36 IEEE A&E SYSTEMS MAGAZINE OCTOBER 2013 INTRODUCTION Researchers, using optimal control theory techniques in solving trajectory problems, have found successful applications in de- sign process. Moreover, many researchers have designed solid propulsion transfer vehicles using various techniques (NASA) and also selected different trajectory analysis techniques such as the Hohmann transfer [1] that obtain minimum fuel for im- pulsive maneuvers between two coplanar and coapsidal orbits. In these types of approximations, many solutions are available, like the “Bi-Elliptical Transfer” [2]. Assuming finite thrust that is more realistic is becoming more popular. Researchers such as Tsien [3], Lawden [4], Biggs [5], Ceballos and Rios-Neto [6], Rios-Neto and Bambace [7], worked on this case. The space system design environment is challenging and chaotic because even very small changes may have sig- nificant effects in unforeseen ways. The design environment contains many interrelated tasks, and the ways these tasks are treated have great consequences on time, cost, and suc- cess of the design process. This article proposes a new design method for a trans- fer vehicle, comprised of a solid propulsion system. Such a vehicle would be used to transfer parts, supplies, and crews between LEO and interplanetary trajectories. Several con- figurations are being investigated, both liquid and solid engines. One concept is an unmanned single-stage transfer vehicle that is solid propulsion powered. Solid propulsion systems are often preferred over liquid engines because of their simplicity, proven design, cost efficiency, and high reli- ability. Unlike liquid engines, solid propulsion systems can- not throttle, and the only way to change thrust magnitude is predesigning the engine in a way that produces the desired thrust profile proportional to the grain burn area. The sub- sequent parts of this article will focus on the integration and role of various disciplines in the design environment. STATEMENT OF PROBLEM The most important factor in designing space systems is trajectory analysis. Traditionally, this task has been accom- plished by using a model with a constant thrust or predes- ignated thrust profile or even considering impulsive ma- neuvers, specifically dedicated to each particular problem. Due to the relatively poor robustness of some assumptions (such as selecting one control variable or considering im- pulsive maneuver), in this article we introduce a new finite burn assumption in trajectory analysis, and other disciplines are modified to compensate this assumption. Propulsion analysis is modified in a way that delivers optimum values of burn time, thrust magnitude history, and thrust direction history. This particular combination, using a thrust finite burn assumption, has two control variables that are more re- alistic in finding a potentially interesting initial position and burn time and presents great novel ideas in the design en- vironment. Designing an upper stage for the interplanetary mission is selected for the test case, and results are compared with the STAR-48B upper stage [8]. Several constraints are considered on the design, such as thrust limitations, feasible thrust profile, and thrust angle limitations. A considered launch vehicle was sent off from a specified launch site and transferred the upper stage and payload to a LEO parking orbit; then at the correct time, the upper-stage engine ignited and transferred the payload to interplanetary orbit. Figure 1 presents a schematic of the considered mis- sion. The transfer vehicle must be able to carry the 1062-kg payload from the LEO parking orbit to hyperbolic orbit. All results are compared with NASA’s Mars Rover mission [9]. Payload mass was selected for comparison between this ap- proach and Star-48B design outputs. DESIGN OBJECTIVE In the design of space propulsion systems, the minimum total mass concept has traditionally been viewed as driv- ing system development and operating cost, which is pro- portional to fuel mass. The goal of this effort is to minimize the fuel consumption of the transfer vehicle through a finite burn assumption under mission constraints and solid pro- pulsion system envelope constraints. Consequently, we try to configure an optimum propulsion system based on op- timized finite burn fuel consumption to achieve our major goal of minimization of total mass in the transfer phase. New Approach in Designing Solid Upper Stage for Interplanetary Missions Using Finite Burn Assumption Javad Amiri Motlagh, Alireza Basohbat Novinzadeh, Mostafa Zakeri K.N. Toosi University of Technology, Iran Authors’ current address: J. A. Motlagh, A. B. Novinzadeh, K.N. Toosi University of Technology, Aerospace Engineering, East Vafadar St., 4th Tehranpars Sq., Tehran, 16765-3381, Iran, e-mail: javadamiri.m@gmail.com, novinzadeh@kntu.ac.ir. M. Nosratollahi, Space Research Institute, Tehran, 16765-3381, Iran, e-mail: m_nostratollahi@sbu.ac.ir. Manuscript SYS- AES-2011-0156 received December 8, 2011, revised May 15, 2012 and October 3, 2013, and ready for publication January 2, 2013. Review handled by M. De Sanctis. 0885/8985/13/ $26.00 © 2013 IEEE