1. MAE 5495: Launch Vehicle Analysis and Design Lesson 8: Liquid Propulsion Systems

2. German V2 The V-2 ballistic missile (known to its designers as the A4) was the world's first operational liquid fuel rocket. It represented an enormous quantum leap in technology, financed by Nazi Germany in a huge development program that cost at least $ 2 billion in 1944 dollars. The British, Americans, and Russians launched a further 86 captured German V-2's in 1945- 1952. Personnel and technology from the V-2 program formed the starting point for post-war rocketry development in America, Russia, and France. Apogee: 85 km. Liftoff Thrust: 270.00 kN. Total Mass: 12,800 kg. Core Diameter: 1.65 m. Total Length: 13.60 m. Stage Number: 1. 1 x A-4 Gross Mass: 12,805 kg. Empty Mass: 4,008 kg. Thrust (vac): 31,800 kgf. Isp: 239 sec. Burn time: 68 sec. Isp(sl): 203 sec. Diameter: 1.65 m. Span: 3.56 m. Length: 12.00 m. Propellants: Lox/Alcohol No Engines: 1. A-4 Other designations: V-2. Status: Out of Production.A-4Lox/AlcoholA-4

3. Saturn V Launch Vehicles

4. 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 kN. 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.Saturn ICLox/KeroseneF-1 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/LH2 No Engines: 5. J-2 Status: Out of Production.Saturn IILox/LH2J-2 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/LH2 No Engines: 1. J-2 Status: Out of Production. Comments: Saturn V version of S-IVB stage.Saturn IVB (S-V)Lox/LH2J-2 USA

5. N1 Launch Vehicles

6. N1 Launches: 4. Failures: 4. Success Rate: 0.000 pct. First Launch Date: 21 February 1969. Last Launch Date: 23 November 1972. LEO Payload: 70,000 kg. to: 225 km Orbit. at: 51.6 degrees. Apogee: 200 km. Liftoff Thrust: 4,400,000 kgf. Liftoff Thrust: 43,000.00 kN. Total Mass: 2,735,000 kg. Core Diameter: 17.00 m. Total Length: 105.00 m. Stage Number: 1. 1 x N1 Block A Gross Mass: 1,880,000 kg. Empty Mass: 130,000 kg. Thrust (vac): 5,130,000 kgf. Isp: 330 sec. Burn time: 125 sec. Isp(sl): 284 sec. Diameter: 10.30 m. Span: 16.90 m. Length: 30.10 m. Propellants: Lox/Kerosene No Engines: 30. NK-15 Status: Out of Production. 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.N1 Block ALox/KeroseneNK-15 Stage Number: 2. 1 x N1 Block B Gross Mass: 560,700 kg. Empty Mass: 55,700 kg. Thrust (vac): 1,431,680 kgf. Isp: 346 sec. Burn time: 120 sec. Diameter: 6.80 m. Span: 9.80 m. Length: 20.50 m. Propellants: Lox/Kerosene No Engines: 8. NK-15V Status: Out of Production. Comments: Includes 3500 kg Stage 2-Stage 3 interstage. Compared to total fuelled mass excludes 1,000 kg in propellants lost to boil-off prior to stage ignition. Values as in draft project as defended on 2-16 July 1962.N1 Block B Lox/KeroseneNK-15V Stage Number: 3. 1 x N1 Block V Gross Mass: 188,700 kg. Empty Mass: 13,700 kg. Thrust (vac): 164,000 kgf. Isp: 353 sec. Burn time: 370 sec. Diameter: 4.80 m. Span: 6.40 m. Length: 14.10 m. Propellants: Lox/Kerosene No Engines: 4. NK-21 Status: Out of Production.N1 Block V Lox/KeroseneNK-21 Stage Number: 4. 1 x N1 Block G Gross Mass: 61,800 kg. Empty Mass: 6,000 kg. Thrust (vac): 45,479 kgf. Isp: 353 sec. Burn time: 443 sec. Diameter: 4.40 m. Span: 4.40 m. Length: 9.10 m. Propellants: Lox/Kerosene No Engines: 1. NK-19 Status: Out of Production. Comments: Empty mass estimated.N1 Block G Lox/KeroseneNK-19 Stage Number: 5. 1 x N1 Block D Gross Mass: 18,200 kg. Empty Mass: 3,500 kg. Thrust (vac): 8,500 kgf. Isp: 349 sec. Burn time: 600 sec. Isp(sl): 0.000 sec. Diameter: 2.90 m. Span: 2.90 m. Length: 5.70 m. Propellants: Lox/Kerosene No Engines: 1. RD-58 Status: Out of Production. Comments: Block D adapted as lunar crasher stage.N1 Block D Lox/KeroseneRD-58

7. 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 kN. 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.Shuttle SRBSolidSRB 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/LH2 No Engines: 0. None Other designations: External Tank. Status: Out of production.Shuttle TankLox/LH2 None 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/LH2 No Engines: 3. SSME Other designations: Shuttle; STS (Space Transportation System). Status: In Production.Shuttle OrbiterLox/LH2 SSME USA

8. Proton Family of Launch Vehicles

9. Proton 8K82M 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 (DirecTV) 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: N2O4/UDMH No Engines: 6. RD-253-11D48 Other designations: 8S810K. Status: In Production.Proton K-1N2O4/UDMHRD-253-11D48 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: N2O4/UDMH No Engines: 4. RD-0210 Other designations: 8S811K. Status: In Production.Proton K-2 N2O4/UDMHRD-0210 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: N2O4/UDMH No Engines: 1. RD-0212 Status: In Production.Proton K-3 N2O4/UDMHRD-0212 Stage Number: 4. 1 x Proton 17S40 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-58M Other designations: Block DM; D-1-e. Status: In Production.Proton 17S40Lox/KeroseneRD-58M RUSSIA

10. 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 kN. 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 UA1207 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. UA1207 Status: In Production.Titan UA1207 SolidUA1207 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: N2O4/Aerozine- 50 No Engines: 2. LR-87-11 Status: In Production.Titan 4-1N2O4/Aerozine- 50LR-87-11 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: N2O4/Aerozine-50 No Engines: 1. LR-91-11 Status: In Production.Titan 4-2N2O4/Aerozine-50LR-91-11 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/LH2 No Engines: 2. RL-10A-3A Status: In Production. Comments: Centaur for Titan 4.Centaur GLox/LH2RL-10A-3A 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 21D. Status: In Production.IUS-1 SolidSRM-1 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.IUS-2 SolidSRM-2

11. Zenit-3 Sea Launch Payload: 3,750 kg. to a: Geosynchronous orbit trajectory. Liftoff Thrust: 740,000 kgf. Liftoff Thrust: 7,300.00 kN. 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.Zenit-1 Lox/KeroseneRD-171 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.Zenit-2 Lox/KeroseneRD-120 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-58M Status: In Production. Comments: Adaptation of Block D for Zenit.Zenit-3Lox/KeroseneRD-58M ENERGIA, UKRAINE

12. 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 (DoD, 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.

13. Atlas Family of Launch Vehicles

14. 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 kN. 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.Atlas V SRBSolidAerojet SRB 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.Atlas CCB Lox/KeroseneRD-180 Stage Number: 2. 1 x Centaur V1 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/LH2 No Engines: 1. RL- 10A-4-2 Status: In production. Centaur is powered by either one or two Pratt & Whitney RL10A-4-2 turbopump-fed engines burning liquid oxygen and liquid hydrogen. For typical, high-energy mission applications, Centaur will be configured with one RL10 engine. For heavy payload, low earth orbit missions, Centaur will use two RL10 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.Centaur V1Lox/LH2RL- 10A-4-2 Lockheed Martin, USA

15. Atlas V ConfigurationLEO 28 degLEO PolarGeosynch TransferGeosynch Atlas V 40112,50010,7505,000N/A Atlas V 50110,3009,0504,1001,500 Atlas V 51112,05010,2004,9001,750 Atlas V 52113,95011,8006,0002,200 Atlas V 53117,25014,6006,9003,000 Atlas V 54118,75015,8507,6003,400 Atlas V 55120,05017,0008,2003,750

16. Delta Family of Launch Vehicles

17. 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, 27deg inclination trajectory. Liftoff Thrust: 884,000 kgf. Liftoff Thrust: 8,670.00 kN. 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/LH2 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%.Delta RS-68Lox/LH2RS-68 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/LH2 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%.Delta RS-68Lox/LH2RS-68 Stage Number: 2. 1 x Delta 4H - 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/LH2 No Engines: 1. RL-10B-2 Status: In production. Comments: Delta 4 second stage with hydrogen tank increased to 5.1 m diameter.Delta 4H - 2Lox/LH2RL-10B-2 Boeing, USA

18. 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.

19. DC-XA The DC-X was an experimental vehicle, 1/3 the size of a planned DC-Y vertical- takeoff/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 kN. 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/LH2 No Engines: 4. RL- 10A-5 Status: Out of Production.DC-XLox/LH2RL- 10A-5 McDonnell Douglas, USA

20. 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 kN. 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/LH2 No Engines: 2. XRS- 2200 Status: Development 2002.X-33Lox/LH2XRS- 2200 LINEAR AEROSPIKE ENGINE: Manufacturer Name: RS-69. Other Designations: J- 2S Linear Aerospike. Designer: Rocketdyne. Developed in: 1998. Application:. Propellants: Lox/LH2 Thrust(vac): 121,600 kgf. Thrust(vac): 1,192.00 kN. 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-2S 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, pump-fed.Lox/LH2

21. Linear Aerospike Engine Unlike conventional rocket engines, which feature a bell nozzle that constricts expanding gasses, the basic aerospike shape is that of a bell turned inside out and upside down. When the reconfigured bell is "unwrapped" and laid flat, it is called a linear aerospike. The linear aerospike features a series of small combustion chambers along the unwrapped bell, also called the ramp, that shoot hot gases along the ramp's outside surface to produce thrust along the length of the ramp, hence the name "linear aerospike." With the aerospike, the ramp serves as the inner wall of the virtual bell nozzle, while atmospheric pressure serves as the "invisible" outer wall. The combustion gasses race along the inner wall (the ramp) and the outer wall (atmospheric pressure) to produce thrust. The key to a conventional bell nozzle's level of performance is its width. At high pressure -- i.e. sea level -- the gasses are more tightly focused, so a bell nozzle with a narrow interior surface works best. At low pressure -- i.e. higher altitudes -- a wider interior works best as the gasses will expand farther. Since the width of the bell nozzles can't change to match the atmospheric pressure as the rocket climbs, bell nozzles are normally designed to provide optimum performance at one certain altitude or pressure. This is called a "point design," and engineers accept the performance loss the nozzle will encounter at any altitude other than the one it was designed for. The aerospike eliminates this loss of performance. Since the combustion gasses only are constrained on one side by a fixed surface -- the ramp -- and constrained on the other side by atmospheric pressure, the aerospike's plume can widen with the decreasing atmospheric pressure as the vehicle climbs, thus maintaining more efficient thrust throughout the vehicle's flight. The X-33's direction of flight will be controlled by varying the thrust side to side and engine to engine, rather than by adjusting the direction of a bell nozzle. The lack of systems for thrust vectoring -- such as gimbals, hydraulics and flex lines -- also will make the aerospike easier to maintain than conventional engines and help keep the X-33's weight down. The X-33 will feature two aerospike engines, each independently supplying 20 of the 40 combustion chambers running along the base of the X-33 to produce a combined 412,000 pounds of thrust at sea level. The specific impulse is estimated at 339.9 seconds at sea level, 429.8 seconds in a vacuum.

22. Aerospike Engines Linear Aerospike A vehicle with an aerospike engine uses 25- 30% less fuel at low altitudes, where most missions have the greatest need for thrust.

23. Pulse Detonation Rocket Engines Pulse detonation rocket engines offer a lightweight, low-cost alternative for space transportation. Pulse detonation rocket engines operate by injecting propellants into long cylinders that are open on one end and closed on the other. When gas fills a cylinder, an igniter -- such as a spark plug -- is activated. Fuel begins to burn and rapidly transitions to a detonation, or powered shock. The shock wave travels through the cylinder at 10 times the speed of sound, so combustion is completed before the gas has time to expand. The explosive pressure of the detonation pushes the exhaust out the open end of the cylinder, providing thrust to the vehicle. A major advantage is that pulse detonation rocket engines boost the fuel and oxidizer to extremely high pressure without a turbopump -- an expensive part of conventional rocket engines. In a typical rocket engine, complex turbopumps must push fuel and oxidizer into the engine chamber at an extremely high pressure of about 2,000 pounds per square inch or the fuel is blown back out. The pulse mode of pulse detonation rocket engines allows the fuel to be injected at a low pressure of about 200 pounds per square inch.