AJCPP 2008 March 6-8, 2008, Gyeongju, Korea Conceptual Design Trade Offs between Solid and Liquid Propulsion for Optimal Stage Configuration of Satellite Launch Vehicle Qasim Zeeshan1 Dong Yunfeng2 1PhD candidate, Beijing University of Aeronautics & Astronautics, BUAA, P.R.China 2Professor, Beijing University of Aeronautics & Astronautics, BUAA, P.R. China School of Astronautics, BUAA, 37 Xue Yuan Road, Beijing 100083, P.R.China [email protected] Keywords: Launch Vehicle, Liquid Propulsion, Mission Design, Optimization, Solid Propulsion, Trajectory Abstract proportional to the GLOW. The boost phase of an orbital space mission is critical to the design process The foremost criterion in the design of a Satellite since the initial weight of the vehicle, which is Launch Vehicle (SLV) is its performance capability to generally related to its cost, is at its maximum at lift- boost the designated payload to the desired mission off. The mpay is a small fraction of the total mass. orbit; it starts from focusing on the SLV configuration Placing an object into orbit is technically to achieve the velocity requirements (ΔV) for the demanding and expensive too. A rough rule of thumb mission. In this paper we review an analytical is that a modern SLV can deliver into orbit a mpay that approach which is suitable enough for preliminary is only a few percent of its overall mass. Since the size conceptual design and is used previously to optimize of the SLV scales with the mpay, there is a tremendous stage configurations for Two Stage to Orbit SLV for incentive to increase the SLV performance by keeping Low Earth Orbit (LEO) Missions; we have extended the mass of SLV as low as possible. We will later this approach to Three Stage to Orbit SLV and discuss that apart from GLOW there are other figures compared different propellant options for the mission. of merit too which strongly dominate and effect the The objective is to minimize the Gross Lift off Weight design and performance of a SLV. (GLOW). The primary performance figures of merit Exiting SLVs can be categorized as: were the total inert weight of the SLV and the payload • Any number of stages, inline and/or piggy-back weight that the SLV could lift into LEO, given • Liquid, solid, hybrid, air breathing propulsion candidate propulsion systems. The optimization is systems. achieved by configuring the ΔV between stages. A • Mono-propellant, bi-propellant, tri-propellant comparison of configurations of single-stage and • Pump-fed or pressure-fed multi-stage SLVs is made for different propellants. • Expendable or recoverable Based upon the optimized stage configurations a The research work is in progress for the design of comparative performance analysis is made between single-stage-to-orbit (SSTO), but no Earth-launched Liquid and Solid fueled SLV. A 3 degree of freedom SSTO SLV has ever been constructed. Current orbital trajectory-analysis program is modeled in SIMULINK launches are either performed by multi-stage fully and used to conduct the performance analysis. expendable launchers or by the Space Shuttle which is Furthermore, a cost analysis is performed on our stage multi-stage and partially reusable. It is extremely optimized SLVs. The cost estimation relationships difficult to design a structure which is strong, safe, (CER) used give us a comparison of development and very light, and economical to build. The problem fabrication costs for the Liquid vs. Solid fueled SLV originally seemed insuperable, and drove all early in man years. The pros and cons of the production, designers to multistage rockets. In this study multi- operation ability, performance, responsiveness, stage expendable SLVs are considered for small mpay logistics, price, shelf life, storage etc of both Solid and to LEO. Liquid fueled SLVs are discussed. The statistics and SLVs are unique forms of transportation since they data are used from existing or historical (real) SLV are the only systems that accelerate continuously stages. throughout their performance envelope. Consequently, velocity is the fundamental measure of performance Introduction for any SLV. The ability of any SLV to achieve orbital velocity comes primarily from its propulsion It is well known that the aerospace industry is more efficiency, with weight and drag acting against it. sensitive to vehicle weight as a primary figure of merit Available technology constrains the specific impulse for vehicle designs than any other industry. This is (Isp), therefore the trade off has to be made. For each because weight (or mass) is a strong driver on vehicle propulsion stage, the amount of work that needs to be performance and cost, and so takes a central role in done is calculated. This is measured in terms of the vehicle design process. In fact, vehicle weight is velocity changes, ΔV. We can estimate the velocity so important that competitive advantage is often that SLV should provide. sought largely or exclusively on the basis of having a lighter weight than the competitor. Experience has ΔVdesign = ΔV burnout + ΔV gravity + ΔV drag (1) shown that the cost of a given class of SLV is roughly 283 AJCPP 2008 March 6-8, 2008, Gyeongju, Korea ΔVburnout is the velocity required for the desired orbit. Table 1 6) We add the losses, ΔVgravity and ΔVdrag, to the ΔVburnout Velocity Budgets to LEO for selected SLVs Σ ΔV to obtain the required ΔVdesign. These losses are V ΔV ΔV ΔV ΔV sensitive to the T/Wo. A low T/Wo ratio causes gravity Vehicle = losses to be high because the vehicle spends more LEO GRAVITY STEERING DRAG ROT* ΔV PROP time in ascent, while high T/W causes drag losses to o Ariane be high because of the higher velocities achieved in 7802 1576 38 135 -413 9138 A-44L the atmosphere. The T/Wo is a key SLV design parameter because it dictates the vibration, acoustic Atlas I 7946 1395 167 110 -375 9243 and dynamic load environment for the spacecraft. If there were no atmosphere and topographical variations, Delta 7842 1150 33 136 -347 8814 an optimum SLV trajectory would be very similar to a Hohmann transfer and gravity losses would be Saturn V 7798 1534 243 40 -348 9267 minimized by thrusting normal to the radius vector. Titan IV/ To accurately estimate gravity losses we need to know 7896 1442 65 156 -352 9207 Centaur a precise ascent profile and time of flight (TOF). But Space for medium-to-large SLVs on nominal trajectories the 7794 1222 358 107 -395 9086 ΔVgravity losses fall between 750 and 1,500 m/s. For Shuttle the current inventory of large, expendable launch vehicles, ΔV drag losses are less than 3% of the total For different propellant types, historical or empirical change in velocity required, about 20 to 40 m/s. The data can relate a vehicle's inert mass to the propellant percentage decreases as the size of the vehicle mass required. To do this, we use the concept of the decreases. Once we know ΔVdesign from mission inert mass fraction requirements, we can estimate the mprop required for 1, 2) the SLV. Several definitions are useful at this finert = ( minert ) / ( mprop + minert ) (2) point. The flight vehicle mass is the sum of the propellant mass, structure mass, including mass of the The inert mass includes everything except the payload fairing, and the mass of everything above the launch and propellant masses. The inert mass usually includes vehicle interface, including mass of the spacecraft bus, tank structure, support structure, engines, the 3) payload, and any upper stages. The total mass of a propellant feed system, fairings, electronics, and any multistage SLV includes the masses of propellants and number of other non-reactive (inert) components. We their tanks, engines, feeding system and mpay. We use use historical data to predict a typical inert mass mass fractions to describe the portion of the flight fraction. Historical numbers only predict what has vehicle devoted to certain sections. The propellant been done in the past. New technologies, new mass fraction is the mass of propellant divided by the technology applications, or new requirements may total flight vehicle mass; the structure mass fraction or change the number drastically. We further assume that deadweight fraction is the structural mass, including the inert mass is only a function of propellant mass. the mass of the fairing, divided by total flight vehicle However, such factors as the thrust level and pressures mass; and the payload mass fraction is the payload (i.e., propellant storage pressure, combustion chamber mass divided by total flight vehicle mass. Typical pressure, pump pressure) can significantly affect the values for propellant, structure, and payload mass mass of the structure and, in particular, the mass of the 4,5) fractions are 0.85, 0.14, and 0.01, respectively. The engine. With all of this, the finert approach is quite structural weight of the vehicle, on the other hand, is powerful and can give us very good preliminary much more difficult to estimate, but it depends on the design results. Using the ideal rocket equation, we get weight of fuel carried. Using the data for stages from operational SLVs with different propellant ⎡ ⎛ ΔV ⎞ ⎤ combinations and including the engines, the vehicle mPAY ⎢exp⎜ ⎟ −1⎥ ⎜ gI ⎟ (3) configuration is estimated. The reduction in GLOW ⎣ ⎝ SP ⎠ ⎦ m prop = (1− f inert ) achieved by using three stages instead of one is 25%, ⎛ ΔV ⎞ − f exp1 ⎜ ⎟ while the reduction in weight compared to two stages inert ⎜ ⎟ ⎝ gI SP ⎠ is about 8%. Whenever we encounter missions requiring a large In the formulation below, we find there is a "not ΔV, we run the risk of not being able to perform the feasible" condition that gives us a relationship missions with certain technologies.