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Launch Vehicle First Stage Reusability a Study to Compare Different Recovery Optionslaunch Vehicle for Am Launch Vehicle First Stage Reusability Launch Vehicle First Stage Reusability a study to compare different Technische Universiteit Delft recovery options for a reusable launch vehicle M. D. Rozemeijer Launch Vehicle First Stage Reusability a study to compare different recovery options for a reusable launch vehicle by M. D. Rozemeijer to obtain the degree of Master of Science at the Delft University of Technology, to be defended publicly on Tuesday December 1st, 2020 at 09:00. Student number: 4141733 Project duration: April 23, 2019 – December 1, 2020 Thesis committee: Ir. B. T. C. Zandbergen, TU Delft, supervisor Dr. A. Cervone, TU Delft Ir. M. C. Naeije, TU Delft An electronic version of this thesis is available at http://repository.tudelft.nl/. Preface This report concludes, and is the culmination of, my Master thesis to complete the master curriculum in the field of Aerospace Engineering. Besides this, it also, hopefully, answers part of the question on how to make access to space easier. I have been working in the field of entry descent and landing since the start of the Stratos III project in 2016. This is also the start of the ParSim tool, which is the base of my thesis tool. Back then it was only able to recover part of the Stratos III rocket. This tool expanded to encompass the demands for the other projects I have been working on within Delft Aerospace Rocket Engineering (DARE), which are Project Aether and the Supersonic Parachute Experiment Aboard Rexus (SPEAR) all within the Parachute Research Group. Besides the tool, working on these projects increased my knowledge of Entry, Descent and Landing tremendously. This all led to the publication of several articles at the AIAA 2018, IAC 2018 and EUCASS 2019. I would like to thank B.T.C. Zandbergen for supervising my thesis, giving me the push to look a bit further sometimes and other times getting me back on the right course with his feedback and ideas. I would like to thank my teammates from the Parachute Research Group for the research over the years. As well as for some of the data used in this research. In particular, I would like to thank Lars Pepermans & Thomas Britting for helping in developing ParSim which is the basis of the tool build for this research. Lastly, I would like to thank my friends and family for supporting me during the whole of my studies and pushing me to where I am now and supporting me through the tough and difficult year that was 2020. M. D. Rozemeijer Delft, November 4th, 2020 iii Summary Going to space is an expensive endeavour, even the cheapest option costs e2719 per kilogram of payload to a low earth orbit (LEO). Since the cost of satellites is decreasing, the demand for low-cost launches is in- creasing. Currently, for most of the launchers, the first stage is expendable, meaning they are only used once. This seems a waste of such an expensive element costing in the millions of euros. Literature indicates that potentially the launch cost be driven down by at least 30% when reusing the first stage at least ten times, but proof is not yet substantial. In addition, although many investigations have been previously performed they mostly focused on only one option. In this research more than one option will be considered. Questions tack- led in this work are: What the best method is to decelerate the stage? What is the best way for a landing the stage? Is it better to only recover the engine or the complete stage and does the target orbit make a difference? To make a stage reusable several elements are needed. First, a deceleration system is needed. In this research four systems are considered including parachutes, both sub- and supersonic, a hypersonic inflatable aerody- namic decelerator, grid fins and/or propulsive landing. For the landing system, both airbags and landing legs are used. To do so, a tool is made that encompasses multiple models. These are mostly mass and cost models for the various deceleration and landing systems. But also models to be able to increase the propellant tank and to predict the thrust and the drag forces accurately. All these models are included in a tool for which a surrogate optimisation is used to determine which deceleration and landing system produces the cheapest option. Thereby taking into account not only the production cost but also refurbishment and retrieval costs. For design optimisation, surrogate optimisation method is used. The surrogate optimisation is a relatively new optimisation algorithm that creates a surrogate function which is minimised and evaluated in order to find the global optimum. Although less accurate than the genetic optimisation, which is the benchmark global optimisation algorithm, it can produce a result in 20% of the time. Although it is significantly faster, it does not produce the same result as the genetic algorithm but a solution within 1% of the genetic optimi- sation algorithm. The difference in the objective function is accepted, because of the difference between the computation time which is 20min-60min for the surrogate optimisation and 2-26hrs for the genetic optimi- sation. Two vehicles are investigated for two different target orbits. These are the Falcon 9, which is designed with reusability in mind, and a modified version of the Delta IV, the Delta IV+, which is an expendable launcher. The two orbits are a circular low earth orbit and a geostationary transfer orbit. The payload mass for the Falcon 9 is 15600 kg and 6500 kg, for LEO and GTO respectively and the payload masses for the Delta IV+ are 9000 kg and 4500 kg for the LEO and GTO respectively. These payloads are the highest proven payloads flown. For the four combinations of vehicle and orbit, three separate cases are investigated. Two cases where the complete stage is recovered, using either retro-propulsion or non-propulsive means. The last case is one where only the engine is recovered using non-propulsive means. For all of the cases, the descent deceleration needs to be lower than 10 g and have a maximum dynamic pressure below 200 kPa. For all the cases, it was found that 30% cost savings could not be achieved. For the Falcon 9 the savings came closest to the 30% but did not reach it. This 28% cost saving was achieved by reusing the complete stage, the difference between non-propulsive and retro-propulsion was minimal, but the non-propulsive way was slightly cheaper. There was no significant difference between LEO and GTO missions since for the Falcon 9 the second stage produced more than 50% of the ¢V of the launcher. This leads to a separation velocity around 2000-3000 m/s where the orbital velocity is 7700 m/s for a LEO and 10100 m/s for a GTO. For the Delta IV it was seen that reusing the complete stage becomes more expensive than reusing only the engine. This is due to the high separation velocity, which is in the range of 3500-5000 m/s, these are signifi- cantly higher than for the Falcon 9. Recovering from these separation velocities, requires a larger and there- fore heavier system to decelerate the first stage. But by making the system larger it becomes more expensive. Furthermore by making it larger the risk of going over the deceleration limit is higher. The maximum savings achieved for the Delta IV is 20%, this is for recovering only the engine. Again here there was little difference between the solution for LEO and GTO. v vi Summary So overall, 30% cost savings could not be achieved. However, there can still be significant savings by reusing the first stage. From the results, it became clear that using a Hypersonic Inflatable Aerodynamic Decelerator and a Ringsail parachute is the best option to decelerate the system. For the landing system, both airbags and mid-air retrieval were used for the optimal solution. Furthermore, it was seen that the deceleration constraint is the leading constraint, and the dynamic pressure limit comes close to the limit when the stage has a small reference area for the drag. There was little difference between LEO and GTO for the two vehicles. However, there were substantial differences between the two vehicles. Interestingly, the maximum savings are achieved between 8-15 reuses meaning that more reuses is not always better. The reason for this minimum is because the expendable case decreases in cost the more launchers are produced. On top of this, the refurbishment cost increases with the number of reuses. This results in the minimum point around ten reuses. Contents Summary v List of Figures xi List of Tables xiii Nomenclature xiv 1 Introduction 1 2 Background 3 2.1 Reusable First Stage Overview...................................3 2.1.1 Launch...........................................3 2.1.2 Deceleration Systems...................................4 2.1.3 Landing & Retrieval....................................5 2.1.4 Refurbishment.......................................6 2.1.5 Production.........................................6 2.1.6 Complete Overview....................................6 2.2 Comparable Launchers......................................7 2.3 Reuse-index............................................8 2.4 Optimisation Algorithm......................................8 3 Models 11 3.1 Extra Tank Mass.......................................... 11 3.1.1 Model........................................... 11 3.1.2 Model Validation...................................... 12 3.2 Landing System.......................................... 15 3.2.1 Landing Legs........................................ 15 3.2.2 Airbags........................................... 16 3.2.3 Flotation Device...................................... 17 3.3 Deceleration Systems....................................... 18 3.3.1 Grid Fins.......................................... 18 3.3.2 Parachute Mass...................................... 19 3.3.3 Hypersonic Inflatable Aerodynamic Decelerator Mass................... 20 3.4 Thrust Model........................................... 21 3.4.1 Model Description..................................... 21 3.4.2 Model Validation...................................... 22 3.5 Drag Model...........................................
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