Encyclopedia Astronautica
Centaur C



centaur.gif
Centaur
Credit: © Mark Wade
Lox/LH2 propellant rocket stage. Loaded/empty mass 15,600/1,996 kg. Thrust 133.45 kN. Vacuum specific impulse 425 seconds. 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.

Cost $ : 20.300 million. No Engines: 2.

Status: Study 1960.
Gross mass: 15,600 kg (34,300 lb).
Unfuelled mass: 1,996 kg (4,400 lb).
Height: 9.14 m (29.98 ft).
Diameter: 3.05 m (10.00 ft).
Span: 3.05 m (10.00 ft).
Thrust: 133.45 kN (30,000 lbf).
Specific impulse: 425 s.
Burn time: 430 s.
Number: 28 .

More... - Chronology...


Associated Countries
Associated Engines
  • RL-10A-1 Pratt and Whitney lox/lh2 rocket engine. 66.7 kN. Isp=425s. Version used on Atlas Centaur LV-3C, and proposed for various early Saturn launch vehicle designs. First flight 1961. More...

Associated Launch Vehicles
  • Juno V-B American orbital launch vehicle. A proposed version of the Juno V for lunar and planetary missions used a Titan I ICBM first stage and a Centaur high-energy third stage atop the basic Juno V cluster. Masses, payload estimated. More...
  • Saturn A-1 American orbital launch vehicle. Projected first version of Saturn I, to be used if necessary before S-IV liquid hydrogen second stage became available. Titan 1 first stage used as second stage, Centaur third stage. Masses, payload estimated. More...
  • Saturn C-2 American orbital launch vehicle. The launch vehicle initially considered for realizing the Apollo lunar landing at the earliest possible date. 15 launches and rendezvous required to assemble direct landing spacecraft in earth orbit. More...
  • Saturn B-1 American orbital launch vehicle. Most powerful version of Saturn I considered. New low energy second stage with four H-1 engines, S-IV third stage, Centaur fourth stage. Masses, payload estimated. More...
  • Saturn A-2 American orbital launch vehicle. More powerful version of Saturn I with low energy second stage consisting of cluster of four IRBM motors and tankage, Centaur third stage. Masses, payload estimated. More...
  • Saturn I American orbital launch vehicle. Von Braun launch vehicle known as 'Cluster's Last Stand' - 8 Redstone tanks around a Jupiter tank core,powered by eight Jupiter engines. Originally intended as the launch vehicle for Apollo manned circumlunar flights. However it was developed so early, no payloads were available for it. More...
  • Saturn C-1 American orbital launch vehicle. Original flight version with dummy upper stages, including dummy Saturn S-V/Centaur (never flown). More...
  • Saturn I Blk2 American orbital launch vehicle. Second Block of Saturn I, with substantially redesigned first stage and large fins to accomodate Dynasoar payload. More...
  • Atlas Centaur American orbital launch vehicle. First test version of Atlas with Centaur upper stage. More...

Associated Propellants
  • Lox/LH2 Liquid oxygen was the earliest, cheapest, safest, and eventually the preferred oxidiser for large space launchers. Its main drawback is that it is moderately cryogenic, and therefore not suitable for military uses where storage of the fuelled missile and quick launch are required. Liquid hydrogen was identified by all the leading rocket visionaries as the theoretically ideal rocket fuel. It had big drawbacks, however - it was highly cryogenic, and it had a very low density, making for large tanks. The United States mastered hydrogen technology for the highly classified Lockheed CL-400 Suntan reconnaissance aircraft in the mid-1950's. The technology was transferred to the Centaur rocket stage program, and by the mid-1960's the United States was flying the Centaur and Saturn upper stages using the fuel. It was adopted for the core of the space shuttle, and Centaur stages still fly today. More...

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