Launch Vehicles Dr Andrew Ketsdever MAE 5595 Lesson

Launch Vehicles Dr Andrew Ketsdever MAE 5595 Lesson 15

Liquid Rocket Engines • Advantages – High Isp – Start / Stop / Restart – Ability to throttle • Disadvantages – Complex • Turbopumps or gas pressurization systems – Lower thrust than solid systems – Lower density storage than solids – Toxic or cryogenic fuels and oxidizers

Solid Rocket Motors • A solid rocket motor is a system that uses solid propellants to produce thrust • Advantages – – High thrust Simple Storability High density Isp • Disadvantages – – Low Isp (compared to liquids) Complex throttling Difficult to stop and restart Safety

Hybrid Rocket Motors ADVANTAGES • SAFETY: Literally no possibility of explosion • Controllable – Throttle – Stop / Re-start • Safe exhaust products • Higher Isp than solids • Higher density Isp than liquids • Lower complexity than liquids • Lower inert mass fraction than liquids DISADVANTAGES • More complex than solids • Lower Isp than liquids • Lower density Isp than solids • Lower combustion efficiency than either liquids or solids • O/F variability • Poor propellant utilization • Higher inert mass fraction than solids

Nuclear Propulsion • ADVANTAGES – High Isp (2 -10 x that of chemical systems) – Low Specific Mass (kg/k. W) – High Power Allows High Thrust – High F/W – Use of Any Propellant – Safety – Reduced Radiation for Some Missions

Saturn V Launch Vehicles

Saturn V • • • Three-stage lunar landing booster. LEO Payload: 118, 000 kg. to: 185 km Orbit. at: 28. 0 degrees. Payload: 47, 000 kg. to a: Translunar trajectory. Liftoff Thrust: 3, 440, 310 kgf. Liftoff Thrust: 33, 737. 90 k. N. Total Mass: 3, 038, 500 kg. Core Diameter: 10. 06 m. Total Length: 102. 00 m. Stage Number: 1. 1 x Saturn IC Gross Mass: 2, 286, 217 kg. Empty Mass: 135, 218 kg. Thrust (vac): 3, 946, 624 kgf. Isp: 304 sec. Burn time: 161 sec. Isp(sl): 265 sec. Diameter: 10. 06 m. Span: 19. 00 m. Length: 42. 06 m. Propellants: Lox/Kerosene No Engines: 5. F-1 Status: Out of Production. Stage Number: 2. 1 x Saturn II Gross Mass: 490, 778 kg. Empty Mass: 39, 048 kg. Thrust (vac): 526, 764 kgf. Isp: 421 sec. Burn time: 390 sec. Isp(sl): 200 sec. Diameter: 10. 06 m. Span: 10. 06 m. Length: 24. 84 m. Propellants: Lox/LH 2 No Engines: 5. J-2 Status: Out of Production. Stage Number: 3. 1 x Saturn IVB (S-V) Gross Mass: 119, 900 kg. Empty Mass: 13, 300 kg. Thrust (vac): 105, 200 kgf. Isp: 421 sec. Burn time: 475 sec. Isp(sl): 200 sec. Diameter: 6. 61 m. Span: 6. 61 m. Length: 17. 80 m. Propellants: Lox/LH 2 No Engines: 1. J-2 Status: Out of Production. Comments: Saturn V version of S-IVB stage. USA

Titan IV • • Launches: 22. Failures: 2. Success Rate: 90. 91% pct. First Launch Date: 14 June 1989. Last Launch Date: 12 August 1998. Launch data is: continuing. LEO Payload: 17, 700 kg. to: 185 km Orbit. Payload: 6, 350 kg. to a: Geosynchronous transfer trajectory. Liftoff Thrust: 1, 307, 380 kgf. Liftoff Thrust: 12, 821. 00 k. N. Total Mass: 886, 420 kg. Core Diameter: 4. 33 m. Total Length: 51. 00 m. Launch Price $: 400. 00 million. in 1997 price dollars. Stage Number: 0. 2 x Titan UA 1207 Gross Mass: 319, 330 kg. Empty Mass: 51, 230 kg. Thrust (vac): 725, 732 kgf. Isp: 272 sec. Burn time: 120 sec. Isp(sl): 245 sec. Diameter: 3. 05 m. Span: 3. 05 m. Length: 34. 14 m. Propellants: Solid No Engines: 1. UA 1207 Status: In Production. Stage Number: 1. 1 x Titan 4 -1 Gross Mass: 163, 000 kg. Empty Mass: 8, 000 kg. Thrust (vac): 247, 619 kgf. Isp: 302 sec. Burn time: 164 sec. Isp(sl): 250 sec. Diameter: 3. 05 m. Span: 3. 05 m. Length: 26. 37 m. Propellants: N 2 O 4/Aerozine 50 No Engines: 2. LR-87 -11 Status: In Production. Stage Number: 2. 1 x Titan 4 -2 Gross Mass: 39, 500 kg. Empty Mass: 4, 500 kg. Thrust (vac): 46, 857 kgf. Isp: 316 sec. Burn time: 223 sec. Isp(sl): 160 sec. Diameter: 3. 05 m. Span: 3. 05 m. Length: 9. 94 m. Propellants: N 2 O 4/Aerozine-50 No Engines: 1. LR-91 -11 Status: In Production. Stage Number: 3. 1 x Centaur G Gross Mass: 23, 880 kg. Empty Mass: 2, 775 kg. Thrust (vac): 14, 970 kgf. Isp: 444 sec. Burn time: 625 sec. Diameter: 4. 33 m. Span: 4. 33 m. Length: 9. 00 m. Propellants: Lox/LH 2 No Engines: 2. RL-10 A-3 A Status: In Production. Comments: Centaur for Titan 4. Stage Number: 3. 1 x IUS-1 Gross Mass: 10, 841 kg. Empty Mass: 1, 134 kg. Thrust (vac): 18, 508 kgf. Isp: 296 sec. Burn time: 152 sec. Isp(sl): 220 sec. Diameter: 2. 34 m. Span: 2. 34 m. Length: 3. 52 m. Propellants: Solid No Engines: 1. SRM-1 Other designations: Orbus 21 D. Status: In Production. Stage Number: 4. 1 x IUS-2 Gross Mass: 3, 919 kg. Empty Mass: 1, 170 kg. Thrust (vac): 7, 996 kgf. Isp: 304 sec. Burn time: 103 sec. Isp(sl): 200 sec. Diameter: 1. 61 m. Span: 1. 61 m. Length: 2. 08 m. Propellants: Solid No Engines: 1. SRM-2 Other designations: TOS. Status: In Production.

Space Shuttle • • • Launches: 117. Failures: 1. Success Rate: 99. 15% pct. First Launch Date: 12 April 1981. Last Launch Date: 26 July 2005. Launch data is: continuing. LEO Payload: 24, 400 kg. to: 204 km Orbit. at: 28. 5 degrees. Payload: 12, 500 kg. to a: space station orbit, 407 km, 51. 6 deg inclination trajectory. Apogee: 600 km. Liftoff Thrust: 2, 625, 932 kgf. Liftoff Thrust: 25, 751. 60 k. N. Total Mass: 2, 029, 633 kg. Core Diameter: 8. 70 m. Total Length: 56. 00 m. Stage Number: 0. 2 x Shuttle SRB Gross Mass: 589, 670 kg. Empty Mass: 86, 183 kg. Thrust (vac): 1, 174, 713 kgf. Isp: 269 sec. Burn time: 124 sec. Isp(sl): 237 sec. Diameter: 3. 71 m. Span: 5. 10 m. Length: 38. 47 m. Propellants: Solid No Engines: 1. SRB Other designations: Solid Rocket Booster. Status: In Production. Stage Number: 1. 1 x Shuttle Tank Gross Mass: 750, 975 kg. Empty Mass: 29, 930 kg. Thrust (vac): 0. 000 kgf. Isp: 455 sec. Burn time: 480 sec. Isp(sl): 363 sec. Diameter: 8. 70 m. Span: 8. 70 m. Length: 46. 88 m. Propellants: Lox/LH 2 No Engines: 0. None Other designations: External Tank. Status: Out of production. Stage Number: 2. 1 x Shuttle Orbiter Gross Mass: 99, 318 kg. Empty Mass: 99, 117 kg. Thrust (vac): 696, 905 kgf. Isp: 455 sec. Burn time: 480 sec. Isp(sl): 363 sec. Diameter: 4. 90 m. Span: 23. 79 m. Length: 37. 24 m. Propellants: Lox/LH 2 No Engines: 3. SSME Other designations: Shuttle; STS (Space Transportation System). Status: In Production. USA

Proton Family of Launch Vehicles

Proton 8 K 82 M • • Manufacturer: Chelomei. Launches: 10. Success Rate: 100. 00% pct. First Launch Date: 7 April 2001. Last Launch Date: 29 December 2005. Launch data is: continuing. LEO Payload: 21, 000 kg. Payload: 4, 500 kg. to a: geosynchronous transfer orbit trajectory. Apogee: 40, 000 km. Total Mass: 712, 800 kg. Core Diameter: 7. 40 m. Total Length: 53. 00 m. 4 out of 10 launches for US companies (Direc. TV) Stage Number: 1. 1 x Proton K-1 Gross Mass: 450, 510 kg. Empty Mass: 31, 100 kg. Thrust (vac): 1, 067, 659 kgf. Isp: 316 sec. Burn time: 124 sec. Isp(sl): 267 sec. Diameter: 4. 15 m. Span: 7. 40 m. Length: 21. 20 m. Propellants: N 2 O 4/UDMH No Engines: 6. RD-253 -11 D 48 Other designations: 8 S 810 K. Status: In Production. Stage Number: 2. 1 x Proton K-2 Gross Mass: 167, 828 kg. Empty Mass: 11, 715 kg. Thrust (vac): 244, 652 kgf. Isp: 327 sec. Burn time: 206 sec. Isp(sl): 230 sec. Diameter: 4. 15 m. Span: 4. 15 m. Length: 14. 00 m. Propellants: N 2 O 4/UDMH No Engines: 4. RD-0210 Other designations: 8 S 811 K. Status: In Production. Stage Number: 3. 1 x Proton K-3 Gross Mass: 50, 747 kg. Empty Mass: 4, 185 kg. Thrust (vac): 64, 260 kgf. Isp: 325 sec. Burn time: 238 sec. Isp(sl): 230 sec. Diameter: 4. 15 m. Span: 4. 15 m. Length: 6. 50 m. Propellants: N 2 O 4/UDMH No Engines: 1. RD-0212 Status: In Production. Stage Number: 4. 1 x Proton 17 S 40 Gross Mass: 14, 600 kg. Empty Mass: 3, 300 kg. Thrust (vac): 8, 670 kgf. Isp: 352 sec. Burn time: 450 sec. Diameter: 3. 70 m. Span: 3. 70 m. Length: 7. 10 m. Propellants: Lox/Kerosene No Engines: 1. RD-58 M Other designations: Block DM; D-1 -e. Status: In Production. RUSSIA

Zenit-3 Sea Launch • • • Payload: 3, 750 kg. to a: Geosynchronous orbit trajectory. Liftoff Thrust: 740, 000 kgf. Liftoff Thrust: 7, 300. 00 k. N. Total Mass: 471, 000 kg. Core Diameter: 3. 90 m. Total Length: 59. 60 m. Stage Number: 1. 1 x Zenit-1 Gross Mass: 354, 300 kg. Empty Mass: 28, 600 kg. Thrust (vac): 834, 243 kgf. Isp: 337 sec. Burn time: 150 sec. Isp(sl): 311 sec. Diameter: 3. 90 m. Span: 3. 90 m. Length: 32. 90 m. Propellants: Lox/Kerosene No Engines: 1. RD-171 Status: In Production. Comments: Modification of same stage used as strap-on for Energia launch vehicle. Stage Number: 2. 1 x Zenit-2 Gross Mass: 90, 600 kg. Empty Mass: 9, 000 kg. Thrust (vac): 93, 000 kgf. Isp: 349 sec. Burn time: 315 sec. Isp(sl): 0. 000 sec. Diameter: 3. 90 m. Span: 3. 90 m. Length: 11. 50 m. Propellants: Lox/Kerosene No Engines: 1. RD-120 Status: In Production. Stage Number: 3. 1 x Zenit-3 Gross Mass: 17, 300 kg. Empty Mass: 2, 720 kg. Thrust (vac): 8, 660 kgf. Isp: 352 sec. Burn time: 650 sec. Diameter: 3. 70 m. Span: 3. 70 m. Length: 5. 60 m. Propellants: Lox/Kerosene No Engines: 1. RD-58 M Status: In Production. Comments: Adaptation of Block D for Zenit. ENERGIA, UKRAINE

Air Force’s Evolved Expendable Launch Vehicle (EELV) Program • EELV is a space launch system development program to replace the current fleet of medium- to heavy-lift expendable vehicles (Titan II, Delta II, Atlas II, and Titan IV) with a more affordable family of vehicles. • The new space launch vehicles must be able to meet the Government’s combined spacelift needs (Do. D, intelligence, and other missions) through at least 2020. • The primary EELV configurations are the Medium-Lift Variant (MLV), required by FY 2002 to support satellite block changes and transitions, and the Heavy-Lift Variant (HLV), required by FY 2005 to assure continued access to space following Titan IV phaseout.

Atlas Family of Launch Vehicles

Atlas V • • • Launches: 6. Success Rate: 100. 00% pct. First Launch Date: 21 August 2002. Last Launch Date: 12 August 2005. LEO Payload: 12, 500 kg. to: 185 km Orbit. at: 28. 5 degrees. Payload: 5, 000 kg. to a: Geosynchronous transfer trajectory. Liftoff Thrust: 875, 900 kgf. Liftoff Thrust: 8, 590. 00 k. N. Total Mass: 546, 700 kg. Core Diameter: 3. 81 m. Total Length: 58. 30 m. Span: 5. 40 m. Launch Price $: 138. 00 million. in 2004 price dollars. Stage Number: 0. 5 x Atlas V SRB Gross Mass: 40, 824 kg. Empty Mass: 4, 000 kg. Thrust (vac): 130, 000 kgf. Isp: 275 sec. Burn time: 94 sec. Isp(sl): 245 sec. Diameter: 1. 55 m. Span: 1. 00 m. Length: 17. 70 m. Propellants: Solid No Engines: 1. Aerojet SRB Status: In production. Comments: New SRB boosters in development for Atlas V. Empty mass, vacuum thrust, sea level Isp estimated. Stage Number: 1. 1 x Atlas CCB Gross Mass: 306, 914 kg. Empty Mass: 22, 461 kg. Thrust (vac): 423, 386 kgf. Isp: 338 sec. Burn time: 253 sec. Isp(sl): 311 sec. Diameter: 3. 81 m. Span: 3. 81 m. Length: 32. 46 m. Propellants: Lox/Kerosene No Engines: 1. RD-180 Status: In production. Comments: Common Core Booster uses Glushko RD-180 engine and new isogrid tanks. Used in Atlas IV/USAF EELV, Atlas V. Includes 272 kg booster interstage adapter and 1297 kg Centaur interstage adapter. Stage Number: 2. 1 x Centaur V 1 Gross Mass: 22, 825 kg. Empty Mass: 2, 026 kg. Thrust (vac): 10, 115 kgf. Isp: 451 sec. Burn time: 894 sec. Diameter: 3. 05 m. Span: 3. 05 m. Length: 12. 68 m. Propellants: Lox/LH 2 No Engines: 1. RL 10 A-4 -2 Status: In production. Centaur is powered by either one or two Pratt & Whitney RL 10 A-4 -2 turbopump-fed engines burning liquid oxygen and liquid hydrogen. For typical, high-energy mission applications, Centaur will be configured with one RL 10 engine. For heavy payload, low earth orbit missions, Centaur will use two RL 10 engines to maximize boost phase mission performance. Guidance, tank pressurization, and propellant usage controls for both Atlas and Centaur phases are provided by the inertial navigation unit (INU) located on the Centaur forward equipment module. Lockheed Martin, USA

Atlas V Configuration LEO 28 deg LEO Polar Geosynch Transfer Geosynch Atlas V 401 12, 500 10, 750 5, 000 N/A Atlas V 501 10, 300 9, 050 4, 100 1, 500 Atlas V 511 12, 050 10, 200 4, 900 1, 750 Atlas V 521 13, 950 11, 800 6, 000 2, 200 Atlas V 531 17, 250 14, 600 6, 900 3, 000 Atlas V 541 18, 750 15, 850 7, 600 3, 400 Atlas V 551 20, 050 17, 000 8, 200 3, 750

Delta Family of Launch Vehicles

Delta IV Heavy • • • Launches: 1. Success Rate: 100. 00% pct. First Launch Date: 21 December 2004. Last Launch Date: 21 December 2004. LEO Payload: 25, 800 kg. to: 185 km Orbit. at: 28. 5 degrees. Payload: 10, 843 kg. to a: Geosynchronous transfer, 27 deg inclination trajectory. Liftoff Thrust: 884, 000 kgf. Liftoff Thrust: 8, 670. 00 k. N. Total Mass: 733, 400 kg. Core Diameter: 5. 00 m. Total Length: 70. 70 m. Span: 15. 00 m. Development Cost $: 500. 00 million. in 2002 average dollars. Launch Price $: 254. 00 million. in 2004 price dollars. Stage Number: 0. 2 x Delta RS-68 Gross Mass: 226, 400 kg. Empty Mass: 26, 760 kg. Thrust (vac): 337, 807 kgf. Isp: 420 sec. Burn time: 249 sec. Isp(sl): 365 sec. Diameter: 5. 10 m. Span: 5. 10 m. Length: 40. 80 m. Propellants: Lox/LH 2 No Engines: 1. RS-68 Status: In production. Comments: Low cost expendable stage using lower performance engine. Used in Delta 4, Boeing EELV. Engine can be throttled to 60%. Stage Number: 1. 1 x Delta RS-68 Gross Mass: 226, 400 kg. Empty Mass: 26, 760 kg. Thrust (vac): 337, 807 kgf. Isp: 420 sec. Burn time: 249 sec. Isp(sl): 365 sec. Diameter: 5. 10 m. Span: 5. 10 m. Length: 40. 80 m. Propellants: Lox/LH 2 No Engines: 1. RS-68 Status: In production. Comments: Low cost expendable stage using lower performance engine. Used in Delta 4, Boeing EELV. Engine can be throttled to 60%. Stage Number: 2. 1 x Delta 4 H - 2 Gross Mass: 30, 710 kg. Empty Mass: 3, 490 kg. Thrust (vac): 11, 222 kgf. Isp: 462 sec. Burn time: 1, 125 sec. Diameter: 2. 44 m. Span: 5. 00 m. Length: 12. 00 m. Propellants: Lox/LH 2 No Engines: 1. RL-10 B-2 Status: In production. Comments: Delta 4 second stage with hydrogen tank increased to 5. 1 m diameter. Boeing, USA

Single Stage to Orbit (SSTO) • The problem with any single-stage-to-orbit concept is that the ability of the launch vehicle to deliver a payload to orbit is extremely sensitive to the empty weight of the final vehicle (Dumbkopf Chart). • The concept of a reusable single-stage-to-orbit Vertical Take-Off Vertical Landing (VTOVL) launch vehicle that would reenter and return to its launch site for turnaround and relaunch.

DC-XA • • The DC-X was an experimental vehicle, 1/3 the size of a planned DC-Y verticaltakeoff/vertical-landing, single stage to orbit prototype. It was not designed as an operational vehicle capable of achieving orbital flight. Its purpose was to test the feasibility of both suborbital and orbital reusable launch vehicles using the VTOVL scheme. The DC-X flew in three test series. Apogee: 10 km. Liftoff Thrust: 242. 00 k. N. Total Mass: 19, 000 kg. Core Diameter: 4. 95 m. Total Length: 12. 60 m. Stage Number: 1. 1 x DC-X Gross Mass: 16, 320 kg. Empty Mass: 7, 200 kg. Thrust (vac): 26, 800 kgf. Isp: 373 sec. Burn time: 127 sec. Isp(sl): 316 sec. Diameter: 3. 05 m. Span: 3. 66 m. Length: 11. 89 m. Propellants: Lox/LH 2 No Engines: 4. RL 10 A-5 Status: Out of Production. Mc. Donnell Douglas, USA

X-33: Venture Star • • NASA-sponsored suborbital unmanned prototype for single stage to orbit winged spacecraft. Lockheed Martin vehicle will use linear aerospike engines, metallic insulation, other features similar to their Starclipper shuttle proposals of 1971. Instrumentation in 1. 5 x 3 m bay to Mach 15. trajectory. Liftoff Thrust: 185, 900 kgf. Liftoff Thrust: 1, 823. 00 k. N. Total Mass: 123, 800 kg. Core Diameter: 20. 70 m. Total Length: 20. 40 m. Stage Number: 1. 1 x X-33 Gross Mass: 123, 800 kg. Empty Mass: 28, 600 kg. Thrust (vac): 233, 000 kgf. Isp: 439 sec. Burn time: 886 sec. Isp(sl): 339 sec. Diameter: 20. 70 m. Span: 20. 70 m. Length: 20. 40 m. Propellants: Lox/LH 2 No Engines: 2. XRS-2200 Status: Development 2002. LINEAR AEROSPIKE ENGINE: Manufacturer Name: RS-69. Other Designations: J 2 S Linear Aerospike. Designer: Rocketdyne. Developed in: 1998. Application: . Propellants: Lox/LH 2 Thrust(vac): 121, 600 kgf. Thrust(vac): 1, 192. 00 k. N. Isp: 439 sec. Isp (sea level): 339 sec. Diameter: 3. 38 m. Length: 2. 01 m. Chambers: 1. Chamber Pressure: 58. 00 bar. Area Ratio: 58. Oxidizer to Fuel Ratio: 5. 5. Country: USA. Status: In Production. Linear aerospike engine for X-33 SSTO technology demonstrator. Based on J-2 S engine developed for improved Saturn launch vehicles in the 1960's. Gas generator cycle; throttling 40% to 119% of nominal thrust; differential thrust between two engines plus-minus 15%. X-33 Advanced Technology Demonstrator Development. Designed for booster applications. Gas generator, pumpfed.

Pegasus • Manufacturer: OSC. Launches: 27. Failures: 3. Success Rate: 88. 89% pct. First Launch Date: 27 June 1994. Last Launch Date: 15 April 2005. Launch data is: continuing. LEO Payload: 443 kg. to: 185 km Orbit. at: 28. 5 degrees. Payload: 190 kg. to a: Sun synchronous, 800 km, 98. 5 deg orbital trajectory. Liftoff Thrust: 49, 623 kgf. Liftoff Thrust: 486. 64 k. N. Total Mass: 24, 000 kg. Core Diameter: 1. 27 m. Total Length: 17. 60 m. Launch Price $: 12. 00 million. in 1994 price dollars.

Pegasus Stage Data - Pegasus XL • • Stage Number: 0. 1 x L-1011 Gross Mass: 156, 000 kg. Empty Mass: 109, 629 kg. Thrust (vac): 57, 300 kgf. Isp: 9, 900 sec. Burn time: 4, 590 sec. Isp(sl): 9, 000 sec. Diameter: 16. 86 m. Span: 47. 00 m. Length: 54. 00 m. Propellants: Air/Kerosene No Engines: 3. RB-211 -22 B Status: In Production. Comments: Lockheed airliner swept wing. Release conditions: Bellymounted, 36, 800 kg, 17. 1 m length x 7. 9 m span at 925 kph at 11, 890 m altitude. Stage Number: 1. 1 x Pegasus XL-1 Gross Mass: 17, 934 kg. Empty Mass: 2, 886 kg. Thrust (vac): 60, 062 kgf. Isp: 293 sec. Burn time: 73 sec. Isp(sl): 180 sec. Diameter: 1. 27 m. Span: 6. 71 m. Length: 8. 88 m. Propellants: Solid No Engines: 1. Pegasus XL-1 Status: In Production. Stage Number: 2. 1 x Pegasus XL-2 Gross Mass: 4, 331 kg. Empty Mass: 416 kg. Thrust (vac): 15, 653 kgf. Isp: 290 sec. Burn time: 73 sec. Isp(sl): 240 sec. Diameter: 1. 27 m. Span: 1. 27 m. Length: 3. 58 m. Propellants: Solid No Engines: 1. Pegasus XL-2 Status: In Production. Stage Number: 3. 1 x Pegasus-3 Gross Mass: 985 kg. Empty Mass: 203 kg. Thrust (vac): 3, 525 kgf. Isp: 293 sec. Burn time: 65 sec. Isp(sl): 240 sec. Diameter: 0. 97 m. Span: 0. 97 m. Length: 2. 08 m. Propellants: Solid No Engines: 1. Pegasus-3 Status: In Production.

Pegasus

Pegasus

Pegasus Capabilities

Minotaur – Minotaur IV • Decommissioned Minuteman (Minotaur) or Peacekeeper (Minotaur IV) boosters • Manufacturer: OSC. Launches: 4 (Minotaur). Success Rate: 100. 00% pct. First Launch Date: 27 January 2000. Last Launch Date: 23 September 2005. Launch data is: continuing. LEO Payload: 640 kg. to: 185 km Orbit. at: 28. 5 degrees. Payload: 335 kg. to a: Sun synchronous, 741 km, 98. 6 deg inclination trajectory. Apogee: 1, 000 km. Liftoff Thrust: 73, 000 kgf. Liftoff Thrust: 720. 00 k. N. Total Mass: 36, 200 kg. Core Diameter: 1. 67 m. Total Length: 19. 21 m. Recurring Price $: 12. 50 million. in 1999 price dollars.

Minotaur • • Stage Number: 1. 1 x Minuteman-1 Gross Mass: 23, 077 kg. Empty Mass: 2, 292 kg. Thrust (vac): 80, 700 kgf. Isp: 262 sec. Burn time: 60 sec. Isp(sl): 237 sec. Diameter: 1. 67 m. Span: 1. 67 m. Length: 7. 49 m. Propellants: Solid No Engines: 1. M 55/TX-55/Tu-122 Status: Out of Production. Comments: First stage of Minuteman I. Proposed as zero stage for various Saturn variants in 1960's. Surplus motors used in ABM SDI tests in 1980's and 1990's. Stage Number: 2. 1 x Minuteman 2 -2 Gross Mass: 7, 032 kg. Empty Mass: 795 kg. Thrust (vac): 27, 300 kgf. Isp: 288 sec. Burn time: 66 sec. Diameter: 1. 33 m. Span: 1. 33 m. Length: 4. 12 m. Propellants: Solid No Engines: 1. SR 19 Status: Out of production. Comments: Second stage of Minuteman 2. Used as second stage of Minotaur launch vehicle and various SDI targets in 1980's. Stage Number: 3. 1 x Pegasus XL-2 Gross Mass: 4, 331 kg. Empty Mass: 416 kg. Thrust (vac): 15, 653 kgf. Isp: 290 sec. Burn time: 73 sec. Isp(sl): 240 sec. Diameter: 1. 27 m. Span: 1. 27 m. Length: 3. 58 m. Propellants: Solid No Engines: 1. Pegasus XL-2 Status: In Production. Stage Number: 4. 1 x Pegasus-3 Gross Mass: 985 kg. Empty Mass: 203 kg. Thrust (vac): 3, 525 kgf. Isp: 293 sec. Burn time: 65 sec. Isp(sl): 240 sec. Diameter: 0. 97 m. Span: 0. 97 m. Length: 2. 08 m. Propellants: Solid No Engines: 1. Pegasus-3 Status: In Production.

Hybrid Propulsion Vehicles

Space Ship One • Designer: Space. Developed in: 2001 -2004. Application: Rocketplane boost. Gross Mass: 2, 700 kg. Empty Mass: 300 kg. Propellants: N 2 O/Solid Thrust(vac): 7, 500 kgf. Isp: 250 sec. Burn time: 80 sec. Chambers: 1. Chamber Pressure: 24. 00 bar. Country: USA. Status: Hardware.

Space. Ship. One • Binnie Date: 4 October 2004 14: 49 GMT. . Landing Date: 4 October 2004. Flight Time: 0. 017 days. Flight Up: Space. Ship. One Flight 17 P. Flight Back: Space. Ship. One Flight 17 P. Program: X-Prize. Firsts: Suborbital altitude record for a manned spaceplane. • Sixth powered flight of Burt Rutan's Space. Ship. One and winner of the $10 million X-Prize by becoming the second flight over 100 km within a week. • Objectives of the flight were to win the Ansari X-Prize and break the rocketplane altitude record set by the X-15 in 1963. The Tier One (White Knight/Space. Ship. One) composite aircraft took off at 06: 49 PST. Drop of the rockeplane was made exactly one hour later at 14. 4 km altitude. Pilot Brian Binnie fired the hybrid rocket motor, which burned for 83 seconds. The engine cut off with Space. Ship. One at Mach 3. 09 (3524 kph) at 65 km altitude. From there it coasted to 112 km altitude. The spacecraft reached Mach 3. 25 G's during re-entry and a peak deceleration of 5. 4 G's at 32 km altitude. • 2004 Oct 4 - Space. Ship. One Flight 17 P - X-Prize Flight 2 Flight Crew: Binnie, Spacecraft: Space. Ship. One. Nation: USA. Launch Site: Mojave. Launch Vehicle: Tier One. Duration: 0. 017 days. Apogee: 112 km.

Nuclear Propulsion Vehicles

Nuclear Thermal Propulsion Systems • Nuclear Thermal Propulsion Systems have been proposed since the late 1940’s. • At the suggestion of Theodore von Kármán and following a request of Gen. H. B. Thatcher, an Ad Hoc Committee of the Scientific Advisory Board met in the Pentagon to consider the application of nuclear energy to missile propulsion. • In its report, the Committee "noted that there was an almost complete hiatus in the study of the nuclear rocket from 1947 following a report by North American Aviation, until a 1953 report by the Oak Ridge National Laboratory. • Because the technical problems appear so severe, and because another 6 years of no progress in this area would seem to be unfortunate, " the Committee felt that a continuing study both analytical and experimental, at a modest level of effort, should be carried on. • NO NTP SYSTEMS HAVE BEEN FLOWN TO DATE

Hyperion - USA • • Hyperion was considered in 1958 as a ca. 1970 Saturn follow-on. It used a small jettisonable chemical booster stage that contained chemical engines and the LOX oxidizer for the conventional engines. This booster stage surrounded the nuclear core vehicle with its large liquid hydrogen tank. The conventional stage would draw fuel from the main hydrogen tank until burnout. Hyperion would have doubled the translunar trajectory performance of the Saturn V and less than one third of the liftoff mass. Manufacturer: Convair. LEO Payload: 145, 000 kg. to: 485 km Orbit. at: 28. 0 degrees. Payload: 82, 000 kg. to a: parabolic escape trajectory. Apogee: 36 km. Liftoff Thrust: 1, 090, 000 kgf. Liftoff Thrust: 10, 700. 00 k. N. Total Mass: 850, 000 kg. Core Diameter: 8. 54 m. Total Length: 85. 40 m. Stage Number: 0. 1 x Hyperion Booster Gross Mass: 394, 625 kg. Empty Mass: 18, 144 kg. Thrust (vac): 1, 400, 000 kgf. Isp: 457 sec. Burn time: 70 sec. Isp(sl): 365 sec. Diameter: 8. 54 m. Span: 13. 00 m. Length: 12. 00 m. Propellants: Lox/LH 2 No Engines: 4. Status: Study 1959. Stage Number: 1. 1 x Hyperion Sustainer Gross Mass: 453, 592 kg. Empty Mass: 110, 000 kg. Thrust (vac): 589, 670 kgf. Isp: 800 sec. Burn time: 460 sec. Diameter: 8. 54 m. Span: 8. 54 m. Length: 51. 00 m. Propellants: Nuclear/LH 2 No Engines: 2. Nerva 12 GW Status: Study 1959.

Orion - USA • The final iteration of the Orion design was a nuclear pulse propulsion module launched into earth orbit by a Saturn V. The 100 tonne unit would have had a diameter of 10 m to match that of the booster. This would limit specific impulse to 1800 to 2500 seconds, still two to three times that of a nuclear thermal system. • A second launch would put a 100 tonne Mars spacecraft with a crew of eight into orbit. After rendezvous and checkout, the combined 200 tonne spacecraft would set out on a round trip to the Mars - total mission duration as little as 125 days!. • Manufacturer: General Atomic. Payload: 100, 000 kg. to a: Mars and back trajectory. Total Mass: 100, 000 kg. Core Diameter: 10. 00 m. Total Length: 50. 00 m.

Ya. RD ICBM - USSR • A 30 June 1958 resolution authorised development of this astounding weapon, and the draft project was completed on 30 December 1959. Perhaps coming under the heading of 'inadvisable rocket science', test launches would have been into an artificial reservoir in the target area to limit contamination by having the reactor crash into water at the end of its trajectory. Interestingly American spy Penkovskiy reported development of this rocket in 1962, but the story was not believed. Only in 1996 was the program revealed. • Manufacturer: Korolev. Liftoff Thrust: 128, 000 kgf. Total Mass: 84, 400 kg. Core Diameter: 3. 33 m. Total Length: 25. 00 m. • Stage Number: 1. 1 x Ya. RD ICBM OKB-670 Gross Mass: 96, 000 kg. Empty Mass: 8, 800 kg. Thrust (vac): 170, 000 kgf. Isp: 470 sec. Burn time: 235 sec. Isp(sl): 430 sec. Diameter: 3. 33 m. Span: 3. 33 m. Length: 23. 00 m. Propellants: Nuclear/Ammonia+Alcohol No Engines: 1. Ya. RD OKB-670 Status: Study 1959. Comments: Nuclear-propelled ICBM with engines in development by Bondayuk. Four expansion nozzles fed by single reactor. Payload 4, 000 kg to 14, 000 km. Empty mass, vehicle length calculated.

N 1 - USSR • • • For a Mars expedition, it was calculated that the AF engine would deliver 40% more payload than a chemical stage. But Korolev’s study also effectively killed the program by noting that his favoured solution, a nuclear electric ion engine, would deliver 70% more payload than the Lox/LH 2 stage. Further investigation of nuclear thermal stages for the N 1 does not seem to have been pursued. Manufacturer: Korolev. LEO Payload: 270, 000 kg. to: 220 km Orbit. at: 51. 6 degrees. Payload: 24, 600 kg. to a: lunar surface trajectory. Liftoff Thrust: 3, 600, 000 kgf. Liftoff Thrust: 35, 000. 00 k. N. Total Mass: 2, 400, 000 kg. Core Diameter: 17. 00 m. Total Length: 180. 00 m. Stage Number: 1. 1 x N 1 1962 - A Gross Mass: 1, 384, 000 kg. Empty Mass: 117, 000 kg. Thrust (vac): 4, 020, 000 kgf. Isp: 331 sec. Burn time: 103 sec. Isp(sl): 296 sec. Diameter: 10. 00 m. Span: 17. 00 m. Length: 30. 00 m. Propellants: Lox/Kerosene No Engines: 24. NK-15 Status: Study 1962. Comments: Includes 14, 000 kg for Stage 1 -2 interstage and payload fairing. Compared to total fuelled mass excludes 15, 000 kg propellant expended in thrust build-up and boil-off prior to liftoff. Values as in draft project as defended on 2 -16 July 1962. Stage Number: 2. 1 x N 1 Nuclear A Gross Mass: 700, 000 kg. Empty Mass: 250, 000 kg. Thrust (vac): 700, 000 kgf. Isp: 900 sec. Burn time: 570 sec. Diameter: 12. 00 m. Span: 12. 00 m. Length: 90. 00 m. Propellants: Nuclear/LH 2 No Engines: 40. Ya. RD Type A Status: Study 1963. Comments: N 1 nuclear upper stage study, 1963. Figures calculated based on given total stage thrust, specific impulse, engine mass.