Encyclopedia Astronautica
Space Tug



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Space Tug
Credit: © Mark Wade
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Tugcrew
Credit: NASA
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OTV Turtle 2
Space Tug. This illustration (from 1984) depicts a manned space tug returning to a space station from geostationary or lunar orbit. The vehicle passes through the Earth's atmosphere to slow down; its aeroshell is heated to thousands of degrees by kinetic friction. The small cylinder is the crew module.
Credit: NASA via Marcus Lindroos
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Space Tug
Credit: NASA
American space tug. Study 1971. The original Boeing Space Tug design of the early 1970's was sized to be flown either in a single shuttle mission or as a Saturn V payload. Optimum mass was found to be 20.6 metric tons regardless.

The Tug could be outfitted with a variety of kits to serve in many roles, including as a manned lunar lander. Aerobraking for recovery in low earth orbit was considered for further study, but the baseline used RL10 engines to brake into earth orbit for refurbishment and refueling at a space station. All further work was cancelled by NASA in 1972, but resurrected as the aerobraking Orbital Transfer Vehicle in the 1980's.

Space Tug Systems had to be compatible for both utilization as (1) upper stages and payload components for the Saturn V vehicle and its derivatives and (2) as upper stages and payload components for the Earth-to-Orbit Shuttle (EOS). Primary applications for the Space Tug/Saturn V Systems would be for transportation of large payloads to lunar orbit and interplanetary missions. The Space Tug systems would be utilized as payload components for the above missions when used in conjunction with the nuclear shuttle. The majority of the Space Tug missions would, however, be in conjunction with the EOS. The baseline EOS considered for selection of the compatible Space Tug inventory was one with a 4.57 m diameter by 18.29 m long cargo bay. The maximum capability of this baseline EOS was specified as 24,500 kg to a 28 deg 185 km circular earth orbit. Later EOS design criteria, however, established the EOS capability to the 185 km. 28-1/2 inclination orbit at 29,500 kg. This larger EOS would allow utilization of a larger Tug propulsion module. The study had shown that the desirability of a larger propulsion module was generally questionable unless the size could be increased to on the order of 40,900 kg. However, if the aerobraking mode was proven feasible, this larger EOS capability could allow either placement or retrieval of 4500 kg of payload to or from geosynchronous orbit with a single EOS launch.

Considering the overall mission requirements and the required compatibility of the Space Tug with the other elements of the Space Transportation System, an inventory of Space Tug elements was selected. This inventory could accomplish, when assembled into the proper configurations, the overall mission spectrum. The selected Tug inventory consisted of the following components:

  • Primary propulsion modules with a 18,000 kg propellant capacity (designed for earth orbit missions).
  • Expendable drop tanks with 18,000 kg propellant capacity.
  • Secondary propulsion modules with a 7,600 kg propellant capacity (designed for earth orbit missions).
  • Astrionics modules (designed for earth orbit missions).
  • All purpose crew modules (outfitted as required for the various missions).
  • Cargo modules which use the shell of the all-purpose crew module.
  • Doughnut cargo modules (to carry experiments for the manned lunar landing missions).
  • Kits as follows:
    • Payload retrieval and placement adapters
    • A manipulator arm kit.
    • Staging adapters and separation mechanisms.
    • Clustering adapters (to provide for clustering of propulsion modules).
    • Plug-in astrionics for specific mission requirements.
    • Insulation and micrometeoroid kits (for increasing the thermal and micrometeoroid protection of the primary propulsion modules for the extended time of lunar landing missions}.
    • Reaction Control System Booster Kit (to increase the reaction control system thrust for the lunar landing mode}.
    • A landing leg kit (for lunar landing).
    • Radar kit for lunar landing.
    • Auxiliary power supply kit (for lunar surface operations. )

The primary propulsion module was designed for earth orbit and planetary missions. This module would use LOX/LH 2 propellant at a nominal mixture ratio, by weight, of 5 to 1, LOX to LH2. The primary thrust would be provided by an uprated RL-10 engine which would provide a maximum thrust of 10,600 kgf at a specific impulse of 460 seconds. The engine was throttleable over a range of from 10 percent maximum thrust to maximum thrust. It was equipped with an extendible nozzle section which could be retracted 1.6 m to minimize length for transport in the EOS or to minimize the interstage length when tandem stages were required or desired. Utilization of this stage for lunar missions would require application of increased insulation and micrometeoroid shielding; a reaction control system booster kit; and auxiliary power kit; and a landing leg kit.

Expendable drop tanks for very high energy missions would be desirable to minimize the size of the required Space Tug configuration. The expendable drop tank consisted of the same tankage arrangement and pressurization systems as that of the primary module. The insulation was the same as that provided to the primary propulsion module. The items deleted from the primary propulsion module to provide drop tanks include the reaction control system, engine, reaction control system, thrust structure, electrical actuation system for engine gimballing and some of the micrometeoroid protection systems.

The Space Tug's basic characteristics included:

Primary Propulsion Module (45,000 lbs Weight)

  • Propellant Weight - 39,800 lbs
  • Inert Weight - 5,610 lbs
  • Stage Weight - 45,410 lbs
  • Mass Fraction - 0.876
  • (1) Engine - 23,300 lbs thrust
  • Lox/LH2 Isp - 460 Sec
  • Expansion Ratio - 225
  • Subsystems & Construction Materials
    • Structures - 2219 T87 Al, 7075 T6 Al
    • Thermal - Superinsulation
    • Micrometeoroid - Hexcel Filler, 2219 T87 Shield
    • Engines - Rl-10-3-8 (Uprated Rl10)
    • Pressurization - Gaseous Helium for Lox; Gaseous Hydrogen For LH2
    • Activation - Electrical Actuators
    • RCS - Supercritical Lox/LH2
    • Electrical - Batteries And Electrical Networks

Drop Tank Module

  • Propellant Weight 39,800 lbs
  • Inert Weight 4,090 lbs
  • Stage Weight 43,890 lbs
  • Mass Fraction 0.906
  • Lox/LH2 Isp 460 Sec
  • Subsystems
    • Structures - 2219 T87 Al (Tanks) 7075 T6 (Load Structure)
    • Thermal - Superinsulation
    • Micrometeoroid - Hexcel Filler, 2219 T87 Shield
    • Pressurization - GHe For Lox, Gh2 For Lh2
    • Electrical - Batteries And Electrical Network
    • Drop Tank Feed System - 7075 T6 Lines

Gross mass: 20,600 kg (45,400 lb).
Unfuelled mass: 2,550 kg (5,620 lb).
Height: 10.80 m (35.40 ft).
Diameter: 4.27 m (14.00 ft).
Thrust: 99.14 kN (22,288 lbf).
Specific impulse: 460 s.

More... - Chronology...


Associated Countries
Associated Engines
  • RL-10A-3A Pratt and Whitney lox/lh2 rocket engine. 73.4 kN. Isp=444s. Used on Centaur stage atop Atlas G, Atlas I, Atlas II, Titan 4. First flight 1984. More...

See also
Associated Launch Vehicles
  • Proton-K Russian orbital launch vehicle. Development of a three-stage version of the UR-500 was authorised in the decree of 3 August 1964. Decrees of 12 October and 11 November 1964 authorised development of the Almaz manned military space station and the manned circumlunar spacecraft LK-1 as payloads for the UR-500K. Remarkably, due to continuing failures, the 8K82K did not satisfactorily complete its state trials until its 61st launch (Salyut 6 / serial number 29501 / 29 September 1977). Thereafter it reached a level of launch reliability comparable to that of other world launch vehicles. More...
  • Shuttle American winged orbital launch vehicle. The manned reusable space system which was designed to slash the cost of space transport and replace all expendable launch vehicles. It did neither, but did keep NASA in the manned space flight business for 30 years. Redesign of the shuttle with reliability in mind after the Challenger disaster reduced maximum payload to low earth orbit from 27,850 kg to 24,400 kg. More...

Associated Manufacturers and Agencies
  • NASA American agency overseeing development of rockets and spacecraft. National Aeronautics and Space Administration, USA, USA. More...
  • Boeing American manufacturer of rockets, spacecraft, and rocket engines. Boeing Aerospace, Seattle, USA. 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...

Bibliography
  • NASA Report, Technical Study for the Use of the Saturn 5, INT-21 and Other Saturn 5 Derivatives to Determine an Optimum Fourth Stage (space tug). Volume 1: Technical Volume, Book 1, Web Address when accessed: here.

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