Launch Vehicle Options Research Articles

Response Surface Equations for Expendable Launch Vehicle Payload Mass Capability

Systems analysis and conceptual design for new spacecraft commonly require the capability to perform rapid, parametric assessments of launch vehicle options. Such assessments allow engineers to incorporate launch vehicle considerations in first-order cost, mass, and orbit performance trades early during conceptual design and development phases. This paper demonstrates an efficient approach to la unch vehicle analysis and selection using response surface equations (RSEs) derived directly from launch vehicle payload planner’s guides. These RSEs model payload capability as a f unction of circular orbit altitude and inclination. Following presentation of the RSE fit ting method and statistical goodness of fit tests, the RSE and model fit error statistics for t he Pegasus XL are derived and presented as an example. In total, 43 RSEs are derived for the following launch vehicles and their derivatives: Pegasus, Taurus, Minotaur, and Falcon series as well as the Delta IV, Atlas V, and the foreign Ariane and Soyuz vehicles. Ranges of validity and model fit error statistics with respect to the original planner’s guide data a re provided for each of the 43 fits. Across all launch vehicles fit, the resulting RSEs have a maximum 90 th percentile model fit error of 4.39% and a mean 90 th percentile model fit error of 0.97%. In addition, of the 43 RSEs, the lowest R≤ value is 0.9715 and the mean is 0.9961. As a result, these equations are sufficiently accurate and well-suited for use in conceptual desi gn trades. Examples of such trades are provided, including demonstrations using the RSEs to (1) select a launch vehicle given an orbit inclination and altitude, (2) visualize orbit altitude and inclination constraints given a spacecraft mass, and (3) calculate the sensitivity of orbital parameters to mass growth. Suited for a variety of applications, the set of RS Es provides a tool to the aerospace engineer allowing efficient, informed launch option trades a nd decisions early during design.

Performance and technological feasibility of rocket powered HTHL-SSTO with take-off assist (aerospace plane/ekranoplane)

It might be said that it is common understanding that rocket-powered single stage to orbit (SSTO) aerospace planes will become feasible with near-term technology as described in [1] (Koelle, D. E. Survey and comparison of winged launch vehicle options, ISTS 94-g-11 V 1994) and [2] (Bekey, I. Why SSTO rocket launch vehicles are now feasible and practical, IAF-94-V.1.524 1994). Among two methods of launching aerospace planes into orbit, vertical take-off (VT) and horizontal take-off (HT), it seems that VT takes the lead from HT [1, 2]. The decision for the X-33 program by NASA, also, seems to favor VT. In retrospect, almost all of the launch vehicles in the past have been VT, mainly because VT solved the problem of exit from atmosphere to space. However, broadening the range of requirements for space transportation systems from military to commercial and unmanned to manned seems to favor the need for HT. In this paper, the authors are going to prove that aerospace plane/ekranoplane system, which is a reusable launch vehicle system based on the HT concept, with ekranoplane as a take-off and possibly, landing assist, could be competitive with the VT concept from both technological and economical view points. Ekranoplane is a wing-in-ground-effect craft (WIG), which moves at a speed of approximately 0.5 M, carrying heavy loads above the sea surface. Combination of high initial velocity and high performance tri-propellant engine for aerospace plane makes it possible to configure an aerospace plane which is competitive with VT. Other specific features of HT in comparison with VT are discussed.

A cost-engineered launch vehicle for space tourism

The paper starts with a set of major requirements for a space tourism vehicle and discusses major vehicle options proposed for this purpose. It seems that the requirements can be met best with a Ballistic SSTO Vehicle which has the additional advantage of lowest development cost compared to other launch vehicle options — important for a commercial development venture. The BETA Ballistic Reusable Vehicle Concept is characterized by the plug nozzle cluster engine configuration where the plug nozzle serves also as base plate and re-entry heat shield. In this case no athmospheric turn maneuver is required (as in case-of the front-entry Delta-Clipper DC-Y concept). In our specific case for space tourism this mode has the avantage that the forces at launch and reentry are in exactly the same direction, easing passenger seating arrangements. The second basic advantage is the large available volume on top of the vehicle providing ample space for passenger accomodation, visibility and volume for zero-g experience (free floating), one of the major passenger mission requirements. An adequate passenger cabin design for 100 passengers is presented, as well as the modern BETA-STV Concept with its mass allocations.

Heliospheric Boundary Exploration Using Ion Propulsion Spacecraft

Anewdeep-spacemissiontotheheliosphericboundaryisanalyzedforspacecraftutilizinglow-thrustionengines. The mission design is performed by optimizing a combination of trajectory variables and propulsion system parameters such that the total trip time to the heliospheric boundary is minimized. Spacecraft utilizing both solar electricpropulsionandnuclearelectricpropulsionareconsidered.Optimalmissiondesignsarepresentedforawide range of launch vehicle options. In addition, a sensitivity analysis of the assumed electric propulsion technology level and gravity-assist e yby conditions is performed. Although electric propulsion is typically associated with payload fraction enhancement, this analysis demonstrates that the use of low-thrust spacecraft results in relatively short trip times for high-energy deep-space missions. Furthermore, it is shown that the performance of both electric-propulsion spacecraft trajectories compare favorably and in many cases show improvement overmissions utilizing all-chemical propulsion systems.

A low cost Mercury orbiter mission

Due to Mercury's small mass and position deep within the solar gravitational well, an orbiter mission poses difficult performance requirements (i.e. ΔV, thermal extremes). However, Yen (JPL) showed that with extended trip times, substantial Mercury missions are feasible. Low cost, quick concept-to-launch, high quality, and reliability are balanced in constraining a small spacecraft mission to Mercury. The primary science objective is high quality multispectral imagery and altimetry. The spacecraft is based on the Clementine spacecraft and sensors. UV/Visual/infrared cameras, star trackers, and a laser altimeter are proposed to determine mineralogical composition, surface structure/morphology, spectral/compositional mapping, and topography. Clementine sensors would require upgrades to enhance thermal dissipation and to increase radiation protection. The first available launch opportunity is in August 1996 and the target orbit is a 300 km polar Mercury orbit. Launch vehicle options are discussed. Options for companion or follow-on spacecraft are presented.

Earth-Mars unmanned transportation systems

This paper discusses the launch vehicle options for the launch of Mars unmanned precursors. More emphasis is put on the possibilities of using China's Long March rockets to make up the Earth-Mars unmanned transportation systems. China's Long March 2E rocket now has reached the payload capability of 8.8 tons to low Earth orbit (LEO). Using existing liquid rocket stages, China has the ability to build up a launch vehicle which could send 20 tons to LEO. The results or recent studies, including performances, trajectory parameters, and preliminary cost estimations are presented. This paper also discusses the concepts of individual Long March stages combined with rocket stages from other nations to build up an international manned Mars mission launch system.