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Supersonic Retro Propulsion Flight Vehicle Engineering of a Human Mission to Mars

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Supersonic Retro Propulsion Flight Vehicle Engineering of a Human Mission to Mars Supersonic Retro Propulsion Flight Vehicle Engineering of a Human Mission to Mars Hanna Marklund Space Engineering, master's level 2019 Luleå University of Technology Department of Computer Science, Electrical and Space Engineering Supervisor Dr. R. Schwane - European Space Agency, Noordwijk, The Netherlands Examiner Dr. S. Larsson - Lule˚aUniversity of Technology, Lulea, Sweden Abstract A manned Mars mission will require a substantial increase in landed mass compared to previous robotic missions, beyond the capabilities of current Entry Descent and Landing, EDL, technologies, such as blunt-body aeroshells and supersonic disk-gap-band parachutes. The heaviest payload successfully landed on Mars to date is the Mars Science Laboratory which delivered the Curiosity rover with an approximate mass of 900 kg. For a human mission, a payload of magnitude 30-50 times heavier will need to reach the surface in a secure manner. According to the Global Exploration Roadmap, GER, a Human Mission to Mars, HMM, is planned to take place after year 2030. To prepare for such an event several technologies need maturing and development, one of them is to be able to use and accurately asses the performance of Supersonic Retro Propulsion, SRP, another is to be able to use inflatable heat shields. This internal study conducted at the European Space Agency, ESA, is a first investigation focusing on the Entry Descent and Landing, EDL, sequence of a manned Mars lander utilising an inflatable heatshield and SRP, which are both potential technologies for enabling future landings of heavy payloads on the planet. The thesis covers the areas of aerodynamics and propulsion coupled together to achieve a design, which considers the flight envelope constraints imposed on human missions. The descent has five different phases and they are defined as circular orbit, hypersonic entry, supersonic retropropulsion, vertical turn manoeuvre and soft landing. The focus of this thesis is on one of the phases, the SRP phase. The study is carried out with the retro-thrust profile and SRP phase initiation Mach number as parameters. Aerodynamic data in the hyper and supersonic regime are generated using Computational Fluid Dynamics, CFD, to accurately assess the retropropulsive performance. The basic concept and initial sizing of the manned Mars lander builds on a preliminary technical report from ESA, the Mission Scenarios and Vehicle Design Document [1]. The overall optimisation process has three parts and is based on iterations between the vehicle design, CFD computations in the software DLR-Tau and trajectory planning in the software ASTOS. Two of those parts are studied, the vehicle design and the CFD, to optimise and evaluate the feasibility of SRP during the descent and test the design parameters of the vehicle. This approach is novel, the efficiency and accuracy of the method itself is discussed and evaluated. Initially the exterior vehicle Computer Aided Design, CAD, model is created, based on the Mission Scenarios and Vehicle Design Document [1], however updated and furthered. The propulsion system is modelled and evaluated using EcosimPRO where the nozzle characteristics, pressure levels and chemistry are defined, and later incorporated in the CAD model. The first iteration of the CFD part has an SRP range between Mach 7 and 2, which results in an evaluation of five points on the trajectory. The thrust levels, the corresponding velocity, altitude and atmospheric properties at those points can then be evaluated and later incorporated in ASTOS. ASTOS, in turn, can simulate the full trajectory from orbit to landing including the CFD data of the SRP phase. Due to time limitation only one iteration of the vehicle design and the SRP range was completed. However, the goals of the study were reached. A first assessment of SRP in Mars atmosphere has been carried out, and the aerodynamic and propulsive data has been collected to be built on in the future. The results indicate that the engines can start at a velocity of Mach 7. They also show consistency with similar studies conducted in Earths atmosphere. The current vehicle design, propulsion system and SRP range can now be furthered, updated and advanced in order to optimise the different descent phases in combination with future results from ASTOS. Acknowledgements I would first and foremost like to thank my supervisor at ESA, Dr. Richard Schwane, for making this thesis happen. His support, his help, his knowledge and enormous experience within this field has been invaluable to me. This thesis has been carried out at ESA and ESTEC at the section TEC-MPA supervised by Section Head Dr. Guillermo Ortega who supported my work and made me feel right at home in his section, thank you. I would also like to thank my colleague at TEC-MPA, Mr. Csaba Jeger, for his very appreciated help and encouragement during my time at ESA, it has been valuable to have such a knowledgeable and professional person to work with. Another thanks to the Space Engineering Department and Fluid Mechanics Section at Lule˚aUniversity of Technology for giving me such good preparations, which eased my first working experience. My family deserves a special thank you for supporting me and helping me realise this dream of moving to the Netherlands to do my Master Thesis at ESA, especially my father Mr. M˚ansMarklund for his never ending encouragement and patience. Contents 1 Glossary ..................................... 1 1.1 List of acronyms . 1 1.2 List of symbols . 2 1.3 Sub and superscripts . 2 1.4 List of figures . 3 1.5 List of tables . 4 2 Introduction .................................. 5 3 Theory ...................................... 6 3.1 Fluid dynamics . 6 3.1.1 Aerodynamic parameters . 6 3.1.2 Mach number and shock wave formation . 6 3.1.3 Rocket nozzle flow . 6 3.1.4 SRP flow . 8 3.1.5 Computational Fluid Dynamics . 8 3.2 Orbital dynamics . 10 3.2.1 Entry, descent and landing . 10 4 Background ................................... 11 4.1 Heavy payload entry vehicles . 12 4.2 Supersonic Retropropulsion . 12 4.3 Inflatable heat shields . 13 5 Method and objectives ............................ 14 5.1 Method . 14 5.2 Objectives . 14 6 Validation of CFD tool DLR-TAU .................... 15 6.1 DLR-TAU . 15 6.1.1 Bow shock location . 15 6.1.2 Pressure distribution over the cone . 17 6.1.3 3D simulation . 18 7 Mission description .............................. 21 7.1 Vehicle design . 21 7.2 Propulsion system . 23 7.3 Atmosphere of Mars . 24 7.4 Simplifications and parameters not taken into account . 25 8 Trajectory planning in ASTOS ....................... 26 9 CFD simulations ................................ 27 9.1 Start-up conditions . 27 9.1.1 Geometry, model and mesh . 27 9.1.2 Simulations - Jet off . 28 9.2 Retro propulsive phase . 31 9.2.1 Geometry, model and mesh . 31 9.2.2 Simulations - Jet on . 32 9.2.3 180 degree model . 37 9.2.4 Similar studies . 39 10 Discussion .................................... 41 10.1 Future work . 42 11 Cited Literature ................................ 44 4 Appendices ..................................... 46 A ........................................... 46 I General 1D rocket nozzle equations . 46 II Validation case data . 47 B ........................................... 48 I Reference atmospheric conditions . 48 II Collected data . 49 1 Glossary 1.1 List of acronyms ASTOS - Analysis, Simulation and Trajectory Optimisation Software CAD - Computer Aided Design CFD - Computational Fluid Dynamics CFL - Courant Freidrichs Lewy CO2 - Carbon dioxide DES - Detached Eddy Simulation DLR - Deutsches Zentrum f¨ur Luft- und Raumfahrt DNS - Direct Numerical Simulations EDL - Entry, Descent and Landing ESA - European Space Agency ESPSS - European Space Propulsion System Simulation ESTEC - European Space Research and Technology Centre EV - Entry vehicle GER - Global Exploration Roadmap GNC - Guidance, Navigation and Control HMM - Human Mission to Mars ISECG - International Space Exploration Coordination Group LES - Large Eddy Simulation LH2 - Liquid Hydrogen LOX - Liquid Oxygen MOLA - Mars Orbiter Laser Altimeter MSL - Mars Science Laboratory RAAN - Right Ascension of the Ascending Node RANS - Reynolds-Averaged Navier-Stokes SRP - Supersonic Retro Propulsion SST - Shear Stress Transport TPS - Thermal Protection System 1 1.2 List of symbols Symbol Description Unit Value A Area m2 - BC Ballistic coefficient kg=m2 - c Speed of sound m=s - CD Coefficient of drag - - CL Coefficient of lift - - F Force N - g Gravity m=s2 - Isp Specific impulse s - L Characteristic length m - M Mach number - - m Mass kg - m_ Mass flow rate kg=s - p Pressure P a - q Dynamic pressure P a - Re Reynolds number - - T Temperature K - u Velocity m=s - w Thermodynamic work J - α Angle of attack ◦ - ρ Density kg/m3 - µ Dynamic viscosity N · s=m2 - γ Flight path angle ◦ - 1.3 Sub and superscripts e - Exit t - Throat T - Thrust 1 - Ambient conditions D - Drag L - Lift 2 1.4 List of figures 1 Rocket nozzle parts and parameters. [2] . 7 2 Rocket nozzle design parameters [3], comparison of Bell and 15◦Cone contours. 7 3 Supersonic retropropulsion flow characteristics. [4] . 8 4 Entry parameters such as flight path angle, lift, drag, height above ground, radius etc. 10 5 Figure summarising the Global Exploration Roadmap, GER. [5]. 11 6 Overview of possible entry mass and diameter criterion's for a HMM. [6] 12 7 Definition of bow shock stand off distance, l. [7] . 16 8 Results from 2D inviscid simulations of bow shock location. 16 9 CFD results of bow shock location of a higher vs. lower thrust coeffi- cient, M1 =3................................ 17 10 Experimental set up by McGhee[7]. 17 11 Result comparison between CFD and experimental data of pressure co- efficient along the cone for maximum pressure ratios at different Mach numbers . 18 12 3D CFD results of bow shock location for CT = 1:4, M1 = 3. 19 13 Comparison of bow shock location between experimental, 2D simula- tions and 3D simulations. 19 14 CFD results of bow shock location for CT = 1:4, M1 = 3.
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