VASIMR VX-200 Performance and Near-term SEP Capability for Unmanned Mars Flight Future In-Space Operations Seminar January 19, 2011 presented by Tim Glover Director of Development [email protected] Ad Astra Rocket Company www.adastrarocket.com 1 Ad Astra Rocket Company Notes and Acronyms Notes: - solar array power values are for 1 AU - a number of publications on VASIMR R&D are available on the company’s website: http://www.adastrarocket.com Acronyms: SEP solar electric propulsion NTR nuclear thermal rocket VASIMR Variable Specific Impulse Magnetoplasma Rocket TMI Trans-Mars Insertion MOI Mars Orbit Insertion SOI sphere of influence IMLEO initial mass in low Earth orbit 2 Ad Astra Rocket Company Outline 1. VASIMR Prototype Performance 2. Simplified Earth‐Mars Trajectories 3. Chemical and NTR Hohmann Transfers 4. SEP: Initial Mass‐to‐Power Ratio and Payload Fraction 5. Near‐term SEP Mars Capabilities 6. Backup slides: • Propellant and transit time variation for actual orbits of Earth and Mars •Atlas V 551 performance curve 3 Ad Astra Rocket Company My Background • B.S., Physics, New Mexico State U. • M.S., Aerospace Engineering (Orbital Mechanics), UT-Austin • M.S., Physics, University of Pittsburgh • five years teaching high school physics, Ethical Culture Schools, New York • Ph.D., Applied Physics, Rice University, 2002 − thesis research at Johnson Space Center: designed and built plasma diagnostics to measure exhaust velocity in early VASIMR prototypes • Research Scientist, MEI Technologies 2003-2005 − continued experimental work on VASIMR up to 50 kW • Director of Development, Ad Astra Rocket Company 2005 – present − business development, external relations 4 Ad Astra Rocket Company VASIMR Operating Principles Superconducting Magnets typical path of an ion Helicon coupler (30 kW) through the rocket Ion cyclotron (170 kW) coupler i gas cold accelerated plasma plasma POWER 1. Ionize 2. Energize 3. Accelerate 4. Detach first stage second stage Caption: Argon gas is fed into the first stage of the VASIMR® engine, which converts the argon from a neutral gas to a fully ionized plasma. At this stage, the plasma is described as “cold” . Although the electrons are hot (10,000 degrees celsius or more), they constitute less than 0.1% of the mass of the plasma; the ions are roughly at room temperature (relatively cold) and make up more than 99.9% of the mass. Once the plasma leaves the first stage, it flows along the magnetic field lines produced by the superconducting magnets and makes no further contact with solid parts of the engine. As the plasma flows through the second stage, the ions are accelerated to much higher energies than the electrons by waves created in the plasma by the ion cyclotron coupler. Due to the strong magnetic field in the second stage, the ions are confined to move in tight helical orbits whose diameter is much smaller than the diameter of the plasma column. As the plasma flows downstream, the nozzle action of the expanding magnetic field accelerates the plasma to a high exhaust velocity. Like the plasma leaving the Sun in a Coronal Mass Ejection, the VASIMR exhaust plasma separates from the magnetic field of the rocket as it accelerates through the magnetic nozzle to high speed. 5 Ad Astra Rocket Company VX-200 Testing 6 Ad Astra Rocket Company VX-200 in Ad Astra Test Facility top view VX-200 A motion table can move the diagnostics through the exhaust plume over an axial range of 500 cm and a diameter range of 200 cm. Exhaust plume diagnostics are located side view through port downstream from the VX-200 exit plane. 7 Ad Astra Rocket Company Exhaust Plasma Diagnostics and Translation Stage Table RPA Force Impact Target #1 Force Impact Target #2 3-axis Magnetometer Langmuir Probe Ion Flux Probe Array #1 Ion Flux Probe Array #2 8 Ad Astra Rocket Company Thrust Target Measurements • We take diameter scans (31 shots, shot to shot in this case) looking at force density directly • Integrate the force density to get a total force number • This method has been found to yield identical results to a thrust stand, by taking simultaneous measurements on a Hall thruster at University of Michigan 9 cm Corresponding size of target 9 1/18/2011 Ad Astra Rocket Company VX-200 Performance in Nov. 2010 Experiment Campaign • In the Thrust graph above, first stage power • The Efficiency graph is derived from the previous Thrust is held fixed at 30 kW for optimal ionization of graph; it does NOT represent VASIMR efficiency as a argon propellant at a flow rate of 130 mg/s, function of exhaust velocity except at the nominal operating and second stage power is raised from zero to point for this propellant flow rate, which is 50 km/s. 170 kW. • a single 400 V power supply powers both stages of the rocket; thruster efficiency is calculated as the jet power in The VX-200 prototype data gives us the plume (determined by the thrust target), divided by the measured power output of the RF amplifiers. confidence that we can build a single 200 kW thruster that achieves 70% thruster • measured DC to RF power conversion efficiency is 95%, efficiency at an Isp of 5000 seconds. leading to a system efficiency of 60% (DC to jet power) for 50 km/s (5000 s Isp) operation. 10 Ad Astra Rocket Company Performance of Chemical, NTR and SEP Systems • to evaluate the performance of SEP relative to chemical and NTR propulsion for Mars cargo/unmanned flights, we consider a simple hypothetical space flight • circular, coplanar orbits for Earth and Mars are used to eliminate variations due to the eccentricity and relative inclination of the planets’ orbits • while not accurate for actual flight studies, this comparison illustrates the relative merits of the three types of propulsion, and estimates the technological requirements needed for SEP to play a useful role 11 Ad Astra Rocket Company A Simplified Problem Relevant to Mars Flights The blue and red circles in the figure at right represent the orbits of Earth and Mars, respectively. Although Mars’ actual orbit is an ellipse, it is represented here by a circle with the same average distance from the Sun as Mars, to remove the effect of launch date on the trajectories. This circular approximation allows us to compare propulsion systems and trajectories for a problem that is independent of launch date and phasing of the planets. This simple situation will provide insights that will be helpful in selecting appropriate SEP vehicle parameters for cargo flights using the actual orbits of the Earth and Mars. To estimate the performance of two propulsion systems considered for Mars flights – chemical and nuclear thermal rockets - we calculate the maximum mass that can be transferred from the Earth’s orbit to an orbit about Mars with two impulsive burns. At Mars, we assume that the cargo vehicle enters a 1-sol Mars orbit. To minimize the delta-v required, this Mars Orbit Insertion (MOI) burn is done at an altitude of only 250 km above the surface of Mars, and results in a highly elliptical orbit with a semi-major axis of 20,000 km. (This is the assumption made in NASA’s Design Reference Architecture for the crewed vehicles envisioned to support human exploration of Mars – see backup slides). A Hohmann maneuver. The transit between the circular approximations to the orbits of Earth and Mars takes 260 For comparison, we calculate the performance of solar days. In the “all-propulsive” case, two burns are necessary. electric propulsion for fixed specific impulse, over a range of If aerocapture at Mars is presumed feasible, only a single power levels. These trajectories have been calculated using burn in LEO is needed. the Copernicus software package, for artificial planets in circular, co-planar orbits having radii equal to the semi-major axes of the orbits of Earth and Mars. 12 Ad Astra Rocket Company Hohmann Transfer Assumptions: • 400 km altitude initial low Earth orbit • 100 mT vehicle in LEO (wet mass) • chemical Isp = 450 s (LOX/LH2) • NTR Isp = 925 s (based on Peewee reactor, final NERVA variant) Transfer orbit: • semi-major axis = 1.89 x 1011 m • transfer time (1/2 period) = 260 days • perihelion velocity = 32,670 m/s • aphelion velocity = 21,490 m/s Two maneuvers: • Trans-Mars Insertion (TMI) burn modeled as a single impulse in the 400 km LEO parking orbit (v1) • Mars Orbit Insertion (MOI) burn modeled as a single impulse performed at the 250 km altitude periapsis of the 1-sol (0.82 eccentricity) Mars parking orbit (v2) 13 Ad Astra Rocket Company TMI Burn for Chemical and NTR • 400 km altitude circular LEO velocity (pre- TMI burn) Excess velocity needed for = 7,676 m/s Hohmann transfer (perihelion) • transfer orbit’s perihelion velocity = 32,670 m/s • Earth’s orbital velocity = 29,750 m/s Earth’s sphere of influence • excess velocity needed from TMI burn: 32,670 – 29,750 = 2,920 m/s at Earth SOI • this corresponds to a post- TMI burn velocity in the 400 km altitude parking orbit 400 km parking orbit of: 2 2 v vLEO GM E E 2 2 rLEO 2GM v v2 E 11,241 m/s v1 LEO rLEO • v1 = 11,240 – 7,676 = 3,565 m/s 14 Ad Astra Rocket Company MOI Burn for Chemical and NTR • transfer orbit’s aphelion velocity = 21,490 m/s • Mars orbital velocity = 24,130 m/s • Mars approach velocity: 24,130 – 21,490 = 2,640 m/s at Mars SOI v2 • this corresponds to a pre-MOI burn velocity at the 250 km periapsis of the 1-sol km parking orbit of: v2 v2 GM E 250 M 2 2 r250 2 2GM M v250 v = 5,513 m/s r250 • 250 km periapsis velocity (post- TMI burn): 1‐sol orbit 1sol GM M ra vperi = 4,619 m/s a rp Mars approach velocity • v2 = 5,513 – 4,619 = 894 m/s 15 Ad Astra Rocket Company Validity of Circular Approximation TMI 3,565 m/s MOI 894 m/s from NASA DRA 2007, B.