Characterizing High-Energy-Density Propellants for Space Propulsion Apllications

57th International Astronautical Congress October 2006, Valencia, Spain.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

Characterizing High-Energy-Density Propellants for Space Propulsion Applications

There exists wide ranging research interest in high-energy-density matter (HEDM) propellants as a potential replacement for existing industry standard fuels for liquid rocket engines. The U.S. Air Force Research Laboratory, the U.S. Army Research Lab, the NASA Marshall Space Flight Center, and the NASA Glenn Research Center each either recently concluded or currently has ongoing programs in the synthesis and development of these potential new propellants. In order to perform conceptual designs using these new propellants, most conceptual rocket engine 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, viscosity, and thermal conductivity. Very little thermophysical property data exists for most of these potential new HEDM propellants. Experimental testing of these properties is both expensive and time consuming and is impractical in a conceptual vehicle design environment. A new technique for determining these thermophysical properties of potential new rocket engine propellants is presented. The technique uses a combination of three different computational methods to determine these properties. Quantum mechanics and molecular dynamics are used to model new propellants at a molecular level in order to calculate density, enthalpy, and entropy. Additivity methods are used to calculate the kinematic viscosity and thermal conductivity of new propellants. This new technique is validated via a series of verification experiments of HEDM compounds. Results are provided for two HEDM propellants: quadricyclane and 2-azido-N, N-dimethylethanamine (DMAZ). In each case, the new technique does a better job than the best current computational methods at accurately matching the experimental data of the HEDM compounds of interest. A case study is provided to help quantify the vehicle level impacts of using HEDM propellants. The case study consists of the National Aeronautics and Space Administrations (NASA) Exploration Systems Architecture Study (ESAS) Lunar Surface Access Module (LSAM). The results of this study show that the use of HEDM propellants instead of hypergolic propellants can lower the gross weight of the LSAM and may be an attractive alternative to the current baseline hypergolic propellant choice.Ph.D.Committee Co-Chair: Jerry M. Seitzman; Committee Co-Chair: John R. Olds; Committee Member: John A. Blevins; Committee Member: Mitchell L. Walker II; Committee Member: Peter J. Ludovic

An Experimental and Analytical Study of High-Energy-Density propellants for Liquid Rocket Engines

41st AIAA/ASME/SAE/ASEE Joint Propulsion Conference and Exhibit July 10-13, 2005, Tucson, AZ.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

Mission Capture Rate versus Turnaround Time and Fleet Size for the Military Spaceplane

38th AIAA/ASME/SAE/ASEE Joint Propulsion Conference And Exhibit Indianapolis, IN, July 7-10, 2002.The United States Air Force Research Laboratory (AFRL) is conducting research into a military spaceplane (MSP) through the Military Spaceplane System Technology Program Office. The goal of this program is to provide the Air Force with safe, reliable, affordable, and routine access to space. An important mission performance metric of the MSP program is the mission capture rate. The mission capture rate is a measure of the MSP’s ability to meet mission sortie requirements. Extending this to a fleet of MSPs, the mission capture rate is defined as the total number of sorties the fleet is capable of divided by the total required number of sorties. This research analyzes the relationship between mission capture rate and both turnaround time and fleet size. The turnaround time is the time between when the vehicle lands and when it can take off again. During this time the vehicle is refueled, maintenance and repair work is done, and the payload is loaded. As turnaround time decreases and fleet size increases, the mission capture rate will increase. A precise definition of this relationship is made in order to determine the necessary fleet size for a given turnaround time subject to a desired mission capture rate. A Monte Carlo simulation is performed to probabilistically analyze the mission capture rates. This analysis takes into account uncertainties in the utilization requirements of the MSP fleet. These uncertainties include the number of wars within the simulation period, the starting date & duration of each war, and each war’s required sortie rate. This analysis utilizes Crystal Ball Pro® along with Microsoft Excel®. This gives the analysis technique compatibility with commonly used computer platforms

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

40th AIAA/ASME/SAE/ASEE Joint Propulsion Conference And Exhibit Fort Lauderdale, FL, July 11 - 14, 2004.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 deg 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 dollars 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 13.6 B dollars. All cost figures are in FY2004 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

The Bearing of Selection Experiments with Drosophila upon the Frequency of Germinal Changes

14th AIAA/AHI Space Planes and Hypersonic Systems and Technologies Conference November 2006, Canberra, AustraliaLazarus 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

Tanker Argus: Re-supply for a LEO Cryogenic Propellant Depot

53rd International Astronautical Congress, The World Space Congress Houston, TX, October 10-19, 2002.The Argus reusable launch vehicle (RLV) concept is a single-stage-to-orbit (SSTO) conical, wingedbodied vehicle powered by two liquid hydrogen (LH2)/liquid oxygen (LOX) supercharged ejector ramjets (SERJ). The 3rd generation Argus launch vehicle utilizes advanced vehicle technologies along with a magnetic levitation (Maglev) launch assist track. A tanker version of the Argus RLV is envisioned to provide an economical means of providing liquid fuel and oxidizer to an orbiting low Earth orbit (LEO) propellant depot. This depot could then provide propellant to various spacecraft, including reusable orbital transfer vehicles used to ferry space solar power (SSP) satellites to geo-stationary orbit. Two different tanker Argus configurations were analyzed. The first simply places additional propellant tanks inside the payload bay of an existing Argus reusable launch vehicle. The second concept is a modified pure tanker version of the Argus RLV in which the payload bay is removed and the vehicle propellant tanks are extended to hold additional propellant. An economic analysis was performed for this study that involved the calculation of the design/development and recurring costs of each vehicle. The goal of this analysis was to determine at what flight rate it would be economically beneficial to spend additional development funds to change an existing, sunk cost, payload bay tanker vehicle into a pure tanker design. The results show that for yearly flight rates greater than ~50 flts/yr it is cheaper, on a $/lb basis , to develop and operate a dedicated tanker