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. Drake, et al.
16 Ad Astra Rocket Company Performance of Chemical System
• for the chemical system, it is assumed that there are separate stages for the TMI and MOI burns.
• for the chemical system, there is a significant TMI S1 TMI S2 MOI PL performance gain from a two-stage TMI burn. Tankage (17%) and stage mass ratio (equal) here are based on NASA Mars DRA 2007. propellant 26.9 26.9 6.8 mass (mT) 29 stage mass 31.5 31.5 1.2 (mT)
2007 DRA: 27% of IMLEO 2 100 vTMI1 (450s)(9.8m / s )ln 1,382 m / s 100 26.9
2 68.5 vTMI 2 (450s)(9.8m / s )ln 2,199 m / s 68.5 26.9 • without accounting for the solar arrays and other mass needed for ZBO cryogenic storage, or the RCS propulsion needed for LEO vTMI 3,581 m / s maneuvers and mid-course corrections, the chemical system can deliver 29 mT of payload 2 37 vMOI (450s)(9.8m / s )ln to a 1-sol orbit about Mars. 37 6.8 896 m / s 17 Ad Astra Rocket Company Performance of Nuclear Thermal System
• tankage for the LH2 propellant is much higher (32%) than for the chemical system’s LOX/LH2 propellant; percentage from NASA Mars DRA 2007.
• for the NTR system, it is assumed that the same engine is used for the TMI and MOI burns, but the tank for the TMI burn is dropped after the TMI burn.
• although approximated here as a single impulsive event, an actual NTR TMI maneuver would be done as two burns separated by an elliptical orbit, to minimize gravity losses due to the gain in altitude during the long (~40 minute) NTR burn.
NTR mass: 6 mT TMI PL MOI (drop tank)
propellant 5.4 32.5 mass (mT) 50.0 –6.0 = 44 2007 DRA: 42% of IMLEO wet tank mass 7.1 42.9 (mT)
• without accounting for the mass needed for ZBO
2 100 cryogenic storage, or the RCS propulsion needed vTMI (925s)(9.8m / s )ln 3,563 m / s 100 32.5 for LEO maneuvers and mid-course corrections, or the additional hydrogen that is run through the reactor to cool it after each burn, the NTR system 57.1 2 can deliver 44 mT to a 1-sol orbit about Mars vMOI (925s)(9.8m / s )ln 900 m / s 57.1 5.4 (assuming NTR mass is same fraction of IMLEO as in DRA). 18 Ad Astra Rocket Company Accuracy of Sphere of Influence (SOI) Approximation
• note that velocity with respect to Earth is set to zero at departure from SOI. • direct integration, for the case below, shows that SOI approximation neglects 870 m/s of velocity relative to the Earth at SOI distance. • vehicle velocity (and specific mechanical energy) at SOI (900,000 km) depend on ratio of vehicle mass to power, but is roughly 1,000 m/s for the ratios considered here.
R = SOI
1 MW 5000 s Earth Departure
Moon’s orbit
19 Ad Astra Rocket Company IMLEO: 100 mT, LEO = 400 km 500 kW Solar Mission to Mars, Thrust // Velocity The power supply is solar, 500 kW at 1 AU. Isp = 5000 s (const), efficiency: 60% When vehicle arrives, its velocity does not precisely match 1) Spiraling from LEO to SOI that of Mars. Thrust schedule to achieve this is more takes 623 days (no shadowing), complicated, but propellant and transit time are nearly 13 mT of propellant, with thrust unchanged. along V
2) Helio-transfer takes 744 days, 11 mT of propellant, Thrust parallel to velocity.
3) Spiraling down to 1-sol orbit about Mars takes 201 days, 2 mT of propellant, with thrust opposite V 1/18/2011 12 20 Ad Astra Rocket Company 500 kW Solar Mission to Mars IMLEO: 100 mT, LEO = 400 km The power supply is solar, 500 kW at 1 AU. Isp = 5000 s (const), efficiency: 60% (Copernicus: guidance modified so that vehicle matches
Mars’ orbital velocity upon arrival) 1) Spiraling from LEO to SOI takes 623 days (no shadowing), 13 mT of propellant, with thrust along V
2) Helio-transfer takes 744 days, 11 mT of propellant
3) Spiraling down to 1-sol orbit about Mars takes 201 days, 2 mT of propellant, with thrust opposite V 1/18/2011 13 21 Ad Astra Rocket Company IMLEO: 100 mT, LEO = 400 km 1 MW Solar Mission to Mars The power supply is solar, 1 MW at 1 AU. Isp = 5000 s (const), efficiency: 60%
1) Spiraling from LEO to SOI takes 314 days (no shadowing), 14 mT of propellant, with thrust along V
2) Helio-transfer takes 399 days, 11 mT of propellant
3) Spiraling down to 1-sol orbit about Mars takes 101 days, 2 mT of propellant, with thrust opposite V
1/18/2011 14 22 Ad Astra Rocket Company IMLEO: 100 mT, LEO = 400 km The power supply is solar, 2 MW at 1 AU. 2 MW Solar Mission to Mars Isp = 5000 s (const), efficiency: 60%
1) Spiraling from LEO to SOI takes 160 days (no shadowing), 14 mT of propellant, with thrust along V
2) Helio-transfer takes 285 days, 16 mT of propellant
3) Spiraling down to 1-sol orbit about Mars takes 51 days, 2 mT of propellant, with thrust opposite V
1/18/2011 15 23 Ad Astra Rocket Company IMLEO: 100 mT, LEO = 400 km The power supply is solar, 6 MW at 1 AU. 6 MW SEP Mission to Mars Isp = 5000 s (const), efficiency: 60% Total time 239 days - less than Hohmann transfer 1) Spiraling from LEO to SOI time for chemical and NTR. takes 56 days (no shadowing), 14 mT of propellant, with thrust along V
2) Helio-transfer takes 166 days, 28 mT of propellant
3) Spiraling down to 1-sol orbit at Mars takes 17 days, 2 mT of propellant, with thrust opposite V
1/18/2011 16 24 Ad Astra Rocket Company 5000 s SEP Transfers from 400 km LEO to Mars 1‐sol Orbit
IMLEO = 100 mT, all cases LEO Power [MW] = 0.5 0.75 1 1.5 2 2.5 6
Initial Mass / Power [kg/kWe] = 200 133 100 67 50 40 17
LEO-ESOI ascent spiral time [days] = 623 411 314 211 160 128 56
Heliocentric Transfer time [days] = 744 467 399 327 285 257 166
Mars descent spiral time [days] = 201 134 101 67 51 41 17
Total Flight time to 1-sol orbit [months] = 52 34 27 20 17 14 8
LEO-ESOI ascent spiral propellant [mT] = 13 13 14 14 14 14 14
Heliocentric transfer propellant [mT] = 11 11 11 13 16 18 28
Mars descent spiral propellant [mT] = 2 2 2 2 2 2 2
Propellant Total [mT] = 26 26 27 29 32 34 44
17% Tankage [mT] = 4.4 4.4 4.6 4.9 5.4 6.0 7.5
Propellant + Tankage [mT] = 30.4 30.4 31.6 33.9 37.4 40 51.5
Array + Thruster : 20/15 /10 kg/kW [mT] = 10/7.5/5 15/ 11.3 / 7.5 20/ 15 / 10 30/ 23 / 15 40/ 30 / 20 50/ 38 / 25 120/ 90 / 60
Mass remaining for payload + S/C [mT] = 60/ 62 / 65 55/ 58 / 62 48/ 53 / 58 36/ 44 / 51 23/ 33 / 43 10/ 22/ 35 0/ 0/ 0
sweet spot? • ‘asymptotic limit’ of heliocentric propellant mass at 5000 s is 11 mT not possible for these as power is reduced from 6 MW to 1 MW, both propellant mass and mass of propulsion system drop values of alpha reducing power below 1 MW doesn’t reduce propellant mass; it only increases the transit time •mass penalty of high alpha P&P becomes negligible at high values of initial mass / power ratio •a ratio of initial mass / power of 100 appears to be a good compromise between payload performance and transit time for cargo flights 25 Ad Astra Rocket Company Estimated Performance Range within Five Years
• below 1 MW for a 100 mT vehicle, transit time grows rapidly •a ratio of initial vehicle mass to power of 100 kg per kW appears to offer the best compromise between ‘payload’ = IMLEO – (propellant + tankage + array + thrusters) transit time and payload
• possible arrays: FAST 7 kg/kW, SOLAROSA 2 kg/kW • rad-shielding impact on array ?
• at present, thruster alphas - VASIMR, Hall, ion - are approximately 10 kg/kW • P+P alpha of 20 kg/kW appears easily achievable with near-term technology (green curve) • 15 kg/kW seems plausible with aggressive development (blue) • 10 kg/kW may be achievable depending on thruster mass scaling with power • an “SEP Freighter” would dramatically amplify the interplanetary payload capacity of present and future launch vehicles 26 Ad Astra Rocket Company Performance of Solar Electric System
• tankage for argon propellant is assumed to be the same as the chemical system’s LOX/LH2 propellant: 17%
• case illustrated here is for 1 MW (@ 1 AU) system with combined specific mass of power + propulsion = 10 kg/kW
• thruster specific impulse = 5,000 s; efficiency = 60%
SEPsolar mass: array 10 mT
Ar tank PL
Element mass Solar Array & Tank & Prop 32 Payload 58 (mT) Engines 10
• without accounting for the mass needed for ZBO cryogenic storage, or the RCS propulsion • propellant mass calculated from Copernicus simulations. needed for LEO maneuvers and mid-course corrections, the NEP system can deliver 58 mT to a 1-sol orbit about Mars.
27 Ad Astra Rocket Company Mars Cargo Propulsion Systems: Comparison
Payload from 400 km LEO Specific Impulse Total Flight Time 100 mT IMLEO System to 1‐sol Mars orbit (seconds) (months) (mT)
Chemical
450 29 9
Nuclear Thermal
925 44 9
1 MW Solar Electric
48 –58 5000 27 ( = 20, = 10)
28 Ad Astra Rocket Company Ratio of IMLEO to Power Determines Payload Fraction
• It is only the ratio of initial vehicle mass to power that determines the fraction of IMLEO that arrives at Mars
• matching the “sweet spot” at 100 kg/kW to the IMLEO capability of the Atlas V 551 (20,000 kg), the corresponding array and thruster power level is: 20,000 kg / (100 kg/kW) = 200 kW 48 %
• assuming the present-day array and thruster alpha of 20 kg/kW, this would result in 48 % of the IMLEO (i.e. 27 mo. approximately 10 tons) arriving in Mars orbit as ‘payload’ (mass in addition to the array and thrusters) in 27 months
• Mars planetary science missions may find the additional transit time acceptable in exchange for doubling the payload • SEP vehicle’s dominant cost element would be the array; this is roughly • note performance relative to $500/watt x 200 kW = $100M chemical and NTR Hohmann transfers (for concentrator arrays) 29 Ad Astra Rocket Company Conclusions
• Ad Astra has demonstrated 60% propulsion system efficiency (DC to jet power) at an Isp of 4800 seconds and an input power of 220 kW DC • The ratio of mass to power (alpha) of solar arrays and thrusters establishes the payload fraction achievable for SEP • Concentrator solar arrays are essential to achieving alpha values low enough for SEP to be attractive for unmanned Mars flights • In the near term (five years?), a combined array + thruster alpha of 20 kg/kW seems within reach, at power levels on the order of hundreds of kW • For this value of alpha, SEP payload fraction is comparable to that for nuclear thermal rocket propulsion, though with a transit time three times longer (27 months) • The estimated cost per kW for concentrating solar arrays ($500/watt) makes SEP an attractive enhancement for near-term Mars science missions, approximately doubling the Mars payload capability of existing launch vehicles, at the expense of longer flight time
30 Ad Astra Rocket Company Back-up Slides
31 Ad Astra Rocket Company Performance Variation with Actual Orbits of Earth and Mars
Mission Requirements and Assumptions for Copernicus Solutions:
• Departure year between 2015 and 2030
• Departure from Earth SOI; Vdep = 0 km/s • Initial Mass 85 mT at Earth SOI • (IMLEO 100 mT; 15 mT propellant used in spiral from LEO to SOI; 2 mT used in spiral down to 1-sol Mars orbit) • Solar Power 1 MW @ 1 AU
• Propulsion System Efficiency (DCe to jet power) 60 %
• Arrival at Mars, Varr = 0 km/s • Fixed Specific Impulse 5,000 sec • Coasting time allowed
• Optimization goal: minimize propellant consumption (Copernicus optimization)
• Question: with this type of propulsion, how does the heliocentric trip time and propellant consumption vary with Mars’ orbital position at S/C arrival?
1/18/2011 32 Ad Astra Rocket Company 1) Arrival at Mars perihelion
73.7 % of IMLEO is placed into 1-sol orbit about Mars
2. Thrust for 130 days 3. Coast for 132 days Propellant used: 5.3 mT
1. Depart from Earth SOI Aug 21, 2026
Mars perihelion 4. Thrust for 204 days Propellant used: 4 mT 5. Arrive at Mars after 466 days Total propellant used: 9.3 + 15 + 2 = 26.3 mT 33 Ad Astra Rocket Company 2) Arrival at Mars aphelion
5. Arrive at Mars SOI after 413 days Total propellant used: 9.7 + 15 + 2 = 24.7 mT 4. Thrust for 249 days 75.3 % of IMLEO is Propellant used 5.2 mT placed into high orbit about Mars
1. Depart from Earth SOI March 21, 2018
2. Thrust for 111 days Propellant used: 4.5 mT 3. Coast for 53 days • 400 kg more propellant than perihelion arrival case (4% difference in propellant used) Mars perihelion • 413 days vs 466 for perihelion case (11% difference in heliocentric transfer time) 34 Ad Astra Rocket Company 3) Arrival 90° from Mars perihelion (intermediate case)
73.8 % of IMLEO is placed into 1-sol orbit 3. Coast for 144 days about Mars
4. Thrust for 149 days Propellant used 2.4 mT
1. Depart from Earth SOI 2. Thrust for 184 days June 18, 2022 Propellant used 6.8 mT
Mars perihelion 5. Arrive at Mars SOI after 477 days propellant used: 9.2 + 15 + 2 = 26.2 mT 35 Ad Astra Rocket Company Summary of variation wrt Mars’ orbital position
Heliocentric Orbital Position of Heliocentric Propellant Departure Date Mars at S/C Transfer Time Consumption Arrival (days) (mT)
August 21, 2026 perihelion 466 9.3
angular position mid-way between June 18, 2022 477 9.2 aphelion and perihelion
March 21, 2018 aphelion 413 9.7
• assuming that the thrust vector schedule determined by Copernicus yields a global minimum for propellant consumption, the results indicate that the variations in heliocentric transfer propellant (5% of mid-range) and time (7% of mid-range) are relatively insensitive to the orbital longitude of Mars at S/C arrival.
• architecture significance: a single vehicle design capable of delivering a fixed payload to Mars requires only a small margin in tankage to accommodate the full range of Mars’ orbital position at S/C arrival. 36 Ad Astra Rocket Company SEP as an Extended TMI Burn
• Propellant for the heliocentric transfer can be significantly reduced if we remove the requirement to match Mars’ orbital velocity vector when arriving at Mars, and assume that the payload can either make a chemical propulsive burn or use aerocapture for MOI, or make a direct entry and landing on Mars, as landing probes have done in the past.
• To evaluate possible enhancements to existing chemical systems, we note that the Atlas V 551 can inject approximately 5000 kg into TMI.
• The Atlas V can put 20,500 kg into LEO.
v = 3.565 km/s, C3 = 13.7 km2/s2 37 Ad Astra Rocket Company