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No description available.

Trajectory optimization is an important part of launch vehicle design and operation. With the high costs of launching payload into orbit, every pound that can be saved increases affordability. One way to save weight in launch vehicle design and operation is by optimizing the ascent trajectory. Launch vehicle trajectory optimization is a field that has been studied since the 1950’s. Originally, analytic solutions were sought because computers were slow and inefficient. With the advent of computers, however, different algorithms were developed for the purpose of trajectory optimization. Computer resources were still limited, and as such the algorithms were limited to local optimization methods, which can get stuck in specific regions of the design space. Local methods for trajectory optimization have been well studied and developed. Computer technology continues to advance, and in recent years global optimization has become available for application to a wide variety of problems, including trajectory optimization. The aim of this thesis is to create a methodology that applies global optimization to the trajectory optimization problem. Using information from a global search, the optimization design space can be reduced and a much smaller design space can be analyzed using already existing local methods. This allows for areas of interest in the design space to be identified and further studied and helps overcome the fact that many local methods can get stuck in local optima. The design space included in trajectory optimization is also considered in this thesis. The typical optimization variables are initial conditions and flight control variables. For direct optimization methods, the trajectory phase structure is currently chosen a priori. Including trajectory phase structure variables in the optimization process can yield better solutions. The methodology and phase structure optimization is demonstrated using an earth-to-orbit trajectory of a Delta IV Medium launch vehicle. Different methods of performing the global search and reducing the design space are compared. Local optimization is performed using the industry standard trajectory optimization tool POST. Finally, methods for varying the trajectory phase structure are presented and the results are compared.

South Africa is one of the few developing countries able to design and build satellites; however it is reliant on other nations to launch them. This research addresses one of the main technological barriers currently limiting an indigenous launch capacity, namely the development of a locally designed liquid fuel turbopump. The turbopump is designed to function in an engine system for a commercial launch vehicle (CLV) with the capacity to launch 50-500 kg payloads to 500 km sun synchronous orbits (SSO) from a South African launch site. This work focuses on the hydrodynamic design of the impeller, vaneless diffuser and volute for a kerosene (RP-1) fuel pump. The design is based on performance analyses conducted using 1D meanline and quasi-3D multi-stream tube (MST) calculations, executed using PUMPAL and AxCent software respectively. Specific concerns that are dealt with include the suction performance, cavitation mitigation, efficiency and stability of the pump. The design is intended to be a relatively simple solution, appropriate for a South African CLV application. For this reason the pump utilises a single impeller stage without a separate inducer element, limiting the design speed. The pump is designed to run at 14500 rpm while generating 889 m of head at a flowrate of 103.3 kg/s and consuming 1127.8 kW of power. The impeller has six blades with an outer diameter of 186.7 mm and axial length of 84.6 mm. The impeller's high speed and power requirement make full scale testing in a laboratory impractical. As testing will be a critical component in the University of KwaZulu-Natal's turbopump research program, this work also addresses the scaling down of the impeller for testing. The revised performance and base dimensions of the scaled impeller are determined using the Buckingham-Pi based scaling rules. The test impeller is designed to run at 5000 rpm with a geometric reduction of 20%, using water as the testing medium. This gives an outer diameter of 147.8 mm and an axial length of 69.9 mm. At its design point the test impeller generates a total dynamic headrise of 67.7 m at a flow rate of 18 kg/s, with a power requirement of 15 kW. A method for maintaining a similar operating characteristic to the full scale design is proposed, whereby the scaled impeller's blade angle distribution is modified to maintain a similar diffusion characteristic and blade loading profile. This technique is validated by MST analysis for off-design conditions with respect to both speed and flowrate. / M.Sc.Eng. University of KwaZulu-Natal, Durban 2013.

This thesis examines the way legal space risks are managed in commercial space transportation provided by major carriers, such as, NASA, the US private launch sector, and Arianespace, as well as in the system envisaged for Australia. Its purpose is to show that even if all systems tend to provide a favorable risk allocation scheme to the space launch industry, there are several alternatives for the telecommunications satellite operators. It also attempts to show that, even if all these risk sharing regimes have been modeled after NASA's, there are certain important differences, which stem from the different political objectives of each of the countries where these carriers are inserted.

During the lift-off of a launch vehicle, the acoustic pressure fluctuations caused by the engine exhaust gases produce high noise levels inside the cavity of the fairing and can damage the payload inside the fairing. Hence reducing the noise transmitted into the payload bay is an important area of research. Work presented in this thesis investigates the external acoustic pressure excitations on the fairing of a launch vehicle during the lift-off acoustic environment. In particular, it investigates the external sound pressure levels in the low frequency range from 50Hz to 400Hz, on the fairing during the lift-off of a launch vehicle. This study establishes theoretical and numerical models for the prediction of external sound pressure loading on composite structures representing launch vehicles, such as a large composite cylinder referred to as a Boeing cylinder and a Representative Small Launch Vehicle Fairing (RSLVF). To predict the external sound pressure loading, various incident wave conditions were investigated, including incident plane waves, oblique plane waves and oblique plane waves with random phases that strike the circumference of the composite structures. For the theoretical model, both the incident and scattered sound pressure fields due to incident plane waves; perpendicular to an idealised long cylinder were investigated. The results show that the scattered sound pressure field plays a major role in determining the total circumferential sound pressure field at the surface of the cylinder and cannot be ignored for the launch case. The theoretical model was developed further for a point source, line source and oblique incident waves, and modified to determine the incident, scattered and total sound pressure fields away from the cylinder. The approach developed overcomes some limitations of previous analytical derivations. An experiment was undertaken to determine the sound pressure patterns at the surface of a cylinder at various frequencies due to a point source positioned at a finite distance from the cylinder surface. The experimental work confirmed the accuracy of the theoretical model for a point source at a finite distance from the cylinder. The Boundary Element Method (BEM), approach was used for the numerical investigation of the acoustic loadings. The numerical analysis was developed for various acoustic loading conditions and verified with the theoretical results, which showed that the numerical and theoretical models agree well. Both models were extended to a Boeing composite cylinder and an RSLVF for various acoustic loading conditions. The complex acoustic environment generated during the lift-off of a launch vehicle was investigated and used as a basis for the acoustic loading on an RSLVF. To predict the acoustic excitations on an RSLVF, two different source allocation techniques were investigated, which considered acoustic sources along the rocket engine exhaust flow. The investigations were conducted both numerically and analytically. Both results agree well and show that it is possible to predict the acoustic loads on the fairing numerically and analytically. / http://proxy.library.adelaide.edu.au/login?url= http://library.adelaide.edu.au/cgi-bin/Pwebrecon.cgi?BBID=1347443 / Thesis (Ph.D.) -- University of Adelaide, School of Mechanical Engineering, 2008

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No description available.

A large fleet (in the hundreds) of high quality telescopes are used for tracking and imaging of launch vehicles during ascent from Cape Canaveral Air Force Station and Kennedy Space Center. A maintenance tool has been development for use with these telescopes. The tool requires rankings of telescope condition in terms of the ability to generate useful imagery. It is thus a case of ranking telescope conditions on the basis of the perceptual image quality of their imagery. Perceptual image quality metrics that are well-correlated to observer opinions of image quality have been available for several decades. However, these are quite limited in their applications, not being designed to compare various optical systems. The perceptual correlation of the metrics implies that a constant image quality curve (such as the boundary between two qualitative categories labeled as excellent and good) would have a constant value of the metric. This is not the case if the optical system parameters (such as object distance or aperture diameter) are varied. No published data on such direct variation is available and this dissertation presents an investigation made into the perceptual metric responses as system parameters are varied. This investigation leads to some non-intuitive conclusions. The perceptual metrics are reviewed as well as more common metrics and their inability to perform in the necessary manner for the research of interest. Perceptual test methods are also reviewed, as is the human visual system. iv Image formation theory is presented in a non-traditional form, yielding the surprising result that perceptual image quality is invariant under changes in focal length if the final displayed image remains constant. Experimental results are presented of changes in perceived image quality as aperture diameter is varied. Results are analyzed and shortcomings in the process and metrics are discussed. Using the test results, predictions are made about the form of the metric response to object distance variations, and subsequent testing was conducted to validate the predictions. The utility of the results, limitations of applicability, and the immediate ability to further generalize the results is presented.

This study investigates the use of a Model Reference Adaptive Control (MRAC) direct approach to solve the attitude control problem of an Expendable Launch Vehicle (ELV) during its boost phase of flight. The adaptive autopilot design is based on Lyapunov Stability Theory and provides a useful means for controlling the ELV in the presence of environmental and dynamical uncertainties. Several different basis functions are employed to approximate the nonlinear parametric uncertainties in the system dynamics. The control system is designed so that the desire dresponse to a reference model would be tracked by the closed-loop system. The reference model is obtained via the feedback linearization technique applied to the nonlinear ELV dynamics. The adaptive control method is then applied to a representative ELV longitudinal motion, specifically the 6th flight of Atlas-Centaur launch vehicle (AC-6) in 1965. The simulation results presented are compared to that of the actual AC-6 post-flight trajectory reconstruction. Recommendations are made for modification and future applications of the method for several other ELV dynamics issues, such as control saturation, engine inertia, flexible body dynamics, and sloshing of liquid fuels.