proposed for small satellite applications, but information regarding the concepts are not available in a single report. Therefore, another objective of this paper is to provide the reader with a review of the current status of solar thermal propulsion. An important finding of this research is the classification of propulsion systems in terms of thrust, specific impulse, cost, and application.
- Abstract: Publication date: Available online 20 November 2018Source: Acta AstronauticaAuthor(s): Fiona Leverone, Angelo Cervone, Eberhard Gill In recent years, satellite design has extended towards miniaturisation to reduce associated cost with launching and conducting space missions. Small satellites provide low-cost platforms for space missions. However, this lower cost comes at the expense of the removal of key sub-systems, such as the propulsion system, due to the small available onboard volume and mass restrictions. For this reason, small, lightweight, high-performing and affordable propulsion systems are necessary. However, there is limited research available on the comparison of propulsion technologies with regards to cost. Motivated by the above challenges the objective of this paper is to provide a comparison of propulsion technologies that are compatible with small satellites with respect to cost and application. The different propulsion systems are investigated for three mission scenarios, a small on-orbit manoeuvre, a station-keeping, and a lunar orbit transfer mission. Each system is evaluated in terms of a total figure of merit which incorporates nine variables such as propellant mass, safety, and hardware price, that affect the total cost of a propulsion system. This figure of merit is used to quantitatively compare the propulsion systems to identify cost-effective solutions as a function of various mission scenarios. Solar thermal propulsion has been proposed for small satellite applications, but information regarding the concepts are not available in a single report. Therefore, another objective of this paper is to provide the reader with a review of the current status of solar thermal propulsion. An important finding of this research is the classification of propulsion systems in terms of thrust, specific impulse, cost, and application.
- Configuration optimization of multi-optical sensors with complex pointing
constraints- Abstract: Publication date: Available online 19 November 2018Source: Acta AstronauticaAuthor(s): Yuchen She, Shuang Li This paper investigates the pointing direction optimization problem of multi-optical sensors with complex pointing constraints. In the spacecraft design area, this problem is characterized as the optical-sensor configuration optimization. The complex pointing requirements and avoidance constraints of different types of optical sensors are taken into account, and a new optimization method is presented. The proposed combinatorial optimization algorithm is divided into three steps. First, an initial guess solution is generated by coordinate projection. The relative directions between all constraint objects and the spacecraft are projected into the spacecraft body-fixed coordinate system. Second, the density-of-presence analysis is conducted for relative directions by introducing the image processing concept into the optimization process. The morphological dilation approach is adopted with a new filter, and the high and low density regions are defined to reduce the dimension of the candidate solution group. Finally, the mathematical model of the optimization problem is developed and solved by the Genetic Algorithm (GA). The computer simulations show that the new method can properly handle the configuration optimization problem with complex requirements and constraints in a relatively short calculation time.
- Abstract: Publication date: Available online 19 November 2018Source: Acta AstronauticaAuthor(s): Yuchen She, Shuang Li This paper investigates the pointing direction optimization problem of multi-optical sensors with complex pointing constraints. In the spacecraft design area, this problem is characterized as the optical-sensor configuration optimization. The complex pointing requirements and avoidance constraints of different types of optical sensors are taken into account, and a new optimization method is presented. The proposed combinatorial optimization algorithm is divided into three steps. First, an initial guess solution is generated by coordinate projection. The relative directions between all constraint objects and the spacecraft are projected into the spacecraft body-fixed coordinate system. Second, the density-of-presence analysis is conducted for relative directions by introducing the image processing concept into the optimization process. The morphological dilation approach is adopted with a new filter, and the high and low density regions are defined to reduce the dimension of the candidate solution group. Finally, the mathematical model of the optimization problem is developed and solved by the Genetic Algorithm (GA). The computer simulations show that the new method can properly handle the configuration optimization problem with complex requirements and constraints in a relatively short calculation time.
- Experimental study on the thermodynamic characteristics of the high
temperature hydrocarbon fuel in the cooling channel of the hypersonic
vehicle- Abstract: Publication date: Available online 17 November 2018Source: Acta AstronauticaAuthor(s): Haowei Li, Jiang Qin, Yuguang Jiang, Weixing Zhou, Wen Bao, Hongyan Huang Cooling is one of the key technologies for the scramjet considering the enormous thermal load and the demand of reusable long-range mission. In the regenerative cooling process, hydrocarbon fuel flows through the cooling channel, gets heated and experiences severe thermal properties changes. The compressibility of fuel becomes non-negligible and possible dynamic process emerges, which affects both the characteristics of fuel cooling and fuel mass flow control. As a result, the dynamic characteristics of the high temperature hydrocarbon fuel within the cooling channel must be carefully studied. In this work, experiments are conducted at different simulated engine working conditions. The settling time of pipe pressure drop is used to characterize the dynamic process of the outlet fuel mass flow, because the outlet temperature of the fuel is too high to measure the fuel mass flow directly. In conditions of the variation of inlet fuel mass flow and backpressure disturbance, overshoot of the pressure drop is observed and the magnitude of which varies with the working conditions. The settling time of outlet fuel temperature and wall temperature channel firstly increases, then decreases with the increase of the outlet fuel temperature. The settling time of both outlet fuel temperature and wall temperature increases with the increase of the heating heat flux. The experimental results in this work are expected to provide supports to the engine control system design.
- Abstract: Publication date: Available online 17 November 2018Source: Acta AstronauticaAuthor(s): Haowei Li, Jiang Qin, Yuguang Jiang, Weixing Zhou, Wen Bao, Hongyan Huang Cooling is one of the key technologies for the scramjet considering the enormous thermal load and the demand of reusable long-range mission. In the regenerative cooling process, hydrocarbon fuel flows through the cooling channel, gets heated and experiences severe thermal properties changes. The compressibility of fuel becomes non-negligible and possible dynamic process emerges, which affects both the characteristics of fuel cooling and fuel mass flow control. As a result, the dynamic characteristics of the high temperature hydrocarbon fuel within the cooling channel must be carefully studied. In this work, experiments are conducted at different simulated engine working conditions. The settling time of pipe pressure drop is used to characterize the dynamic process of the outlet fuel mass flow, because the outlet temperature of the fuel is too high to measure the fuel mass flow directly. In conditions of the variation of inlet fuel mass flow and backpressure disturbance, overshoot of the pressure drop is observed and the magnitude of which varies with the working conditions. The settling time of outlet fuel temperature and wall temperature channel firstly increases, then decreases with the increase of the outlet fuel temperature. The settling time of both outlet fuel temperature and wall temperature increases with the increase of the heating heat flux. The experimental results in this work are expected to provide supports to the engine control system design.
-
SiO 2
exposed to vacuum ultraviolet radiation- Abstract: Publication date: Available online 16 November 2018Source: Acta AstronauticaAuthor(s): V.I. Pavlenko, N.I. Cherkashina Vacuum ultraviolet (VUV) radiation produced by the Sun in the space environment can cause degradation to thermoregulating coatings producing changes in optical, mechanical, and chemical properties. These effects are particularly important for polymers. The purpose of this study was to study the effect of VUV radiation on polymeric composites based on polyalkaneimide. Amorphous and crystalline SiO2 were used as the filler. A comparative characterization of physico-mechanical properties of composites is presented depending on the polymorphic structure of the filler being introduced. According to the microhardness of the surface of composites with amorphous and crystalline SiO2, an optimum filler content of 60–65 wt% is established. The exposure of VUV radiation to polymeric composites with SiO2 leads to a mass loss and a change in their near-surface layers, which leads to a deterioration in the optical properties. At an elevated temperature (125oC), the mass loss of all composites with amorphous crystalline SiO2 does not exceed 1.33%, and with crystalline SiO2, the maximum mass loss is 0.52%. The mass loss of composites with crystalline SiO2 of all compositions did not exceed the allowable value for spacecraft products of 1%. For high-filled (50 wt% filler or more) composites with amorphous SiO2, mass loss after VUV treatment is more than 1%, which does not allow them to be used in outer space. The change in the main optical characteristics of thermoregulating coatings after VUV treatment was studied. The smoothing of the surface of composites after VUV treatment is established.
- Abstract: Publication date: Available online 16 November 2018Source: Acta AstronauticaAuthor(s): V.I. Pavlenko, N.I. Cherkashina Vacuum ultraviolet (VUV) radiation produced by the Sun in the space environment can cause degradation to thermoregulating coatings producing changes in optical, mechanical, and chemical properties. These effects are particularly important for polymers. The purpose of this study was to study the effect of VUV radiation on polymeric composites based on polyalkaneimide. Amorphous and crystalline SiO2 were used as the filler. A comparative characterization of physico-mechanical properties of composites is presented depending on the polymorphic structure of the filler being introduced. According to the microhardness of the surface of composites with amorphous and crystalline SiO2, an optimum filler content of 60–65 wt% is established. The exposure of VUV radiation to polymeric composites with SiO2 leads to a mass loss and a change in their near-surface layers, which leads to a deterioration in the optical properties. At an elevated temperature (125oC), the mass loss of all composites with amorphous crystalline SiO2 does not exceed 1.33%, and with crystalline SiO2, the maximum mass loss is 0.52%. The mass loss of composites with crystalline SiO2 of all compositions did not exceed the allowable value for spacecraft products of 1%. For high-filled (50 wt% filler or more) composites with amorphous SiO2, mass loss after VUV treatment is more than 1%, which does not allow them to be used in outer space. The change in the main optical characteristics of thermoregulating coatings after VUV treatment was studied. The smoothing of the surface of composites after VUV treatment is established.
- Numerical and experimental study of the thermochemical erosion of a
graphite nozzle in a hybrid rocket motor with a star grain- Abstract: Publication date: Available online 13 November 2018Source: Acta AstronauticaAuthor(s): Tian Hui, Yu Ruipeng, Li Chengen, Zhao Sheng, Zhu Hao Hybrid rocket motors is a promising propulsion system because of its intrinsic advantages over a conventional solid rocket motor and liquid rocket engine. However, serious nozzle erosion is a key problem that prevents hybrid rocket motors from being widely used, especially for propulsion systems with long operating times. In this paper, the erosion of a graphite-based nozzle coupled with a combustion flow field is studied in a hybrid rocket motor with a star grain. As the oxidizer and fuel, 90% hydrogen peroxide and hydroxide-terminated polybutadiene are adopted, respectively. The nozzle erosion was simulated coupled with the flow field in a typical hybrid rocket motor through three-dimensional numerical simulations. The simulations are based on a pure-gas steady numerical model considering turbulence, fuel pyrolysis, oxidizer/fuel reactions, thermal conduction and solid-gas boundary interactions on the fuel and nozzle surfaces. The results indicate that the nozzle erosion is greatly influenced by the inner flow field. The flame near the grain trough is thicker than that near the grain peak. Therefore, the maximum erosion rate (0.042 mm/s) occurs near the nozzle throat corresponding to the grain trough. The OH and H2O contribute 49.8% and 45.5% to the erosion rate, respectively, in this area. Furthermore, 56.6% and 31.9% contributions are made by OH and H2O, respectively, in the area corresponding to the grain peak. The O, CO2 and O2 make much lower contributions to the total erosion. In addition, a firing test is carried out to characterize the graphite nozzle erosion on a full-scale hybrid rocket motor with star grain. The nozzle inner profiles before and after test show that the erosion behavior of the graphite material is strictly related to the fuel shape.
- Abstract: Publication date: Available online 13 November 2018Source: Acta AstronauticaAuthor(s): Tian Hui, Yu Ruipeng, Li Chengen, Zhao Sheng, Zhu Hao Hybrid rocket motors is a promising propulsion system because of its intrinsic advantages over a conventional solid rocket motor and liquid rocket engine. However, serious nozzle erosion is a key problem that prevents hybrid rocket motors from being widely used, especially for propulsion systems with long operating times. In this paper, the erosion of a graphite-based nozzle coupled with a combustion flow field is studied in a hybrid rocket motor with a star grain. As the oxidizer and fuel, 90% hydrogen peroxide and hydroxide-terminated polybutadiene are adopted, respectively. The nozzle erosion was simulated coupled with the flow field in a typical hybrid rocket motor through three-dimensional numerical simulations. The simulations are based on a pure-gas steady numerical model considering turbulence, fuel pyrolysis, oxidizer/fuel reactions, thermal conduction and solid-gas boundary interactions on the fuel and nozzle surfaces. The results indicate that the nozzle erosion is greatly influenced by the inner flow field. The flame near the grain trough is thicker than that near the grain peak. Therefore, the maximum erosion rate (0.042 mm/s) occurs near the nozzle throat corresponding to the grain trough. The OH and H2O contribute 49.8% and 45.5% to the erosion rate, respectively, in this area. Furthermore, 56.6% and 31.9% contributions are made by OH and H2O, respectively, in the area corresponding to the grain peak. The O, CO2 and O2 make much lower contributions to the total erosion. In addition, a firing test is carried out to characterize the graphite nozzle erosion on a full-scale hybrid rocket motor with star grain. The nozzle inner profiles before and after test show that the erosion behavior of the graphite material is strictly related to the fuel shape.
- Effects of fueling distance on combustion stabilization modes in a
cavity-based scramjet combustor- Abstract: Publication date: Available online 10 November 2018Source: Acta AstronauticaAuthor(s): Yanan Wang, Zhenguo Wang, Mingbo Sun, Hongbo Wang, Zun Cai The effects of distance from the injector upstream to the cavity leading edge on combustion stabilization modes were investigated experimentally and numerically. High-speed flame luminosity and schlieren images were utilized to reveal the combustion characteristics. Under a Mach 2.52 supersonic inflow condition with a stagnation temperature of 1629 K, hydrogen was injected with injection distances of 30 mm, 100 mm and 160 mm at global equivalence ratios from 0.20 to 0.32. The flame in cavity stabilized mode is observed at a relatively low equivalence ratio. Increasing the global equivalence ratio, the combustion stabilization mode transfers to the jet-wake stabilized mode with injection distances of 100 mm and 160 mm. It is surprising that the mode transition is the easiest for the medium injection distance of 100 mm. Large eddy simulations were then performed to explain the easiest mode-transition. Mixing characteristics, jet-cavity interactions and the jet-wake features upstream of the cavity for three injection distances were analyzed. With the injection distance of 100 mm, the reflected shock wave impacts on the cavity shear layer, and the interactions of the shock wave structures and the cavity induce the beginning phase of the jet-wake stabilized flame and help the flame propagate upstream around the cavity. Relatively larger amount of the fuel, lower streamwise velocity and lower static pressure was also found in the jet wake upstream of the cavity, which further supports the flame to propagate upstream and stabilize in the jet wake.
- Abstract: Publication date: Available online 10 November 2018Source: Acta AstronauticaAuthor(s): Yanan Wang, Zhenguo Wang, Mingbo Sun, Hongbo Wang, Zun Cai The effects of distance from the injector upstream to the cavity leading edge on combustion stabilization modes were investigated experimentally and numerically. High-speed flame luminosity and schlieren images were utilized to reveal the combustion characteristics. Under a Mach 2.52 supersonic inflow condition with a stagnation temperature of 1629 K, hydrogen was injected with injection distances of 30 mm, 100 mm and 160 mm at global equivalence ratios from 0.20 to 0.32. The flame in cavity stabilized mode is observed at a relatively low equivalence ratio. Increasing the global equivalence ratio, the combustion stabilization mode transfers to the jet-wake stabilized mode with injection distances of 100 mm and 160 mm. It is surprising that the mode transition is the easiest for the medium injection distance of 100 mm. Large eddy simulations were then performed to explain the easiest mode-transition. Mixing characteristics, jet-cavity interactions and the jet-wake features upstream of the cavity for three injection distances were analyzed. With the injection distance of 100 mm, the reflected shock wave impacts on the cavity shear layer, and the interactions of the shock wave structures and the cavity induce the beginning phase of the jet-wake stabilized flame and help the flame propagate upstream around the cavity. Relatively larger amount of the fuel, lower streamwise velocity and lower static pressure was also found in the jet wake upstream of the cavity, which further supports the flame to propagate upstream and stabilize in the jet wake.
