Satellite has various deployable mechanisms for positioning different spacecraft appendages like antennae, arrays, etc. These mechanisms are stowed during launch and deployed on-orbit. During orbital phase of the satellite, mechanisms need to withstand different loads experienced by the satellite. In the present paper, an earth observation spacecraft consisting of a very large deployable antenna along with solar array and dual-axis gimbal drives has been considered. Unlike the conventional satellite loads like thermal excursion and thruster firing, the mechanisms in this satellite experience large induced loads due to deployment of the large antenna. These loads need to be estimated in order to establish adequate design margins in the mechanisms. The induced load is a function of different parameters like location of the appendage joint, center of mass of appendage, mass of appendage, spacecraft mass. In order to take care of these loads, appendage location and center of mass position can be varied by suitably orienting and locating the appendage on satellite. It is essential to take these factors especially the latter, which is the position of appendage, into consideration during the configuration design. The present study caters to determination of these loads induced on solar array and dual-axis gimbal antenna due to deployment of large antenna so as to ensure design adequacy.