Exploiting technological synergies for future launch vehicles

Two launch vehicle concepts based on technologies available today or in a short term future in Western Europe are presented. The design of both launchers has the goal of exploiting synergies with current European programs to limit development and operational costs. Technologies of particular interest here are the high performance solid rocket motors with carbon-epoxy filament wound monolithic motor cases and the future high performance cryogenic expander cycle engine Vinci. The first concept dubbed ANGELA (A New GEneration LAuncher) is a study financed with funds of the German Ministry of Economics and managed by the DLR Space Administration. The project, which started in the summer of 2012 aims at designing a low cost versatile launcher able to place payloads between 2 and 5 tons into GTO. Three architectures have been considered during the first phase of the study. This phase was concluded in March 2013 with the preliminary stagings, which will be the starting point of more detailed analyses. The first architecture is made out of an H110 (stage with 110 tons of LOx/LH2) equipped with two Vulcain 2 engines with shortened nozzles and an H29 propelled by a Vinci engine. In addition the variation of the number of P36 solid rocket boosters allow to reach the entire range of payload performance. The second architecture differs from the first one only by the use of a new staged-combustion engine instead of two Vulcain 2 engines. The new engine, which should deliver 1800 kN in vacuum, allows a reduction of the size of the stages to H90-H24, enhanced with P34 boosters. The third and last architecture is a so called Multi PPH. The first stage is a bundle of 2 or 3 P120 solid rocket motors. The second stage is made out of one single P120, strictly similar to those used for the first stage. Finally the upper stage is an H23 equipped with a Vinci engine, the same as the two other architectures. The second launcher concept described in this paper is the small TSTO launch vehicle. It consists of a large solid rocket motor first stage P175 and a cryogenic upper stage propelled by the Vinci engine, H26. The preliminary design performed at DLR-SART considers two target performances. The light version of the small TSTO shall perform Galileo satellite replacement single launch missions to MTO corresponding to a payload performance of about 1400 kg in GTO. A heavy version of the launch vehicle shall be able to launch payloads up to 3000 kg in GTO. The performance increase for the heavy version is made possible by the addition of two pairs of P23 boosters, the second pair being ignited with a delay

Capsule Rescue Motors for SpaceLiner 7.1

The hypersonic point-to-point passenger transport SpaceLiner 7 under development in DLR Bremen utilizes a capsule design for its emergency escape system and is activated in case of a catastrophic failure of its propulsion systems. In this study, separation motors for the capsule is designed and dimensioned with regards to a worst case scenario. This scenario is identified to occur at the launch pad where maximum fuel is present and the atmospheric pressure is highest resulting in the greatest amount of thrust losses, highest propagation speed and strength of the resulting explosive shockwave. During the design process, the maximum sustainable acceleration on the human body and any geometrical limitations are also considered. Furthermore an estimation of the structural mass of the motors is performed. The resulting separation system utilizes a five motor configuration that is designed for a nominal chamber pressure of 150 Bars and a burn time of two seconds. The motor consists of an end-burning grain geometry, an 80 % length bell nozzle consisting of a carbon-carbon material and a Kevlar based composite chamber. The resulting Structural Indexes are 11.8 % or 13.3 % for each motor depending on chosen overpressure limit. The overall conclusion is that the motors designed in this study fulfill all considered requirements but suggests more work can be performed. This includes considering dangers of fragmentation debris, more detailed structural mass analysis and further investigation of the propellant composition

Optimization of Solid Rocket Grain Geometries

Solid Rocket Motors (SRM) are employed in many space launch applications from the booster rockets on the now retired Space Shuttle to the new European launch vehicle Vega. Preliminary design of these complex three dimensional SRM, given certain requirements and limitations can be considered as an optimization process where a best geometrical solution is sought resulting in a desirable thrust profile. In this project, a derivate free direct search package titled NOMAD is employed together with the internally developed numerical burnback analysis tool SRP-GEO and ballistic solver SRP. An analytical model for the burnback analysis tool is also developed to take advantage of the support for surrogate functions within NOMAD. Due to the local nature of the optimizer, the results for complex geometries are shown to converge toward configurations different from the globally optimal geometries. Yet in most instances, the resulting calculated thrust profiles are shown to correlate well with the desired counterpart. This highlights the importance of carefully chosen initial values and boundaries while also emphasize the many possible solutions to a single problem

Critical Analysis of ESA’s Ariane 6 Linear Multi-Booster Configuration

In 2013 the European Space Agency released a document specifying the requirements of Key Launcher Elements (KLE) for the proposed Ariane 6 launch vehicle based on predominantly solid motor architecture. As a response, this study aims to conduct a critical analysis of the trajectory and payload performance with the given KLE. An aerodynamic and mass model is created for this purpose with the largest/smallest permitted (boundary) values given by ESA. Additionally, the solid motor performance is also evaluated with available data. Critical unavailable data are filled in based on related existing studies or from currently reached consensuses. The resulting payload performance to a Geostationary Transfer Orbit is calculated. A sensitivity analysis on the upper stage propellant loading is performed and compared to the 32 tons stipulated by ESA. The impact on payload by both ESA’s upper stage target mass and a structural index derived from Ariane 5 Mid-Life Evolution upper stage is assessed

Visualization of Solid Rocket Motors with VRML 2.0

The Solid Rocket Propulsion (SRP) software package developed within DLR Bremen is utilized for modeling and optimizing three dimensional solid rocket motors. As part of the computation process, a conceptual 3D geometrical representation of the motor is created. However, this geometry has so far existed solely as a set of parameters in a text based file. This study thus aims to create an additional program module, written in object oriented Fortran 2003 that can convert the text file containing arbitrary SRP geometries into three dimensional models following the VRML 2.0 format. A central part of the study is the development and verification of novel mathematical equations for converting cylindrical and star propellant grain cross sections into Cartesian coordinates. These equations are incorporated into the resulting program module. The program then utilizes an existing meshing tool named GGH to generate the VRML 2.0 compliant input files from the given Cartesian coordinates. These input files can be opened through any VRML compliant viewers or imported into CAD programs such as CATIA. In conclusion, the program module was successfully developed with all core requirements fulfilled. Thorough documentations of the source code and user instructions have been created and a potential future development direction of the program has been identified

Process Chain Development for Iterative, Concurrent Design of Advanced Space Transportation Systems

Development of advanced Space Transportation Systems (STS) at an early definition phase requires an iterative and multidisciplinary approach. High sensitivity and coupling of subsystem parameters with overall system characteristics necessitates close collaboration between domains. Rapid transfer of large amounts of data is necessary while errors and inconsistencies must be avoided. Concurrent Engineering (CE) is one approach through which to automate data sharing and minimise communication overheads. Previous effective implementation of this method at the German Aerospace Centre (DLR) Concurrent Engineering Facility (CEF) was limited to early-phase satellite design activities. Applying CE methods to STS design have been met with limited success. This paper reports on a new initiative from DLR titled Collaborative Launch vhicle Analysis (CLaVA). Within the scope of CLaVA, a flexible design environment for STS is investigated applied to the DLR CEF. The investigation led to the development of a high-level design cycle for STS. In collaboration with the aeronautical divisions of DLR a solution was identified, employing a central data model based on Model Based Systems Engineering (MBSE) methodology, implemented within a distributed environment. With modern revision control systems, it is envisioned that multiple STS concepts will be analysed in parallel in the CE environment. A preliminary test of these concepts is planned through a mock CE study

Next Generation Launcher-Potential Ariane 5ME and Ariane 6 Common Upper Stage Options

This study compares the performances of Ariane 5ME and Ariane 6 (for two different configurations: PPH and PH) with common upper stages. In that perspective, the investigated upper stages are ESC-B (Perigee Decrease and Direct Deorbiting Versions) and newly defined H31 and H34. Ariane 6 is modeled based on one or two P135 stages (serial) and similarly, two P135 boosters

Reverse Engineering of Athena II Launcher

This report gives details about the model of the Athena II launch vehicle that was created with the goal to use it in the frame of the AtILa study. The small and medium payload launch vehicle Athena II is analyzed with respect to its performance, mass, launch trajectory and orbit characteristics. The motor efficiencies of the first and second stage Castor 120 motor are calculated and compared to available public data. In addition, the third stage thrust and chamber pressure histories are also reverse engineered. A simplified mass estimation is then performed for each stage. Lastly, utilizing the calculated mass breakdown and performance characteristics, a trajectory and orbit analysis is conducted assuming a circular sun-synchronous orbit at 687 km (corresponding to the particular mission studied in the frame of the AtILa study). The results show that the vehicle is capable of launching 1198 kg of payload to this orbit, a difference of 4 % compared to value from available literature and where part of the discrepancy can be attributed to some inaccuracy and discrepancies in the published data. The analysis conclusively proves that the available data for the Athena II launch vehicle are indeed accurate and the reverse engineering process is successful. The calculated trajectory is also considered as representative of the real flight and is aimed to be used for the AtILa project

NewFex Mission Analysis

After the end of the Shefex III project the reentry experiment required a new assessment of its scientific goals that are compatible with the preparation of future launch systems a role that may be assumed by reusable launch systems. In order to prepare decisions on the main orientation for a new reentry project, named NewFex as a working name, some technical work was conducted. The present report summarizes the work conducted since the second half of 2014 within the Launcher System Analysis department SART concerning mission analysis and, when required and possible, on trim capability for NewFex

Critical Analyses and Performance Validation of the SKYLON D1 Space Plane

DLR-SART has performed a critical analysis of the Reaction Engines Ltd. single-stage-to-orbit (SSTO) re-usable space plane concept, the SKYLON D1 vehicle, and its innovative dual-mode engine, the Synergistic Air Breathing Rocket Engine (SABRE). This analysis focused on the validation of the SKYLON mass budget, system performance and characteristics. A preliminary mass breakdown of the SKYLON D1 vehicle was determined using parametric methods, based on vehicle performance parameters and geometric data provided by REL. The margin philosophy provided by REL was found to be insufficient, and it should be re-iterated. A mass breakdown of the SKYLON Upper Stage (SUS) was also generated using empirical methods based on performance and geometric data provided by REL. Subsystem masses were scaled from the Ariane 6 PPH studies. The ascent trajectory of the SKYLON vehicle was modelled in three parts; the air-breathing ascent, using a comparable turbo/ram engine; the rocket ascent using the SABRE engine; and the SUS flight. It was determined that the SKYLON vehicle with the DLR-SART mass estimates could not achieve orbit, however an optimized trajectory was found for the vehicle using REL mass assumptions. Finally, the performance of the SABRE engine was determined. The conventional rocket-mode performance of the SABRE was found to be highly consistent with REL-provided data, however due to convergence issues related to the control of the innovative pre-cooler used in the air-breathing mode, the full air-breathing cycle could not be modelled. An assessment of the development cost, production cost and launch service cost has been performed using a parametric method and considering different scenarios. Both the SKYLON and SABRE are highly non-conventional, and the need has been highlighted to create analytical models and tools that are compatible with the dual-mode mission profile, space-frame structure and innovative propulsion system for future analyses. It was also seen that the DLR-SART methods for determining sub-system and component masses are based on existing aircraft and launch vehicles and studies, and therefore do not always consider new solutions or technologies. The structure and thermal protection systems are key enabling technologies, not just the propulsion system. New solutions will need to be developed and implemented for these systems