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NASAâ s VAriable Specific Impulse Magnetoplasma Rocket Engine (VASIMR) will provide a highly efficient propulsion source that can dramatically reduce Martian transit times, provide for more abort contingencies, and protect astronauts from space radiation with its highly radiation-absorbent hydrogen fuel. The VASIMR is still in its developmental infancy and requires many years of research before its initial operational capability. Much is still unknown about the complex plasma interactions in the exhaust. A Langmuir probe was designed, constructed, and operated to determine current density radial profiles and total particle flux at various stages in the exhaust of VASIMR. The Langmuir probe results proved that the exhaustâ s radial profile is Gaussian, experimentally validated predictions of magnetic field line dragging, and verified the ionization efficiency of VASIMR.

Picosatellites are defined as satellites with a mass between 0.1 and 1kg (Miniaturized satellite). Picosatellites are typically designed to work together or function in formations (Miniaturized satellite). A specific type of Picosatellite known as CubeSats were introduced in 1999 and since then have increased in popularity so that there are now over 80 CubeSat programs around the world. CubeSats are defined as cubic units 10cm on each side and no more than 1kg in mass. CubeSats are required to conform to the CubeSat Standard created by California Polytechnic State University and Stanford University and be compatible with Cal Poly’s P-POD deployment system (Toorian, 2005). Some CubeSat uses include earth imaging, communications projects and various scientific experiments. CubeSats currently require attitude control and in the future, may require, maintaining a specific orbit, or changing orbit. With this ability many new activities may be possible for CubeSats. These activities could include rendezvous, vehicle inspection, formation flying and de-orbiting. For these activities to be possible, a high performance propulsion system is required. The goal of this thesis is to design and test an affordable, safe, and effective micro-propulsion system for CubeSats.

Hybrid rockets motors provide distinct safety advantages when compared to traditional liquid or solid propellant systems, due to the inherent stability and relative inertness of the propellants prior to established combustion. As a result of this inherent propellant stability, hybrid motors have historically proven dicult to ignite. State of the art hybrid igniter designs continue to require solid or liquid reactants distinct from the main propellants. These ignition methods however, reintroduce to the hybrid propulsion system the safety and complexity disadvantages associated with traditional liquid or solid propellants. The results of this study demonstrate the feasibility of a novel direct electrostatic arc ignition method for hybrid motors. A series of small prototype stand-alone thrusters demonstrating this technology were successfully designed and tested using Acrylonitrile Butadiene Styrene (ABS) plastic and Gaseous Oxygen (GOX) as propellants. Measurements of input voltage and current demonstrated that arc-ignition will occur using as little as 10 watts peak power and less than 5 joules total energy. The motor developed for the stand-alone small thruster was adapted as a gas generator to ignite a medium-scale hybrid rocket motor using nitrous oxide /and HTPB as propellants. Multiple consecutive ignitions were performed. A large data set as well as a collection of development 'lessons learned' were compiled to guide future development and research. Since the completion of this original groundwork research, the concept has been developed into a reliable, operational igniter system for a 75mm hybrid motor using both gaseous oxygen and liquid nitrous oxide as oxidizers. A development map of the direct spark ignition concept is presented showing the ow of key lessons learned between this original work and later follow on development.

The Auroral Spatial Structures Probe (ASSP) mission is a sounding rocket mission studying solar energy input to space weather. ASSP requires the high velocity ejection (up to 50 m/s) of 6 secondary payloads, spin stabilized perpendicular to the ejection velocity. The proposed scientific instrumentation depends on a high degree of spin stability, requiring a maximum coning angle of less than 5º. It also requires that the spin axis be aligned within 25º of the local magnetic field lines. The maximum velocities of current ejection methods are typically less than 10m/s, and often produce coning angles in excess of 20º. Because of this they do not meet the ASSP mission requirements. To meet these requirements a new ejection method is being developed by NASA Wallops Flight Facility. Success of the technique in meeting coning angle and B-field alignment requirements is evaluated herein by modeling secondary payload dynamic behavior using a 6-DOF dynamic simulation employing state space integration written in MATLAB. Simulation results showed that secondary payload mass balancing is the most important factor in meeting stability requirements. Secondary mass payload properties will be measured using an inverted torsion pendulum. If moment of inertia measurement errors can be reduced to 0.5%, it is possible to achieve mean coning and B-field alignment angles of 2.16º and 2.71º, respectively.

In an attempt to generalize previous models of the bidirectional vortex mean flow, a new solution is presented that can cope with arbitrary injections and outlet conditions. In the process, the steady, inviscid and axisymmetric equations of motions are reduced to one partial differential equation for the stream function, known as the Bragg-Hawthorne equation, which is solved exactly. The solution is shown to be highly dependent on the imposed boundary conditions: the mean flow changes according to the manner by which the fluid is injected or extracted from the vortex chamber. From the stream function, the velocity is obtained along with the vorticity and pressure distributions which are carefully derived and analyzed. The results are then compared to several inviscid models found in the literature. After determining an exact inviscid solution to the problem, viscous effects at the core are added to overcome the known singularity that arises at the centerline. The governing equations are hence revisited while keeping the viscous diffusion term in the tangential momentum equation. The core region, where viscous effects lead to the onset of a forced vortex, is rescaled using appropriate transformations. An asymptotic approximation is then applied to linearize and solve the resulting ODE for the tangential vi velocity. The inner viscous solution is then matched to the outer inviscid result using Prandtl’s Matching Principle. Finally, the viscous correction is passed onto the vorticity and pressure formulations.

Space vehicle is a kind of tool and weapon for transporting materials and personnel, conveying energy and execution work in space. The war of 21st century is a Hi-Tech war to fight for the power over space control and information transmission. Space power means the control of information transmission, and therefore the victory of this war and quality of space vehicle is the key factor in determining Whether one has the power space in. China was developed space vehicle technology since 1956 After great efforts for half a century, it has already established certain foundation. China like U.S.A and Russia, has invested huge funds and manpower in developing space science and technology. China hopes to improve military strength, drive the development of relevant industries and economy, and increase national competitiveness and overall strength through the development in space vehicle science and technology, and further conquer this expects to new high land of space. This paper aims at studying the development, capabilities, restrictions, and breakthroughs the space vehicle technology of China in terms of China¡¦s position on the control over space and its strategic intention in developing space vehicles. This paper reveals the direction of space vehicle development in China and suggests measures for maintaining Taiwan security. It will help to provide a relevant background for the future study of space vehicle development in China.

Thesis (Aeronautical and Astronautical Engineer)--Naval Postgraduate School, June 2003. / Thesis advisor(s): I. Michael Ross, Steve Matousek. Includes bibliographical references (p. 149-150). Also available online.

An experiment was performed to measure the impulse of the A8-3 Estes model rocket engine. The measurements were obtained by mounting an engine vertically onto a frame and measuring the changing weight of the assembly during the firing period for six trials. Thrust was calculated by the difference in measured weight from the rest weight while accounting for the mass loss rate. The impulse of each of the six thrust curves was averaged to find the final impulse measurement: 2.36 (+/-) .16 N-s. Estimates of the average thrust and exhaust velocity were also calculated from the impulse measurement.

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