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«DESIGN OF AN INTEGRAL THERMAL PROTECTION SYSTEM FOR FUTURE SPACE VEHICLES By SATISH KUMAR BAPANAPALLI A DISSERTATION PRESENTED TO THE GRADUATE SCHOOL ...»

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DESIGN OF AN INTEGRAL THERMAL PROTECTION SYSTEM FOR FUTURE SPACE

VEHICLES

By

SATISH KUMAR BAPANAPALLI

A DISSERTATION PRESENTED TO THE GRADUATE SCHOOL

OF THE UNIVERSITY OF FLORIDA IN PARTIAL FULFILLMENT

OF THE REQUIREMENTS FOR THE DEGREE OF

DOCTOR OF PHILOSOPHY

UNIVERSITY OF FLORIDA

© 2007 Satish Kumar Bapanapalli To my loving wife Debamitra, my parents Nagasurya and Adinarayana Bapanapalli, brother Gopi Krishna and sister Lavanya

ACKNOWLEDGMENTS

I would like to express my sincere gratitude to my advisor and mentor Dr. Bhavani Sankar for his constant support (financial and otherwise) and motivation throughout my PhD studies. He allowed me to work with freedom, was always supportive of my ideas and provided constant motivation for my research work, which helped me grow into a mature and confident researcher under his tutelage. I also thank my committee co-chair Dr. Rafi Haftka for his invaluable inputs and guidance, which have been instrumental for my research work. I am also grateful to him for getting my interest into the field of Structural Optimization, which I hope would be a huge part of all my future research endeavors. I am also thankful to Dr. Max Blosser (NASA Langley) for his crucial inputs in my research work, which kept us on track with the expectations of NASA.

I sincerely thank my dissertation committee members Dr. Ashok Kumar and Dr. Gary Consolazio for evaluating my research work and my candidature for the PhD degree. I also would like to acknowledge Dr. Peter Ifju and Dr. Nam-Ho Kim for their useful inputs and comments. I am thankful to NASA, for their financial support through the CUIP Project, and the program manager Claudia Meyer.

Thanks also go to all the past and current members of Center of Advanced Composites:

Ryan, Jongyoon, Huadong, Oscar, Jianlong, Thi, Sujith and Ben, and other graduate students in the department: Christian, Tushar, Vijay, and Palani.

TABLE OF CONTENTS

page ACKNOWLEDGMENTS

LIST OF TABLES

LIST OF FIGURES

ABSTRACT

CHAPTER 1 INTRODUCTION AND OBJECTIVES

1.1 Introduction

1.2 Requirements of a Thermal Protection System

1.3 Objectives

1.4 Approach

1.4.1 The Optimization Problem

1.4.2 Geometry and Design Variables

1.4.3 Analysis: Finite Element Methods

1.4.4 Loads and Boundary Conditions

1.4.5 Formulation of Constraints

1.4.6 Optimized Designs

2 BACKGROUND

2.1 Approaches to Thermal Protection

2.1.1 Active TPS

2.1.2 Semi-passive TPS

2.1.3 Passive TPS

2.2 TPS and NASA

2.3 Need for a load-bearing TPS: ITPS

2.4 ITPS Design Challenges

2.5 Choice of Constraints

2.6 Background on Corrugated-Core and Truss-Core Sandwich Structures

2.7 Background on Multi-Disciplinary Optimization and Response Surface Approximation Techniques

3 FINITE ELEMENT MODELS AND ANALYSES

3.1 Finite Element Analysis for Heat Transfer

3.1.1 Incident Heat Flux and Radiation Equilibrium Temperature

3.1.2 Loads, Boundary Conditions and Assumptions

3.1.3 One-Dimensional and Two-Dimensional FE Models

3.1.4 Temperature vs. Reentry Time and Temperature Distribution

3.1.5 Obtaining Temperature Data from the FE Analysis

3.2 Finite Element Buckling Analysis

3.3 Stress and Deflection Analysis

4 RESPONSE SURFACE APPROXIMATIONS AND OPTIMIZATION PROCEDURE

4.1 Response Surface Approximations

4.1.1 Response Surface Approximations for Maximum Bottom Face Sheet Temperature

4.1.2 Response Surface Approximations for Buckling

4.1.3 Response Surface Approximations for Stress and Deflection

4.2 Procedure for Generation of Response Surfaces Approximations

4.3 Optimization Procedure

5 ITPS DESIGNS

5.1 Selection of Loads, Boundary Conditions and Other Input Parameters

5.2 Corrugated-Core Designs

5.2.1 Accuracy of Response Surface Approximations

5.2.2 Optimized Corrugated-Core Panel Designs

5.2.3 Buckling Eigen Values, Deflections and Stresses at Different Reentry Times......90 5.2.4 Optimized Designs with Changed Boundary Conditions

5.3 Truss-Core Structures for ITPS

6 EFFECT OF INPUT PARAMETERS ON THE ITPS DESIGN

6.1 Sensitivity of ITPS Designs to Heat Transfer Parameters

6.1.1 Changing the Emissivity of the Top Surface of ITPS

6.1.2 Allowing Heat Loss from the Bottom Face Sheet

6.1.3 Increasing the initial temperature of the structure

6.2 Sensitivity of ITPS Designs to Loads and Boundary Conditions

6.2.1 Effect of Boundary Conditions

6.2.2 Increasing the Pressure Load on the Top Surface





6.2.3 Increasing the In-Plane Load

7 CONCLUSIONS AND FUTURE WORK

7.1 Conclusions

7.2 Future Work

APPENDIX: MATERIAL PROPERTIES USED FOR THE RESEARCH WORK..................116 LIST OF REFERENCES

BIOGRAPHICAL SKETCH

–  –  –

3-1 Peak radiation equilibrium temperatures for reentry heat fluxes..

3-2 Load steps in the FE heat transfer analysis

5-1 Summary of mechanical loads applied on the ITPS panel for buckling and stress analysis cases

5-2 Ranges of the 7 design variables for corrugated-core ITPS panels

5-3 Accuracy of the response surface approximations for peak bottom face sheet temperature and top face sheet deflections.

5-4 Accuracy of the response surface approximations for top face sheet and web buckling eigenvalues

5-5 Table showing the accuracy of the response surface approximations for stresses in the ITPS panel

5-6 Table listing the optimized designs for corrugated-core panels

5-7 Table listing the optimized designs for corrugated-core panels with relaxed boundary conditions on the bottom face sheet

5-8 Comparison of values predicted by response surface approximations and finite element analysis for Designs I and IV

5-9 Comparison of the stresses in beam model and shell model..

6-1 Changes in ITPS design due to decrease in emissivity value to 0.7

6-2 Changes in ITPS design due to loss of heat from the bottom face sheet corresponding to a bottom face sheet emissivity value of 0.2..

6-3 ITPS designs with increase in initial temperature to 395 K. Initial temperature in the baseline model was 295 K.

6-4 ITPS designs with increase in pressure load to 2 atmospheres..

6-5 ITPS designs with increase in in-plane load to 150,000 N/m..

A-1 Temperature dependent material properties of Ti-6Al-4V

A-2 Temperature dependent material properties of Inconel-718

A-3 Temperature dependent material properties of Beryllium alloy

A-4 Temperature dependent material properties of SAFFIL®

–  –  –

1-1 A unit cell of the corrugated-core ITPS design..

2-1 Pictures of the space capsules

2-2 Apollo heat shield structure

2-3 Distribution of TPS on the Space Shuttle Orbiter

2-5 ARMOR TPS construction.

2-6 Prepackaged superalloy honeycomb panel

2-7 Corrugated-core sandwich structure construction for ITPS applications

3-1 Heating profiles of a Shuttle-like vehicle..

3-2 Typical heating profile used for the ITPS design.

3-3 Schematic of loading and boundary conditions for the heat transfer problem

3-4 Typical mesh for 2-d heat transfer problem..

3-5 Schematic representation of 1-D heat transfer model.

3-6 Comparison of 1-d and 2-d heat transfer analyses..

3-7 Temperature variation vs. reentry times for top and bottom surfaces and web mid-point obtained from 1-d heat transfer analysis

3-8 Temperature distribution through the thickness of the ITPS panel at different reentry times

3-9 Typical FE shell element mesh for buckling analysis...

3-10 Typical ITPS panel illustrating the manner in which the webs are partitioned into 10 regions to impose approximate temperature dependent material properties.

3-11 Typical buckling modes of the ITPS panel.

4-1 Flowchart illustrating the procedure followed by the ITPS Optimizer..

5-1 Heat flux input used for the design of corrugated-core structures.

5-2 Aerodynamic pressure load on the TPS for a Space Shuttle-like design.

5-3 A unit cell of a corrugated-core ITPS panel illustrating the 6 design variables

5-4 A typical FE contour plot illustrating the stresses at the panel edges at the junction between the face sheets and the webs.

5-5 Buckling modes for optimized designs.

5-6 Typical web deformation tendency due to the temperature gradient in the panel..................90 5-7 Von Mises stress distribution in the ITPS panel, Design 1, tmax∆T.

5-7 ITPS panel behavior with respect to reentry time for Design 1.

5-8 ITPS panel behavior with respect to reentry time for Design 2.

5-9 Stress singularity at beam-plate junction points..

5-10 FE experiment to compare beam model to a more realistic shell model.

5-11 Truss-core model with only shell elements.

6-1 Heat flux entering the ITPS through the top surface for two different emissivity values imposed on Design I..

–  –  –

Chair: Bhavani V. Sankar Cochair: Raphael T. Haftka Major: Mechanical Engineering Thermal protection systems (TPS) are the features incorporated into a spacecraft’s design to protect it from severe aerodynamic heating during high-speed travel through planetary atmospheres. The ablative TPS on the space capsule Apollo and ceramic tiles and blankets on the Space Shuttle Orbiter were designed as add-ons to the main load-bearing structure of the vehicles. They are usually incompatible with the structure due to mismatch in coefficient of thermal expansion and as a result the robustness of the external surface of the spacecraft is compromised. This could potentially lead to catastrophic consequences because the TPS forms the external surface of the vehicle and is subjected to numerous other loads like aerodynamic pressure loads, small object high-speed impacts and handling damages during maintenance. In order to make the spacecraft external surface robust, an Integral Thermal Protection System (ITPS) concept has been proposed in this research in which the load-bearing structure and the TPS are combined into one single structure.

The design of an ITPS is a formidable task because the requirement of a load-bearing structure and a TPS are often contradictory to one another. The design process has been formulated as an optimization problem with mass per unit area of the ITPS as the objective function and the various functions of the ITPS were formulated as constraints. This is a multidisciplinary design optimization problem involving heat transfer and structural analysis fields.

The constraints were expressed as response surface approximations obtained from a large number of finite element analyses, which were carried out with combinations of design variables obtained from an optimized Latin-Hypercube sampling scheme. A MATLAB® code has been developed to carry out these FE analyses automatically in conjunction with ABAQUS®.

Corrugated-core structures were designed for ITPS applications with loads and boundary conditions similar to that of a Space Shuttle-like vehicle. Temperature, buckling, deflection and stress constraints were considered for the design process. An optimized mass ranging between 3.5–5 lb/ft2 was achieved by the design. This is considerably heavier when compared to conventional TPS designs. However, the ITPS can withstand substantially large mechanical loads when compared to the conventional TPS. Truss-core geometries used for ITPS design in this research were found to be unsuitable as they could not withstand large thermal gradients frequently encountered in ITPS applications.

The corrugated-core design was used for further studying the influence of the various input parameters on the final design weight of the ITPS. It was observed that boundary conditions not only significantly influence the ITPS design but also have a major impact on the effect of various input parameters. It was found that even a small amount of heat loss from bottom face sheet leads to significant reduction in ITPS weight. Aluminum and Beryllium are the most suitable materials for bottom face sheet with Beryllium having considerable advantages in terms of heat capacity, stiffness and density. Although ceramic matrix composites have many superior properties when compared to metal alloys (Titanium alloys and Inconel), their low tensile strength presents difficulties in ITPS applications.

–  –  –

One of the most important goals of the space industry in the new century is to reduce the cost of access to space. Large amount of money is being invested into launching satellites for various civilian and military purposes, building the International Space Station and planning future missions to Moon and Mars. According to estimates [1], the cost of delivering a pound of payload into space in the period 1990 to 2000 varied from $4000 to $20,000∗. There has been no significant improvement over this period in the form of development of a new vehicle or implementation of new concepts for space launch. Therefore, it can be assumed that the cost of launching a pound of payload has increased considerably. One of the long-standing goals of National Aeronautics and Space Administration (NASA) has been to reduce the launch cost by an order of magnitude (to about $1000 per pound [2]).



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