Preliminary assessment of using gel and hybrid propellant propulsion for VTOL/SSTO launch systems

Pelaccio, Dennis G.; Palaszewski, Bryan; O'Leary, Robert

doi:10.2514/6.1997-3216

Cited by 8 publications

(4 citation statements)

References 8 publications

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“…In this case, the primary benefit of the gelled propellant can be its improved safety characteristics due to resistance to spilling or leaking. 6 Ultra-High I sp HEDM Systems. In these systems, a HEDM propellant is produced by adding a high-energy atom or molecule to a cryogenic solid.…”

Section: Cryogenic Propellant Thermal Controlmentioning

confidence: 99%

“…For example, the cost of ground-test facilities for NTR testing scale with engine thrust (because of the need to "scrub" the engine exhaust of any nuclear materials); this approach can enable lowcost testing of a small NTR engine capable of producing high thrust in LANTR mode. (The LANTR mode could be tested in a non-nuclear facility 6 American Institute of Aeronautics and Astronautics separately from the NTR engine testing by using resistively-heated [i.e., non-nuclear heated] H 2 .) Additional systems-level benefits include reduced gravity-losses for liftoff from the Moon or for escape/capture in low Earth orbit.…”

Section: American Institute Of Aeronautics and Astronauticsmentioning

confidence: 99%

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Advanced Propulsion for the XXIst Century

Frisbee

2003

AIAA International Air and Space Symposium and Exposition: The Next 100 Years

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This paper presents an overview of advanced space propulsion concepts and their research activities at the beginning of the 21 st Century. If successfully developed, these concepts will enable in the coming decades the implementation of propulsion technologies that will reduce the cost of access to space and/or enable aggressive exploration, and ultimately exploitation of the Solar System and beyond. This paper focuses primarily on mid-to far-term in-space propulsion applications, although advanced launch concepts (e.g., launch assist catapults) are also discussed briefly.Examples of advanced propulsion concepts to be discussed include advanced chemical propulsion (high-energy density materials, gelled propellants, etc.), nuclear propulsion (fission, fusion, and antimatter annihilation), electric propulsion (solar-and nuclearelectric propulsion power systems and electric propulsion thrusters), micropropulsion, beamed-energy propulsion (e.g., solar/laser/microwave thermal/electric combinations), beamed-momentum propulsion (e.g., solar/laser/microwave sails, electromagnetic/plasma sails), aero/gravity assist, launch assist catapults, tethers (momentum exchange and electrodynamic), extraterrestrial resource utilization, and breakthrough physics propulsion.

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Section: Cryogenic Propellant Thermal Controlmentioning

confidence: 99%

Section: American Institute Of Aeronautics and Astronauticsmentioning

confidence: 99%

Advanced Propulsion for the XXIst Century

Frisbee

2003

AIAA International Air and Space Symposium and Exposition: The Next 100 Years

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“…Experiments in the formation of solid hydrogen particles in liquid helium have been performed 3 . Studies using gelled hydrogen and metallized gelled hydrogen fuels have shown potential in significantly increasing payload delivery capability and/or reducing GLOW 4,5 . Gelled hydrogen fuel consists of liquid hydrogen with solid, frozen particles of a different fuel added to form a gel structure in the hydrogen.…”

Section: Introductionmentioning

confidence: 99%

Aztec: A TSTO Hypersonic Vehicle Concept Utilizing TBCC and HEDM Propulsion Technologies

Kokan

Olds

Hutchinson

et al. 2004

40th AIAA/ASME/SAE/ASEE Joint Propulsion Conference and Exhibit

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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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“…Experiments in the formation of solid hydrogen particles in liquid helium have been performed 1 . Studies using gelled hydrogen and metallized gelled hydrogen fuels have shown potential in significantly increasing payload delivery capability and/or reducing GLOW 2,3 . Gelled hydrogen fuel consists of liquid hydrogen with solid, frozen particles of a different fuel added to form a gel structure in the hydrogen.…”

Section: Introductionmentioning

confidence: 99%

Lazarus: A SSTO Hypersonic Vehicle Concept Utilizing RBCC and HEDM Propulsion Technologies

Young

Kokan²,

Clark³

et al. 2006

14th AIAA/AHI Space Planes and Hypersonic Systems and Technologies Conference

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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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Preliminary assessment of using gel and hybrid propellant propulsion for VTOL/SSTO launch systems

Cited by 8 publications

References 8 publications

Advanced Propulsion for the XXIst Century

Advanced Propulsion for the XXIst Century

Aztec: A TSTO Hypersonic Vehicle Concept Utilizing TBCC and HEDM Propulsion Technologies

Lazarus: A SSTO Hypersonic Vehicle Concept Utilizing RBCC and HEDM Propulsion Technologies

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