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

SupervisorDr. R. Schwane - European Space Agency,

Noordwijk, The Netherlands

ExaminerDr. S. Larsson - Lulea University of

Technology, Lulea, Sweden

Abstract

A manned Mars mission will require a substantial increase in landed mass comparedto previous robotic missions, beyond the capabilities of current Entry Descent andLanding, EDL, technologies, such as blunt-body aeroshells and supersonicdisk-gap-band parachutes. The heaviest payload successfully landed on Mars to dateis the Mars Science Laboratory which delivered the Curiosity rover with anapproximate mass of 900 kg. For a human mission, a payload of magnitude 30-50times heavier will need to reach the surface in a secure manner. According to theGlobal Exploration Roadmap, GER, a Human Mission to Mars, HMM, is plannedto take place after year 2030. To prepare for such an event several technologies needmaturing and development, one of them is to be able to use and accurately asses theperformance of Supersonic Retro Propulsion, SRP, another is to be able to useinflatable heat shields.

This internal study conducted at the European Space Agency, ESA, is a firstinvestigation focusing on the Entry Descent and Landing, EDL, sequence of amanned Mars lander utilising an inflatable heatshield and SRP, which are bothpotential technologies for enabling future landings of heavy payloads on the planet.The thesis covers the areas of aerodynamics and propulsion coupled together toachieve a design, which considers the flight envelope constraints imposed on humanmissions. The descent has five different phases and they are defined as circularorbit, hypersonic entry, supersonic retropropulsion, vertical turn manoeuvre and softlanding. 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 Machnumber as parameters. Aerodynamic data in the hyper and supersonic regime aregenerated using Computational Fluid Dynamics, CFD, to accurately assess theretropropulsive performance. The basic concept and initial sizing of the mannedMars lander builds on a preliminary technical report from ESA, the MissionScenarios and Vehicle Design Document [1]. The overall optimisation process hasthree parts and is based on iterations between the vehicle design, CFDcomputations in the software DLR-Tau and trajectory planning in the softwareASTOS. Two of those parts are studied, the vehicle design and the CFD, tooptimise and evaluate the feasibility of SRP during the descent and test the designparameters of the vehicle. This approach is novel, the efficiency and accuracy of themethod itself is discussed and evaluated. Initially the exterior vehicle ComputerAided Design, CAD, model is created, based on the Mission Scenarios and VehicleDesign Document [1], however updated and furthered. The propulsion system ismodelled and evaluated using EcosimPRO where the nozzle characteristics, pressurelevels 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, whichresults in an evaluation of five points on the trajectory. The thrust levels, thecorresponding velocity, altitude and atmospheric properties at those points can thenbe evaluated and later incorporated in ASTOS. ASTOS, in turn, can simulate thefull trajectory from orbit to landing including the CFD data of the SRP phase. Dueto time limitation only one iteration of the vehicle design and the SRP range wascompleted. However, the goals of the study were reached. A first assessment of SRPin Mars atmosphere has been carried out, and the aerodynamic and propulsive datahas been collected to be built on in the future. The results indicate that the enginescan start at a velocity of Mach 7. They also show consistency with similar studiesconducted in Earths atmosphere. The current vehicle design, propulsion system andSRP range can now be furthered, updated and advanced in order to optimise thedifferent descent phases in combination with future results from ASTOS.

Acknowledgements

I would first and foremost like to thank my supervisor at ESA, Dr. RichardSchwane, for making this thesis happen. His support, his help, his knowledge andenormous experience within this field has been invaluable to me.

This thesis has been carried out at ESA and ESTEC at the section TEC-MPAsupervised by Section Head Dr. Guillermo Ortega who supported my work andmade 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 veryappreciated help and encouragement during my time at ESA, it has been valuableto have such a knowledgeable and professional person to work with.

Another thanks to the Space Engineering Department and Fluid Mechanics Sectionat Lulea University of Technology for giving me such good preparations, whicheased my first working experience.

My family deserves a special thank you for supporting me and helping me realisethis dream of moving to the Netherlands to do my Master Thesis at ESA, especiallymy father Mr. Mans Marklund for his never ending encouragement and patience.

Contents

1 Glossary . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 11.1 List of acronyms . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 11.2 List of symbols . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 21.3 Sub and superscripts . . . . . . . . . . . . . . . . . . . . . . . . . . . . 21.4 List of figures . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 31.5 List of tables . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 4

2 Introduction . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 5

3 Theory . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 63.1 Fluid dynamics . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 6

3.1.1 Aerodynamic parameters . . . . . . . . . . . . . . . . . . . . . 63.1.2 Mach number and shock wave formation . . . . . . . . . . . . . 63.1.3 Rocket nozzle flow . . . . . . . . . . . . . . . . . . . . . . . . . 63.1.4 SRP flow . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 83.1.5 Computational Fluid Dynamics . . . . . . . . . . . . . . . . . . 8

3.2 Orbital dynamics . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 103.2.1 Entry, descent and landing . . . . . . . . . . . . . . . . . . . . 10

4 Background . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 114.1 Heavy payload entry vehicles . . . . . . . . . . . . . . . . . . . . . . . 124.2 Supersonic Retropropulsion . . . . . . . . . . . . . . . . . . . . . . . . 124.3 Inflatable heat shields . . . . . . . . . . . . . . . . . . . . . . . . . . . 13

5 Method and objectives . . . . . . . . . . . . . . . . . . . . . . . . . . . . 145.1 Method . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 145.2 Objectives . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 14

6 Validation of CFD tool DLR-TAU . . . . . . . . . . . . . . . . . . . . 156.1 DLR-TAU . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 15

6.1.1 Bow shock location . . . . . . . . . . . . . . . . . . . . . . . . . 156.1.2 Pressure distribution over the cone . . . . . . . . . . . . . . . . 176.1.3 3D simulation . . . . . . . . . . . . . . . . . . . . . . . . . . . . 18

7 Mission description . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 217.1 Vehicle design . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 217.2 Propulsion system . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 237.3 Atmosphere of Mars . . . . . . . . . . . . . . . . . . . . . . . . . . . . 247.4 Simplifications and parameters not taken into account . . . . . . . . . 25

8 Trajectory planning in ASTOS . . . . . . . . . . . . . . . . . . . . . . . 26

9 CFD simulations . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 279.1 Start-up conditions . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 27

9.1.1 Geometry, model and mesh . . . . . . . . . . . . . . . . . . . . 279.1.2 Simulations - Jet off . . . . . . . . . . . . . . . . . . . . . . . . 28

9.2 Retro propulsive phase . . . . . . . . . . . . . . . . . . . . . . . . . . . 319.2.1 Geometry, model and mesh . . . . . . . . . . . . . . . . . . . . 319.2.2 Simulations - Jet on . . . . . . . . . . . . . . . . . . . . . . . . 329.2.3 180 degree model . . . . . . . . . . . . . . . . . . . . . . . . . . 379.2.4 Similar studies . . . . . . . . . . . . . . . . . . . . . . . . . . . 39

10 Discussion . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 4110.1 Future work . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 42

11 Cited Literature . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 44

4

Appendices . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 46

A . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 46I General 1D rocket nozzle equations . . . . . . . . . . . . . . . . . . . . 46II Validation case data . . . . . . . . . . . . . . . . . . . . . . . . . . . . 47

B . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 48I Reference atmospheric conditions . . . . . . . . . . . . . . . . . . . . . 48II 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 fur 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 Pa -

q Dynamic pressure Pa -

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 - Exitt - ThroatT - Thrust∞ - Ambient conditionsD - DragL - Lift

2

1.4 List of figures

1 Rocket nozzle parts and parameters. [2] . . . . . . . . . . . . . . . . . 72 Rocket nozzle design parameters [3], comparison of Bell and 15◦Cone

contours. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 73 Supersonic retropropulsion flow characteristics. [4] . . . . . . . . . . . 84 Entry parameters such as flight path angle, lift, drag, height above

ground, radius etc. . . . . . . . . . . . . . . . . . . . . . . . . . . . . 105 Figure summarising the Global Exploration Roadmap, GER. [5]. . . . 116 Overview of possible entry mass and diameter criterion’s for a HMM. [6] 127 Definition of bow shock stand off distance, l. [7] . . . . . . . . . . . . . 168 Results from 2D inviscid simulations of bow shock location. . . . . . . 169 CFD results of bow shock location of a higher vs. lower thrust coeffi-

cient, M∞ = 3. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1710 Experimental set up by McGhee[7]. . . . . . . . . . . . . . . . . . . . . 1711 Result comparison between CFD and experimental data of pressure co-

efficient along the cone for maximum pressure ratios at different Machnumbers . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 18

12 3D CFD results of bow shock location for CT = 1.4, M∞ = 3. . . . . . 1913 Comparison of bow shock location between experimental, 2D simula-

tions and 3D simulations. . . . . . . . . . . . . . . . . . . . . . . . . . 1914 CFD results of bow shock location for CT = 1.4, M∞ = 3. . . . . . . . 2015 Vehicle design of HMM vehicle proposed by ESA. [1] . . . . . . . . . . 2116 Updated vehicle design with the nozzles integrated in the heat cap. . . 2217 Schematic of liquid propulsion system in EcosimPro. . . . . . . . . . . 2318 Atmosphere levels of temperature and pressure above 7 km. . . . . . . 2419 Atmosphere levels of temperature and pressure below 7 km. . . . . . . 2520 CAD model for simulating start-up conditions, nozzles are sealed. . . . 2821 2D snapshots of the 45 degree slice of the 3D geometry and the finest

mesh used including adaptation, in total 12e6 nodes. . . . . . . . . . . 2822 Pressure outside and inside bow shock, M∞ = 7. . . . . . . . . . . . . 2923 Comparison of bow shock location vs. total free stream pressure up-

stream of the bow shock, M∞ = 7. . . . . . . . . . . . . . . . . . . . . 2924 Comparison of surface pressure on the heat shield behind the bow

shock, M∞ = 7. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 3025 CAD model and example of adapted mesh for jet on simulations. . . . 3126 Pressure and Mach number curve extending along the tilted nozzle axis

from the throat of the nozzle and forward, M∞ = 7. . . . . . . . . . . 3227 Pressure and Mach number plumes extending along x [mm] from the

nozzle and forward, M∞ = 7. . . . . . . . . . . . . . . . . . . . . . . . 3228 Mach number and velocity contour, M∞ = 6. . . . . . . . . . . . . . . 3329 Mach number plumes, M∞ = 5. . . . . . . . . . . . . . . . . . . . . . . 3430 Pressure and Mach number plumes extending along x from the nozzle

and forward, M∞ = 4. . . . . . . . . . . . . . . . . . . . . . . . . . . . 3431 Chemistry difference and distribution between exhaust gases and the

ambient CO2 dominant atmosphere , M∞ = 4. . . . . . . . . . . . . . 3532 Mach number and pressure, M∞ = 3. . . . . . . . . . . . . . . . . . . . 3533 , M∞ = 2. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 3634 Mach number plume and stream traces of the flow field, M∞ = 7. . . . 3735 Pressure and temperature contours, M∞ = 7. . . . . . . . . . . . . . . 3736 Mach number and temperature, M∞ = 7. . . . . . . . . . . . . . . . . 3837 Comparison of nozzle temperature between this study and the study

carried out by DLR of a Falcon 9 re-entry scenario. . . . . . . . . . . . 39

3

38 Comparison of Mach number levels in the nozzle between this studyand the study carried out by DLR of a Falcon 9 re-entry scenario. . . 39

39 Comparison of temperature levels around the vehicles between thisstudy and the study carried out by DLR of a Falcon 9 re-entry scenarioin Earths atmosphere [8] . . . . . . . . . . . . . . . . . . . . . . . . . . 40

40 Mach and temperature contours from this study and the study carriedout by DLR of a Falcon 9 re-entry scenario in Earths atmosphere [8] . 40

1.5 List of tables

1 Summary of size parameters for the updated vehicle design. . . . . . . 222 Parameters of the liquid propulsion system simulation in EcosimPro. . 233 Chemistry data file parameters for the exhaust gases in DLR-TAU. . . 244 Chemistry data file parameters in DLR-TAU for Mars atmosphere. . . 245 Circular orbit elements of phase 1 implemented in ASTOS. . . . . . . 266 Summary of outputs from the inflatable heat shield simulations. . . . . 267 Atmospheric reference values of the first point in the SRP phase, start-

up of the engines. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 278 Summary of outputs from the simulations of the start-up conditions in

DLR-TAU. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 309 Difference in atmospheric conditions for the points of interest at differ-

ent altitude and velocity. . . . . . . . . . . . . . . . . . . . . . . . . . . 3310 Summary of outputs of CFD simulations for the retropropulsive phase

of 45 degree model. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 36

4

2 Introduction

A manned Mars mission will require a substantial increase in landed mass comparedto previous robotic missions, beyond the capabilities of current EDL, Entry Descentand Landing, technologies, such as blunt-body aeroshells and supersonicdisk-gap-band parachutes. The heaviest payload successfully landed on Mars to dateis the Mars Science Laboratory which delivered the Curiosity rover with a mass of900 kg. For a human mission, a payload of magnitude 30-50 times heavier will needto reach the surface in a secure manner. According to the GER, Global ExplorationRoadmap, a HMM, human mission to Mars, is planned to take place after the year2030. To prepare for such an event several technologies need maturing anddevelopment, one of them is to be able to use and accurately asses the performanceof SRP, supersonic retro propulsion.

This internal study carried out at ESA is conducted as a first investigation of SRPin a low density atmosphere. SRP is a technology for braking a spacecraft by firingrocket thrusters in the opposite direction of the velocity vector of the vehicle. Theusage of SRP rockets is still a novel, but promising, technology for slowing downthrough an atmosphere in an efficient way, which has been shown by SpaceX inrecent years [9] for applications here on Earth. While the SRP technology is makingprogress here on Earth with re-usability and reduced cost as its main objectives, itis also a critical area of development for further exploration of the solar system.Mars has a very low density atmosphere compared to Earth, additionally thepayload of a human mission is large. This in combination with the difficulties inhandling large parachutes have led to the conclusion that SRP is one of the criticaltechnologies to study further when aiming to carry out pin point landings of heavypayloads on the red planet.

The goal of this study is to introduce a first set of results of SRP from CFD inDLR-TAU as a reference to build on. The task is focused on the EDL sequence of amanned Mars lander utilising an inflatable heatshield and supersonic retropropulsion, which are both potential technologies for enabling future landings ofheavy payloads on the planet. The parameters used for the optimisation are theSRP initiation Mach number, the retro-thrust profile and the design of the vehicle.The initial design and concept of the mission is based on a technical note from ESA[1] which comprises a payload of 23 metric tonnes, an inflatable heat shield surfacearea of 240 square meters and four main engines and nozzles for SRP.

5

3 Theory

3.1 Fluid dynamics

Fluid dynamics is a subarea within fluid mechanics and describes the motion ofliquids and gases. Included in the fluid dynamics discipline is the area ofaerodynamics, which traditionally described only the flow of air, but was laterextended to the flow of other mediums as well. The basic fluid and aerodynamicconcepts used in this study will be shortly described in this chapter.

3.1.1 Aerodynamic parameters

The lift force, FL, is defined as the force perpendicular to the flow direction, whilethe drag force, FD is defined in the direction parallel and opposite to the flowdirection. For a zero angle of attack the generated lift force of a symmetric body iszero, since the flow then surrounds the body with an even pressure. Together, theperpendicular forces FD and FL, expressed below in their respective aerodynamiccoefficients, make up the total aerodynamic force on a vehicle.

CD =2 · FDρu2A

(1)

CL =2 · FLρu2A

(2)

3.1.2 Mach number and shock wave formation

The Mach number is a dimensionless quantity which is used for high velocity flows.It is defined by the speed of sound, c, in the medium that it is travelling through.The different regions defined by the Mach number are subsonic M < 1 , sonicM = 1, supersonic M > 1 and hypersonic M > 5.

M =u

c(3)

Shock waves appear when an object propagates faster than the speed of sound in amedium. The properties of the shock varies depending on the object shape and themedium. Commonly used terms for different parts of a shock waves include normalshock and oblique shock, an oblique shock is inclined with respect to the direction ofthe free stream and the normal shock is not. An example of shock waves present ina supersonic retro propulsion scenario is shown in Figure 3.

3.1.3 Rocket nozzle flow

Rocket nozzles are designed to create large amounts of thrust which implies a highvelocity increase through the nozzle and a high mass flow, m, at the exit, equation 4shows Tsiolkovsky’s rocket equation [10]. This study only considers liquidpropulsion, and hence only liquid fuel turned gaseous in the process flows throughthe nozzles. Figure 1 shows a schematic of a rocket nozzle, the design variables ofthe nozzle determines the flow acceleration, Mach number, mass flow rate etc. Thecomplete 1D rocket nozzle equations which determines the basic characteristics ofthe flow can be found in Appendix A.

F = mue +Ae(pe − p∞) = mIsp (4)

6

Figure 1: Rocket nozzle parts and parameters. [2]

The rocket nozzle contour can be calculated for a specific pressure or area ratiobetween the chamber, throat and the exit, or for a desired thrust level. One problemthat often arises when trying to maximise the thrust and mass flow, such as forrocket applications, is that the ideal equations may result in an inadequately longnozzle which then has to be cut short to fit with the sizing of the spacecraft. Anoverview of nozzle performance and the full set of equations deciding the nozzleshape can be found in Sutton [3]. Two commonly used shapes are the Bell nozzleand the 15◦ cone nozzle, described in Figure 2.

Figure 2: Rocket nozzle design parameters [3], comparison of Bell and 15◦Cone contours.

7

3.1.4 SRP flow

SRP is defined as when a rocket engine is fired into an opposing supersonic freestream. Hence, SRP flow consists of the combined characteristics of the the ambientsupersonic retro flow and the rocket plume flowing in the opposite direction. Theplume pushes the bow shock away from the body and enlarges it, which in turnincreases the braking force. A fully developed jet plume extending from a nozzlewhich is integrated in a vehicle body and travelling through a flow at supersonicspeeds is schematically explained in Figure 3. The general flow properties torecognise is the bow shock, jet shock, re-circulation regions and the jet plume.

Figure 3: Supersonic retropropulsion flow characteristics. [4]

3.1.5 Computational Fluid Dynamics

To model and analyse a flow field and its behaviour, a computational software canbe used. This is a relatively new field witch needs maturing and furtheradvancement to keep up with the applications it can be used for. One of thoseapplications are SRP for spacecrafts landing on Earth or other bodies such as theMoon or Mars. There are several different computational schemes that can be usedfor a CFD simulation and it is important to choose and customise it according tothe application studied. When inviscid cases are simulated in this study the solutionis computed using the Euler equations, shown below. An inviscid solution issimplified and contains many assumptions, however it can still be useful tounderstand the basic flow field. The letter u denotes the velocity, t the time, wstands for work and g describes the gravity (or other acceleration).

∂~u

∂t+ ~u · ∇~u = −∇w + ~g (5)

∇ · ~u = 0 (6)

Other computational schemes that can be used are for example a Direct NumericalSimulation, DNS, which relies on the Navier-Stokes equations and does not includea turbulence model. Another widely used computation in engineering applications isthe Large Eddy Simulations, LES, which is a model for simulating turbulent flows.Yet another is the Detached Eddy Simulation, DES, which is a hybrid model

8

combining the LES and the Reynolds Averaged Navier Stokes, RANS, equations.These simulations each have their benefits, and are appropriate for differentapplications. However, for computing the solution of the viscous and turbulentsimulations in this study the RANS equations are used alone. It is a well tested andstraight forward computation, feasible to use in a first study of a partly unknownflow field as in this thesis. Below, the compressible RANS equations for ideal gas aredescribed. Equation 7 shows the RANS equations for conservation of mass, equation8 shows the conservation of momentum and equation 9 shows the conservation ofenergy. P is the stress tensor, ρ the density, q is the heat flux and E is the energydensity.

∂ρ

∂t+

∂

∂xi(ρui) = 0 (7)

∂

∂t(ρuj) +

∂

∂xi(ρuiuj − Pij) = 0 (8)

∂E

∂t+

∂

∂xi(uiE − ujPij + qi) = 0 (9)

The RANS equations are time averaged solutions of the Navier-Stokes equations. Tocomplete the solution a turbulence model is required, many different turbulencemodels exists today and can be chosen depending on the application. Furthercalibration and development of turbulence modelling is a critical step in theadvancement and reliability of CFD tools.

In this thesis, hyper and supersonic flows are studied and the CFD code used iscalled DLR-Tau, developed by DLR - the German Aerospace Center. To run a CFDsimulation a geometry, a mesh and a CFD code calculating the viscid or inviscidequations are needed. The geometry and enclosure around the geometry is specifiedby boundary conditions on all surfaces. Examples of common boundary conditionsare walls, inlets, symmetries and outlets. The boundary conditions apply rules interms of equations to the computational domain and enables a solution. CFD ofvery high velocities are more uncertain and require more computer power, orcomputation time, than low velocity flow simulations. This is due to the mesh whichhas to be made up of smaller elements, compared to low velocity applications, inorder to be able to capture the flow field of the high velocity flow. To check if themesh is fine enough to catch the boundary layer near the walls, a measurementcalled y+ describing a dimensionless wall distance is often used. It is defined inequation 10 where u∗ is the velocity at the wall, y∗ is the cell length at the wall andν is the kinematic viscosity. In this study the aim was to obtain a y+ ofapproximately 1 or lower.

y+ =u∗y∗ν

(10)

It should be noted that all CFD simulations contain approximations. Quality andtrust in CFD is an important topic to be discussed in order to evaluate how far froma ’real life’ solution, an experimental result, the outputs are. The errors, ordeviations from experimental results, can arise from incomplete problem definition,turbulence modelling, numerical differences and discretization etc. When running aCFD simulation, error residuals of certain parameters can be monitored to check ifthe errors of those parameters are decreasing or increasing. If the monitoredresiduals reach a low value after a certain amount of iterations it is an indicationthat the solution is correct, how low below one that residual value should be differsfrom case to case. The grid quality, as mentioned above for y+, is another importantmeasurement. The wall boundary layer, for viscous simulations, should be meshedaccording to the y+ equation and is the finest meshed area in the grid. In thisthesis, a study of quality and trust in CFD is carried out to some extent. The mainfocus is validation and verification based on the best practice guidelines described

9

by the fluid dynamics laboratory at Sulzer Innotech [11]. Since SRP in Marsatmosphere has not been done yet outside of simulations, no experimental data canbe used as a comparison. However, a validation case with experimental data of asimilar case is set up prior to continuing with the CFD simulations of the spacecraftof interest in this study. This is done to validate the DLR-Tau code for SRP flows.

3.2 Orbital dynamics

Orbital dynamics is a field that describes the ballistic motion of a body which canbe propelled or in relation to another body’s gravity, such as trajectories forspacecrafts. Some of the key variables and concepts concerning the descenttrajectory planning relevant for this mission will be shortly described in this chapter.

3.2.1 Entry, descent and landing

From orbit to landing there are several phases that needs to be specified anddesigned to ensure that the braking procedures are efficient, safe and powerfulenough. The ballistic coefficient is a size to mass measurement of how a bodybehaves in a flow in terms of negative acceleration, the value of the ballisticcoefficient impacts the entry corridor and the flyable domain for the spacecraft. Fora vehicle with a high ballistic coefficient, more forceful braking capabilities areneeded, such as SRP, since it does not slow down enough due to its shape, whichdetermines CD, and high mass when travelling through the medium.

BC =m

CD ·A(11)

For a human mission, which has a large mass m, the ballistic coefficient is highcompared to robotic missions. To have enough time to slow down during thedescent, the entry angle, or flight path angle γ, can be chosen as shallow to extendthe trajectory length. Both these features are included in the characteristics of aballistic entry. For a ballistic entry of a axisymmetric vehicle, the local angle ofattack is zero and there is only forces directed in parallel to the velocity vector.However, that is only exactly true if the center of pressure coincides with the centerof mass, which may not always be the case, but could be used as a generalassumption for a generic ballistic entry case.

Figure 4: Entry parameters such as flight path angle, lift, drag, height above ground, radiusetc.

10

4 Background

Spacecrafts have been sent towards Mars since the 1960’s, initially to orbit theplanet but early on some spacecrafts carried probes to be deployed towards thesurface. The first successful orbiter around Mars was Mariner 9 in 1971 sent byNASA, and the first successful landing took place only weeks after by the spacecraftcalled Mars 3, sent by the Soviet Union [12][13]. The Mars 3 descent module had atotal wet mass of about 1200 kg which comprised the spherical landing capsule ofabout 350 kg, the braking aids of the vehicle consisted of a 2.9 m conical shield,parachute system and retrorockets [12]. The first European mission to Mars wascalled Mars Express, launched in 2003, consisting of both an orbiter and a lander.The orbiter was inserted in its desired trajectory around the planet without faults,the lander was however not able to send out any signals after arriving at the surface[14]. The latest successful landing on Mars to date is NASAs InSight lander whichhad a wet mass of about 600 kg entering the atmosphere, the braking systemconsisted of a parachute and retrorockets to obtain a soft pinpoint landing [15].

The Global Exploration Roadmap [5] is a document where 14 Space Agencies,including ESA, summarise their common goals for future space exploration. A cleargoal stated in the 3rd Edition is to put humans on the surface on Mars after year2030. In the GER, one critical technology gap highlighted to investigate is how tosafely descent and land large robotic missions with a payload of more than 1000 kgand human missions of about 40 000 kg, i.e when parachutes and aero-braking areno longer viable EDL technologies. As a step in that direction it was decided tostart evaluating and prepare the capabilities at ESA ESTEC for assessment of SRPin combination with an inflatable heatshield for such a human mission to Mars.

Figure 5: Figure summarising the Global Exploration Roadmap, GER. [5].

11

4.1 Heavy payload entry vehicles

The mission with highest payload mass which has successfully landed on the surfaceof Mars is the MSL, which reached the planet in 2012. The MSL contained theCuriosity Rover as payload, which had a mass of about 900 kg. The MSL entryvehicle was slowed down from 5.8 km/s in orbit to 400 m/s by aero-braking with its70 degree cone shape during 259 seconds. The MSL Entry Vehicle, EV, comprisedthe largest disk gap-band parachute ever made, which was deployed at a velocity of400 m/s, 1.7 Mach, and at an altitude of 12 km. The powered descent was initiated250 meters above ground in a vertical manner and ended with a procedure namedthe sky crane, where the rover was lowered down with wires to the ground from thehovering EV module 20 meters above ground [16]. For future Human Mars missionsthe payload will need to contain a habitat and a crew, and the mass of that payloadwill increase from Curiosity’s 900 kg up to 20-40 tonnes. A study at the Departmentof Aeronautics, Imperial College of London, have presented a paper highlighting thetechnology gaps between today’s missions and future heavy human or roboticsmissions [6], this is envisioned in figure 6 below.

Figure 6: Overview of possible entry mass and diameter criterion’s for a HMM. [6]

4.2 Supersonic Retropropulsion

In 2013, SpaceX was the first company to fire a rocket engine into an opposingsupersonic free stream, in this case Earths atmosphere, and in 2015 the firstsuccessful vertical landing of the first stage of the Falcon 9 rocket was achieved [9].This became the start of the re-usable era in space exploration, and since then theboosters and rocket stages of Falcon Heavy have also been successfully landed withSRP, and re-used on Earth. SRP has been a technology of interest for future humanMars missions since the 1960’s, hence the need and interest to accurately assess andsimulate the performance of SRP only grew once it was shown to be a game changeron Earth. A study carried out in 2017 by NASA and partners [17] compares flightdata with CFD codes from Falcon 9’s first stage at re-entry at Mars relevantconditions in Earths atmosphere. The study shows that CFD of SRP is animportant tool and gives sufficiently correct solutions in many cases, but themethods needs further advancing. Since SRP is such a novel technology there is notyet a large amount of gathered flight data. Especially not for different vehicles andnozzle configurations, since Falcon 9 and Falcon Heavy are the only large scale

12

spacecrafts successfully using this technology at present. The study comparing flightdata and CFD [17] is an important step to establish a base to build on, in a similarmanner as this ESA study gathers and validates CFD data for SRP.

4.3 Inflatable heat shields

Since the size of the descent module is restricted by the fairing diameter of thelaunch vehicle, there is a design limit for spacecrafts and their heat shields. Inaddition parachutes are no longer an option due to the mass. Larger spacecrafts inthis sense indicate future robotic missions with payloads above 1000 kg or humanmissions. However, the development of inflatable heat shields could solve the sizerestriction of the launch vehicle fairing by being deflated until needed for descent[18]. An inflatable structure can utilise a flexible TPS which can withstand re-entryheats, as shown by NASA [19], and have a sufficient size once deployed to slow downa large and heavy spacecraft in its initial descent trajectory. The inflation of the heatshield would take place in orbit prior to the initiation of the descent. To combineSRP and an inflatable heat shield in an EDL sequence on Mars has been suggestedby both ESA and NASA as a feasible approach for future heavy missions [20] [1].

13

5 Method and objectives

5.1 Method

The overall optimisation process is iterative between CFD simulations in DLR-Tau,trajectory planning in ASTOS and updates of the vehicle model using the CADsoftware CATIA. In this thesis two parts of the optimisation process are featured,the vehicle design and CFD simulations. The general goal is to start up theiteration process by creating an initial vehicle design and to collect data from CFDsimulations of the complex aerothermodynamics during the retro propulsive phase.The results from this thesis could then serve as a base for continuing to completethe whole optimisation process, including ASTOS simulations. The CFD results ofthe SRP phase determines the available SRP range of the engines and the operatingconditions, therefore a range of Mach numbers are tested. A critical aspect isdetermining when and where the SRP phase can begin, the start up behaviour ofthe engines in high Mach number retro flows is one of the parameters studied indetail. After there is an initial vehicle design and corresponding CFD results, theSRP performance can be incorporated in ASTOS to complete the first iteration ofthe full process and further evaluate the results. The combined results can then beassessed, and the vehicle design updated for the next iteration.

5.2 Objectives

The objectives and expected outputs of this master thesis are summarised below.

• Deliver internal study, preliminary to a phase A, of the SRP part of an EDLsequence for a HMM.

• Create first design of the vehicle and propulsion system.

• Gather aerodynamic data of SRP phase in Mars atmosphere.

• Find the preliminary thrust profile.

14

6 Validation of CFD tool DLR-TAU

The complex flow fields of SRP and nozzle exhausts are a challenge to predict anddescribe precisely. The lack of a exact mathematical description of turbulence incombination with the said complexity of the flow field results in a problem set upwith many uncertainties. Initially, the CFD software ANSYS Fluent was used incomparison to DLR-TAU for a nozzle flow problem. It was shown that theDLR-TAU code was more sophisticated for these particular cases, and that ANSYSFluent was more time consuming. Hence, it was decided to continue only withDLR-TAU. In agreement with the Best Practice Guidelines in CFD [11], avalidation case is set up. Due to the lack of experimental data of SRP in Marsatmosphere, the flow field investigated for the SRP phase of the the EDL sequencein this study can not be fully validated. However, a similar experimental set up waschosen to validate SRP flows in DLR-Tau.

6.1 DLR-TAU

DLR-TAU is a code developed by the German space agency, DLR, especiallyfocused towards the aerospace sector. To verify the CFD tool DLR-Tau for SRPapplications a validation case was conducted. The experimental data can be foundin a technical note from NASA Langley Research Centre by Robert J. McGhee [7],where a nozzle plume situated at the apex of a 140 ◦ cone is studied in retro flows ofMach 3.0, 4.5 and 6.0. This particular experimental study was chosen both since ithas a reasonable SRP interval and since it includes a cone body similar to aheatshield. The experimental set up, which is thoroughly described in McGhee, wastranslated into an axisymmetric 2D model in the CAD software CATIA which wereused in the verification simulations. The following validation case aims to highlightthe problematic areas and to find the most valid approach. To ensure thecorrectness of the solution and set up of the 2D cases, a final step was to create a3D model and repeat one of the simulations.

The result parameters chosen for comparison were bow shock location and pressurecoefficient distribution over the cone surface. The general flow field of thesimulations should contain the features shown in figure 3. In McGhee, several casesper opposing Mach number were tested, to scale down the validation case amounttwo pressure ratios per Mach number were chosen. Expressed in thrust coefficientsthe first area is around CT = 0.4 and the second around CT = 1.4. Both areas ofinterest were found to be stable by McGhee, since the jet pressure is higher than theambient. The simulations were set up as inviscid and steady with the Eulerequations to minimise the calculation time, and also to compare how well theinviscid cases could compare to the experimental data. The mesh were created asunstructured in the software Centaur, and contained about 600 000 nodes and noboundary layer refinement due to in inviscid solver. A coarse outer mesh enclosed arefined area around the nozzle and cone which had a cell length limit of 0.4millimetres. The boundary conditions used were Euler walls, supersonic inflow,supersonic outflow and farfield.

6.1.1 Bow shock location

Figure 7 shows the definition of bow shock location in reference to the cone bodyand nozzle plume. The paper states that the experimental measurements areaccurate within 2 %, and the 2D simulations show results within 5 % of theexperimental results. All error bars in the following plots indicate 2 % deviation andthe angle of attack is 0 degrees.

15

Figure 7: Definition of bow shock stand off distance, l. [7]

Figure 8 shows the comparison of bow shock stand off distance, l, over the cone basediameter, Db, as a function of thrust coefficient, CT , and can be compared to Figure12 in McGhee, which can also be found in appendix B section I. To compare theCFD results with the experimental results the Mach number distribution wasextracted along the symmetry axis and the distance from the exit of the nozzle tothe bow shock front was measured.

CT =FTq∞A

(12)

Figure 8: Results from 2D inviscid simulations of bow shock location.

Figure 9 shows the difference in bow shock location when applying a different inletpressure at the nozzle, different thrust coefficient. As expected, and seen in theexperimental data, the bow shock location moves further away from the nozzle exitwhen the thrust is increased. Figure 9 shows the difference in bow shock locationbetween a higher and a lower thrust coefficient. Due to the inviscid simplificationand the mesh refinement the CFD figures show some discrepancy and unsteadinessin the flow field, which can be expected. However, the general flow properties arewell represented with the bow shock, jet shock and re-attachment shock.

16

(a) CT = 1.4 (b) CT = 0.4

Figure 9: CFD results of bow shock location of a higher vs. lower thrust coefficient,M∞ = 3.

Figures 9a and 9b can be compared with figure 3d in McGhee [7], where additionalcases of different pressure ratios and thrust coefficients are represented as well.

6.1.2 Pressure distribution over the cone

The pressure distribution over the cone is experimentally measured at orifices alongthe surface. The results are then translated into dimensionless quantities Cp andr/Db, where r is the radius to each orifice and Db is the length of the cone base.The simulated results were obtained by extracting the pressure along a line at thecone surface, then interpolated linearly and plotted against a length axis and madenon dimensional. The experimental set up can be seen in figure 10.

Cp =P − P∞q∞

(13)

Figure 10: Experimental set up by McGhee[7].

17

(a) M∞ = 3 (b) M∞ = 4.5

(c) M∞ = 6

Figure 11: Result comparison between CFD and experimental data of pressure coefficientalong the cone for maximum pressure ratios at different Mach numbers

The simulated results plotted together with the experimental data can be seen inFigure 11. The full set of experimental results of pressure distribution can be foundin Figures 3a, 4a and 5a in McGhee [7]. To be able to compare the data in anefficient and more exact way the figures given in McGhee were converted to digitalform, and the desired data could be extracted. Experimental data of the validationcase can be found in Appendix A.

6.1.3 3D simulation

The results of the 3D simulation, shown in Figure 12, for a quarter axisymmetricdomain of the CT = 1.4 case shows a similar bow shock location and pressuredistribution as the 2D case. Additionally Figure 12b roughly shows the pattern ofSRP flow fields, which is schematically explained in figure 3.

18

(a) (b)

Figure 12: 3D CFD results of bow shock location for CT = 1.4, M∞ = 3.

Due to boundary effects of a supersonic outflow there is an incorrect increase invelocity and Mach number behind the shield, which can be seen in Figure 12a. Thiscould be corrected by changing the boundary condition to an inviscid wall, Eulerwall, in the future. The general flow field is correct but not resolved with a goodresolution, which could be fixed with a finer mesh and adaptation in futuresimulations.

Figure 13: Comparison of bow shock location between experimental, 2D simulations and3D simulations.

The bow shock location is similar between the experimental, 2D and 3D results,

19

which is expected since the experimental data and the 2D data agreed well. Figure14 shows that a 3D simulation may catch the spectrum of pressure levels of the conesurface more precisely than a 2D simulation.

Figure 14: CFD results of bow shock location for CT = 1.4, M∞ = 3.

In conclusion the validation case indicates that a finer mesh, mesh adaptation andmore iterations could help to sharpen the results of future SRP simulations. Theinviscid simulations show good agreement with the experimental data for theparameters that were compared, however an even better agreement is of courseexpected for viscous simulations. Between 2D and 3D there is no large differencebetween the results, however a 3D model in a quarter enclosure gives a more correctpressure curve, and is perhaps to prefer in order to visualise the results in a niceway. The general flow field agrees with the theory, but should be resolved moreclearly in the coming simulations.

20

7 Mission description

7.1 Vehicle design

This study considers the retro propulsive phase of an EDL sequence of a mannedMars lander. The design of the Mars crew transfer vehicle is based on a technicalnote from ESA [1], section 5.1.7, where the payload is estimated to 23 metric tonnesconsisting of the habitat and the crew. The suggested design from the technical notecan be seen in figure 15 which has a cylindrical structure, contains four tanks, fourrocket nozzles, an inflatable unit, a solid front heat cap and a parachute. In theearly stages of this investigation it was concluded that a parachute wont beappropriate for such a heavy vehicle, and it was hence removed from the study. Toavoid a 180 degree turn in the descent trajectory, it was also concluded that thepropulsion system and nozzles must be located behind the solid heat cap and not atthe top part of the vehicle.

Figure 15: Vehicle design of HMM vehicle proposed by ESA. [1]

To begin the first iteration of the design process, the initial guess of the vehicledesign was created in CATIA. The inflatable heat shield has a diameter of 17.5meters and the cylinder is 8.2 meters long, which corresponds to the proposed sizeby ESA [1]. The four nozzles are conical with an area ratio of 16, has an initiallength of 1.5 meters and are tilted with a 7 degree angle outwards. The nozzles donot protrude but are flush against the curvature of the heat cap, this makes thelength of the nozzle differ in different sections between 1.3-1.5 meters. The nozzlesare treated as sealed until ignition when the protective caps are blown off by the jetsand the retro propulsive phase is initiated. Once the retropropulsive phase isinitiated the shield is no longer affecting the braking of the vehicle. It can theneither be deflated and lie flat against the cylinder surface as an extra heat protectionlayer, or it can stay inflated during the thrust phase which is the chosen method inthis case.

21

(a) Tilted view. (b) Side view.

Figure 16: Updated vehicle design with the nozzles integrated in the heat cap.

Figure 16 shows the updated design used in this study with the nozzles situated inthe rigid heat cap. The size parameters of the spacecraft model are summarised intable 1.

Table 1: Summary of size parameters for the updated vehicle design.

Spacecraft size parametersShield radius [m] 8.750Cylinder radius[m] 2.580Cylinder length [m] 8.260Shield area [m2] 240Total estimated mass [kg] 35000Nozzle shape [-] ConicalNozzle tilt angle [◦] 7Nozzle length [m] 1.3 - 1.5 (Curved surface)Nozzle throat radius [m] 0.089Nozzle exit radius [m] 0.87

The moment of inertia matrix and the centre of gravity position for theconfiguration is presented in Equation 14 and 15 respectively. Ixx = 7.7e5 −Ixy = 9.5 −Ixz = −48.7

−Ixy = 9.5 Iyy = 6.9e5 −Iyz = 0−Ixz = −48.704 −Iyz = 0 Izz = 6.9e5

(14)

Gx = 0.064Gy = 0

Gz = −0.762

(15)

22

7.2 Propulsion system

The propulsion system was assumed to use liquid hydrogen, LH2, as fuel and liquidoxygen, LOX, as oxidiser. It is known that it is a difficult task to start a cryogenicengine, and that it requires flushing with LH2. However it was discussed that aliquid cryogenic engine would be the only option for this study, and it had to beassumed that the engines would start. A model of such a propulsion system wasrealised in the software ECOsimPro and ESPSS, ESPSS is a toolkit withinEcosimPro developed by ESA. The main parameters to analyse was the weight andsize to ensure that the dimensions agreed with the size and mass estimations of thespacecraft stated in the technical note. The schematic created in ECOsimPro isshown in figure 17 below.

Figure 17: Schematic of liquid propulsion system in EcosimPro.

Table 2 shows the inputs and outputs of the simulation in EcosimPro. The totalmass of payload and propulsion system is estimated to 33 tonnes, which is then notinclusive of the spacecraft structure mass. As stated earlier a total mass of about35-45 tonnes is expected and it can be concluded that the propulsion system massfits that estimation.

Table 2: Parameters of the liquid propulsion system simulation in EcosimPro.

Propulsion system estimationsP0 [Pa] 50e5LOX mass fraction [%] 85.7LH2 mass fraction [%] 14.3Payload mass [kg] 23000Fuel + Oxidiser mass [kg] 8540Propulsion structure mass [kg] 1450Nozzles + chambers mass [kg] 266Propulsion + PL total mass [kg] 33256Thrust/nozzle [N] 7.3e4

To be able to simulate the correct exhaust mix of the LOX and LH2 in DLR-TAU, achemical file was created and added to the set up of the chamber in the simulations.The data was obtained in EcosimPro and re-calculated to fit the input variables forchemistry files, the used parameters for the exhaust gas can be found in Table 3.

23

Table 3: Chemistry data file parameters for the exhaust gases in DLR-TAU.

Exhaust chemical parametersMolecular weight [kg/mol] 0.027Gas constant gamma [-] 1.3967Prandtl number [-] 0.64Sutherland constant [-] 119.14Sutherland viscosity [kg/m s] 8.1135e-5Sutherland temperature [K] 3000

7.3 Atmosphere of Mars

The atmosphere of Mars consists mostly of CO2 and is very thin compared to theEarths. The chemistry file incorporated in DLR-TAU as a reference of the generalatmosphere is used in the simulations to make the ambient environment as realisticas possible, values can be found in table 4. The atmosphere has a strongexponentially decreasing pressure curve, and hence density curve, while thetemperature decreases linearly with height above the surface. Graphs showing thelevels of pressure and temperature above and below 7 km altitude is shown infigures 18 and 19.

Table 4: Chemistry data file parameters in DLR-TAU for Mars atmosphere.

Atmospheric input parametersMolecular weight [kg/mol] 0.044009Gas constant gamma [-] 1.298Gas constant R [J/K mol] 188Prandtl number [-] 0.7Sutherland constant [-] 240Sutherland viscosity [kg/m s] 8.17e-6Sutherland temperature [K] 164

Figure 18: Atmosphere levels of temperature and pressure above 7 km.

24

Figure 19: Atmosphere levels of temperature and pressure below 7 km.

7.4 Simplifications and parameters not taken into account

The flush surface of the heat cap gives slanted nozzles with different lengthsegments. This will result in different side loads and may cause problems forstability and separation. The reason for angling the nozzles with a 7 degree angle isto separate the plumes slightly and increase stability of the vehicle. These side loadsand the issue of the different length segments has not been studied and is notincluded in the current scope.

The engines are set as cryogenic, operating with liquid oxygen as oxidiser and liquidhydrogen as fuel. This requires flushing with LH2 before start up. That is notpossible if the nozzles are sealed until ignition, which is how they are treated in thisstudy to simplify the start up conditions.

Throttling of the engines is a necessity, but not included in this first study, and theengine performance parameters and estimations at steady state are simplified. It isnot investigated how to recover from an angle of attack, since a ballistic path is usedfor the majority of the trajectory. The possibility of gimbaling the engines, puttingthe system on a pivoted support to be able to change the thrust direction of thenozzles has not been investigated, but could be beneficial. The overall GNC has notbeen studied and would have to be included in future iterations of the designprocess. The idea of a four nozzle vehicle in a propelled descent has however beendiscussed and approved as a first geometry by a GNC expert at ESA.

25

8 Trajectory planning in ASTOS

To envision the descent from a circular orbit around Mars to touchdown using thetrajectory software ASTOS, the different descent phases has to be defined. The onlyphase studied in detail in this thesis is the SRP phase. However, to decide on theconstraints and set up of the SRP phase a preliminary trajectory was used as a baseline for the altitude vs. velocity range and the corresponding atmospheric propertiesetc. This trajectory could then serve as a guide when choosing SRP initiation Machnumber and Mach number range.

• Phase 1: Circular orbit

• Phase 2: Hypersonic entry

• Phase 3: Supersonic Retro Propulsion

• Phase 4: Manoeuvre to vertical position

• Phase 5: Soft pinpoint landing

The orbit is set as circular, the orbital elements are described below.

Table 5: Circular orbit elements of phase 1 implemented in ASTOS.

Orbital parameters

Apoapsis altitude [km] 500

Periapsis altitude [km] 500

Inclination [deg] 27

RAAN [deg] 120

Mean anomaly [deg] 70

Argument of periapsis [deg] 0

A short CFD evaluation of the inflatable heat shield properties at three differentpoints were initially computed to get some preliminary data of the aerodynamics ofthe shield. These values can then be insert into the ASTOS set up file of phase 2,depending on where the SRP phase is set to begin.

Table 6: Summary of outputs from the inflatable heat shield simulations.

CFD results of the inflatable heat shield

Mach number [-] 15 10 5

Altitude [km] 45 40 28

Cd [-] 1.551 1.564 1.618

The third phase is the SRP phase which is defined by the CFD simulations. Theresults of the CFD, especially the thrust levels at each point between Mach 7 and 2,can then be incorporated in ASTOS together with the propulsion systemparameters. The fourth phase is specified to be a manoeuvre to position thespacecraft vertically above the surface, unfold landing legs or other gear, anddescent the very last bit of the trajectory and finally, phase five, touch down gently.

26

9 CFD simulations

To set up and run a CFD simulations a geometry, a mesh and a CFD code is needed.In this study, the mesh surrounding the spacecraft and its enclosure is set up usingCENTAUR, developed by CentaurSoft. The mesh is then converted into a DLR-Taureadable grid file where the boundary conditions can be re-specified in the DLR-Taulanguage. The CFD simulations are computed with the DLR-Tau code developed bythe German Aerospace Center, DLR, and used at ESA. The program has nointerface and is user controlled by programming in python and by manipulation oftext files which the software can read and compute. All parameters, including filesof boundary conditions and chemistry, are specified in an extensive parameter file.The parameter file containing all the information can the be run in a Linuxterminal. The general approach during the viscous CFD simulations is starting thecomputation with a first order upwind scheme and stepping up from a low CFLnumber, sometimes as low as 0.01, up to 1 during a few thousand iterations. Afterobtaining low and steady residuals with the initial first order solution, the upwindscheme is switched to second order. The second order computation is the run forover 100 000 iterations in total. The results from DLR-Tau are converted intoreadable files and the post processing is done in Tecplot and Matlab.

9.1 Start-up conditions

One of the most interesting phases to investigate is the start up of the engines in ahigh Mach number retro flow. The first assessment of the SRP phase had a Machrange of 7 to 2, where Mach 7 indicates ignition of the engines and Mach 2 indicatesthat the vehicle has slowed down enough to perform a vertical manoeuvre and latersoftly land. Mach 7 in the Martian atmosphere and at an altitude of 35 kmcorresponds to a velocity of about 1400 m/s, the full set of atmospheric propertiesat this altitude can be found in table 7. The nozzles are treated as sealed untilignition, when the protective caps covering the nozzles are blown of by the jet. Themain issue is to identify if the jet flow is pushed back into the nozzle, or if the jetstream can overcome the ambient pressure caused by the retro flow of theatmosphere and brake the spacecraft in a steady manner. The reference atmosphericconditions were given by a previous trajectory calculated at ESA, the data can befound in Appendix B.

Table 7: Atmospheric reference values of the first point in the SRP phase, start-up of theengines.

Atmospheric start-up conditionsMach number [-] 7Altitude [km] 35Ambient static pressure [Pa] 26Temperature [K] 164Density [kg/m3] 8.4e-4Velocity [m/s] 1400Speed of sound [m/s] 203

9.1.1 Geometry, model and mesh

To evaluate the jet off conditions, a CAD model with a flushed heat cap and hencesealed nozzles, shown in figure 20, was simulated in a flow of Mach 7.

27

(a) Tilted view. (b) Side view.

Figure 20: CAD model for simulating start-up conditions, nozzles are sealed.

The simulations were carried out with a 3D axisymmetric 45 degree slice of the 3Dmodel enclosed in a cylindrical wedge domain. 2D snapshots of the 45 degree modelenclosed in the domain with and without mesh are shown in figure 21.

(a) Geometry used in DLR-TAU. (b) Fine mesh with adaptation.

Figure 21: 2D snapshots of the 45 degree slice of the 3D geometry and the finest meshused including adaptation, in total 12e6 nodes.

9.1.2 Simulations - Jet off

To determine the jet off conditions four cases were compared in DLR-TAU, oneinviscid case and 3 viscous cases with increasing mesh size from 1 million nodes to12 million nodes, the turbulence model used for the viscous jet off cases were k-wSST. The parameters of interest were the pressure and the bow shock location. Thebow shock location determines how far the jet plume will travel in the lowerpressure environment behind the bow shock once ignition is initiated, before hittingthe largely increased pressure upstream of the bow shock. Figures 22a and 22bshows the solution of the overall dynamic pressure and the pressure on the heatshield behind the bow shock.

28

(a) Pressure distribution. (b) Surface pressure over the heat shield.

Figure 22: Pressure outside and inside bow shock, M∞ = 7.

The results shows a decreasing bow shock stand off distance with increasing numberof mesh nodes, the difference between the cases are approximately 10 cm as seen infigure 23. The free stream pressure does not show large differences between thetested cases, which can also be seen in figure 23. The maximum total pressure of thefree stream at a velocity of Mach 7 was determined to 2.88e5 Pa, which then givesthe minimum jet pressure needed to extend a plume in the opposite direction of thisflow measured in the next chapter.

Figure 23: Comparison of bow shock location vs. total free stream pressure upstream ofthe bow shock, M∞ = 7.

29

The surface pressure, figure 24, shows less difference between the cases. Thedifferent mesh size of the viscous cases does not seem to have a significant effect onthe shield surface pressure behind the bow shock. This is due to a similarly meshedwall region. The only exception is shown at the very nose, x = 0, of the shield wherethe pressure varies from 1450 Pa to 1550 Pa with increasing mesh density, and thisis in correlation with a more closely located bow shock. All simulations converged toa maximum density residual of 1e-3.

Figure 24: Comparison of surface pressure on the heat shield behind the bow shock,M∞ = 7.

Table 8 shows the obtained outputs of the start-up simulation of the viscous casewith the finest grid, that will be used with the jet on case at the Mach 7 trajectorypoint. Moving forward, the boundary layer mesh of the viscous and finest meshedcase will be used. The y+ value of the viscous and finest meshed cased variedbetween 1e-3 and 1. However, since the results of this mesh study did not varymuch in terms of flow field, the mesh of the domain will initially be set as coarseand then refined heavily in areas with sharp Mach number changes and pressuregradients to save computational time in future simulations.

Table 8: Summary of outputs from the simulations of the start-up conditions in DLR-TAU.

Simulation results of start-up conditionsBow shock stand off distance [m] 0.52Ambient dynamic pressure [Pa] 2.8e5Heat shield surface pressure [Pa] 1600Lift coefficient [-] 0Drag coefficient [-] 1.56Ballistic coeff. [-] 93Force x-direction (Drag force) [N] 3.9e4

30

9.2 Retro propulsive phase

The retro-propulsive phase starts at ignition of the engines and ends when thespacecraft has reached a velocity of Mach 2 where a soft landing is within reach. Tosimplify the task, the nozzles are treated as sealed until ignition, even though this isa method used for solid rocket engines and not liquid ones which requires flushingwith cold hydrogen to start. The covers sealing the nozzles are blown off by the jetplume once ignited, and the plumes travel through the bow shock locatedapproximately half a meter upstream of the centre of the heat cap and starts thebraking. To ensure that the jet plume is powerful enough to develop in the retroflow of maximum pressure 2.88e5 Pa as shown in the previous section, and not bepushed back into the nozzle, the jet on conditions in Mach 7 at an altitude of 35 kmin Mars atmosphere was initially studied before continuing with the rest of the retropropulsive phase. The ambient atmospheric parameters described in table 7 areused again, since the point of ignition and development of an exhaust plume ismodelled as instant.

9.2.1 Geometry, model and mesh

The jet on simulations are using a cylindrical enclosure which contains an 45 degreeslice of the 3D axis-symmetric vehicle model and one of the nozzles. The nozzles arenow open, and the model used is shown in figure 25a. The chemical files of thecomposition of the exhaust and the atmosphere of Mars, shown in section 7.2, wereincorporated in the set up of the DLR-TAU code to get the correct species in thechamber and in the ambient flow. The boundary conditions of the set up were asfollows, supersonic inflow at the inlet of the ambient atmosphere, farfield andsupersonic outflow as surrounding walls, symmetry at the two wedge planes creatingthe quarter geometry, pressure inflow at the nozzle inlet and viscous walls for allremaining parts of the spacecraft. The simulations are viscous using theSpalart-Allmaras one equation turbulence model, which is specially designed foraerospace applications. The initial mesh contains 6e5 nodes, it is refined usingDLR-TAU several times during each simulation according to the current flow field toget a more accurate solution, ending up with a final mesh containing approximately1.3e6 nodes. The refinement tool is set to refine areas with a high pressure gradientor areas with a large difference in Mach number, the result can be seen in figure 25b.The wanted y+ value is set to one but ranges between 7e-4 to 7.7 between thedifferent simulations.

(a) CAD model with uncovered nozzles. (b) Mesh with adaptation for SRP flow.

Figure 25: CAD model and example of adapted mesh for jet on simulations.

31

9.2.2 Simulations - Jet on

The initial jet on case at a velocity of Mach 7 was set up as a steady state, viscoussimulation. All atmospheric parameters such as altitude, pressure and density etc.are equal to the jet off simulation. The most interesting parameters are the pressureof the plume and the distance the high pressure extends, since those values can becompared with the jet off results and determine if the jet flow will be pushed backinto the nozzle or if it can overcome the dynamic pressure of the atmosphere.

(a) Mach number curve. (b) Pressure curve.

Figure 26: Pressure and Mach number curve extending along the tilted nozzle axis fromthe throat of the nozzle and forward, M∞ = 7.

In figure 26, the pressure which extends from the throat is shown to be of amagnitude 10 higher than the ambient dynamic pressure of 2.8e5 Pascals at jet offconditions. The jet pressure can also be seen to extend for approximately 32 meters,which is the jet shock location and stagnation point at a velocity of Mach 7. Thebow shock location is 40 meters measured from the exit of the nozzle. In summarythese results show that the pressure force of the engines is high enough to operate atsteady state, in a retro flow of Mach 7. The question the actual ignition and startup moment is more complicated but these results are seen as a positive indicator.

(a) Mach number plume contour, [-]. (b) Pressure plume contour, [Pa].

Figure 27: Pressure and Mach number plumes extending along x [mm] from the nozzle andforward, M∞ = 7.

32

To initiate the first iteration of the design of the retropropulsive phase, pointsbetween 36 and 8 kilometer in altitude on the trajectory were chosen to beexamined further. The last point was set as close to ground level as possible to testa worst case scenario with the available fuel, at an altitude of 8 kilometer and with avelocity of 475 meters per second. The atmospheric properties at the differentpoints of interest are presented in table 9 below.

Table 9: Difference in atmospheric conditions for the points of interest at different altitudeand velocity.

Atmospheric parameters

Mach number [-] 7 6 5 4 3 2

Velocity [m/s] 1400 1252 1069 875 700 475

Altitude [km] 36 32 28 24 12 8

Density [kg/m3] 8.4e-4 1.14e-3 1.6e-3 2.1e-3 5.6e-3 7.6e-3

Temperature [K] 169 179 188 197 223 232

Pressure [Pa] 26 38 56 81 237 340

Dynamic pressure [Pa] 2.8e5 1.3e5 5.2e4 1.6e4 9.5e3 2.6e3

Speed of sound [m/s] 203 209 214 219 233 238

(a) Mach number plume and bow shock, [-]. (b) Velocity contour, [m/s].

Figure 28: Mach number and velocity contour, M∞ = 6.

Figure 28 shows the Mach number and velocity contour at a velocity of Mach 6.The angled nozzle gives a tilted plume from the x axis.

33

(a) Mach number plume contour, [-]. (b) Mach number plume and bow shock, [-].

Figure 29: Mach number plumes, M∞ = 5.

Figure 29 shows the Mach number plume and contour at the trajectory point withvelocity Mach 5. The results does not vary much from the previous trajectory pointat a velocity of Mach 6.

(a) Mach number plume and bow shock, [-]. (b) Pressure plume contour, [Pa].

Figure 30: Pressure and Mach number plumes extending along x from the nozzle andforward, M∞ = 4.

Figure 30 shows the Mach number and pressure contour at the trajectory point withvelocity Mach 4. The bow shock is moving further away from the spacecraft sincethe pressure of the opposing flow is decreasing with the velocity.

34

(a) Mass fraction of exhaust gases. (b) Mass fraction of ambient atmosphere.

Figure 31: Chemistry difference and distribution between exhaust gases and the ambientCO2 dominant atmosphere , M∞ = 4.

Figure 31 shows the difference in the boundary conditions of the nozzle inlet andenclosure inlet, which are modelled with different chemistry files.

(a) Mach number plume and bow shock,[-]. (b) Pressure plume contour, [Pa].

Figure 32: Mach number and pressure, M∞ = 3.

Figure 32 shows the Mach number and pressure contour at a velocity of Mach 3.The bow shock distance from the spacecraft is further increasing which allows theplume to become more elongated since the dynamic pressure against it is decreasing.

35

(a) Mach number plume and bow shock,[-]. (b) Temperature distribution, [K].

Figure 33: , M∞ = 2.

Figure 33 shows the Mach number and temperature contour at the last point in theSRP range. The results at the different points shows an increasing bow shock standoff distance together with a more elongated plume shape and distance withdecreasing velocity, as expected. The results of the six points are summarised intable 10.

Table 10: Summary of outputs of CFD simulations for the retropropulsive phase of 45degree model.

CFD results of retropropulsive phase

Mach number [-] 7 6 5 4 3 2

Altitude [km] 36 32 28 24 12 8

Drag force [N] 3.48e5 3.19e5 3.11e5 3.59e5 2.14e5 3.45e5

Total axial force [N] 4.52e4 4.50e4 4.49e5 4.56e4 4.57e4 4.69e4

Bow shock dist. [m] 40 40 40 50 55 65

Ct [-] 1.831 1.674 1.635 1.892 1.124 1.815

The y plus values of the different cases varies between 7e-4 to 7.7, the desired y plusis set to 1 in all cases. In table 10 the total axial force, Fx, is aligned with the centeraxis of the spacecraft. As mentioned the nozzles are however tilted with an angle of7 degrees outwards from the center axis and have a thrust of 7.3e4 N per enginealong the nozzle axis. All values are for one nozzle.

36

9.2.3 180 degree model

To start an investigation of plume interaction between the nozzles, and to betterasses the flow field, a 180 degree case were set up. The figures below shows thevehicle at the Mach 7 point in the trajectory with two nozzles active. Thesesimulations with two active nozzles confirmed the output data in table 10.

(a) Mach number plume, [-]. (b) Stream traces of the flow field.

Figure 34: Mach number plume and stream traces of the flow field, M∞ = 7.

Figure 34 shows the Mach number plume and stream traces of the flow field withtwo active nozzles. Figure 34b shows the characteristic SRP flow field including thebow shock, plume shape and re-circulation regions also mentioned in the TheoryChapter.

(a) Pressure contour, [Pa], with bow shock. (b) Temperature distribution, [K].

Figure 35: Pressure and temperature contours, M∞ = 7.

Figure 35 shows the pressure and temperature contours with two active nozzles.Figure 35a shows the bow shock contour in 3D surrounding the spacecraft like abowl. Figure 35b shows the temperature distribution with two active nozzles at avelocity of Mach 7. The walls of the spacecraft appears to be at very hightemperatures, which is noted as an error to be discussed in terms of Quality andTrust in CFD.

37

(a) Mach number slice, [-]. (b) Temperature slice, [K].

Figure 36: Mach number and temperature, M∞ = 7.

Figure 36 shows contour slices of Mach number and temperature at a velocity ofMach 7 with two active nozzles. It can be seen in figure 36a that the plume of twonozzles has a different shape compared to the results using one nozzle. Thecombined plume is more elongated and has a shape more similar to a large centrednozzle than a single tilted, which is expected.

38

9.2.4 Similar studies

Since SRP is a technology not yet used in any other atmosphere other than theEarth’s, experimental data for low density and pressure environments is difficult tocome by. However, one study conducted by Karl T. Edquist et al. [17] have done acomparative CFD study including flight data from the first stage of Falcon 9 inMars relevant conditions. However, this study has the interesting and meaningfuldata edited out and only flow characteristics of SRP can be compared. Another,even more relevant CFD study is from the German Aerospace Center [8] alsoanalysing Falcon 9 at re-entry, using the CFD code DLR-TAU. It should be notedthat similarity between this study and other studies in Earths atmosphere is not thegoal of the comparison. It is a comparison to show differences and similarities alike,since no experimental data of SRP is available from Mars atmosphere at this time.

(a) Nozzle temperature, [K] . (b) Nozzle temperature - study by DLR [8].

Figure 37: Comparison of nozzle temperature between this study and the study carriedout by DLR of a Falcon 9 re-entry scenario.

Figure 37a shows the temperature distribution in the converging-diverging nozzleused by DLR in their Falcon 9 study [8]. The conical nozzle used in this studyshows a similar temperature distribution in the nozzle.

(a) Nozzle mach number, [-]. (b) Nozzle Mach number - study by DLR [8].

Figure 38: Comparison of Mach number levels in the nozzle between this study and thestudy carried out by DLR of a Falcon 9 re-entry scenario.

39

Figure 38a shows the Mach number distribution in the converging-diverging nozzlesimulated by DLR [8] compared to the Mach number distribution in the conicalnozzle of this study. The study from DLR [8] focuses on the flow field from 3, of thetotal 9 nozzles, on Falcon 9. None of the nozzles in the Falcon 9 configuration areangled, and the studied Mach number SRP range by DLR is 9.45 to 5.09.

(a) Temperature, [K], at Mach 5.(b) Temperature, [K], 3 nozzles at Mach 5 -study by DLR [8].

Figure 39: Comparison of temperature levels around the vehicles between this study andthe study carried out by DLR of a Falcon 9 re-entry scenario in Earths atmosphere [8] .

Figure 39 shows plume shape and temperature distribution of this study and the onefrom DLR [8]. The influence of the inflated heat shield can be noted, even thoughthe temperatures of this study seems to be higher at the back body/cylinder eitherway. Figure 40 shows the plume shape of the two nozzles in this study and theshape from three nozzles on the Falcon 9. It can be seen that the plume shapes aremore similar between two and three nozzles, than for one and three which is seen tobe quite different in the previous figure, figure 39, however an expected observation.

(a) Mach contour with 2 active nozzles,M∞ = 7.

(b) Temperature, [K], 3 nozzles at Mach 5-study by DLR [8].

Figure 40: Mach and temperature contours from this study and the study carried out byDLR of a Falcon 9 re-entry scenario in Earths atmosphere [8] .

40

10 Discussion

This study has the main focus of obtaining aerodynamic data of SRP in Marsatmosphere, and this goal has been achieved. It was clear after the literature studythat the research within this field is about to grow exponentially, there are manynew ideas on how to model and simulate SRP in the most accurate and efficientway. The fact that SRP is making waves on Earth right now with re-usability andhigher profit as main objectives is of course a reason why this technology is makinga comeback in the discussions of future space exploration.

The initial challenge of this project was to scale down the mission, make reasonableassumptions and set up mission constraints to be able to begin the process. First upwas the design, which is based on a vehicle design in a previous study by ESA [1] in2002, but it had to be redone to ensure a safer and more simple descent trajectory.The old vehicle had the SRP engines positioned at the opposite end of the inflatableheatshield, which would indicate that a 180 degree turn would have to happensomewhere along the descent trajectory, this option was ruled out in discussionswith ESA experts for such a heavy vehicle. The design created in this study isgeometrically simple and could be furthered to be much more detailed in futureiterations. The new design of the vehicle in this thesis, however, gave new questionssuch as the slanted nozzles positioned in the heat cap, which may give rise toundesired side loads. The side loads were not studied in this thesis but should benoted as a possible future problem. Perhaps there would be smaller side loads with8, 12 or 16 smaller engines tilted with less than 7 degrees as in this study, or with noangle at all. The chosen configuration is just one of many imaginable, but the scopewas to start as simple as possible not to rule anything possibilities or problems toosoon. To ensure that the lander created in this study was not impossible from aGuidance, Navigation and Control, GNC, point of view. A GNC expert at ESA wasshortly consulted and gave the green light of the design with SRP capabilities as afirst configuration.

Another issue early on was deciding on how to model the cryogenic propulsionsystem and the nozzles, which in the end had to be simplified in order to speed upthe process due to the time limitation. As seen in section 8.2.4, using conical nozzlesdoes not give a significantly different nozzle flow even though it is not likely to beused in a real scenario. The simplification of treating the nozzles as sealed untilignition worked well for the steady state cases, which was the goal. Even though itis well known that the nozzles would need to be flushed with LH2 before igniting thecryogenic engine, it was not in the scope to investigate the start up in a moredetailed manner than on/off conditions. It was discussed what to do with theinflatable heat shield after the SRP phase begins, it could still be attached butdeflated and folded flat against the cylinder, or it could stay inflated or it could beblown off. It was decided to keep it inflated as this is a simple solution and since itwas seen as extra protection. Additionally, it was discussed that it might notinterfere much once the SRP phase begins and the rocket engines take over thebraking.

Some problematic areas that should be noted and improved in the set up of theCFD simulations is first of all the wall temperature of the spacecraft which is notspecified in this study, but which could be set to some reasonable value to avoidhigh wall temperatures in the results. Secondly, the symmetry boundary conditionfor the 45 degree model has to be meshed finer along the symmetry axis to avoidclustering of heat which can be seen in some temperature contours, or otherdiscrepancies that should not appear there, it also affects the convergence and cancause a simulation to crash in some cases. This was done later on in the CFDprocess when the problem was identified, and should be noted for future mesh

41

generation of similar bodies. The results of the CFD simulations are deemed correctenough for this initial study. Some quality and trust parameters of the CFD can bediscussed, the y+ value increases with the Mach number as expected but is stillwithin reasonable limits at the highest flow velocity which indicates a well meshedboundary layer. As mentioned before, the discretization of the rest of the domainshould be studied in more detail, especially close to the nozzle around the symmetryaxis. Misunderstandings or mistakes in the setup can be a source of error since thisis a very complex task. The lack of previous experience in DLR-Tau is also a factorthat should be taken into account.

The data obtained can be compared to CFD studies of SRP in Earth atmosphere,such as the DLR study mentioned [8] in section 8.2.4, where comparable results interms of flow fields etc. can be found even though the atmospheres are verydifferent. Since there is very limited amounts of published flight data for SRPavailable, it is difficult to find experimental data to compare with. The paper fromNASA in collaboration with SpaceX [17] claims to examine flight data from Falcon 9in Mars relevant conditions, in the Earths atmosphere, but since the actual data isedited out in that particular paper not many conclusions can be drawn.

The work process has to a large content consisted of CFD set up and simulations,and due to the complex flow and high velocities it has taken large amounts of bothtime and computer power to get correct results. Many attempts finally resulted in awork flow where one steady simulation of a point on the trajectory could be fullycalculated within 24 hours. The thrust level and outputs of the CFD, which can befound in table 10, can for example be updated by changing the velocity at a pointand running the simulation again. The method of iterating between CFD inDLR-TAU, trajectory calculations in ASTOS and CAD design updates in CATIAhas worked well even though the ASTOS part was minimal in this case due to thetime limitation. However, since the set up and simulation of the CFD part is themost extensive and time consuming it took quite some time and research just toobtain the first results. Since there was no real trajectory to follow from thebeginning, a lot of assumptions and good guesses had to be set in order to moveforward and start running the tests. Now, that there is a more clear understandingof the problematic areas in the set up of the CFD, and the first set of results aredone, the process can be accelerated. All data, results, set ups of simulations,vehicle designs, propulsion models etc. are stored at ESA and used as a steppingstone for continuing this optimisation process. The first iteration is the mostdifficult due to the many mission constraints and ideas that need to be discussedand decided on. Since this was first and foremost an aerodynamic study,simplifications at the expenses of other areas had to be made.

10.1 Future work

The engine start up is the most critical stage, and it would therefore be the mostuseful and interesting part to investigate in more detail. More results of engine startup conditions would also affect the available SRP range, which could be altered in afuture study. It would also be useful to further the model in EcosimPRO, testdifferent parameters of the propulsion system to optimise the fuel, mass andpressure levels in the chamber to minimise the weight. The CAD model could thenbe refined with a more detailed propulsion system integrated in the body, withchambers, tanks etc. This would make it possible to simulate sloshing and otherengine properties in both the CFD environment and in EcosimPRO. The next thingto add in the optimisation process, after ASTOS is completed for this iteration, isinput from GNC is some way. The question of gimbaling of the nozzles must beaddressed. Also, the configuration of the engines and nozzles would have to belooked into, maybe a set of 12 or 16 smaller nozzles are advantageous in terms ofsteering, recovering from angle of attack and safety. The GNC requirements will

42

have a large influence on the vehicle design and propulsion system. The heat shieldcould both be put in the design optimisation, to optimise the diameter, or it couldbe held constant and instead investigate how to deflate or jettison it once the SRPphase begins, or conclude that it should be kept inflated until landing.

43

11 Cited Literature

References

[1] Guillermo Ortega Hernando. Mission Scenarios and Vehicle DesignDocument - 2002. ESA ESTEC.

[2] Jan Ostlund. SUPERSONIC FLOW SEPARATION WITHAPPLICATION TO ROCKET ENGINE NOZZLES - 2004. StockholmRoyal Institute of Technology.

[3] George P. Sutton and Oscar Biblarz. Rocket Propulsion Elements. 7thEdition.

[4] Ashley M. Korzun1, Robert D. Braun1, and Juan R. Cruz2. Survey ofSupersonic Retropropulsion Technology for Mars Entry, Descent,and Landing - 2008. (1) Georgia Institute of Technology and (2) NASALangley Research Center.

[5] ISECG International Space Exploration Coordination Group. GlobalExploration Roadmap - 3rd Edition - 2018.

[6] Max Braun, Paul Bruce, and Errikos Levis. Strategies to Utilise AdvancedHeat Shield Technology for High-Payload Mars Atmospheric EntryMissions - 2017. Imperial College London.

[7] Robert J. McGhee. EFFECTS OF A RETRONOZZLE LOCATED ATTHE APEX OF A 140 BLUNT CONE AT MACH NUMBERS OF3.00, 4.50, AND 6.00 - 1971. NASA Langley Research Center.

[8] Tobias Eckery 1, Franziska Zilkery 1, Etienne Dumont 1, Sebastian Karly 1, andKlaus Hannemanny 1. AEROTHERMAL ANALYSIS OF REUSABLELAUNCHER SYSTEMS DURING RETRO-PROPULSIONREENTRY AND LANDING - 2018. (1) German Aerospace Center(DLR).

[9] Robert D. Braun1, Brandon Sforzo2, and Charles H. Campbell3. AdvancingSupersonic Retropropulsion Using Mars-Relevant Flight Data: AnOverview - 2017. (1)University of Colorado at Boulder, (2) Georgia Instituteof Technology, (3) Lyndon B. Johnson Space Center.

[10] Jan Ostlund 1. Flow Processes in Rocket Engine Nozzles with Focuson Flow Separation and Side Loads. - 2002. (1) KTH Royal Institute ofTechnology.

[11] Torsten Wintergerste 1 Michel Casey 1. Flow Processes in Rocket EngineNozzles with Focus on Flow Separation and Side Loads. - 2002. (1)Fluid Dynamics Laboratory Sulzer Innotech.

[12] V.G Perminov. The difficult road to Mars - A brief history of Marsexploration in the Soviet Union - 1999. NASA History division.

[13] Robert J. Parks. Astronautical Research - Mariner 9 and theExploration of Mars - 1972. NASA - Jet Propulsion Laboratory.

[14] A. Chicarro, P. Martin, and R. Trautner. The Mars Express Mission: AnOverview - 2003. ESA ESTEC.

44

[15] NASA. InSight - Entry, Descent and Landing.https://mars.nasa.gov/insight/timeline/landing/entry-descent-landing/.

[16] David W. Way1, Richard W. Powell1, Allen Chen2, Adam D. Steltzner2,A. Miguel San Martin2, P. Daniel Burkhart2, and Gavin F. Mendeck3. MarsScience Laboratory: Entry, Descent, and Landing SystemPerformance - 2006. (1)NASA Langley Research Center, (2) NASA JetPropulsion Laboratory, (3) NASA Johnson Space Flight Center .

[17] Karl T. Edquist1, Ashley M. Korzun1, Karen L. Bibb1, Daniel G.Schauerhamer 2 Edward C. Ma3, Peter L. McCloud 4, Grant E. Palmer 5, andJoshua D. Monk5. Comparison of Navier-Stokes Flow Solvers to Falcon9 Supersonic Retropropulsion Flight Data - 2017. (1)NASA LangleyResearch Center, (2) MRI Technologies, (3) Jacobs Technology, (4) ERC Inc.,(5) Analytical Mechanics Associates Inc.

[18] Dr. Guillermo Ortega Hernando1, Alvaro Martinez Barrio1, and MichaelKhan2. TRAJECTORY ANALYSIS OF MARTIAN ENTRYPROBES FOR THE ESA EXPLORATION PROGRAM - 2002. (1)ESA ESTEC, (2) ESA ESOC.

[19] Walter E. Bruce1, Nathaniel J. Mesick1, Paul G. Ferlemann1, Paul M.Siemers2, Joseph A. Del Corso1 Stephen J. Hughes1, Steven A. Tobin1, andMatthew P. Kardell3. AEROTHERMAL GROUND TESTING OFFLEXIBLE THERMAL PROTECTION SYSTEMS FORHYPERSONIC INFLATABLE AERODYNAMIC DECELERATORS- 2012. (1) NASA, (2) Analytical Services and Materials, Inc., (3) BoeingTechnology Services.

[20] Alicia D. Cianciolo1, Thomas A. Zang1, Ronald R. Sostaric2, and M. KathyMcguire3. OVERVIEW OF THE NASA ENTRY, DESCENT ANDLANDING SYSTEMS ANALYSIS EXPLORATIONFEED-FORWARD STUDY - 2011. (1)NASA Langley Research Center,(2)NASA Johnson Space Center, (3) NASA Ames Research Center.

45

Appendices

A

I General 1D rocket nozzle equations

Expr =AeAt

m =Atp0√T0

√γ

R(γ + 1

2)−

γ+12(γ−1)

AeAt

= (γ + 1

2)−

γ+12(γ−1) (

1 + γ−12 M2

e

Me)

γ+12(γ−1)

TeTtot

= (1 +γ − 1

2M2e )−1

peptot

= (1 +γ − 1

2M2e )−

γγ−1

ue = Me

√γRTe

F = mue + (pe − p∞)Ae

Isp =F

mg

mfuel =

∫ landing

ignition

FtIsp · g

dt

46

II Validation case data

47

B

I Reference atmospheric conditions

48

II Collected data

49

50

Inertia matrix and axis system in relation to the body.

51