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An Overview of Advanced Concepts for Space Access Andrew Ketsdever Assistant Professor University of Colorado at Colorado Springs and Marcus Young Deppyuty Prog ram Manag g,er, Advanced Concep ts Group Air Force Research Laboratory Distribution A: Approved for public release; distribution unlimited. Early Attempts to Defy Gravity Dreams Applications Ideas Break- Research throughs 2 Advances in Flight Turbojet c. 1940 Ramjet C. 1960 Scramjet C. 1990 Advanced Turbofan 3 Rocketry Hero Engine ~62 AD (?) Chinese Rocketry ~First Century AD 1926 1976 4 Introduction • Every man-made object launched to space has been by chemical combustion • Liquid propulsion systems (bi-propellants) • Solid propulsion systems • Hybrid propulsion systems (suborbital) • Energy to reach low-Earth orbit (LEO) • 45 MJ/kg or 12. 5 kW-h/khr/kg • At current Colorado Springs Utilities rates: $1.75/kg • Room for improvement? YES Thrust Payload Mass Cost/kg Launch Cost/ Efficiency Fraction ($1000/kg) Energy Cost 97% (SSME) .025 10s~5500 Trillions spent to increase chemical propulsion efficiency have led to minimal advances – e.g. Isp increases of mere 10’s seconds out of 400. Æ Need a breakthrough 5 Today’s Launch Outlook: Dismal • New vehicles coming on line are looking at advances in the business model of launch Æ not necessarily technology advancement • SXSpaceX (COTS pro duc tion ) • Blue Origins • NASA has cut all funding for Advanced Concept development • Ares depends on existing and in many cases Apollo-era technology • Not necessarily a bad thing • The Air Force has minimal funding currently going into Advanced Propulsion technology for space access • Advanced Concepts Group, Edwards AFB, CA ($1M) • Hyperson ic VhilVehicles B ranch hWihtPtt, Wright Patterson AFB AFBOH($5M), OH (<$5M) • Black programs (?) – probably not in space access technologies directly • Department of Defense • DARPA – Falcon program 6 Essence of the Problem 7 Introduction • AFRL Advanced Concepts Group performed critical review of advanced technologies for space access. • Technologies Considered: Using Propellant Propellantless •NlNuclear •Electromagneti c (R ail) •Space Tug •Elevator •Beamed Energy •Space Platforms and Towers •Advanced Chemical •Gravity Modification and Breakthrough Physics •Hypersonic Air Breathing •Launch Assist • Analysis perffformed for advanced concepts (15-50)0 years) is not su fficiently accurate for more than semi-qualitative comparisons. • Qualitatively consider known missions: microsat to LEO and large comsat to GEO. Distribution A: Approved for public release; distribution unlimited. 8 Existing State of the Art • Advanced launch concept must be more than just a new solution. • Must yield system level performance improvements over SOA. Microsat to LEO Large Comsat to GEO Orbital Minotaur IV Boeing Delta IV Heavy •Reduces microsat launch costs by •Developed as part of EELV program. reusing Peacekeeper boosters. •Reduce costs by 25%. •4 stage all solid propellant rocket. •Increase simplicity and reliability. •First flight scheduled for Dec. 2008. •Increase standardization. •7 successful Minotaur I flights… •Decrease parts count. •Stage 1: 3 CBCs RS-68 (LH2/LO2). •Stage 2: 1 RL -10B-2 (LH2/LO2) . •First flight Nov. 20, 2002. Performance: Performance: •Thrust: I: 2.2MN, II: 1.2MN, III: .29MN. •Stage 1: Sea Level: 8.673MN @ 410s •1750kg to LEO. •Stage 2: At Altitude: 110kN @ 462s •Minotaur I ~ $30,000/kg. •22,950 kg to LEO. •~$10, 000/kg. •“Advanced Concepts” have not aided most recent generation! Distribution A: Approved for public release; distribution unlimited. 9 Launch Costs • Technologically feasible to launch 130,000kg to LEO (Ares V). • What else is important? •Isp: Propppellant cost represents small fraction of overall… • Responsiveness: Years/months Æ Weeks/days? • Cost($/kg): Limitation on type and amount of payload. • Major focus on reducing launch cost (1/10). • Improved performance (STS): Not successful. • Reduced performance (EELV): Not quite successful. Distribution A: Approved for public release; distribution unlimited. 10 Other Considerations • Reliability: Likelihood that launch vehicle will perform as expected and deliver payload into required orbit. •Typically 0.91-0.95 (Sauvageau,,) Allen JPC 1998). • 2/3 due to propulsion elements. • Upper stages less reliable. • Increasing decreases insurance costs, improve RLV competitiveness. • Availability: Fraction of desired launch dates that can be used. • Responsiveness: Time from determination of desired launch to actual launch. • Currently measured in months/years. • DtDesert Storm: Sep t. 1990 Æ Launc h Fe b. 1992! • Ideal to have weeks/days/hours capability. • EtExtreme MitdMagnitudes 2 2 • SSME: P=6GW dthroat=600cm Æ 10MW/cm . • Saturn V: Height: 116m, Diameter: 10m, Mass: 6.7 million pounds. Distribution A: Approved for public release; distribution unlimited. 11 Propellant: Nuclear • Nuclear materials have extremely high energy densities. • Fission: 7 x 1013 J/kg at 100% efficiency. • Fusion: 6 x 1014 J/kggy at 100% efficiency. •~107 –108 J/kg Æ chemical •Benefit practical launch systems? Nuclear powered His tory upper stage •Nuclear fission rockets first proposed in the late 1940s. •Variety of concepts exist with Isp from 800s to > 5000s. •Typically use hydrogen working gas. •Nuclear propulsion enabling for large interstellar Orion missions. •Launch concepts exist. •NERVA upper s tage. •Primary concerns: system mass, system cost, allowable temperatures, socio-political. •Larggppgpye size limits applications to large payloads. Distribution A: Approved for public release; distribution unlimited. 12 Propellant: Nuclear Tug • Nuclear fission propulsion can enable space tugs. • Reduce the requirements for launch systems? • Example: mtug (no payload) of 22, 000kg, ΔV = 4. 178km/s. Where is breakeven? 200000 Finert = 010.1 180000 Finert = 0.3 160000 Finert = 0.5 Finert = 0.7 140000 (kg) 120000 O EE G to 100000 80000 Payload 60000 40000 20000 0 0 500 1000 1500 2000 2500 3000 3500 4000 4500 Specific Impulse (sec) Significant investments required to reduce specific mass of nuclear systems. Distribution A: Approved for public release; distribution unlimited. 13 Propellant: Laser Beamed Energy • Chemical Propulsion: energy and ejecta same material (neither fully optimized). • Beamed Propulsion: energy stored remotely so ejecta could be optimized. • Lasers and microwaves are both proposed for beamed energy launch. • Both lasers and microwave sources are under continuous development. • More emphasis on laser propulsion. • Laser propulsion was first introduced by Kantrowitz in 1972 1. Heat Laser Æ heat exchanger Æ flow Exotic heat exchangers are required. Exchange 2. Plasma Form plasma in a nozzle to reach high Have high accuracy pointing requirements. Formation operating temperatures. 3. Laser Removal and acceleration of pppropellant via More thrust than PLT,,y but must carry Ablation laser ablation. propellant. 4. Photon Pressure from photons directly used for Bae’s PLT has shown 3000x amplification. Pressure propulsion. Still requires higher powered lasers. Generation: 1MW Æ 1GW Laser beamed propulsion will take significant money to develop and deploy and will only service μSat launches in foreseeable future due to required power levels. Distribution A: Approved for public release; distribution unlimited. 14 Propellant: μwave Beamed Energy Source: Parkin and Culick (2004): • 300 gyrotron sources (140GHz,1MW) Æ 1000kg to LEO. • Transmission: Frequency very important. • Atmospheric Propagation. • Breakdown. • Coupling Efficiency. • Generator Size. •Coupling • Plasma FtiFormation • (Oda et al, 2006) Gas discharge formed at focus of beam. Plasma absorbs beam energy. • Heat Exchanger • (Parkin and Culick) Heat exchanger & hydrogen propellant yield 1000s, payload mass fraction 5- 15%. • Both laser & microw av e beamed energy propu lsion sy stems requ ire significant sou rce (>1GW) and coupling development to yield viable systems for microsatellite launches. • Overlap with other source applications. Distribution A: Approved for public release; distribution unlimited. 15 Microwave Thermal LV Parkin, 2006 16 Propellant: HEDM • Performance of chemical rocket is critically dependent on propellant properties. 1000.0 10000 R ⎛ m ⎞ T 900.0 ⎜ i ⎟ 8000 R ΔV = I sp g ln I sp ∝ 6000 R ⎜ ⎟ 800.0 m f m ⎝ ⎠ 700.0 • Problem: High I typically low density. 600.0 sp 500.0 Impulse (sec) • Goal: Find high Isp,,ypp density propellant cc 400.0 1. Strained ring hydrocarbons. 300.0 Specifi 2. Polynitrogen 200.0 Theoretical Isp Gamma = 1.15 3. Metallic Hydrogen (216MJ/kg). 100.0 P1/P2 = 750 000.0 0 5 10 15 20 25 30 Exhaust Molecular Weight Difficulties • Molecules containingggp high potential ener gy are t ypyypically less stable. • Dramatically more expensive (difficult to manufacture, less alternative uses). • Require new nozzle materials/techniques. •Wide range of potential materials yielding both near-term and far-term potential improvements, but with similar technological challenges: less stable, higher operating temperatures. Distribution A: Approved for public release; distribution unlimited. 17 Propellant: Hypersonic Air Breathing Vehicles • Oxidizer mass fraction >> payload mass fraction for existing launch systems (30% vs. 1.2% for STS). • Can atmosppygheric oxygen be used instead? Thrust-to-Weegight •SSME: 73.12 •Scramjet ~ 2 • Alternative technologies show significantly higher Isp, but over a limited range of Mach number. • Multi-stage systems are required. • Parallel systems suffer from volume and mass constraints.
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