Payload deployment from shuttle with linear and angular velocity

#1. Field of the Invention

#This invention relates generally to arrangements for launching payloads, such as spacecraft or satellites from launch vehicles and, more particularly to launching arrangements in which the payload is gyroscopically stabilized by spinning at the time it is launched or ejected from the launch vehicle.

#2. Description of the Prior Art

#The invention, including its various objects, features, and advantages, may be more readily understood with reference to the following detailed description of two embodiments thereof, taken in conjunction with the accompanying drawings, wherein like reference numerals designate like structural elements, and in which:

#FIG. 1 is a drawing of a space shuttle showing the payload bay doors open, and having a portion broken away to show a spacecraft in a supporting cradle mounted therein;

#FIGS. 2a and 2b schematically illustrate the ejection of the spacecraft from the payload bay of the space shuttle;

#Referring to FIG. 1 of the drawings, there is shown a space shuttle 50 developed by the National Aeronautics and Space Administration. The space shuttle 50 is shown with its payload bay doors open, and a portion is broken away to show a spacecraft 51 mounted in the payload bay by means of a cradle 52. The cradle 52 has trunnions on both sides adjacent its upper ends. Only trunnions 65a and 65b are visible in FIG. 1. These trunnions are carried by the longerons 49 (shown in dot-dash outline) of the space shuttle, supporting the cradle 52 in the payload bay of the space shuttle 50.

#The spacecraft 51 is stowed in the bay of the space shuttle 50 with the spin axis parallel to the roll axis of the shuttle 50. Thus, the attitude of the spacecraft 51 is set by the attitude of the shuttle 50 and is maintained during ejection by simultaneously imparting both linear and angular momentum to the spacecraft 51. The spin provides gyroscopic stabilization. FIGS. 2a and 2b show the spacecraft 51 before and during ejection from the cradle. Integration of the spacecraft and cradle provides a pivot axis 53 (see FIG. 3) at one side of the spacecraft 51, paralleling the spacecraft's central axis, and provides an ejection mechanism 54, a spring, or piston or other thruster, at the other side of the spacecraft 51.

#As will be seen in FIG. 3 the specific arrangement involves a three-point ejection system including two longitudinally spaced pivot points 160 and 161, defining pivot axis 53, supported by the upper ends of the lefthand side of the cradle 52 as viewed. A single push off point is provided at the upper end of the right side of the cradle 52 at the point 162. The three points 160, 161 and 162 define a triangle which includes or encloses the center of mass CM as indicated, the single push off point being the vertex or apex of the triangle at 162. The longitudinally spaced pivot points 160 and 161 provide spaced reaction points which carry the reaction loads during the satellite ejection interval and thus remain engaged as long as an ejection thrust force exists at the push off point 162, which lifts that side of the spacecraft from the cradle. After the deployment force ceases, the spacecraft contact force at its pivot points 160 and 161 also ceases and the spacecraft simply lifts off while simultaneously rotating, the direction of separation being normal to the plane defined by the pivot points and the spacecraft center of mass at force termination. Free body spacecraft rotation is about its center of mass. Separation velocity and rotational speed depend upon satellite intertia, ejection force and stroke length as well as spacecraft diameteral dimensions. The attitude precision of the spacecraft after separation from the shuttle depends upon shuttle stability during spacecraft ejection, structural rigidity of pivots and their supports and fuel motion.