Thermal Protection System Analysis and Sizing for Spaceplane Conﬁgurations

; Thermal Protection System Analysis AND Sizing FOR Spaceplane ConfiGURATIONS PrELIMINARY DESIGN OF THE TPS AND OPTIMIZATION OF THE INSULATION LAYER EvELYNE Roorda ii 1 1Cover image from http://www.cbsnews.com/pictures/space-plane-of-tomorrow/8/, visited on 20/06/2017 Thermal Protection System Analysis and Sizing for Spaceplane Configurations Preliminary design of the TPS and optimization of the insulation layer by Evelyne Roorda to obtain the degree of Master of Science at the Delft University of Technology, to be defended publicly on Monday July 17, 2017 at 14:00 AM. Student number: 4044398 Project duration: June 1, 2016 – July 17, 2017 Thesis committee: Dr. ir. E. Mooij, TU Delft, supervisor Ir. A. Kopp, DLR, supervisor Prof. Dr. P. Visser, TU Delft Dr. Ir. D.I. Gransden, TU Delft Ir. K.J. Sudmeijer, TU Delft An electronic version of this thesis is available at http://repository.tudelft.nl/. ii Preface I would like to take this opportunity to thank my two supervisors, Erwin Mooij and Alexander Kopp. Erwin thank you for your understanding and patience during my thesis, and your useful (though sometimes frus- trating) input and comments. Alexander, I would also like to thank you for your patience, as well as your willingness to always help me think about the problems I encountered and your useful suggestions. Furthermore I would like to sincerely thank all who have taken the time to given me their expert opinions and help during my thesis, especially Dominic Dirkx, Kees Sudmeijer, Javad Fatemi and Martin Lemmen. My friends and fellow students have helped keep me sane throughout my study and the thesis process. Thanks for giving me the confidence and courage to keep going. I would specifically like to thank the students of the 9th floor student room for their company and support. Especially Tim, you let me ramble on about my problems and gave me advice that let me in the right direction. And of course Hanneke, for being my friend throughout my entire studies and always helping me out. My friends of JC Babylon have been one of the great things I have gained from my student life in Delft. Kirsten and Nina, I look forward to concurring parts of South-America with you. Lastly I would like to thank my parents and sister. Thank you for your encouragements and understanding. Vati, you have been an inspiration for me throughout my studies. Without your support, financially but mostly mentally, I would not be where I am today. Mutti thank you for always being there for me, to talk to or if I just wanted to come home and be taken care of for a few days. After almost eight years I have come to the end of my studies. It has not been an easy journey for me, and I am proud to have pulled through and make it to the end. The result of my hard work, sweat, and tears is this report. I can only hope that the effort that has gone into it shines through. Evelyne Roorda Delft, July 2017 iii iv Preface Abstract During its lifetime a space launcher is subjected to extreme heat loads. These can dramatically influence the design of the launcher, when parts of the vehicle heat up to very high temperatures. Every part of the vehicle has a temperature range in which it is functional, thus it is important that these limits are respected. There- fore, space vehicles are equipped with a Thermal Protection System (TPS). A properly designed TPS is vital to every reusable launcher vehicle, as an under-designed system will lead to mission failure, and an over- designed system to an increase in mass, resulting in an increase in cost. Heat loads are typically the largest during the ascent or decent phase, where the spacecraft flies through the atmosphere. The large velocities combined with the atmosphere causes atmospheric drag, which results in aerodynamic heating. In this thesis the design of a TPS is considered, focused on spaceplane vehicles. The accompanying research question is: how robust is a thermal protection system design for a spaceplane wing-body configuration to variations with respect to the design parameters or trajectory taking into account heat transfer through radiation and conduction in three dimensions? To answer the research question a tool was developed which is capable of designing a TPS and optimizing the insulation layer thickness. Furthermore, a trajectory simulation was made for the reentry phase. For rocket powered launchers the largest heat loads typically occur during reentry, as in this mission phase the longest time is spent in the atmosphere. From the trajectory specifics the heat flux over the vehicle over time could be generated, which is the source of the temperature increase. In the TPS design tool the first major task is to divide the vehicle into different TPS areas, based on the temperatures that are experienced at each of these areas of the skin surface. Five different TPS areas will be defined, all assigned a passive TPS type. The TPS types are FRSI, AFRSI, TABI, AETB TUFI and CMC. A thermal analysis is performed, taking into account both conduction in three dimensions and radiation to outer space as well as to the inner subsystems of the vehicle. From this analysis the maximum experienced temperature can be deduced for each area. When the TPS design is found, it is aimed to optimize the insulation layer thicknesses in all TPS areas. The goal is to find a design that is as light as possible, thus with a minimum insulation layer thickness, while still meeting the given constraints. These constraints are the maximum reusable temperatures of the TPS types, and the limit functional temperature of the underlying structure. The structure is made of aluminum, and has a maximum functional temperature of 450 K. The maximum temperatures of the TPS types are 644 K, 922 K, 1400 K, 1600 K, and 1850 K respectively. The TPS design tool was applied to a reference vehicle. A sensitivity study was performed to investigate the robustness of the TPS design resulting from the tool. The performance of the TPS design was tested when small changes were made to it, for the nominal reentry trajectory of the reference vehicle. Furthermore the TPS designs performance was analyzed for small changes in the trajectory. From the analysis of the results of the developed tool it was found that a TPS design can be developed for a simple wing-body configuration, under the specified conditions. However, the functionality of the TPS design is limited. Improvements must be made to the developed tool to increase its performance, so that it can come to an acceptable TPS design. It is suspected that with suggested improvements the tool will work properly, and a well functioning TPS design can be made. However, further research is required to ensure this. v vi Abstract Abbreviations 1D one dimensional 2D two dimensional 3D three dimensional AETB Alumina Enhanced Thermal Barrier AFRSI Advanced Flexible Reusable Surface Insulation c.o.g. center of gravity c.o.m. center of mass C/SiC Carbon Fibre reinforced Silicon Carbide CFRP Carbon Fiber Reinforced Polymer CMC Ceramic Matrix Composite DLR German Aerospace Center ECEF Earth-Centered, Earth-Fixed ECI Earth-Centered Inertial EoM Equations of Motion ESA European Space Agency FEI Flexible External Insulation FEM Finite Element Model FESTIP Future European Space Transportation Investiga- tions Program FRSI Flexible Reusable Surface Insulation HOTOL HOrizontal TakeOff and Landing HOTSOSE Hot Second Order Shock Expansion LEO Low Earth Orbit MLI Multilayer Insulation NASP National AeroSpace Plane ODE ordinary differential equation RK4 fourth-order Runge-Kutta RKF45 Runge-Kutta-Fehlberg RLV Reusable Launch Vehicle SART Space Launcher Systems Analysis SOH Suborbital Hopper SPFI Surface Protected Flexible Insulation vii viii Abbreviations SSTO Single-Stage-To-Orbit TABI Tailorable Advanced Blanket Insulation TBD To Be Determined TOP2 Thermal protection system Optimization Program Version 2 TPS Thermal Protection System TSTO Two-Stage-To-Orbit TUDAT TU Delft Astodynamics Toolbox TUFI Toughened Uni-Piece Fibrous Insulation US United States US76 US Standard Atmosphere 1976 USSR Soviet Union VD Voronoi Diagram Nomenclature Greek Symbols Symbol Description Units ® angle of attack rad 1 ®T thermal expansion K¡ Â heading angle rad ± latitude rad ² emissivity - ² error - ²¤ desired accuracy - ²T strain - γ Flight path angle rad γ ratio of specific heats - ¹ gravitational Parameter m3/s Á Euler angle rad Á golden ratio - ½ density kg/m3 3 ½0 density at sea level kg/m σ stress N/m2 8 2 4 σ Boltzmann’s constant 5.667 10¡ W/m K B ¢ ¿ longitude rad Roman Symbols Symbol Description Units A transfer area m2 a speed of sound m/s 2 Ae exit area m CD aerodynamic drag coefficient - CL aerodynamic lift coefficient - Cm aerodynamic pitch moment coefficient - Cp specific heat at constant pressure J/K CS aerodynamic side coefficient - ix x Nomenclature Cv specific heat at constant volume J/K Cmb aerodynamic pitch moment coefficient body flap - D aerodynamic drag force N E modulus of elasticity N/m2 F geometric form factor - F force vector N FA aerodynamic reference frame FB body-fixed reference frame FI inertial planetocentric reference frame FR rotating planetocentric reference frame FV vertical reference frame 3 1 2 G universal gravity constant m kg¡ s¡ g gravitational acceleration m/s2 2 g0 gravitational acceleration at sea level m/s h (geopotential) altitude m hc convective heat transer coefficient - K conductance matrix kc thermal conductivity W/mK L aerodynamic lift force N L0 original length m M Mach number - M molecular mass kg/mole m mass kg m˙ 0 free

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