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Space Systems Design Laboratory (SSDL)
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ItemLazarus: A SSTO Hypersonic Vehicle Concept Utilizing RBCC and HEDM Propulsion Technologies(Georgia Institute of Technology, 2006-11) Young, David Anthony ; Kokan, Timothy Salim ; Clark, Ian G. ; Tanner, Christopher ; Wilhite, Alan W.Lazarus is an unmanned single stage reusable launch vehicle concept utilizing advanced propulsion concepts such as rocket based combined cycle engine (RBCC) and high energy density material (HEDM) propellants. These advanced propulsion elements make the Lazarus launch vehicle both feasible and viable in today's highly competitive market. The Lazarus concept is powered by six rocket based combined cycle engines. These engines are designed to operate with HEDM fuel and liquid oxygen (LOX). During atmospheric flight the LOX is augmented by air traveling through the engines and the resulting propellant mass fractions make single stage to orbit (SSTO) possible. A typical hindrance to SSTO vehicles are the large wings and landing gear necessary for takeoff of a fully fueled vehicle. The Lazarus concept addresses this problem by using a sled to take off horizontally. This sled accelerates the vehicle to over 500 mph using the launch vehicle engines and a propellant cross feed system. This propellant feed system allows the vehicle to accelerate using its own propulsion system without carrying the necessary fuel required while it is attached to the sled. Lazarus is designed to deliver 5,000 lbs of payload to a 100 nmi x 100 nmi x 28.5° orbit due East out of Kennedy Space Center (KSC). This mission design allows for rapid redeployment of small orbital assets with little launch preparation. Lazarus is also designed for a secondary strike mission. The high speed and long range inherent in a SSTO launch vehicle make it an ideal global strike platform. Details of the conceptual design process used for Lazarus are included in this paper. The disciplines used in the design include aerodynamics, configuration, propulsion design, trajectory, mass properties, cost, operations, reliability and safety. Each of these disciplines was computed using a conceptual design tool similar to that used in industry. These disciplines were then combined into an integrated design process and used to minimize the gross weight of the Lazarus design.
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ItemMars Entry, Descent, and Landing Parametric Sizing and Design Space Visualization Trades(Georgia Institute of Technology, 2006-08) Alemany, Kristina ; Wells, Grant William ; Theisinger, John ; Clark, Ian G. ; Braun, Robert D.Entry, descent, and landing (EDL) is a multidimensional, complex problem, which is difficult to visualize in simple plots. The purpose of this work is to develop a systematic visualization scheme that could capture Mars EDL trades as a function of a limited number of variables, such that programmatic design decisions could be effectively made with insight of the design space. Using the Mars Science Laboratory (MSL) as a basis, contour plots have been generated for key EDL figures of merit, such as maximum landed elevation and landed mass as a function of four input parameters: entry mass, entry velocity, entry flight path angle, and vehicle L/D. Additionally, sensitivity plots have been generated in an attempt to capture the effects of varying the fixed input parameters. This set of EDL visualization data has been compiled into a Mars EDL handbook to aid in pre-phase A design space exploration and decision making.
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ItemAn Evaluation of Ballute Entry Systems for Lunar Return Missions(Georgia Institute of Technology, 2006-08) Clark, Ian G. ; Braun, Robert D. ; Theisinger, John ; Wells, Grant WilliamThis study investigates the advantages and feasibility of using ballutes for Earth entry at lunar return velocities. Using analysis methods suitable for conceptual design and assuming a CEV type entry vehicle, multiple entry strategies were investigated. Entries that jettison the ballute after achieving low Earth orbit conditions were shown to reduce heating rates to within reusable thermal protection system limits. Deceleration was mitigated to approximately four g's when a moderate amount of lift was applied subsequent to ballute jettison. Primary ballute size drivers are the thermal limitations and areal densities of the ballute material. Performance requirements for both of those metrics were generated over a range of total ballute system masses. Lastly, preliminary investigation of a lower mass cargo variant of the CEV allowed for additional reduction of ballute system mass. However, ballute system mass as a percentage of the total entry mass was shown to be relatively independent of the entry mass.
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ItemTempest: Crew Exploration Vehicle Concept(Georgia Institute of Technology, 2005-07) Hutchinson, Virgil L., Jr. ; Olds, John R. ; Alemany, Kristina ; Christian, John A., III ; Clark, Ian G. ; Crowley, John ; Krevor, Zachary C. ; Rohrschneider, Reuben R. ; Thompson, Robert W. ; Young, David Anthony ; Young, James J.Tempest is a reusable crew exploration vehicle (CEV) for transferring crew from the Earth to the lunar surface and back. Tempest serves as a crew transfer module that supports a 4-person crew for a mission duration of 18 days, which consists of 8 days total transit duration and 10-day surface duration. Primary electrical power generation and on-orbit maneuvering for Tempest is provided by an attached Power and Propulsion Module (PPM). Hydrogen (H2)/oxygen (O2) fuel cells and a high energy-density matter (HEDM)/liquid oxygen (LOX) propellant reaction control system (RCS) provide power and reaction control respectively during Tempest’s separation from the PPM. Tempest is designed for a lifting entry and is equipped with parachutes for a soft landing. Tempest is part of an overall lunar transportation architecture. The 60,731 lbs combination of Tempest and the PPM are launched atop the notional Centurion C-1 heavylift launch vehicle (HLLV) and delivered to a 162 nmi, 28.5º circular orbit. After separating from the C-1 upper stage, the Tempest/PPM autonomously rendezvous with Manticore, an expendable trans-lunar injection (TLI) stage pre-positioned in the current orbit, and transfer to a lunar trajectory. After entering a 54 nmi polar circular lunar orbit, the Tempest/PPM separate from Manticore. Tempest separates from the PPM and is ferried to/from the lunar surface by Artemis, a reusable lunar lander. Upon return from the lunar surface, Tempest reconnects with the PPM, and the PPM provides the trans-earth injection (TEI) burn required to return to low earth orbit (LEO). Prior to atmospheric entry, Tempest separates from the PPM and subsequently executes a lifting entry trajectory. Crushable thermal foam attached to the lower surface of Tempest serves as an ablative thermal protection system (TPS) and the impact absorber of the parachute landing. Details of the conceptual design process used for Tempest are included in this paper. The disciplines used in the design include: configuration, aerodynamics, propulsion, trajectory, mass properties, environmental control life support system (ECLSS), entry aeroheating and TPS, terminal landing system (TLS), cost, operations, and reliability & safety. Each of these disciplines was computed using a conceptual design tool similar to that used in industry. These disciplines were then combined and optimized for the minimum gross weight of the Tempest CEV. The total development cost including the design, development, testing and evaluation (DDT&E) cost was determined to be $2.9 B FY’04. The theoretical first unit (TFU) cost for the Tempest CEV was $479 M FY’04. A summary of design disciplines as well as the economic results are included.
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ItemReusable Exploration Vehicle (REV): Orbital Space Tourism Concept(Georgia Institute of Technology, 2005-05) Clark, Ian G. ; Francis, Scott R. ; Otero, Richard E. ; Wells, Grant WilliamOn the heels of the recent success of the X-Prize, sub-orbital space tourism is nearly a reality. Though the requirements are significantly tougher, orbital space tourism is the next logical step. The Reusable Exploration Vehicle (REV) concept is an economically feasible design capable of making this next step. Centered around a lenticular lifting body, the REV concept relies on commercial launch vehicles to reduce DDT&E expenditures. Capable of ferrying five passengers and one crew member for three orbits, the REV is shown to be capable of keeping maximum debt exposure to less than $250M while attaining an IRR of 70% with an estimated market capture of 66%.