SpaceX Propulsion

1 SpaceX Propulsion Tom Markusic Space Exploration Technologies 46th AIAA/ASME/SAE/ASEE. Joint Propulsion Conference July 28, 2010. Friday, August 6, 2010. SpaceX Propulsion Tom Markusic Space Exploration Technologies 46th AIAA/ASME/SAE/ASEE. Joint Propulsion Conference July 28, 2010. Friday, August 6, 2010. Overview Inverse Hyperbolic Bessel Functions Friday, August 6, 2010. Near-term Propulsion Needs Friday, August 6, 2010. Near-term Propulsion Needs HLLV Propulsion Merlin 2 uses scaled-up, flight proven Merlin 1 design J-2X SpaceX can develop and flight qualify the Merlin 2 engine in ~3. years at a cost of ~$1B. Production: ~$50M/engine J-2X development already in progress under Constellation Merlin 2 J-2X. program Propellant LOX/RP LOX/LH2. Merlin 2 Thrust (vac) [klbf] 1,700 292. Isp (vac) [sec] 322 448. T/W [lbf/lbm] 150 55. Friday, August 6, 2010. Near-term Propulsion Needs HLLV Propulsion Solar Electric Propulsion for Cargo Tug Merlin 2 uses scaled-up, flight Cluster of ~5 high TRL thrusters proven Merlin 1 design NEXT process 100 kWe solar power J-2X SpaceX can develop and flight Ion Thruster Next generation tug uses single qualify the Merlin 2 engine in ~3 high power thruster, such as NASA.

2 Years at a cost of ~$1B. 457M. Production: ~$50M/engine Third generation tug uses nuclear J-2X development already in electric Propulsion at megawatt progress under Constellation levels NEXT BHT-20 457M. Busek BHT-20K k Merlin 2 J-2X Propellant Xenon Xenon Xenon program Hall Thruster Propellant LOX/RP LOX/LH2 Power [kWe] 7 20 96. Merlin 2 Thrust (vac) [klbf] 1,700 292. Thrust [mN] 236 1080 3300. Isp (vac) [sec] 322 448. NASA 457M Isp [sec] 4100 2750 3500. T/W [lbf/lbm] 150 55 Hall Thruster Efficiency [%] 70 72 58. Friday, August 6, 2010. Near-term Propulsion Needs HLLV Propulsion Solar Electric Propulsion for Cargo Tug Merlin 2 uses scaled-up, flight Cluster of ~5 high TRL thrusters proven Merlin 1 design NEXT process 100 kWe solar power J-2X SpaceX can develop and flight Ion Thruster Next generation tug uses single qualify the Merlin 2 engine in ~3 high power thruster, such as NASA.

3 Years at a cost of ~$1B. 457M. Production: ~$50M/engine Third generation tug uses nuclear J-2X development already in electric Propulsion at megawatt progress under Constellation levels NEXT BHT-20 457M. Busek BHT-20K k Merlin 2 J-2X Propellant Xenon Xenon Xenon program Hall Thruster Propellant LOX/RP LOX/LH2 Power [kWe] 7 20 96. Merlin 2 Thrust (vac) [klbf] 1,700 292. Thrust [mN] 236 1080 3300. Isp (vac) [sec] 322 448. NASA 457M Isp [sec] 4100 2750 3500. T/W [lbf/lbm] 150 55 Hall Thruster Efficiency [%] 70 72 58. Nuclear Thermal Propulsion for Mars Stage 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 >17. Hours of hot-fire tests, including restarts. No additional developmental, terrestrial tests (with nuclear) fuel are required. Extensive Russian knowledge can be leveraged.

4 Friday, August 6, 2010. Near-term Propulsion Needs HLLV Propulsion Solar Electric Propulsion for Cargo Tug Merlin 2 uses scaled-up, flight Cluster of ~5 high TRL thrusters proven Merlin 1 design NEXT process 100 kWe solar power J-2X SpaceX can develop and flight Ion Thruster Next generation tug uses single qualify the Merlin 2 engine in ~3 high power thruster, such as NASA. years at a cost of ~$1B. 457M. Production: ~$50M/engine Third generation tug uses nuclear J-2X development already in electric Propulsion at megawatt progress under Constellation levels NEXT BHT-20 457M. Busek BHT-20K k Merlin 2 J-2X Propellant Xenon Xenon Xenon program Hall Thruster Propellant LOX/RP LOX/LH2 Power [kWe] 7 20 96. Merlin 2 Thrust (vac) [klbf] 1,700 292. Thrust [mN] 236 1080 3300. Isp (vac) [sec] 322 448. NASA 457M Isp [sec] 4100 2750 3500. T/W [lbf/lbm] 150 55 Hall Thruster Efficiency [%] 70 72 58.

5 Nuclear Thermal Propulsion for Mars LOX/Methane Propulsion for Ascent/Desc Stage NERVA derived technology ISRU-derived methane will be used for ascent/descent Total thrust ~ 60 klbf, using 2 to 6 Propulsion NDR Strong developmental programs currently underway at Propellant: hydrogen, Isp ~ 930 sec Aerojet, ATK/XCOR. ISRU or pre-deployed propellant for return SpaceX Merlin 1 engine may be reconfigurable to for LOX/. mission methane, providing a large (~100 klbf) GG cycle engine for Technology has been verified with >17 ascent/descent Hours of hot-fire tests, including restarts. No additional developmental, terrestrial tests (with nuclear) fuel are required. Extensive Russian knowledge can be leveraged. Aerojet, T = k-lbf, Isp = ATK/XCOR, T = k-lbf, Isp 350 sec =? Friday, August 6, 2010. Friday, August 6, 2010. Testing Survey This slide may contain SpaceX proprietary and/or ITAR sensitive content.

6 Friday, August 6, 2010. Friday, August 6, 2010. Raptor Friday, August 6, 2010. HLLV 1st 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 V ( 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 for both cases. Ballistic trajectory. Equations of motion again integrated using assumptions and boundary conditions above, but with gravity and aerodynamic drag included. Friday, August 6, 2010. HLLV 1st Stage Propulsion LOX/RP versus LOX/LH2 Booster Fundamentals Trade Studies Simple 1-D dynamic model used to compare LOX/RP and Recent NASA-led Heavy Lift Launch Vehicle Study . LOX/LH2 first stage performance for a HLLV compared many configurations of LOX/LH2, LOX/RP, SRB.

7 First, for both propellants, propellant mass was Propulsion for a HLLV. chosen to yield the same V ( km/s) for a given Configuration with 6 Lox/RP engine first stage payload ( 750 MT), consistent with Saturn V, but competitive with all concepts in performance and with no external forces. mission capture metrics Typical engine performance and tank mass Configuration with 6 Lox/RP engine first stage fractions assumed. shown to provide benefits in safety and annual Initial T/W fixed at for both cases. Ballistic Operations recurring cost metrics above all LOX/LH2 and SRB. trajectory. Handling. configurations Equations of motion again integrated using Deep cryogenic (-432 F) vs room temperature for RP. assumptions and boundary conditions above, but LH2 has high infrastructure investment for test and launch with gravity and aerodynamic drag included. Safety. LH2 leaks lead to detonation risk extensive monitoring required RP leaks are easily (visually) detectable, low explosion risk Friday, August 6, 2010.

8 HLLV 1st Stage Propulsion LOX/RP versus LOX/LH2 Booster Fundamentals Trade Studies Simple 1-D dynamic model used to compare LOX/RP and Recent NASA-led Heavy Lift Launch Vehicle Study . LOX/LH2 first stage performance for a HLLV compared many configurations of LOX/LH2, LOX/RP, SRB. First, for both propellants, propellant mass was Propulsion for a HLLV. chosen to yield the same V ( km/s) for a given Configuration with 6 Lox/RP engine first stage payload ( 750 MT), consistent with Saturn V, but competitive with all concepts in performance and with no external forces. mission capture metrics Typical engine performance and tank mass Configuration with 6 Lox/RP engine first stage fractions assumed. shown to provide benefits in safety and annual Initial T/W fixed at for both cases. Ballistic Operations recurring cost metrics above all LOX/LH2 and SRB. trajectory. Handling. configurations Equations of motion again integrated using Deep cryogenic (-432 F) vs room temperature for RP.

9 Assumptions and boundary conditions above, but LH2 has high infrastructure investment for test and launch with gravity and aerodynamic drag included. Safety. LH2 leaks lead to detonation risk extensive monitoring required RP leaks are easily (visually) detectable, low explosion risk RP staged combustion versus GG. cycle First Stage V simplified model compared Merlin 2 gas generator cycle engine with scaled up RS-84 derived staged combustion engine. Mass of Merlin 2 based on current design (sea level thrust = Mlbf). Mass of RS-84 derived engine estimated by linearly scaling thrust and assuming T/W. is constant. Merlin 2 vac Isp = sec, RS-84 derived vac Isp =. sec. Modeled Falcon X with F9 flight trajectory (250 km x deg). Found burnout velocity for Merlin 2 stage and RS-84. derived stages to be 3526 m/sec and 3527 m/sec, Friday, August 6, 2010 respectively. Dead Sea Scrolls Black water shall elevate thy children to the heavens.

10 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.. Friday, August 6, 2010. Backup Assumptions for Mission and Vehicle Sizing HLLV T/W SEP Isp 2750 s 1st Stage Payload 750 MT SEP thrust per engine N. RP-1 inert mass fraction Xenon tank mass fraction LH2 inert mass fraction SEP structural and margin mass RP-1 Isp 300 s fraction LH2 Isp 420 s Solar Arrays and PPU mass fraction kg/kW. RP O/F ratio Low-thrust Delta V LEO to Phobos km/s LH2 O/F ratio Stage height, excluding 36 m engines NTR Isp 930 s RP-1 GLOM 3040 MT. Delta V LEO to TMI km/s LH2 GLOM 2060 MT. RP-1 Burnout time 177 s Delta V TMI to MOC km/s LH2 Burnout time 205 s Delta V MOC to Phobos Capture km/s RP-1 Stage diameter m NTR 15k lbf-thrust engine mass 2600 kg LH2 Stage diameter m NTR tank mass fraction Earth Aerocapture Delta V savings km/s Friday, August 6, 2010.