University of Tennessee, Knoxville TRACE: Tennessee Research and Creative Exchange Doctoral Dissertations Graduate School 12-2002 Design considerations for a pressure-driven multi-stage rocket Steven Craig Sauerwein University of Tennessee Follow this and additional works at: https://trace.tennessee.edu/utk_graddiss Recommended Citation Sauerwein, Steven Craig, "Design considerations for a pressure-driven multi-stage rocket. " PhD diss., University of Tennessee, 2002. https://trace.tennessee.edu/utk_graddiss/6301 This Dissertation 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 Doctoral Dissertations 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 dissertation written by Steven Craig Sauerwein entitled "Design considerations for a pressure-driven multi-stage rocket." I have examined the final electronic copy of this dissertation for form and content and recommend that it be accepted in partial fulfillment of the equirr ements for the degree of Doctor of Philosophy, with a major in Aerospace Engineering. Gary A. Flandro, Major Professor We have read this dissertation and recommend its acceptance: 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.) To the Graduate Council: I am submitting herewith a dissertation written by Steven Craig Sauerwein entitled "Design Considerations for a Pressure-Driven Multi-Stage Rocket." I have examined the final paper copy of this dissertation for form and content and recommend that it be accepted in partial fulfillment of the requirements for the degree of Doctor of Philosophy, with a major in Aerospace Engineering. We have read this dissertation and recommend its acceptance: Accepted for the Council: Graduate Studies DESIGN CONSIDERATIONS FOR A PRESSURE-DRIVEN MULTI-STAGE ROCKET A Dissertation Presented for the Doctor of Philosophy Degree The University of Tennessee, Knoxville Steven Craig Sauerwein December 2002 ABSTRACT The purpose of this study was to examine the feasibility of using propellant tank pressurization to eliminate the use of high-pressure turbopumps in multi-stage liquid-fueled satellite launchers. Several new technologies were examined to reduce the mass of such a rocket. Composite materials have a greater strength-to-weight ratio than metals and can be used to reduce the weight of rocket propellant tanks and structure. Catalytically combined hydrogen and oxygen can be used to heat pressurization gas, greatly reducing the amount of gas required. Ablatively cooled rocket engines can reduce the complexity and cost of the rocket. Methods were derived to estimate the mass of the various rocket components. These included a method to calculate the amount of gas needed to pressurize a propellant tank by modeling the behavior ofthe pressurization gas as the liquid propellant flows out of the tank. A way to estimate the mass and size of a ablatively cooled composite cased rocket engine. And a method to model the flight of such a rocket through the atmosphere in conjunction with optimization of the rockets trajectory. The results show that while a liquid propellant rocket using tank pressurization are larger than solid propellant rockets and turbopump driven liquid propellant rockets, they are not impractically large. 11 TABLE OF CONTENTS CHAPTER PAGE I. INTRODUCTION............................................................................................ 1 II. TANK MATERIAL.......................................................................................... 16 III. TANK CONFIGURATION.............................................................................. 25 IV. PROPELLANTTANKPRESSURIZATION................................................... 31 V. PRESSURIZATION SYSTEM......................................................................... 47 VI. ROCKET ENGINE DESIGN............................................................................ 60 VII. ROCKET ENGINE CHAMBER PRESSURE.................................................. 69 VIII. NUMBER OF STAGES.................................................................................... 76 IX. BOOSTER PERFORMANCE.......................................................................... 82 X. BOOSTER COMPARISON............................................................................. 86 XI. CONCLUSION................................................................................................ 89 LIST OF REFERENCES.............................................................................................. 96 APPENDICES.............................................................................................................. 101 APPENDIX A. FIGURES............................................................................................. 102 APPENDIX B. TRAJECTORY SIMULATION............................................................ 153 APPENDIXC. TRAJECTORYOPTIMIZATION........................................................ 170 VITA ................................................................................................................................ 196 l1l LIST OF TABLES TABLE PAGE 1. Material Properties............................................................................................ 23 2. Configuration 4 Length to Width Ratio............................................................ 29 3. Experimental Values for Lewis and Lockheed H2 Experiments....................... 44 4. Experimental Values for Lewis and Lockheed He Experiments....................... 44 5. Pressurization Gas Mass.................................................................................... 45 6. Liquid Hydrogen/Liquid Oxygen O/F Ratio...................................................... 66 7. Mass Comparison.............................................................................................. 67 8. Optimum Engine Chamber Pressure, Thrust to Weight Ratio= 1.5.................. 73 9. Optimum Engine Chamber Pressure, Thrust to Weight Ratio= 2.0.................. 73 10. Mass Penalty ofNon-optimal Chamber Pressure, Thrust to Weight Ratio= 1.5 74 11. Mass Penalty ofNon-optimal Chamber Pressure, Thrust to Weight Ratio= 2.0 75 12. RocketNozzleOptimumAltitude..................................................................... 78 13. 1st Stage Burnout Altitude.................................................................................. 79 14. Hydrogen Fueled Rocket with 100 kg Payload................................................... 84 15. Hydrogen Fueled Rocket with 1000 kg Payload................................................. 84 16. Hydrogen Fueled Rocket with 10,000 kg Payload.............................................. 84 17. Kerosene Fueled Rocket with 100 kg Payload.................................................. 84 18. Kerosene Fueled Rocket with 1000 kg Payload................................................ 85 19. Kerosene Fueled Rocket with 10,000 kg Payload............................................. 85 20. Booster Mass...................................................................................................... 87 IV 21. Booster Comparison........................................................................................... 87 B-1 Drag Coefficients.............................................................................................. 15 8 C-1 Typical Convergence Behavior............................................................................ 181 C-2 Final Burn Magnitude versus Scalar k................................................................ 181 V LIST OF FIGURES FIGURE PAGE 1. Tank Liner Mass............................................................................................... 103 2. Spherical Tank Mass vs. Pressure.................................................................... 103 3. Spherical Tank Mass vs. Volume...................................................................... 104 4. Cylindrical Tank Mass vs. Pressure.................................................................. 104 5. Cylindrical Tank Mass vs. Volume.................................................................. 105 6. Tank Configuration.......................................................................................... I 06 7. Cylindrical Tank Mass versus Radius, 10,000 kg Stage.................................... 107 8. Cylindrical Tank Mass versus Radius, 100,000 kg Stage.................................. 107 9. Tank Configuration, 10,000 kg Stage, 2 mm Thick Aeroshell.......................... 108 10. Tank Configuration, 10,000 kg Stage, 5 mm ThickAeroshell.......................... 108 11. Tank Configuration, 10,000 kg Stage, 11 mm Thick Aeroshell........................ 109 12. Tank Configuration, 100,000 kg Stage, 2 mm Thick Aeroshell........................ 109 13. Tank Configuration, 100,000 kg Stage, 5 mm Thick Aeroshell........................ 110 14. Tank Configuration, 100,000 kg Stage, 11 mm Thick Aeroshell...................... 110 15. Multiple Cylindrical Tank Configurations....................................................... 111 16. Multiple Cylindrical Tank Mass....................................................................... 112 17. Massvs.NumberofCylindricalTanks............................................................