University of Tennessee, Knoxville TRACE: Tennessee Research and Creative Exchange Masters Theses Graduate School 5-2012 Optimization and Design for Heavy Lift Launch Vehicles Paul Andreas Ritter University of Tennessee - Knoxville, [email protected] Follow this and additional works at: https://trace.tennessee.edu/utk_gradthes Part of the Aerodynamics and Fluid Mechanics Commons, Astrodynamics Commons, Propulsion and Power Commons, Space Vehicles Commons, and the Systems Engineering and Multidisciplinary Design Optimization Commons Recommended Citation Ritter, Paul Andreas, "Optimization and Design for Heavy Lift Launch Vehicles. " Master's Thesis, University of Tennessee, 2012. https://trace.tennessee.edu/utk_gradthes/1197 This Thesis is brought to you for free and open access by the Graduate School at TRACE: Tennessee Research and Creative Exchange. It has been accepted for inclusion in Masters Theses by an authorized administrator of TRACE: Tennessee Research and Creative Exchange. For more information, please contact [email protected]. To the Graduate Council: I am submitting herewith a thesis written by Paul Andreas Ritter entitled "Optimization and Design for Heavy Lift Launch Vehicles." I have examined the final electronic copy of this thesis for form and content and recommend that it be accepted in partial fulfillment of the requirements for the degree of Master of Science, with a major in Aerospace Engineering. James E. Lyne, Major Professor We have read this thesis and recommend its acceptance: Robert W. McAmis, Kivanc Ekici Accepted for the Council: Carolyn R. Hodges Vice Provost and Dean of the Graduate School (Original signatures are on file with official studentecor r ds.) Optimization and Design for Heavy Lift Launch Vehicles A Thesis Presented for a Masters of Science The University of Tennessee, Knoxville Paul Andreas Ritter May 2012 Copyright © by Paul Andreas Ritter All rights reserved ii DEDICATION This thesis is dedicated to both Jesus Christ and my wonderful family who have both made completing this project possible. iii ACKNOWLEDGEMENTS I would like to thank Dr. Evans Lyne for inspiring and pushing me to work on this project. Without this encouragement, I would not have developed a deeper understanding of the components of launch vehicles and the tools necessary to accurately model such a vehicle. These ideas and inspiration have provided a methodology and work ethic that I will be able to use the rest of my life. I would also like to thank Dr. Robert McAmis and Dr. Kivanc Ekici who serve on the thesis committee. I would also like to thank NASA for providing me with the opportunity to work with them. Specifically, I am very grateful to Dr. Jonathon Jones and Jason Campbell for their direction and willingness to answer the many questions that I had when developing the modeling code. There are also numerous people that answered many questions for me and provided encouragement when things got tough. I would like to thank my Mom, Dad, and brother, the Hixson family, Dr. John Blevins, Stephen Phillips, Dale Jackson, Rachel Johnson, Tim Kibbey, and to anyone else who might have offered an encouraging word. iv Abstract The simulation and evaluation of an orbital launch vehicle requires consideration of numerous factors. These factors include, but are not limited to the propulsion system, aerodynamic effects, rotation of the earth, oblateness, and gravity. A trajectory simulation that considers these different factors is generated by a code developed for this thesis titled Trajectories for Heavy-lift Evaluation and Optimization (THEO). THEO is a validated trajectory simulation code with the ability to model numerous launch configurations. THEO also has the capability to provide the means for an optimization objective. Optimization of a launch vehicle can be specified in terms of many different variables. For a heavy lift launch vehicle in this thesis, the goal of optimization is to minimize Gross Lift Off Weight (GLOW). THEO provides the capability to optimize by simulating hundreds of thousands of trajectories for a single configuration through the variation of preset independent variables. The sheer volume of these trajectories provides the means to locate configurations that minimize GLOW. Optimization can also be performed by determining the minimum amount of energy necessary to reach target burnout conditions. The energy requirements are then correlated to the propellant mass which can be used to estimate GLOW. This thesis first discusses the validation of THEO as a simulation program and the properties associated with accurately modeling a trajectory. It then relates how THEO and other developed tools can be utilized to determine a configuration that is optimized to minimize GLOW to orbit for adaptable payload sizes. v Table of Contents Chapter 1 – Introduction 1 1.1. Original Thesis Objective 2 1.2. Revised Objective 2 1.3. Thesis Significance 3 Chapter 2 – Trajectory Tool Development and Validation 4 2.1. Options for Trajectory Modeling 5 2.1.1. Matlab 5 2.1.2. POST 5 2.1.3. Introducing a New Code 6 2.2. THEO Objectives 6 2.2.1. Programming Language Selection 6 2.3. Launch Trajectory Description 7 2.3.1. Relating Pitch and Flight Path Angle 7 2.3.2. Pitch Angle Conditions 7 2.3.3. Vehicle Properties 8 2.4. THEO Validation 8 2.4.1. Validation Constraints 8 2.4.2. Validation Configuration 9 2.4.3. Gravity and Planet Properties 10 2.4.4. Atmospheric Model 10 2.4.5. Level of Trajectory Complexity 12 2.4.6. Equation of Motion 13 2.4.7. Method of trajectory Modeling within THEO 14 2.4.8. Revised Equations of Motion 15 2.4.9. Aerodynamic Coefficients 16 2.4.10. Thrust Profile 18 2.4.11. Validation of Output Data 19 2.5. THEO as an Optimization Tool 25 2.5.1. Defining a Successful Case 25 2.5.2. Additional Independent Variables 26 2.6. Data Output Files 26 2.7. Completed FORTRAN Improvements 27 Chapter 3 – Aerodynamic Considerations and Evaluation 29 3.1. Atmospheric Model 30 3.2. Aerodynamic Effects 30 3.2.1. Lift 30 3.2.2. Drag 31 3.2.3. Dynamic Pressure 33 3.2.4. Aerodynamic Heating 33 3.2.5. Aerodynamic Acoustics 33 3.2.6. Static Stability 34 3.3. Missile DATCOM 34 3.3.1. Purpose and Capability 34 vi 3.3.2. Limitations 35 3.4. Nose Cone Analysis 35 3.4.1. Shroud Consderations 35 3.4.2. Nose Cone Types 37 3.4.3. Nose Cone Reference Model 38 3.4.4. Nose Cone Performance and Properties 39 3.4.5. Final Fairing Selection 42 3.5. Aerodynamic Coefficient Estimation 43 Chapter 4 – Rocket Propulsion 48 4.1. Rocket Propulsion Systems 49 4.1.1. Liquid 49 4.1.2. Solid 50 4.1.3. Hybrid 50 4.1.4. Nuclear 50 4.1.5. Electric 50 4.2. Rocket Performance 51 4.2.1. Evaluating Thrust 51 4.2.2. Nozzle Inefficiencies 51 4.2.3. Specific Impulse 52 4.2.4. The Rocket Equation 52 4.3. Propulsion System Advantages and Disadvantages 54 4.3.1. Solid Propulsion Systems 54 4.3.2. Liquid Propulsion Systems 56 4.4. Optimization Engine Selections 58 4.4.1. Solid Rocket Boosters 58 4.4.2. Liquid Rocket Engines 59 Chapter 5 – Independent Variables and Computation Considerations 61 5.1. Independent Variables 62 5.1.1. Primary Pitch Event Start Time 62 5.1.2. Length of Primary Pitch Event 63 5.1.3. Primary Pitch Event Pitchrate 66 5.1.4. Stage Propellant Mass 66 5.2. THEO Time Step 69 Chapter 6 – Analysis Methodology and Secondary Tools 71 6.1.Vehicle Optimization Utilizing THEO 72 6.1.1. THEO Decision Logic 72 6.1.2. Three Dimensional Visualization 72 6.1.3. Intersection Curves 74 6.1.4. Variable to Optimize 75 6.2. Rocket Equation Optimization 76 6.2.1. Mathematical Approach 76 6.2.2. Velocity Budget Analysis Approach 77 6.3. Optimization Methods Comparison 80 vii Chapter 7 – Launch Vehicle Analysis and Results 82 7.1. Configuration Types 83 7.2. 2.5 Stage RSRM Vehicle 85 7.2.1. 2.5 Stage RSRM Aerodynamic Coefficients 85 7.2.2. Inert Mass Correction 86 7.2.3. 2.5 Stage RSRM General Analysis 87 7.2.4. 2.5 Stage RSRM Refined Analysis 90 7.2.5. 100 mt Payload Requirement for 2.5 Stage RSRM 95 7.2.6. 60 mt Payload Requirement for 2.5 Stage RSRM 97 7.3. 2.5 Stage RSRM V Vehicle 98 7.3.1. 2.5 Stage RSRM V Aerodynamic Coefficients 98 7.3.2. 2.5 Stage RSRM V General Analysis 99 7.3.3. 2.5 Stage RSRM V Refined Analysis 102 7.3.4. 100 mt Payload Requirement for 2.5 Stage RSRM V 104 7.3.5. 60 mt Payload Requirement for 2.5 Stage RSRM V 107 7.4. 2.5 Stage M550 FW3 Vehicle 108 7.4.1. 2.5 Stage M550 Aerodynamic Coefficients 109 7.4.2. 2.5 Stage M550 General Analysis 109 7.4.3. 2.5 Stage RM550 Refined Analysis 111 7.4.4. 100 mt Payload Requirement for 2.5 Stage M550 117 7.4.5. 60 mt Payload Requirement for 2.5 Stage M550 119 7.5. 2.5 Stage P80 Vehicle 121 7.5.1. 2.5 Stage P80 Aerodynamic Coefficients 121 7.5.2. 2.5 Stage P80 General Analysis 121 7.5.3. 2.5 Stage P80 Refined Analysis 124 7.5.4. 100 mt Payload Requirement for 2.5 Stage P80 129 7.5.5. 60 mt Payload Requirement for 2.5 Stage P80 130 7.6. 2 Stage Vehicle 131 7.6.1.