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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. Nomenclature α = angle-of-attack, ° AFRSI = Advanced Flexible Reusable Surface Insulation CAD = computer aided design CER = cost estimating relationship c L = coefficient of lift DDT&E = design, development, test, & evaluation DoD = Department of Defense DSM = Design Structure Matrix EMA = electro-mechanical actuators

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A technique for computationally 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. In order to accurately model rocket engine performance, cost and weight in a conceptual design environment, several thermodynamic and physical properties are required over a range of temperatures and pressures. The approach presented here combines quantum mechanical and molecular dynamic (MD) calculations and group additivity methods. A method for improving the force field model coefficients used in the MD is included. This approach is used to determine thermophysical properties for two HEDM compounds of interest: quadricyclane and 2-azido-N,N-dimethylethanamine (DMAZ). The modified force field approach provides results that more accurately match experimental data than the unmodified approach. Launch vehicle and Lunar lander case studies are presented to quantify the system level impact of employing quadricyclane and DMAZ rather than industry standard propellants. In both cases, the use of HEDM propellants provides reductions in vehicle mass compared to industry standard propellants. The results demonstrate that HEDM propellants can be an attractive technology for future launch vehicle and Lunar lander applications.

The Aztec reusable launch vehicle (RLV) concept is a two-stage-to-orbit (TSTO) horizontal takeoff, horizontal landing (HTHL) vehicle. The first stage is powered by ten JP-5 fueled turbine-based combined-cycle (TBCC) engines. The second stage is powered by three high energy density matter (HEDM)/liquid oxygen (LOX) staged-combustion rocket engines. The HEDM fuel is a liquid hydrogen-based propellant with a solid aluminum and methane gel additive.Aztec is designed to deliver 20,000 lbs of payload to a 100 nmi x 100 nmi x 28.5° orbit due East out of Kennedy Space Center (KSC). The second stage separates at Mach 8 and continues to the target orbit while the first stage flies back to KSC in ramjet mode. For the above payload and target orbit, the gross lift-off weight (GLOW) is estimated to be 690,000 lbs and the total dry weight for both stages is estimated to be 230,000 lbs. Economic analysis indicates that the Aztec recurring launch costs will be approximately $590 per lb. of payload delivered to the target orbit. The total non-recurring cost including design, development, testing and evaluation (DDT&E), acquisition of the first vehicle, and the construction of launch and processing facilities is expected to be $13.6B. All cost figures are in FY$2004 unless otherwise noted.Details of the Aztec design including external and internal configuration, aerodynamics, mass properties, first and second stage engine performance, ascent and flyback trajectory, aeroheating results and thermal protection system (TPS), vehicle ground operations, vehicle safety and reliability, and a cost and economics assessment are provided. Nomenclature α = angle-of-attack, ° AFRSI = Advanced Flexible Reusable Surface Insulation ASTP = Advanced Space Transportation Program CAD = computer aided design CER = cost estimating relationship c L = coefficient of lift DDT&E = design, development, test, & evaluation DSM = Design Structure Matrix EMA = electro-mechanical actuators GLOW = gross lift-off weight GRC = Glenn Research Center HEDM = high energy density matter HTHL = horizontal takeoff, horizontal landing IOC = initial operating capability 2 Isp = specific impulse, sec KSC = Kennedy Space Center LCC = life cycle cost LOX = liquid oxygen LRU = line replacement unit MECO = main engine cutoff MER = mass estimating relationship MR = mass ratio (gross weight / burnout weight) MSFC = Marshall Space Flight Center OMS = orbital maneuvering system PEF = propellant packaging efficiency q = dynamic pressure, psf RCS = reaction control system RLV = reusable launch vehicle RTA = Revolutionary Turbine Accelerator TBCC = turbine-based combined-cycle TFU = theoretical first unit TPS = thermal protection system TSTO = two-stage-to-orbit TUFI = Toughened Unipiece Fibrous Insulation UHTC = Ultra-High Temperature Ceramic

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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 1,2 , the U.S. Army Research Lab 3,4 , and the NASA Marshall Space Flight Center 5 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 6 and molecular dynamics 7) and experimental investigations. Details of this technique and its application to an example HEDM fuel currently of interest are provided.