Space Debris Affirmative
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- Sub-Orbital Payload Solvency
- Electro Dynamic Tether System Solvency
Space Vacuum SolvencyUsing a space vacuum solvesKiger 09 (co-author of two books, “Poplorica: A Popular History of the Fads, Mavericks, Inventions and Lore that Shaped Modern America," and “Oops: 20 Life Lessons From the Fiascoes That Shaped America.”, Science Good Ideas, A Space Debris Dustbuster?, March 27, 2009, http://blogs.discovery.com/good_idea/2009/03/a-space-debris-dustbuster.html, MS) What if NASA launched a spacecraft specially designed not for research or space exploration, but to pick up the increasing amount of trash accumulating in orbit and increasingly endangering satellites and astronauts? The spacecraft would be the metaphorical equivalent of a gigantic Dustbuster -- except, that, of course, an actual vacuum sweeper wouldn’t do much good in the vacuum of space, so the device instead would use lasers to redirect pieces of orbital junk into its path and then deploy a powerful electromagnet to suction them up. The space trash would be gathered into the vehicle’s compartment, and then transported back to Earth for recycling or disposal in landfills. Such a garbage-collecting spacecraft—or rather, a fleet of them—might be able to eliminate what is turning into a huge, potentially catastrophic problem for our spacefaring civilization. A space debris Dustbuster would also help establish a new ethic of off-world environmentalism for the exploration and commercial use of space. It would help make clear that we don’t regard orbital space, the Moon, and other planets merely as natural resources to be exploited—or trashed, depending upon human convenience or whim. Instead, we would take responsibility for cleaning up our own mess, and hopefully do a better job of it than we’ve done on Sub-Orbital Payload SolvencyThe aff can solve cheaply – studies show that a sub orbital payload can be used to reduce space debris.Hollopeter no date (James E., Director of Technology Development at GIT Satellite Communications, The X-Journals [a blog exploring revolutionary new technologies], Development of a ballistic orbital debris removal system, http://x-journals.com/2009/development-of-a-ballistic-orbital-debris-removal-system/, SP) GIT’s proposal is to attack the problem using a sub-orbital approach that cannot add to the orbital junk problem. Based on studies done under the Space Defense Initiative in the ‘80’s and on previous anti-satellite studies, GIT proposes a sub-orbital payload lofted to the appropriate altitude that could clear or reduce existing debris from selected areas of low earth orbit. By using a ballistic launch profile, there is no chance of adding to the existing debris problem. The payload would re-enter at the end of its mission, as well as all of its lower propulsive stages. There have been many suggestions to orbit a vehicle to collect debris and then de-orbit the debris using onboard propulsion systems. This is a very expensive approach. It would require all the associated ground control systems that are needed for any orbital missions today. By using a sub-orbital launch profile and existing sounding rockets in use today, a small ground based infrastructure, which presently exists could easily handle the launch load. There are many launch sites all over the world to support this type of mission. Since this debris problem exists for all space faring nations, the task could be shared among all users. Payload: Many payloads have been suggested to de-orbit the space debris. Most collect the debris and then de-orbit, while others such as tethers, would slowly lower the orbits until atmospheric drag takes over to de-orbit the debris. GIT’s approach is to use water, H2O, as the passive payload. It has the highest volumetric efficiency in the payload space. It can easily and predictably be deployed and has significant mass that will be used to reduce the debris orbital momentum. The payload would be launched retrograde to the target debris orbits. The resulting collisions would easily reduce the velocity of the smaller debris. The dispersion pattern of the water in space could be easily adjusted to accommodate the required velocity reduction for the target debris. Widely dispersed for very small objects of interest or narrowly dispersed for a focused collision of larger objects. Electro Dynamic Tether System SolvencyUsing electro-dynamic tethers solves – preliminary experiments show that with high thrust capacity 42 out of 42 large debris pieces can be collected per year.Barbee et. al 11 (Brent William, NASA Goddard Space Flight Center, with Elfego Pinon III, Emergent Space Technologies, Inc., Kenn Gold, Emergent Space Technologies, Inc., David Gaylor, Emergent Space Technologies, Inc., and Salvatore Alfano, Center for Space Standards and Innovation, Aerospace Conference 2011 IEEE, Design of spacecraft missions to remove multiple orbital debris objects, March 2011, http://ieeexplore.ieee.org/xpls/abs_all.jsp?arnumber=5747303&tag=1, SP) The set of 42 debris pieces defined previously was subjected to multi-target rendezvous trajectory analysis via the Series Method algorithm just described. This yielded an array of required total spacecraft launch masses as a function of the number of debris pieces visited and the specific impulse of the debris removal spacecrafts thruster. These required launch masses were then compared to the published capabilities of existing launch vehicles to assess the feasibility of removing various quantities of debris. The orbits of the 42 debris pieces were propagated using two body dynamics and making the deliberate approximation of treating the two-line elements as osculating elements just to give a sense of the debris piece orbits under consideration. A plot showing these orbits is presented in Figure 5. It is clear that these orbits mostly occupy different planes. This is because while they all have similar inclination angles (between 80.5° and 82°), they all have different right ascen sions of the ascending node. The differences between their orbit planes will clearly have a strong impact on the � V re quired to travel between them. It would be fortuitous if large groups of debris pieces tended to reside in nearly similar orbit planes, but natural orbit perturbations tend to preclude this, particularly in LEO. Figure 6 presents the relationship between the semimajor axes and eccentricities of the debris piece orbits. All of the debris pieces have similar semimajor axes and eccentricities, except for a few with slightly higher, but still small, eccentricities. Most of the orbits are not far from circular and have an alti tude near 845 km. Figure 7 makes clear the nature of the debris piece orbit planes suggested by Figure 5; all the debris pieces have an inclination angle near 81 ° but have different right ascensions, clustered mostly between 210° and 310°. W hile the natural nodal drift caused by J2 might be exploited in the design of multi-target rendezvous trajectories, we did not have the resources to pursue that idea in this study and have relegated it to future work. Table 4 presents the launch mass capabilities of various launch vehicles in the Delta series of rockets to a circular orbit of 845 km altitude at 81.2° inclination as this corresponds to the first debris piece in the itinerary identified by applying the Series Method. T hese launch vehicle performance data are approximate and were derived from the Kennedy Space Cen ter (KSC) NASA Launch Services Programs Launch Vehicle Performance Web Site (http://elvperf.ksc.nasa. gov / e 1 vMap I). Table 4 includes several variants of the Delta II and Delta IV launch vehicles. T he Series Method algorithm was executed on the 42 debris pieces using Lambert targeting to compute the rendezvous trajectories between debris pieces. The algorithm was set to construct an itinerary for visiting 32 of the debris pieces and to try each of the 42 pieces as the first target, selecting the choice of first target that served to minimize the total � V. The total � V computed for the rendezvous sequence for 32 target objects was approximately 12 km/s, which is substan tial, and was driven, as predicted, by the plane changes re quired to travel between the debris pieces. The total time re quired for the rendezvous sequence for 32 debris pieces was 260 days and includes some stay time at each debris piece for terminal rendezvous, proximity operations, capture, tether at tachment, release, and departure (though the � V associated with these activities was not computed). The trajectory design results were then subjected to a post processing step in which the total required launch mass was computed as a function of the spacecraft dry mass, which is (continued) a function of the number of debris pieces to visit as specified in (1), and the spacecraft thruster specific impulse. Six repre sentative values of specific impulse were selected: 200, 300, 450, 1, 600, 2, 200, and 3, 000 seconds. The values of 200 and 300 seconds serve to bound the typical performance of high-thrust conventional propulsion systems while 450 sec onds represents the upper bound on conventional propulsion. The values of 1, 600,2, 200, and 3, 000 seconds serve to rep resent low, medium, and high performance low-thrust propul sion systems, respectively. While no low-thrust trajectory de sign was actually performed in this study, the total � V re quirements will generally be similar (total � V for low-thrust trajectories does tend to be somewhat higher than for ballistic trajectories due to kinematic inefficiencies) to what we com puted for the impulsive maneuver ballistic trajectories and so utilizing the low-thrust propulsion system performance pa rameters in this study sheds light on how low-thrust tech nology might aid the debris removal effort. However, it is worth noting that the required flight times for the low-thrust trajectories would tend to be substantially longer than what we computed here for the high-thrust trajectories. Table 5 presents the post-processing results in terms of re quired launch mass to rendezvous with each of 5 to 32 de bris pieces using spacecraft thruster specific impulses rang ing from 200 to 3, 000 seconds. Table 5 is color-coded to indicate which combinations of number of debris pieces and thruster specific impulse can be handled by particular groups of launch vehicles. Some solutions can be handled by ei ther the smaller Delta II rockets or the larger Delta IV rockets (colored green in the table), some only by the larger Delta IV rockets (colored yellow in the table), and some solutions are not possible even with the largest rocket, the Delta IV Heavy (colored red in the table). As described previously, the dry mass is a direct function of the number of objects to be vis ited as this affects the number of EDTs required for a given mission. The results shown in Table 5 are presented in a different way in Table 6, showing how many debris pieces can be de-orbited in less than a year via EDTs for each combination of launch vehicle and thruster. These results are fairly promising; even the least capable launch vehicle and thruster are capable of de-orbiting 6 pieces of debris. The smallest of the Delta IV series of launch vehicles is capable of de-orbiting 11 pieces of debris even with the least capable thruster. The largest (and most expensive) launch vehicle, the Delta IV Heavy, is ca pable of de-orbiting 20 pieces of debris, with the spacecraft using a moderately capable conventional thruster. Note that increasing the specific impulse, as would be the case if low thrust spacecraft propulsion was used, shows that it might be possible to de-orbit all 42 pieces of debris when used with one of the Delta IV launch vehicles (not necessarily the Delta IV Heavy). Detailed low-thrust trajectory design needs to be performed to verify this result and determine how much mis sion time would be required, but even the preliminary results obtained here are profound. Directory: files files -> Answer True False 2 points Question 2 files -> Integer programming and game theory files -> 4 Integer Programming files -> Bpa vehicle Window Repair Scenario #1 task: Procure vehicle window relacement. 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