63 research outputs found
Staavia capitella Sond.
No full textStaavia capitella (Thunb.) Sond. in Harvey & Sonder, Fl. Cap. 2: 323. 1862 ≡ Brunia capitella Thunb., Brunia: 5. 1804 – Type: e Cap. b. spei, sinecoll. [= Thunberg] (UPS-sheet 5735, holo – microfiche!). = Staavia comosa sensu Pillans in J. S. African Bot. 13: 174 (1947) non (Thunb.) Colozza (= Berzelia lanuginosa).Published as part of Claßen-Bockhoff, Regine, Oliver, Edward G. H., Hall, Anthony V. & Quint, Marcus, 2019, A new classification of the South African endemic family Bruniaceae based on molecular and morphological data, pp. 1138-1155 in TAXON 60 (4) on page 1147, DOI: 10.1002/tax.604016, http://zenodo.org/record/759043
Berzelia lanuginosa Brongn
No full textBerzelia lanuginosa (L.) Brongn. in Ann. Sci. Nat. (Paris) 8: 372. 1826; PillansinJ. S. African Bot. 13: 197. 1947 ≡ Brunia lanuginosa L., Sp. pl.: 199. 1753 – Lectotype: sine loc., sinecoll. (Herb. Clifford: 71), Brunia 2 (BM-000558118, designated by Powrie in J. S. African Bot. 38: 302. 1972). = Bruniacomosa Thunb. in Phytogr. Blätt. 1: 17. 1803 ≡ Staavia comosa (Thunb.) Colozza in Ann. Bot. (Rome) 2: 22, 35, t. 1. 1905; PillansinJ. S. African Bot. 13: 174. 1947 – Type: e Cap. b. spei, sinecoll. [= Thunberg] (UPS-sheet 5736 holo – microfiche!).Published as part of Claßen-Bockhoff, Regine, Oliver, Edward G. H., Hall, Anthony V. & Quint, Marcus, 2019, A new classification of the South African endemic family Bruniaceae based on molecular and morphological data, pp. 1138-1155 in TAXON 60 (4) on page 1148, DOI: 10.1002/tax.604016, http://zenodo.org/record/759043
Preliminary design of a long-endurance Mars aircraft
No full textThe preliminary design requirements of a long endurance aircraft capable of flight within the Martian environment was determined. Both radioisotope/heat engine and PV solar array power production systems were considered. Various cases for each power system were analyzed in order to determine the necessary size, weight and power requirements of the aircraft. The analysis method used was an adaptation of the method developed by Youngblood and Talay of NASA-Langley used to design a high altitude earth based aircraft. The analysis is set up to design an aircraft which, for the given conditions, has a minimum wingspan and maximum endurance parameter. The results showed that, for a first approximation, a long endurance aircraft is feasible within the Martian environment. The size and weight of the most efficient solar aircraft were comparable to the radioisotope powered one
Analysis of lunar regolith thermal energy storage
No full textThe concept of using lunar regolith as a thermal energy storage medium was evaluated. The concept was examined by mathematically modeling the absorption and transfer of heat by the lunar regolith. Regolith thermal and physical properties were established through various sources as functions of temperature. Two cases were considered: a semi-infinite, constant temperature, cylindrical heat source embedded in a continuum of lunar regolith and a spherically shaped molten zone of lunar regolith set with an initial temperature profile. The cylindrical analysis was performed in order to examine the amount of energy which can be stored in the regolith during the day. At night, the cylinder acted as a perfect insulator. This cycling was performed until a steady state situation was reached in the surrounding regolith. It was determined that a cycling steady state occurs after approximately 15 day/night cycles. Results were obtained for cylinders of various diameters. The spherical molten zone analysis was performed to establish the amount of thermal energy, within the regolith, necessary to maintain some molten material throughout a nighttime period. This surrounding temperature profile was modeled after the cycling steady state temperature profile established by the cylindrical analysis. It was determined that a molten sphere diameter of 4.76 m is needed to maintain a core temperature near the low end of the melting temperature range throughout one nighttime period
Effect of Date and Location on Maximum Achievable Altitude for a Solar Powered Aircraft
No full textThe maximum altitude attainable for a solar powered aircraft without any energy storage capability is examined. Mission profiles for a solar powered aircraft were generated over a range of latitudes and dates. These profiles were used to determine which latitude-date combinations produced the highest achieavable altitude. Based on the presented analysis the results have shown that for a given time of year lower latitudes produced higher maximum altitudes. For all the cases examined the time and date which produced the highest altitude was around March at the equator
Design and optimization of a self-deploying single axis tracking PV array
No full textThis study was performed in order to design a tracking photovoltaic (PV) array and optimize the design for maximum specific power. The design considerations were minimal deployment time, high reliability, and small stowage volume. The array design was self-deployable, from a compact stowage configuration, using a passive pressurized gas deployment mechanism. The array structural components consist of a combination of beams, columns, and cables used to deploy and orient a flexible PV blanket. Each structural component of the design was analyzed to determine the size necessary to withstand the various forces to which it would be subjected. An optimization was performed to determine the array dimensions and blanket geometry which produce the maximum specific power. The optimization was performed for both lunar and Martian environments with 4 types of PV blankets (silicon, GaAs/Ge, GaAs CLEFT, and amorphous silicon). For the lunar environment, the amorphous silicon array produced the highest specific power, whereas, for Mars the GaAs CLEFT array produced the highest specific power. A comparison was made to a fixed PV tent array of similar design. The tracking array produced a higher specific power with all types of the PV blankets examined except amorphous silicon at both locations
APEX 3D Propeller Test Preliminary Design
No full textA low Reynolds number, high subsonic mach number flight regime is fairly uncommon in aeronautics. Most flight vehicles do not fly under these aerodynamic conditions. However, recently there have been a number of proposed aircraft applications (such as high altitude observation platforms and Mars aircraft) that require flight within this regime. One of the main obstacles to flight under these conditions is the ability to reliably generate sufficient thrust for the aircraft. For a conventional propulsion system, the operation and design of the propeller is the key aspect to its operation. Due to the difficulty in experimentally modeling the flight conditions in ground-based facilities, it has been proposed to conduct propeller experiments from a high altitude gliding platform (APEX). A preliminary design of a propeller experiment under the low Reynolds number, high mach number flight conditions has been devised. The details of the design are described as well as the potential data that will be collected
Small Lunar Base Camp and in Situ Resource Utilization Oxygen Production Facility Power System Comparison
No full textThis report examines the power requirements for operating an in situ resource utilization (ISRU) oxygen production system on the lunar surface and a small six-person base camp. The baseline ISRU system produced 1.63 kg/h for a total day and night production rate of 1,154 kg. It was estimated that this plant would require 25.83 kW of power to operate. The base camp power includes auxiliary equipment as well as a communications system. The required power estimate for the base camp was 28.05 kW. This estimation was used to size a power system and determine its mass for meeting these requirements. Three types of power systems were considered: a solar photovoltaic (PV) array system using batteries for energy storage, a PV array system using a regenerative fuel cell (RFC) for energy storage, and a modular 10-kW electrical output power Kilopower reactor system. Three separate cases were examined: a stand-alone ISRU oxygen production system, a base camp, and a combined ISRU oxygen production system and base camp. For the PV array-based system, the RFC energy storage method had a mass advantage over a battery- based energy storage system. For higher power nighttime power operation for all three cases, the RFC systems specific energy was just over 830 Wh/kg. For the lower power nighttime keep-alive level used as part of the Case 1 analysis, the specific energy for the RFC was 456 Wh/kg. Both of these levels are significantly above the specific energy of 200 Wh/kg for the battery. Because of this higher specific energy, the RFC-based system provided significant mass advantages over the battery-based energy storage system. The baseline reactor system utilized shielding and separation distance to meet the desired maximum radiation dose level of 5 rem/yr for personnel operating within the vicinity of the power loads, base camp, and oxygen production facility. There are methods that could potentially be utilized to reduce the shielding requirements and separation distance. Implementing these would reduce the overall system mass for the reactor. Also, optimizing the reactor output to a specific mission would provide benefits in mass at the expense of modularity. The results of the power system comparison between a solar PV array-based system and a Kilopower reactor-based system has shown that for missions required to operate throughout the lunar night at power levels comparable to those used during the day, the reactor-based system provides a significant mass advantage. However, for applications that can meet their mission requirements while only having to operate during the daytime with minimal power required to survive the nighttime, the PV array-based system provides a mass advantage
Comparison of Mars Aircraft Propulsion Systems
No full textThe propulsion system is a critical aspect of the performance and feasibility of a Mars aircraft. Propulsion system mass and performance greatly influence the aircraft s design and mission capabilities. Various propulsion systems were analyzed to estimate the system mass necessary for producing 35N of thrust within the Mars environment. Three main categories of propulsion systems were considered: electric systems, combustion engine systems and rocket systems. Also, the system masses were compared for mission durations of 1, 2, and 4 h
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