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10-Meter Mars Capture Mission Orion by William-Black 10-Meter Mars Capture Mission Orion by William-Black
Source: GA-5009 Volume 3 Nuclear Pulse Space Vehicle Study Vol. III Conceptual Vehicle Designs and Operational Systems

This is the first in a new extended series on Project Orion which will include the large and small Orion's, the USAF military application Orion's, interplanetary exploration spacecraft, and chemical launch vehicles. The military application series will include both real, worked designs, weapons systems, along with speculative designs based on existing studies, along with a full work up of the orbital fleet of nuclear pulse and fleet support spacecraft and stations as the USAF Space Plan 1962 might have required. There will be diagram style images, such as this rendering, along with artistic renders. For full access to the entire series sign up today at WilliamBlack @ Patreon.  

Individual full resolution art is available for $10.00, see my profile page for details or write wblackarts@gmail.com

The exploration missions considered in GA-5009 have durations varying from 150 days for a fast Mars round trip to 910 days for an exploration of Jupiter. The more typical Mars and Venus missions varied from 300 to 450 days.

The vehicle is shown ready to depart earth orbit on a 450-day Mars orbital-capture mission, which requires a velocity increment of 72,850 fps for the complete round trip. This configuration carries only 750 kg of destination payload (mapping equipment plus a data-handling and storage system) and therefore will not support a Mars landing.

The two space taxis carried by each vehicle were intended for liaison between vehicles in a two-vehicle convoy and for inspection of the nuclear pulse engine pusher plate between propulsive burns. On most of the configurations they are located atop the personnel accommodations compartment where entrance airlocks protrude from the access-ways. Placing them at this higher location avoids using space required for propellant magazines or external payload on the more heavily loaded vehicles.

The reentry vehicle shown in the illustration is assumed to be capable of a successful earth reentry from an approach speed of 50,000 fps. For this capability it carries 3,000 kg more structure, ablative material, etc., than a similar reentry vehicle intended for a 36,300-fps approach (approximately Apollo reentry speed). Below the reentry vehicle is a stubby maneuver stage (capable of approximately 1,000-fps AV) which (when attached to the vehicle) is used to initiate the reentry.

The payload spine has a minimum length of 12 m (39.4 ft), since this length provides a minimum radius of 50 ft from the normally manned personnel accommodations to the composite CG for artificial gravity purposes. The first designs of this concept during the study had a shorter spine whenever the space required for propellant magazines or external payload permitted and employed a coast-period spine extension to attain the 50-ft minimum radius. Provisions for the extension, however, were estimated to be about as heavy as a longer fixed spine and introduced another reliability problem (the longer spine also provides a radiation shielding advantage because of the increased separation from the pulse source).

Note: Propulsion module is sized for suborbital start-up. Two factors affect propulsion module length and geometry, payload mass and initial thrust. Scott Lowther provides an in-depth description of different boost modes in Aerospace Projects Review ev1n5. I’ve illustrated a Mode I propulsion module (25.9 meters (85 feet) in length, vehicle stack 53 meters (173.8 feet)). In Mode I a single liquid or solid chemical rocket stage (solid rocket launch would require a cluster of large, 156 inch, diameter boosters) would loft the fully loaded crewed Orion onto a suborbital “lob,” to an altitude of about 90 km and velocity of about 900 meters/second. The Orion engine would then either boost the vehicle into orbit or directly onto an interplanetary trajectory.  

By the time of this September 1964 study earlier plans for nuclear ground launch of large nuclear pulse spacecraft had been set aside in favor of rocket boosted launch. As Scott Lowther notes, in Aerospace Projects Review ev1n5, studies underway at the time by Boeing and NASA to make the Saturn V first stage fully recoverable/reusable might have made this an economically attractive option. Development of the large 4,000 to 10,000 ton Orion’s was deferred in favor of the lower developmental risk (which came at the greater engineering challenge) of 10, 12, and 20 meter Orion’s. The Convair NEXUS, or a booster of equivalent class, would be required to loft the 20-meter and larger 4,000 tone Orion’s. (Note: the meter designation refers to pusher plate diameter which is how Orion spacecraft are classed.)  The political and financial cost of failure with a spacecraft the size of a naval battleship would be high, it made a certain amount of sense to initiate the program with the smaller diameter spacecraft before proceeding to development of the more ambitious vehicles.

In discussion regarding this post Winchell Chung of Atomic Rockets, pointed me to this page, 10-Meter Mars Mission, which links to a study by Paul R. Shipps Manned Planetary Exploration Capability Using Nuclear Pulse Propulsion. Basically it shows how an Orion-powered Mars mission is so superior to a chemically powered mission that it just isn't funny.

Related Art:

20-Meter Orion Jupiter Moon Landing Mission
Nuclear Pulse Engine Operation Diagram
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