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

Many conceptual launch vehicles are designed through the integration of various disciplines, such as aerodynamics, propulsion, trajectory, weights, and aeroheating. In the determination of the total vehicle weight, a large percentage of the vehicle weight is composed of the structural weight of the vehicle subsystems, such as propellant tanks. Empirical mass estimating relations (MERs) and multi-dimensional finite element analysis (FEA) are two methods commonly used by the aerospace industry to estimate the loadbearing structural weight. MERs rapidly estimate the weight by evaluating empirical equations and the high-fidelity techniques of FEA accurately calculates the structural weight. The extreme inability for either method to provide both rapid and accurate weight estimations warrants an investigation into developing an improved, intermediate method.A methodology based on fundamental beam structural analysis has been developed for the rapid estimation of the load-bearing structural weight of the launch vehicle fuselage and integral propellant tanks. By creating a simplified beam approximation model of the vehicle, the method utilizes the vehicle component weights, load conditions, and basic material properties to analytically estimate the structural shell and stability frame weight. Implementation of this methodology into a fast-acting software tool allowed for rapid estimation of the component structural weight. Using statistical techniques, an empirical relationship between the estimated and actual load-bearing structure weights was determined. The method was utilized to estimate the liquid hydrogen (LH 2 ) and liquid oxygen (LOX) propellant tanks for an existing Evolved Expendable Launch Vehicle (EELV) and the Space Shuttle External Tank (ET) for verification and correlation. Nomenclature A= cross-sectional area (in 2 ) a = semi-major axis (in) A f = stability frame cross-sectional area (in 2 ) axial_accel = axial acceleration (g's) b = semi-minor axis (in) c = farthest from the neutral axis along the y-axis (in) C f = Shanley constant (1/16,000) I y = Area Moment of Inertia with respect to y-axis norm_accel = normal acceleration (g's) P ell = perimeter of ellipse (in) p head = head pressure (lb/in 2 ) prop_ullage = ullage pressure (lb/in 2 ) p ullage = ullage pressure (lb/in 2 ) t f = smeared equivalent stability frame thickness (in) t s = equivalent shell thickness (in) x1, x2 = simple support reaction location (in) f = density of stability frame material (lb/in 3 ) = propellant density (lb/in 3 ) s = density of shell material (lb/in 3 ) axial = axial stress (lb/in 2 ) bend = bending stress (lb/in 2 ) hhead = normal stress due to head pressure in hoop (circumferential) direction (lb/in 2 ) hullage = normal stress due to ullage pressure in hoop (circumferential) direction (lb/in 2 ) lhead = normal stress due to head pressure in axial (longitudinal) direction (lb/in 2 ) lullage = normal stress due to ullage pressure in axial (longitudinal) direction (lb/in 2 ) UTS = ultimate tensile strength (lb/in 2 ) YS =...

Centurion is an expendable heavy lift launch vehicle (HLLV) family for launching lunar exploration missions. Each vehicle in the family is built around a common two-stage core. The first stage of the core uses kerosene (RP-1) fuel and utilizes four staged-combustion RD-180 rocket engines. The upper stages consist of liquid oxygen (LOX)/liquid hydrogen (LH2) propellant with three 220,000 lb thrust-class expander rocket engines. The larger variants in the Centurion family will also use either one or two pairs of five-segment solid rocket motors which are now being developed by ATK Thiokol. The Centurion family consists of three vehicles denoted as C-1, C-2, and C-3. The first vehicle (C-1) is a four RD-180 core with a LOX/LH2 upper stage. The C-1 is designed to deliver a 35 metric ton (MT) CEV to a 300 km X 1000 km highly elliptical orbit (HEO). This HEO allows the CEV to more easily transfer to a lunar trajectory, while still having the ability to abort after one revolution. The C-1 also is designed to meet mission requirements with a failure of both one RD-180 and one upper stage engine. The C-2 and C-3 Centurions are both cargo carrying variants which carry 100 MT and 142 MT of cargo to a 407 km low earth orbit (LEO) respectively. The C-2 utilizes two five-segment solid rocket boosters (SRB), while the C-3 uses four SRBs. Details of the conceptual design process used for Centurion are included in this paper. The disciplines used in the design include configuration, aerodynamics, propulsion design and selection, trajectory, mass properties, structural design, aeroheating and thermal protection systems (TPS), 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 C-1 variant. The C-2 and C-3 variants were simulated using the C-1 optimized core with different configurations of SRBs. Each of the variants recurring and non-recurring costs were computed. The total development cost including the design, development, testing and evaluation (DDT&E) cost and a new launch pad at Kennedy Space Center (KSC), was $7.98 B FY'04. The theoretical first unit (TFU) cost for the C-2 variant was $532 M FY'04. A summary of design disciplines as well as the economic results are included. Nomenclature Al-Li = Aluminum-Lithium CAD = computer aided design CER = cost estimating relationship

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. Nomenclature CAD = computer aided design CER = cost estimating relationship

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