In this work the conditions for which a (0,90,90,0,0,90)s BMI-carbon fiber laminate will initiate transverse microcracking are determined for the fabrication of a cryogenic fuel tank for use in a Reusable Launch Vehicle (RLV). This is accomplished using a quadratic interaction criterion failure analysis on the total stress state at possible launch conditions. There are three major sources of stress, that is, thermal residual stress, internal pressure stress, and applied load stress, that are evaluated at the launch stage to determine the total stress state. To assess the accuracy of the analysis the well known X-33 cryogenic fuel tank failure was analyzed as an example. The results of the X-33 example show that the analysis accurately portrays the failure of the X-33 and provides evidence that the analysis can be used to provide reliable conditions for the initiation of microcracking. The final result of this study is a range of launch conditions that can be used without the initiation of microcracking and a limiting range of conditions that cause complete microcracking throughout the laminate.

In 2004, the NASA Astronaut Office produced a memo regarding the safety of next generation launch vehicles. The memo requested that these vehicles have a probability of loss of crew of at most 1 in 1000 flights, which represents nearly an order of magnitude decrease from current vehicles. The goal of LOC of 1 in 1000 flights has since been adopted by the launch vehicle design community as a requirement for the safety of future vehicles. This research addresses the gap between current vehicles and future goals by improving the capture of vehicle architecture effects on reliability and safety. Vehicle architecture pertains to the physical description of the vehicle itself, which includes manned or unmanned, number of stages, number of engines per stage, engine cycle types, redundancy, etc. During the operations phase of the vehicle life-cycle it is clear that each of these parameters will have an inherent effect on the reliability and safety of the vehicle. However, the vehicle architecture is typically determined during the early conceptual design phase when a baseline vehicle is selected. Unless a great amount of money and effort is spent, the architecture will remain relatively constant from conceptual design through operations. Due to the fact that the vehicle architecture is essentially “locked-in” during early design, it is expected that much of the vehicle's reliability potential will also be locked-in. This observation leads to the conclusion that improvement of vehicle reliability and safety in the area of vehicle architecture must be completed during early design. Evaluation of the effects of different architecture decisions must be performed prior to baseline selection, which helps to identify a vehicle that is most likely to meet the reliability and safety requirements when it reaches operations. Although methods exist for evaluating reliability and safety during early design, weaknesses exist when trying to evaluate all architecture effects simultaneously. The goal of this research was therefore to formulate and implement a method that is capable of quantitatively evaluating vehicle architecture effects on reliability and safety during early conceptual design. The ConcepTual Reliability Growth Approach for CompariSon of Launch Vehicle ArchiTectures (CONTRAST) was developed to meet this goal. Using the strengths of existing techniques a hybrid approach was developed, which utilizes a reliability growth projection to evaluate the vehicles. The growth models are first applied at the subsystem level and then a vehicle level projection is generated using a simple system level fault tree. This approach allows for the capture of all trades of interest at the subsystem level as well as many possible trades at the assembly level. The CONTRAST method is first tested on an example problem, which compares the method output to actual data from the Space Transportation System (STS). This example problem illustrates the ability of the CONTRAST method to capture reliability growth trends seen during vehicle operations. It also serves as a validation for the development of the reliability growth model assumptions for future applications of the method. The final chapter of the thesis applies the CONTRAST method to a relevant launch vehicle, the Space Launch System (SLS), which is currently under development. Within the application problem, the output of the method is first used to check that the primary research objective has been met. Next, the output is compared to a state-of-the-art tool in order to demonstrate the ability of the CONTRAST method to alleviate one of the primary consequences of using existing techniques. The final section within this chapter presents an analysis of the booster and upper stage block upgrade options for the SLS vehicle. A study of the upgrade options was carried out because the CONTRAST method is uniquely suited to look at the effects of such strategies. The results from the study of SLS block upgrades give interesting observations regarding the desired development order and upgrade strategy. Ultimately this application problem demonstrates the merits of applying the CONTRAST method during early design. This approach provides the designer with more information in regard to the expected reliability of the vehicle, which will ultimately enable the selection of a vehicle baseline that is most likely to meet the future requirements.

Generally, a ground telemetry station for launch vehicle (LV) has tracking function only; therefore, position measurements depend on radar. Time of arrival (TOA) and angle of arrival (AOA) are typical location techniques for emitting targets. In this paper, we propose a Combined TOA and AOA localization method for LV using two ground stations. When transmitter (Tx) time is not known, it is necessary to make virtual onboard timer for TOA estimation. The virtual onboard timer generates time stamps of streaming frame according to data rate. First station which is located in space center has no tracking function. But it can generate the virtual onboard timer. Second station has tracking function, so it generates AOA information. By solving sphere equation(s) of TOA from at least one station and a line equation of AOA, target position in three-dimensions (3D) can be obtained. We confirm the localization performance by means of comparison with an on-board GPS of a real launch mission.

Stability is an important criterion in the design and performance of launch vehicles. Present day launch vehicles have become more and more flexible due to the constraints of weight reduction, necessarily imposed for enhanced performance of the vehicle. Due to higher flexibility, the launch vehicle stability becomes a concern. Instability in the launch vehicles has been noticed due to three major sources: thrust, aerodynamic forces and combustion induced instabilities. Instability in the launch vehicles may pose problem to the structural integrity leading to structural failure or it may lead to the deviation in the trajectory of the vehicle. Several structural failures of launch vehicles due to instabilities have been reported in the literature. The prediction of the structural response due to various excitations such as thrust and aerodynamic loading is essential to identify any failure scenarios and to limit the vibrations transmitted to the payload. Therefore, determination of dynamic and stability characteristics of a launch vehicle under the influence of different parameters, is of vital importance. Disciplines such as, flight mechanics (dynamics), structural dynamics, aerodynamics, propulsion, guidance and control are closely related in the design and analysis of launch vehicles. Typically, flight mechanics, guidance and control problems consider a rigid vehicle for modeling and simulation purposes. The disciplines of structural dynamics and aeroelasticity consider a flexible vehicle. In order to bring in the effect of flexibility on the flight dynamics of the launch vehicle, structural dynamics and aeroelasticity aspects need to be effected. The preliminary design of a new launch vehicle requires inputs from different disciplines and parametric studies are required to finalise the vehicle configuration. The study of the effect of different parameters on the dynamics and stability of launch vehicles is required. In this context, there is a need to develop an integrated approach that provides tools for the design and analysis of a launch vehicle. The availability of integrated modeling and simulation tools will reduce the requirement of costly prototype development and testing. In the present thesis, an attempt has been made to develop a numerical tool to conduct parametric studies for launch vehicle dynamics and stability. The developed tool is suitable for prediction of onset of instabilities under the influence of different parameters. The approach developed in this thesis is also well suited for specialized analysis of problems involving vertical launch, stage separation, engine shutdown and internal stress wave propagation related to structural integrity. Stability problems due to thrust and the aerodynamic forces (aeroelastic stability) in the launch vehicles/ missiles have been reported in the literature. Most of these works have modeled the vehicle as a beam or by using discrete degrees of freedom. In these works, the effect of thrust or aerodynamic forces on the flexible body modes is investigated and it is shown that the instability may occur in one of the bending modes due to change in the parameters such as thrust or aerodynamic forces. Traditionally, the dynamic characteristics are obtained in a body-fixed coordinate system, whereas the prediction of trajectory (rigid body dynamics) is carried out in an inertial frame of reference. Only few works have addressed the coupling of the rigid body motion and the flexible body dynamics of a vehicle. But these works also, do not consider the total derivative of displacements with respect to an inertial frame of reference. When the integrated equations of motion are derived in an inertial frame of reference, the rigid body motion and the elastic displacements are highly coupled. In this thesis, the rigid body motion and the flexible body dynamics is studied in an inertial frame of reference. The flexible body dynamics of the moving vehicle is studied in an inertial frame of reference, including velocity induced curvature effects, which have not been considered so far in the published literature. A detailed mechanics based model is developed to analyze the problem of structural instabilities in launch vehicles. Coupling among the rigid-body modes, the longitudinal vibrational modes and the transverse vibrational modes due to asymmetric lifting-body cross-section are considered. The model also incorporates the effects of aerodynamic forces and the propulsive thrust of the vehicle. The propulsive thrust is considered as a follower force. The model is one-dimensional, and it can be employed to idealized slender vehicles with complex shapes. The governing differential equations along with the boundary conditions are derived using Extended Hamilton’s principle. Subsequently, the modeling of the propulsive thrust and the aerodynamic forces are included in the formulation. In the literature, the propulsive thrust has generally been modeled as a follower force applied at the nozzle end. Few of the works in the literature have modeled the combustion process in the solid rocket motor and the liquid propellant engine in detail. This is required to understand the combustion induced instabilities. In the present thesis, the propulsive thrust is considered as a follower force and few of the combustion parameters affecting the thrust are considered. In the literature, the modeling of the aerodynamic forces acting on a launch vehicle has been carried out using general purpose computational fluid dynamics (CFD) codes or by using empirical methods. CFD codes are used to obtain the pressure and the shear stress distribution on the vehicle surface by the solution of Navier Stokes/ Euler equations. The empirical methods have been used to obtain the distributed aerodynamic forces acting on the vehicle. The aerodynamic forces are expressed in terms of distributed aerodynamic coefficients. In the present work, the modeling of the aerodynamic forces has been carried out in two different ways: using a CFD package and by using empirical methods. The stability of a system can be studied by determining the system response with time. Eigenvalue analysis is another tool to investigate the stability of a linear system. To study the stability characteristics of the system using eigenvalue analysis, a computational framework has been developed. For this purpose, the finite element discretization of the system is carried out. Further to that, two different methods are utilized for finite element discretization of the vehicle structure: Fourier Transform based Spectral Finite Element method (SFEM) and an hp Finite Element method (FEM). The conventional FEM is a versatile tool for modeling complicated structures and to obtain the solution of the system of equations for a variety of forcing functions. The SFEM is more suitable for obtaining the solution for simple 1D and 2D structures subjected to shock and transient loads, having high frequency content. In this thesis, the spectral finite element model is developed for a vehicle subjected to the propulsive thrust and the aerodynamic forces. Prediction of instability using SFEM, means solving a nonlinear eigenvalue problem. Standard computer codes or routines are not available for solving a nonlinear eigenvalue problem. A computer code has been written to solve the nonlinear eigenvalue problem using one of the algorithms available in the literature. An hp finite element model is also developed for launch vehicle. The finite element stiffness and damping matrices due to the thrust, the aerodynamic forces and the rigid body velocity and acceleration are derived using Lagrange’s equations of motion. A standard linear eigenvalue problem and a polynomial eigenvalue problem is formulated for determination of instability regimes of the vehicle. It is important to understand the influence of different parameters such as thrust, velocity, angle of attack etc. on the stability of a launch vehicle. Parametric studies are important during the preliminary design phase of a vehicle to identify the instability regimes. The design parameters can be changed to reduce the possibility of instabilities. Numerical simulations are carried out to determine the unstable regimes of a slender launch vehicle for propulsive thrust and velocity as the parameters, neglecting the aerodynamic forces. Comparison between the results based on a Fourier spectral finite element model and a hp finite element model are carried out. Phenomenon of static instability (divergence) and dynamic instability (flutter) are observed. Determination of mode shapes of the vehicle is important for deciding the placement of sensors and actuators on the vehicle. In this context, eigenvectors (mode shapes) for different end thrust and speed are analyzed. Further, numerical simulations are also carried out to determine the instabilities in a slender launch vehicle considering the combined effects of propulsive thrust, aerodynamic forces and mass variation. The finite element model simulation results for aeroelastic effects are compared with the published literature. Stability of a vehicle is analysed for velocity (free stream Mach number) as a parameter, at maximum propulsive thrust, including the effect of aerodynamic forces and mass variation. Phenomenon of static instability (divergence) and dynamic instability (flutter) are observed. With the increase in the Mach number, branching (splitting) and merging of the modes is observed. At higher Mach numbers, divergence and flutter are observed in different modes simultaneously. Numerical simulations are carried out for a typical nosecone launch vehicle configuration to analyse the aeroelastic stability at two different Mach numbers using empirical aerodynamic data. The phenomenon of flow separation and reattachment is observed at the cone-cylinder junction. The stability of a typical vehicle under propulsive thrust and aerodynamic forces is investigated using CFD derived aerodynamic data. The aerodynamic pressure and shear stress distribution for a launch vehicle are obtained from the CFD analysis. The effect of different parameters such as combustion chamber pressure, tip mass and slenderness ratio on the stability of a vehicle is studied. In the later part of the thesis, solution methodology for the time domain response for a coupled axial and transverse motion of a vehicle is developed. The axial responses (displacements and velocities) of a typical vehicle subjected to axial thrust are determined using direct integration of the equations of motion. The axial displacements due to two different thrust histories are compared. The axial velocities with time at different locations are determined. The time domain and the frequency domain responses for a representative vehicle subjected to a transverse shock force are determined using Spectral Finite Element method (SFEM). The system of equations for a coupled axial and transverse motion of a vehicle is developed. Numerical simulations are carried out to determine the coupled axial and transverse response of a vehicle subjected to axial and transverse forces. The coupling of rigid body motion with the elastic displacements is illustrated. The thesis is comprised of seven chapters. The first chapter gives a detailed introduction to launch vehicles and covers literature survey of launch vehicle dynamics and stability. The dynamics and stability related aspects of flexible structures are also discussed. In chapter 2, a detailed mathematical model of a slender launch vehicle is developed to analyze the problem of structural instabilities. Chapter 3 deals with the finite element discretization of the vehicle structure using two different methods: Fourier spectral finite element method and an hp finite element method. In chapters 4 and 5, numerical simulations are carried out to determine the instabilities in a slender launch vehicle considering the effects of propulsive thrust, aerodynamic forces and mass variation. In chapter 6, solution methodology for the time domain response for a coupled axial and transverse motion of a vehicle is developed. The last chapter gives the conclusions and the future scope of work. To summarize, this thesis is a comprehensive document, that not only describes some detailed mathematical models for launch vehicle stability studies, but also presents the effect of aerodynamic, propulsion and structural loads on the launch vehicle stability. Linear stability analysis of a representative vehicle is carried out for prediction of onset of the instabilities under the influence of different parameters such as velocity, thrust, combustion factors etc. The correlation between the stability analysis and the time domain response is established. In short, the matter presented in this thesis can serve as a useful design aide for those working in the launch vehicle design.

An analytical modal response tool was developed to investigate the characteristics of and to estimate static and dynamic launch vehicle responses to ground wind loads (GWL). The motivation of this study was to estimate the magnitude of response of the Ares I-X launch vehicle to ground winds and wind-induced oscillation (WIO) during roll-out and on the pad. This method can be extended to other launch vehicle designs or structures that possess a nearly cylindrical cross-section. Presented in this thesis is an overview of the theory used, a comparison of the theory with wind tunnel data, further investigation of the data to support the assumptions used within the analysis, and a prediction of the full-scale Ares I-X response. Additionally, an analytical investigation is presented that estimates the effect of atmospheric turbulence on WIO response. Most of the wind tunnel data presented in this report is taken from the GWL Checkout Model tested in the NASA Langley Transonic Dynamics Tunnel (TDT) in April 2007. The objective of the GWL Checkout Model was to reestablish and evaluate the capability of the facility to conduct GWL testing and to operate the associated equipment. This wind tunnel test was not necessarily intended to predict the full scale Ares vehicle response to GWL; however, it can be used to help validate the newly developed analytical method described in this thesis. A detailed GWL test incorporating updated vehicle designs and launch pad configurations of the Ares I-X flight test vehicle was also conducted in the TDT during the fall of 2008. This test provides more accurate predictions of the second bending mode response of the Ares I-X, and it models effects of the nearby tower and support structures. The proposed analytical method is also compared to select data from the Ares I-X GWL test; however, it is presented as normalized values to protect the sensitivity of the data. Results of the proposed analytical method show reasonable correlation to wind tunnel data. Also, this method was the first to determine that second bending mode WIO response was not only possible for the Ares I-X, but will also produce the most critical loads. Finally, an explanation is offered in this thesis regarding discrepancies between wind tunnel and full-scale WIO response data. / Master of Science

ITC/USA 2005 Conference Proceedings / The Forty-First Annual International Telemetering Conference and Technical Exhibition / October 24-27, 2005 / Riviera Hotel & Convention Center, Las Vegas, Nevada / This paper presents the dynamic RF link estimation result for telemetry system of KSLV (Korea Space Launch Vehicle)-I. In particular, it utilizes the parameters of the instantaneous vehicle antenna gain pattern in three dimensions, the improvement by polarization diversity combiner at the ground receiver, and the free space propagation loss. The structural transformation and discontinuity of ground plane after the separation events of nose fairing, stage, and spacecraft, are also included in this analysis. As a consequence, the prediction of link variation has been performed in accordance with ARDP (Antenna Radiation Distribution Plot) and look angle trace of vehicle. In addition, the optimum position of onboard antennas has been investigated to provide better RF link margin in the nominal trajectory.

International Telemetering Conference Proceedings / October 26-29, 1998 / Town & Country Resort Hotel and Convention Center, San Diego, California / Development of an extended test range, with range instrumentation providing continuous vehicle communications, is required to flight-test the X-33, a scaled version of a reusable launch vehicle. The extended test range provides vehicle communications coverage from California to landing at Montana or Utah. This paper provides an overview of the approaches used to meet X-33 program requirements, including using multiple ground stations, and methods to reduce problems caused by reentry plasma radio frequency blackout. The advances used to develop the extended test range show other hypersonic and access-to-space programs can benefit from the development of the extended test range.

International Telemetering Conference Proceedings / October 23-26, 2000 / Town & Country Hotel and Conference Center, San Diego, California / NASA Marshall Space Flight Center’s Bantam Robust Guidance Navigation & Control Project is investigating off the shelf navigation sensors that may be inexpensively combined into Kalman filters specifically tuned for launch and orbital vehicles. For this purpose, Marshall has purchased several GPS receivers and is evaluating them for these applications. The paper will discuss the receiver selection criteria and the test equipment used for evaluation. An overview of the analysis will be presented including the evaluation used to determine their success or failure. It will conclude with goals of the program and a recommendation for all GPS users.

ITC/USA 2011 Conference Proceedings / The Forty-Seventh Annual International Telemetering Conference and Technical Exhibition / October 24-27, 2011 / Bally's Las Vegas, Las Vegas, Nevada / Spacecraft and launch vehicle reliability is dominated by premature equipment failures and surprise equipment failures that increase risk and decrease safety, mission assurance and effectiveness. Large, complex aerospace systems such as aircraft, launch vehicle and satellites are first subjected to most exhaustive and comprehensive acceptance testing program used in any industry and yet suffer from the highest premature failure rates. Desired/required spacecraft equipment performance is confirmed during factory testing using telemetry, however equipment mission life requirement is not measured but calculated manually and so the equipment that will fail prematurely are not identified and replaced before use. Spacecraft equipment mission-life is not measured and confirmed before launch as performance is but calculated using stochastic equations from probability reliability analysis engineering standards such as MIL STD 217. The change in the engineering practices used to manufacture and test spacecraft necessary to identify the equipment that will fail prematurely include using a prognostic and health management (PHM) program. A PHM includes using predictive algorithms to convert equipment telemetry into a measurement of equipment remaining usable life. A PHM makes the generation, collection, storage and engineering and scientific analysis of equipment performance data "mission critical" rather than just nice-to-have engineering information.

Recent developments in technology are changing the trend both in satellite design and application of that technology. As the number of small satellites built by experts from academia and private companies increases, more effective ways of inserting those satellites into orbit is needed. Among the various studies that focus on the launch of such small satellites, research on design of Launch Vehicle tailored for nano-satellites attracts special attention. In this thesis, Multiple Cooling Multi Objective Simulated Annealing algorithm is applied for the conceptual design of Launch vehicle for nano-satellites. A set of fitness functions are cooled individually, and acceptance is based on the maximum value of the acceptance probabilities calculated. Angle of attack and propulsion characteristics are employed as optimization parameters. Algorithm finds the optimum trajectory as well as the design parameters that satisfies user defined constraints. In this study burnout velocity, and payload mass are defined as objectives. The methodolgy is applied for different design scenarios including multistage, air and ground launch vehicles.