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ItemCharacterizing High-Energy-Density Propellants for Space Propulsion Apllications(Georgia Institute of Technology, 2006-10) Kokan, Timothy Salim ; Olds, John R.A technique for determining the thermophysical properties of high-energy-density matter (HEDM) propellants is presented. HEDM compounds are of interest in the liquid rocket engine industry due to their high density and high energy content relative to existing industry standard propellants (liquid hydrogen, kerosene, and hydrazine). In order to model rocket engine performance, cost, and weight in a conceptual design environment, several thermodynamic and physical properties are needed. These properties include enthalpy, entropy, density, viscosity, and thermal conductivity. These properties need to be known over a wide range of temperature and pressure. A technique using a combination of quantum mechanics and molecular dynamics is used to determine these properties for quadricyclane, a HEDM compound of interest. Good agreement is shown with experimentally measured thermophysical properties. A vehicle case study is provided to quantify the system level benefits of using quadricyclane instead of hydrazine for the lunar lander ascent stage of the Exploration Systems Architecture Study. The results show that the use of HEDM propellants can significantly reduce the lunar lander mass and indicate that HEDM propellants are an attractive technology to pursue for future lunar missions.
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ItemProbabilistic Cost, Risk, and Throughput Analysis of Lunar Transportation Architectures(Georgia Institute of Technology, 2006-03) Alemany, Kristina ; Olds, John R.The President's Vision for Space Exploration presents a need to determine the best architecture and set of vehicle elements in order to achieve a sustained human lunar exploration program. The Lunar Architecture Stochastic Simulator and Optimizer (LASSO), a new simulation-based capability based on discrete-event simulation, was created to address this question by probabilistically simulating lunar transportation architecture based on cost, reliability, and throughput figures of merit. In this study, two competing lunar transportation architectures are examined for a variety of launch vehicle scenarios to determine the best approach for human lunar exploration. Additionally, the two architectures are also compared for varying available ground infrastructure and desired flight rates. It is concluded that an expendable architecture is favored, using man-rated versions of existing evolved expendable launch vehicles (EELVs) for crew launches and developing a heavy-lift launch vehicle for cargo launches.
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ItemAn Experimental and Analytical Study of High-Energy-Density propellants for Liquid Rocket Engines(Georgia Institute of Technology, 2005-07) Kokan, Timothy Salim ; Olds, John R.There exists wide ranging research interest in high-energy-density matter (HEDM) propellants as a potential replacement of existing industry standard fuels (LH2, RP-1, MMH, UDMH) for liquid rocket engines. The U.S. Air Force Research Laboratory, the U.S. Army Research Lab, and the NASA Marshall Space Flight Center each have ongoing programs in the synthesis and development of these potential new propellants. The thermophysical understanding of HEDM propellants is necessary to model their performance in the conceptual design of liquid rocket engines. Most industry standard powerhead design tools (e.g. NPSS, ROCETS, and REDTOP-2) require several thermophysical properties of a given propellant over a wide range of temperature and pressure. These properties include enthalpy, entropy, density, internal energy, specific heat, viscosity, and thermal conductivity. For most of these potential new HEDM propellants, this thermophysical data either does not exist or is incomplete over the range of temperature and pressure necessary for liquid rocket engine design and analysis. The work presented is a technique for obtaining enthalpy and density data for new propellants through the use of a combination of analytical/computational methods (quantum mechanics and molecular dynamics) and experimental investigations. Details of this technique and its application to an example HEDM fuel currently of interest are provided.
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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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ItemArtemis: A Reusable Excursion Vehicle Concept for Lunar Exploration(Georgia Institute of Technology, 2005-07) Young, David Anthony ; Olds, John R. ; Hutchinson, Virgil L., Jr. ; Krevor, Zachary C. ; Young, James J.Artemis is a reusable excursion vehicle for lunar landing missions. It is intended to transport a notional CEV vehicle from low lunar orbit (LLO) to the lunar surface. It can be reused by refueling the vehicle in LLO. Artemis is nominally sized to carry a 10 MT payload to the lunar surface and then return it to LLO. Artemis is powered by four liquid oxygen and liquid hydrogen fueled RL-10 engines. These RL-10 engines provide the necessary thrust and allow the Artemis lander to complete its nominal mission with two engines inoperative. The Artemis lander has volume margin built into its propellant tanks. This volume margin combined with an innovative cross-feed system allows Artemis to complete its ascent from the lunar surface with a propellant tank failure. This cross-feed system also allows Artemis to adjust the center of gravity (cg) of the vehicle by transferring propellant among the propellant tanks. Artemis lands on the moon with six articulating legs. This provides redundancy against a leg failure on landing and provides Artemis with the ability to land on uneven terrain. This vehicle is designed to be launched by a heavy-lift evolved expendable launch vehicle (EELV). This design constraint results in the distinct shape of the lander. Artemis is launched as a compact cylinder in the EELV payload shroud, and then autonomously assembles itself via robotic arms similar to those currently used by the shuttle program. Details of the conceptual design process used for Artemis are included in this paper. The disciplines used in the design include configuration, propulsion design and selection, trajectory, mass properties, structural design, cost, operations, and 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 team process and used to minimize the gross weight of the Artemis. Once the design process was completed, a parametric Excel based model was created from the point design. This model can be used to resize Artemis for changing system metrics (such as payload) as well as changing technologies. The Artemis recurring and non-recurring costs were also computed. The total development cost including the design, development, testing and evaluation (DDT&E) cost is $2.17 B FY'04. The theoretical first unit (TFU) cost is $303 M FY'04. Trade studies on life cycle costs (LCC) vs. fuel cost to LLO as well as flight rate are also discussed. A summary of design disciplines as well as the economic results are included.
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ItemResponsive Access Small Cargo Affordable Launch (RASCAL) Independent Performance Evaluation(Georgia Institute of Technology, 2005-05) Young, David Anthony ; Olds, John R.RASCAL is a United States Defense Department initiative that stands for Responsive Access, Small Cargo, Affordable Launch. The overall launch concept involves three stages. The first stage will consist of a reusable aircraft similar to a large scale Air Force fighter. The first stage will also utilize Mass Injection Pre-Compressor Cooling (MIPCC) turbojet engines that will propel the stage to approximately two hundred thousand feet before releasing the second and third rocket stages. The first stage will be similar to a large fighter of the F-22 class, although the turbofans will be that of the more available F100 class. The MIPCC system will be a plug-in addition to the engines to help high altitude performance. This stage will act as a first stage in the RASCAL architecture and will contribute significantly to the overall acceleration of the vehicle. The second and third stages of the RASCAL concept consist of expendable rockets. Releasing the upper stages outside the atmosphere will reduce the loads on the stages as well as the risk of staging. Also by relying on the reusable portion for all atmospheric flight, the expendable stages can be designed simpler and therefore cheaper. The purpose of this project is to compare the published RASCAL numbers with those computed using a design methodology currently used in the Space System Design Laboratory (SSDL) at The Georgia Institute of Technology. When the initial Space Launch Corporation design was evaluated using the SSDL methodology it was found to fall short of the performance as well as the cost goals set by DARPA for the RASCAL program. The baseline vehicle was found to only carry 52 lbs to the 270 nmi sun synchronous orbit. Several alternatives were evaluated off of the baseline design. The best of these alternatives can meet DARPA’s performance goals and reach the cost goals of $5,000 per pound of payload with eight first stage vehicles flying 46 times per year for a total of 363 flights per year. Different economic cases were also evaluated to try and meet the cost goals in a less ambitious number of flights per year. It was found that if the DDT&E was paid for by another party (NASA, DOD, etc.) the cost goals can be met with just three vehicles flying 42 times per year for a total of 125 flights per year.
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ItemDaedalon: A Revolutionary Morphing Spacecraft Design for Planetary Exploration(Georgia Institute of Technology, 2005-01) Lafleur, Jarret M. ; Olds, John R. ; Braun, Robert D.The product of a study sponsored by the NASA Institute for Advanced Concepts (NIAC), Daedalon is a spacecraft design baselined for Mars which utilizes morphing wing technology to achieve the design objective of a standard, flexible architecture for unmanned planetary exploration. This design encompasses a detailed vehicle mass and power sizing study for the Daedalon lander as well as its cruise stage and entry backshell. A cost estimation and comparison study is also performed, and qualitative functionality comparisons are made between Daedalon and other Mars lander and airplane designs. Quantitative comparisons of gross mass and range are also made, including the results of scaling an existing Mars aerial vehicle design to match Daedalon functionality. Altogether, the Daedalon launch mass is found to be 896 kg for a 12 kg payload capacity. If five such vehicles are produced, it is found that the per-mission cost can be as low as $224 million. Given the necessary morphing wing technology development, it is concluded that the Daedalon design may be a feasible and cost-effective approach to planetary exploration 20-40 years in the future.
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ItemDevelopment of SAMURAI - Simulation and Animation Model Used for Rockets with Adjustable Isp.(Georgia Institute of Technology, 2005-01) Sakai, Tadashi ; Olds, John R. ; Alemany, KristinaAn interplanetary trajectory calculation application SAMURAI - Simulation and Animation Model Used for Rockets with Adjustable Isp - has been developed. SAMURAI is written in C++ and calculates transfer trajectories with variable thrust, variable Isp (VASIMR type) engines as well as conventional constant low thrust, constant Isp engines and high thrust engines. SAMURAI utilizes a calculus of variations algorithm to evaluate the thrust history that minimizes the fuel consumption from one planet to another. A trajectory with a planetary swing-by can also be calculated. After calculation, a 3D animation of the resulting transfer trajectory is created and can be viewed with a web browser using VRML. In this paper, equations and methods used in SAMURAI, and the capabilities of this application are presented. A few examples including a round trip from Earth to Mars have been analyzed, and trajectories with variable Isp engines, constant Isp, engines, and high thrust engines are compared.
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ItemA Quantitative Methodology for Identifying Evolvable Space Systems(Georgia Institute of Technology, 2005-01) Christian, John A., III ; Olds, John R.With the growing emphasis on spiral development, a system’s ability to evolve is becoming increasingly critical. This is especially true in systems designed for the exploration of space. While returning to the Moon is widely regarded as the next step in space exploration, our journey does not end there. Therefore, the technologies, vehicles, and systems created for near-term lunar missions should be selected and designed with the future in mind. Intelligently selecting evolvable systems requires a method for quantitatively measuring evolvability and a procedure for comparing these measurements. This paper provides a brief discussion of a quantitative methodology for evaluating space system evolvability and an in-depth application of this methodology to an example case study.
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ItemFlight System Options for a Long Duration Mars Airplane(Georgia Institute of Technology, 2004-09) Rohrschneider, Reuben R. ; Olds, John R. ; Kuhl, Christopher A. ; Braun, Robert D. ; Steffes, Stephen R. ; Hutchinson, Virgil L., Jr.The goal of this study was to explore the flight system options for the design of a long endurance Mars airplane mission. The mission model was built in the design framework ModelCenter and a combination of a hybrid and user-driven fixed point iteration optimization method was used to determine the maximum endurance solution of each configuration. Five different propulsion systems were examined: a bipropellant rocket, a battery powered propeller, a direct methanol fuel cell powered propeller, and beamed solar and microwave powered propeller systems. Five airplane configurations were also studied. The best configuration has a straight wing with two vertical tails. The direct methanol fuel cell proved to be the best onboard power system for a long endurance airplane and the solar beamed power system showed potential for indefinite flight. The combination of the best configuration and the methanol fuel cell resulted in an airplane capable of cruising for 17.8 hours on Mars.