SpaceX Propulsion, Summaries of Design

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SpaceX Propulsion

Tom Markusic

Space Exploration Technologies

46th AIAA/ASME/SAE/ASEE

Joint Propulsion Conference

July 28, 2010

SpaceX Propulsion

Tom Markusic

Space Exploration Technologies

46th AIAA/ASME/SAE/ASEE

Joint Propulsion Conference

July 28, 2010

Near-term Propulsion Needs

HLLV Propulsion

J-2X

Merlin 2 Merlin 2 J-2X Propellant LOX/RP LOX/LH 2 Thrust (vac) [klbf] 1,700 292 Isp (vac) [sec] 322 448 T/W [lbf/lbm] 150 55

Near-term Propulsion Needs

• (^) Merlin 2 uses scaled-up, flight proven Merlin 1 design
• (^) SpaceX can develop and flight qualify the Merlin 2 engine in ~ years at a cost of ~$1B. Production: ~$50M/engine
• (^) J-2X development already in progress under Constellation program

HLLV Propulsion

Nuclear Thermal Propulsion for Mars

Stage

Solar Electric Propulsion for Cargo Tug

J-2X

Merlin 2

NEXT

Ion Thruster Busek BHT-20K Hall Thruster NASA 457M Hall Thruster Merlin 2 J-2X Propellant LOX/RP LOX/LH 2 Thrust (vac) [klbf] 1,700 292 Isp (vac) [sec] 322 448 T/W [lbf/lbm] 150 55

Near-term Propulsion Needs

• (^) Merlin 2 uses scaled-up, flight proven Merlin 1 design
• (^) SpaceX can develop and flight qualify the Merlin 2 engine in ~ years at a cost of ~$1B. Production: ~$50M/engine
• (^) J-2X development already in progress under Constellation program NEXT BHT- k 457M Propellant Xenon Xenon Xenon Power [kWe] 7 20 96 Thrust [mN] 236 1080 3300 Isp [sec] 4100 2750 3500 Efficiency [%] 70 72 58
• (^) Cluster of ~5 high TRL thrusters process 100 kWe solar power
• (^) Next generation tug uses single high power thruster, such as NASA 457M
• (^) Third generation tug uses nuclear electric propulsion at megawatt levels
• (^) NERVA derived technology
• (^) Total thrust ~ 60 klbf, using 2 to 6 NDR
• (^) Propellant: hydrogen, Isp ~ 930 sec
• (^) ISRU or pre-deployed propellant for return mission
• (^) Technology has been verified with > Hours of hot-fire tests, including restarts. No additional developmental, terrestrial tests (with nuclear) fuel are required.
• (^) Extensive Russian knowledge can be leveraged.

HLLV Propulsion

Nuclear Thermal Propulsion for Mars

Stage

Solar Electric Propulsion for Cargo Tug

J-2X

Merlin 2

NEXT

Ion Thruster Busek BHT-20K Hall Thruster NASA 457M Hall Thruster Merlin 2 J-2X Propellant LOX/RP LOX/LH 2 Thrust (vac) [klbf] 1,700 292 Isp (vac) [sec] 322 448 T/W [lbf/lbm] 150 55

LOX/Methane Propulsion for Ascent/Desc

Aerojet, T = 5.5 k-lbf, Isp = 350 sec ATK/XCOR, T =7.5 k-lbf, Isp =?

Near-term Propulsion Needs

• (^) Merlin 2 uses scaled-up, flight proven Merlin 1 design
• (^) SpaceX can develop and flight qualify the Merlin 2 engine in ~ years at a cost of ~$1B. Production: ~$50M/engine
• (^) J-2X development already in progress under Constellation program NEXT BHT- k 457M Propellant Xenon Xenon Xenon Power [kWe] 7 20 96 Thrust [mN] 236 1080 3300 Isp [sec] 4100 2750 3500 Efficiency [%] 70 72 58
• (^) Cluster of ~5 high TRL thrusters process 100 kWe solar power
• (^) Next generation tug uses single high power thruster, such as NASA 457M
• (^) Third generation tug uses nuclear electric propulsion at megawatt levels
• (^) NERVA derived technology
• (^) Total thrust ~ 60 klbf, using 2 to 6 NDR
• (^) Propellant: hydrogen, Isp ~ 930 sec
• (^) ISRU or pre-deployed propellant for return mission
• (^) Technology has been verified with > Hours of hot-fire tests, including restarts. No additional developmental, terrestrial tests (with nuclear) fuel are required.
• (^) Extensive Russian knowledge can be leveraged.
• (^) ISRU-derived methane will be used for ascent/descent propulsion
• (^) Strong developmental programs currently underway at Aerojet, ATK/XCOR
• (^) SpaceX Merlin 1 engine may be reconfigurable to for LOX/ methane, providing a large (~100 klbf) GG cycle engine for ascent/descent

This slide may contain SpaceX proprietary and/or ITAR sensitive content.

Testing Survey

HLLV 1

st

Stage Propulsion

LOX/RP versus LOX/LH2 Booster

Fundamentals

• (^) Simple 1-D dynamic model used to compare LOX/RP and LOX/LH2 first stage performance for a HLLV - (^) First, for both propellants, propellant mass was chosen to yield the same ΔV (3.6 km/s) for a given payload ( 750 MT), consistent with Saturn V, but with no external forces. - (^) Typical engine performance and tank mass fractions assumed. - (^) Initial T/W fixed at 1.2 for both cases. Ballistic trajectory. - (^) Equations of motion again integrated using assumptions and boundary conditions above, but with gravity and aerodynamic drag included.

HLLV 1

st

Stage Propulsion

LOX/RP versus LOX/LH2 Booster

Fundamentals

• (^) Simple 1-D dynamic model used to compare LOX/RP and LOX/LH2 first stage performance for a HLLV - (^) First, for both propellants, propellant mass was chosen to yield the same ΔV (3.6 km/s) for a given payload ( 750 MT), consistent with Saturn V, but with no external forces. - (^) Typical engine performance and tank mass fractions assumed. - (^) Initial T/W fixed at 1.2 for both cases. Ballistic trajectory. - (^) Equations of motion again integrated using assumptions and boundary conditions above, but with gravity and aerodynamic drag included.

Trade Studies

• (^) Recent NASA-led “Heavy Lift Launch Vehicle Study” compared many configurations of LOX/LH2, LOX/RP, SRB propulsion for a HLLV. - (^) Configuration with 6 Lox/RP engine first stage competitive with all concepts in performance and mission capture metrics - (^) Configuration with 6 Lox/RP engine first stage shown to provide benefits in safety and annual recurring cost metrics above all LOX/LH2 and SRB configurations

Operations

Handling.

• Deep cryogenic (-432 F) vs room temperature for RP
• LH 2 has high infrastructure investment for test and launch Safety.
• LH2 leaks lead to detonation risk—extensive monitoring required
• RP leaks are easily (visually) detectable, low explosion risk

Dead Sea Scrolls

“Black water shall

elevate thy children to

the heavens. Purify it.

But thou shalt not

combine it in a ratio

greater than one

kikkar to twenty

shekkels, nor shalt

thou burn rocks.

Thus saith the lord.”

• Assumptions for Mission and Vehicle Sizing

Backup

SEP Isp 2750 s SEP thrust per engine 1.08 N Xenon tank mass fraction 0. SEP structural and margin mass fraction

Solar Arrays and PPU mass fraction 3.5 kg/kW Low-thrust Delta V LEO to Phobos 11.2 km/s NTR Isp 930 s Delta V LEO to TMI 4.2 km/s Delta V TMI to MOC 2.5 km/s Delta V MOC to Phobos Capture 0.4 km/s NTR 15k lbf-thrust engine mass 2600 kg NTR tank mass fraction 0. Earth Aerocapture Delta V savings 3.2 km/s

HLLV T/W 1.

1st Stage Payload 750 MT RP-1 inert mass fraction 0. LH2 inert mass fraction 0. RP-1 Isp 300 s LH2 Isp 420 s RP O/F ratio 2. LH2 O/F ratio 5. Stage height, excluding engines 36 m RP-1 GLOM 3040 MT LH2 GLOM 2060 MT RP-1 Burnout time 177 s LH2 Burnout time 205 s RP-1 Stage diameter 8.7 m LH2 Stage diameter 11.3 m