Centaur home
topic index
Centaur

Credit - © Mark Wade
Model: Centaur C. Gross Mass: 15,600 kg (34,300 lb). Empty Mass: 1,996 kg (4,400 lb). Thrust (vac): 133.448 kN (30,000 lbf). Isp: 425 sec. Burn time: 430 sec. Propellants: Lox/LH2. Isp(sl): 0 sec. Diameter: 3.05 m (10.00 ft). Span: 3.05 m (10.00 ft). Length: 9.14 m (29.98 ft). Country: USA. No Engines: 2. Motor: RL-10A-1. Status: Study 1960. Cost $ : 20.300 million. First Flight: 1958. Last Flight: 1965. No Launched: 28. No Failed: 2.

The first high-energy liquid oxygen/liquid hydrogen propellant stage in history. Despite initial development problems, the Centaur is entering its sixth decade of development and production.

Early Centaur Guidance System

The Centaur guidance system was all-inertial, consisting primarily of a four-gimbal all-attitude inertial platform and a general purpose serial digital computer with a magnetic drum memory. The airborne guidance program was written onto the drum memory from a punched paper tape along with a pre-flight calibration and alignment program for trimming and aligning the platform prior to launch.

For the geosynchronous equatorial orbit mission the Centaur guidance system performed the following functions:

During the Atlas booster phase, the vehicle pitch program was generated by the Atlas autopilot; however, the guidance system monitored the vehicle position and velocity and generated the booster staging discrete as a function of vehicle acceleration. For the Atlas sustainer stage the guidance system generated vehicle steering signals, which were used to orient the thrust vector so as to reduce the position and velocity dispersions generated during the open-loop booster stage. The sustalner engine cutoff command was also given by the guidance system.

After separation of the Centaur stage from the Atlas booster, the Centaur guidance system controlled the vehicle during each of the succeeding three phases of powered flight necessary to place the vehicle in its final orbit. The guidance system provided steering and cutoff signals to the Centaur autopilot during the powered phases of flight and also provided an attitude reference to the autopilot prior to the second and third firings of Centaur in order that the vehicle assumed the proper attitude prior to thrust initiation. The following was a typical 3 start flight sequence:

  • T+0 to 15 sec. Vertical rise and roll to desired azimuth
  • Time dependent pitch program to booster staging (booster staging initiated by an accelerometer when acceleration reached 5.8 g's.
  • At Beco (Booster engine cut-off) + 15 seconds, the Centaur tank insulation panels were jettisoned. The sustainer phase was flown at a constant inertial attitude
  • At Beco +63 seconds, the payload shroud was jettisoned. The sustainer phase was terminated by propellant depletion. At Seco (Sustainer engine cut-off) the vehicle continued in a constant inertial attitude while the Atlas continued a low acceleration in the vernier solo phase.
  • At Seco + 9.5 seconds the Centaur main engine prestart (chilldown) was initiated
  • Centaur separation and ullage rocket firing was initiated at Veco (Vernier engine out-off). This first ullage rocket firing period was 14.5 sec.
  • First Centaur main engine firing; a constant pitch rate was maintained until main engine cutoff, at which point Centaur and its payload were in a low earth parking orbit.
  • The Centaur was orientated "tail to sun" in parking orbit for the first coast period. Approximately 300 seconds prior to the second main engine start, the vehicle was re-oriented with the firing direction rockets starting 42 seconds prior to the main engine. Engine prestart was initiated 20 seconds prior to main engine start.
  • Prior to the second main engine burn, the vehicle was again oriented "tail to sun" for the second coast period. The vehicle was re-oriented for the third firing direction, with the ullage rockets starting 50 seconds prior to the main engine. Prestart was initiated 20 seconds prior to the main engine.
  • Third main engine burn was followed by payload separation

Structural considerations of the configuration limited the product of the angle of attack and dynamic pressure, q, to approximately 67 kN/m^2 with 2-sigma winds at Cape Canaveral. The maximum permissible longitudinal and lateral acceleration factors were 7.0 g and 1.0 g, respectively.

Model: Centaur D/E. Gross Mass: 16,258 kg (35,842 lb). Empty Mass: 2,631 kg (5,800 lb). Thrust (vac): 131.222 kN (29,500 lbf). Isp: 444 sec. Burn time: 470 sec. Propellants: Lox/LH2. Isp(sl): 0 sec. Diameter: 3.05 m (10.00 ft). Span: 3.05 m (10.00 ft). Length: 9.60 m (31.40 ft). Country: USA. No Engines: 2. Motor: RL-10A-3. Status: Study 1968. Cost $ : 20.300 million. First Flight: 1960. Last Flight: 1983. No Launched: 71. No Failed: 5.

Model: Centaur I. Gross Mass: 15,600 kg (34,300 lb). Empty Mass: 1,700 kg (3,700 lb). Thrust (vac): 146.800 kN (33,002 lbf). Isp: 444 sec. Burn time: 402 sec. Propellants: Lox/LH2. Isp(sl): 0 sec. Diameter: 3.05 m (10.00 ft). Span: 3.05 m (10.00 ft). Length: 9.15 m (30.01 ft). Country: USA. No Engines: 2. Motor: RL-10A-3A. Status: Out of production. Cost $ : 20.300 million. First Flight: 1984. Last Flight: 1997. No Launched: 18. No Failed: 3.

Model: Centaur G Prime. Gross Mass: 19,501 kg (42,992 lb). Empty Mass: 3,000 kg (6,600 lb). Thrust (vac): 146.800 kN (33,002 lbf). Isp: 444 sec. Burn time: 550 sec. Propellants: Lox/LH2. Isp(sl): 0 sec. Diameter: 4.33 m (14.20 ft). Span: 4.33 m (14.20 ft). Length: 8.87 m (29.10 ft). Country: USA. No Engines: 2. Motor: RL-10A-3A. Status: Out of production. Cost $ : 22.000 million. First Flight: 1987.

Centaur for Shuttle payload bay. Cancelled after Challenger disaster on safety grounds.

Model: Centaur G STS. Gross Mass: 16,327 kg (35,994 lb). Empty Mass: 2,600 kg (5,700 lb). Thrust (vac): 146.800 kN (33,002 lbf). Isp: 444 sec. Burn time: 420 sec. Propellants: Lox/LH2. Isp(sl): 0 sec. Diameter: 4.33 m (14.20 ft). Span: 4.33 m (14.20 ft). Length: 5.95 m (19.52 ft). Country: USA. No Engines: 2. Motor: RL-10A-3A. Status: Out of production. Cost $ : 21.000 million. First Flight: 1987.

Model: Centaur II. Gross Mass: 18,833 kg (41,519 lb). Empty Mass: 2,053 kg (4,526 lb). Thrust (vac): 146.800 kN (33,002 lbf). Isp: 444 sec. Burn time: 488 sec. Propellants: Lox/LH2. Isp(sl): 0 sec. Diameter: 3.05 m (10.00 ft). Span: 3.05 m (10.00 ft). Length: 10.10 m (33.10 ft). Country: USA. No Engines: 2. Motor: RL-10A-3A. Status: Out of production. Cost $ : 25.000 million. First Flight: 1991. Last Flight: 1998. No Launched: 10.

Model: Centaur IIA. Gross Mass: 19,073 kg (42,048 lb). Empty Mass: 2,293 kg (5,055 lb). Thrust (vac): 185.012 kN (41,592 lbf). Isp: 449 sec. Burn time: 392 sec. Propellants: Lox/LH2. Isp(sl): 0 sec. Diameter: 3.05 m (10.00 ft). Span: 3.05 m (10.00 ft). Length: 10.10 m (33.10 ft). Country: USA. No Engines: 2. Motor: RL-10A-4. Status: Out of production. Cost $ : 30.000 million. First Flight: 1992. Last Flight: 2004. No Launched: 56.

Model: Centaur G. Gross Mass: 23,880 kg (52,640 lb). Empty Mass: 2,775 kg (6,117 lb). Thrust (vac): 146.800 kN (33,002 lbf). Isp: 444 sec. Burn time: 625 sec. Propellants: Lox/LH2. Diameter: 4.33 m (14.20 ft). Span: 4.33 m (14.20 ft). Length: 9.00 m (29.50 ft). Country: USA. No Engines: 2. Motor: RL-10A-3A. Status: Retired 2005. Cost $ : 23.000 million. First Flight: 1994. Last Flight: 2005. No Launched: 21. No Failed: 1.

Centaur for Titan 4

Model: Centaur B-X. Gross Mass: 19,138 kg (42,192 lb). Empty Mass: 2,358 kg (5,198 lb). Thrust (vac): 186.797 kN (41,994 lbf). Isp: 470 sec. Burn time: 408 sec. Propellants: Lox/LH2. Isp(sl): 0 sec. Diameter: 3.05 m (10.00 ft). Span: 3.05 m (10.00 ft). Length: 10.10 m (33.10 ft). Country: USA. No Engines: 2. Motor: RL-10B-X. Status: Development 1998. Cost $ : 21.000 million. First Flight: 2000.

Conceptual design. Not put into production.

Model: Centaur C-X. Gross Mass: 19,138 kg (42,192 lb). Empty Mass: 2,358 kg (5,198 lb). Thrust (vac): 110.800 kN (24,909 lbf). Isp: 450 sec. Burn time: 660 sec. Propellants: Lox/LH2. Isp(sl): 0 sec. Diameter: 3.05 m (10.00 ft). Span: 3.05 m (10.00 ft). Length: 10.10 m (33.10 ft). Country: USA. No Engines: 1. Motor: RL-10C-X. Status: Development 1998. Cost $ : 21.000 million. First Flight: 2001.

Conceptual design. Not put into production.

Model: Centaur IIIA. Gross Mass: 18,710 kg (41,240 lb). Empty Mass: 1,905 kg (4,199 lb). Thrust (vac): 99.155 kN (22,291 lbf). Isp: 451 sec. Burn time: 738 sec. Propellants: Lox/LH2. Isp(sl): 0 sec. Diameter: 3.05 m (10.00 ft). Span: 3.05 m (10.00 ft). Length: 10.00 m (32.00 ft). Country: USA. No Engines: 1. Motor: RL-10A-4-1. Status: Retired 2004. Cost $ : 25.000 million. First Flight: 2000. Last Flight: 2004. No Launched: 2.

Single-engine Centaur for Atlas IIIA. The Lockheed Martin manufactured Centaur IIIA upper stage is powered by one Pratt & Whitney RL10A-4-1 turbopump-fed engine burning liquid oxygen and liquid hydrogen. Using the proven flight design of the Centaur IIAS stage, the only changes to Centaur for IIIA are in the aft region of the stage. For Centaur IIIA, one of Centaur IIAS's two RL10 engines is removed. The remaining engine is re-positioned to a center-mount, and electro-mechanical thrust vector control actuators replace the hydraulically actuated system previously in use. Guidance, tank pressurization, and propellant usage controls for both Atlas and Centaur phases are provided by the inertial navigation unit (INU) located on the forward equipment module. The first Centaur burn lasts about nine minutes after which the Centaur and its payload coast in a parking orbit. During the first burn, approximately ten seconds after ignition, the payload fairing is jettisoned. The second Centaur ignition occurs about 23 minutes into the flight, continues for about three minutes, and is followed several minutes later by the separation of the spacecraft from Centaur.

Model: Centaur IIIB. Gross Mass: 22,960 kg (50,610 lb). Empty Mass: 2,130 kg (4,690 lb). Thrust (vac): 198.319 kN (44,584 lbf). Isp: 451 sec. Burn time: 460 sec. Propellants: Lox/LH2. Isp(sl): 0 sec. Diameter: 3.05 m (10.00 ft). Span: 3.05 m (10.00 ft). Length: 11.68 m (38.32 ft). Country: USA. No Engines: 2. Motor: RL-10A-4-2. Status: Out of production. First Flight: 2002. Last Flight: 2005. No Launched: 4.

Dual-engine Centaur for Atlas IIIB. The Lockheed Martin manufactured Centaur IIIB upper stage is powered by two Pratt & Whitney RL10A-4-2 turbopump-fed engines burning liquid oxygen and liquid hydrogen. The changes to Centaur for Atlas IIIB are a stretched tank (1.68 m) and the addition of the second engine. Guidance, tank pressurization, and propellant usage controls for both Atlas and Centaur phases are provided by the inertial navigation unit (INU) located on the forward equipment module. The first Centaur burn lasts about five minutes, after which the Centaur and its payload coast in a parking orbit. During the first burn, approximately eight seconds after ignition, the payload fairing is jettisoned. The second Centaur ignition occurs 27 minutes into the flight, continues for about three minutes, and is followed several minutes later by the separation of the spacecraft from Centaur.

Model: Centaur V1. Gross Mass: 22,825 kg (50,320 lb). Empty Mass: 2,026 kg (4,466 lb). Thrust (vac): 99.194 kN (22,300 lbf). Isp: 451 sec. Burn time: 894 sec. Propellants: Lox/LH2. Diameter: 3.05 m (10.00 ft). Span: 3.05 m (10.00 ft). Length: 12.68 m (41.60 ft). Country: USA. No Engines: 1. Motor: RL-10A-4-2. Status: Active. First Flight: 2002. Last Flight: 2007. No Launched: 21.

Single-engine Centaur for Atlas V. 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. 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.

Model: Centaur V2. Gross Mass: 23,050 kg (50,810 lb). Empty Mass: 2,250 kg (4,960 lb). Thrust (vac): 198.398 kN (44,602 lbf). Isp: 451 sec. Burn time: 435 sec. Propellants: Lox/LH2. Diameter: 3.05 m (10.00 ft). Span: 3.05 m (10.00 ft). Length: 12.68 m (41.60 ft). Country: USA. No Engines: 2. Motor: RL-10A-4-2. Status: Active. First Flight: 2002. Last Flight: 2002. No Launched: 1.

Dual-engine Centaur for Atlas V. For heavy payload, low earth orbit missions, Centaur will use two RL10 engines to maximize boost phase mission performance.

Used on Launch Vehicles:
Contact us with any corrections, additions, or comments.
Conditions for use of drawings, pictures, or other materials from this site..
To contact astronauts or cosmonauts.

© Mark Wade, 1997 - 2008 except where otherwise noted.