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Cold-flow tests were conducted in mediums of air and water to investigate the internal flow field about the nozzle region of a proposed solid rocket motor (SRM) configuration that would potentially replace the current external boosters on NASA's Space Shuttle. One-eighth-scale clear acrylic models of the proposed submerged aft-dome and aft finned grain elements were constructed to simulate the aft segment of the SRM at ignition and 35 seconds into the firing sequence. Pressure, velocity, and turbulence profiles were obtained during cold air testing, while air bubbles and dye were used for flow visualization during water tunnel testing. The flow visualization experiments indicated the presence of strong inlet vortices, alternating vortex shedding from both grain models' fins, circumferential flow in the aft-dome and around the nozzle, and recirculatory flow in the aft-dome and near an upstream portion of the 35-second grain model. Data acquired during cold air testing showed a turbulent low-velocity flow field in the aft-dome for both grain models. With respect to pressure and mean velocity virtually the entire nozzle/aft-dome region exhibited a minimal sensitivity to nozzle vectoring. / Master of Science

The ADP rocket pump inducer was studied computationally using a 3-D Navier-Stokes solver, The Moore Elliptic Flow Program. Design and off-design flow rates were simulated to qualitatively and quantitatively study the effects of flow rate on the flow and performance. Several views of the results were created to aid flow visualization. The 3-D laser measurements made by Rocketdyne were used for comparison. The velocity magnitudes as well as the flow patterns within the inducer match well between the calculated and measured results. The axial velocity distribution and the rotary stagnation pressure, losses, are predicted very well by the calculation. The internal flow patterns developed in the simulation as expected, with radial outflow in the blade boundary layers. The tip leakage flow formed a recirculation region, a toroidal shaped vortex at the tip leading edge of the blades. The associated backflow forms a blockage that varies with flow rate. The thermodynamic performance was evaluated by calculating the contributions to pressure rise, the pump characteristic, the contributions to moment of momentum, and the efficiency. The centrifugal effect and relative velocity effect were found to vary with flow rate. The effective inlet throat radius, which governs these two effects, changes with flow rate because of the recirculation blockage. The shear on the blades was found to produce a small fraction of the work in the inducer, and most was produced by the pressure difference across the blade. The inducer efficiency was about 89%, and increased with decreasing flow rate in the range of flow rates considered, from 89% to 110% of the design flow rate. / Master of Science

A mixed variational statement and corresponding finite element model are developed for an axisymmetric solid body under external symmetric loads using the updated Lagrangian formulation. The mixed finite element formulation treats the nodal displacements and stresses as the variables that can be approximated independently. The method of static condensation is used to keep some stresses across interfaces of a solid of revolution discontinuous. The stiffness matrix is transformed from semi-positive definite to positive definite. A rocket motor is composed of (1) case (2) propellant and (3) hollow air core and is modelled as an axisymmetric solid. The propellant of a rocket motor is treated as a viscoelastic material. Static and dynamic analyses are performed under (1) two opposite line loads (2) two opposite patch loads and (3) one line and one patch load combination. The modified Newton-Raphson method is used in the solutions of nonlinear algebraic equations. The analysis of free vibration is executed first and then the Newmark direct integration method is used in a transient analysis. Results of these analyses are compared with solutions obtained from different methods that are independent of the finite element method. / Ph. D.

Thesis (MScEng)--University of Stellenbosch, 2003. / ENGLISH ABSTRACT: The focus of this thesis is on the investigation of the functioning of a liquefied-gas thruster. Such a thruster could be used to provide secondary propulsion to a microsatellite in orbit. A general overview of the need for thrusters in micro-satellites is put forward in the introduction. Motivation for deciding to investigate a liquefied-gas system is presented. Recent developments in the field of micro-satellites are discussed as well as their relevance to the project undertaken. Fundamental background theory relevant to the engineering problems associated with the development and analysis of such a system is also presented. Computer programs were written to simulate such a liquefied-gas thruster system. The experimental work carried out to analyse the system from a practical view-point is documented. Attention is also given to the measurement and calibration techniques used to obtain experimental data. One-dimensional fully explicit transient mathematical models of the thruster system were developed to model the system using both compressed air and butane as propellants. These models were incorporated into computer programs used to simulate the transient behaviour of the system. Although it is intended to use butane as the propellant onboard a satellite, the reason for modelling and simulating a system using compressed air is because air is a convenient fluid to work with from both a theoretical and practical point of view. An experimental model of a thruster system was designed, built and tested using air and butane as propellants. Most of the model was built using perspex to allow for the observation of the two-phase behaviour of the propellant inside the system. Locally purchased components were used for the solenoid and fill valves. Readily available butane lighter fluid was used for butane testing. Self-made heating elements were used to provide heat input to the propellant. Testing was done at different back pressures ranging from 100 kPa down to 20 kPa in a vacuum chamber. Good comparison between theoretical and experimental results was obtained for air. Theoretical results for peak thrusts tended to over predict experimental results by approximately 15 % for a system exhausting to a pressure of 100 kPa. Peak thrusts as high as 0.2 N were obtained for vacuum tests conducted at an absolute pressure of 20 kPa. Peak thrusts of approximately 50 mN were achieved for experimental testing III atmospheric conditions using butane with a starting pressure of between 270 and 290 kPa. Typical average thrusts of between 20 mN and 30 mN were noted for butane testing with initial pressure of between 200 to 300 kPa. Peak thrusts of over 0.1 N were observed for vacuum testing at an absolute pressure of 20 kPa. An equation to correlate the experimentally determined average thrust as a function of pulse duration and starting pressure was developed. This correlated most of the experimental data to within ±25 %. Theoretical results for butane testing are able to predict peak thrusts within approximately 20 % for starting pressures in the range of 200 to 300 kPa. Since the project was an exploratory investigation into a liquefied-gas thruster, some additional aspects relating to such systems were also given attention. The effect of liquid propellant motion or sloshing was considered and recommendations regarding the design and placement of the propellant tanks were made. The use of heat pipes as an alternative to electrical heating elements was investigated and some elementary design aspects are presented graphically. The management of the liquid propellant using surface tension devices was examined qualitatively. Recommendations relating to future projects in the field of simple, low-cost propulsion systems for micro-satellites are put forward. More specifically these recommendations are with regard to: thermo-fluid modelling of the propellant, future experimental work to be done, techniques to measure small thrusts and vacuum chamber testing. / AFRIKAANSE OPSOMMING: Die tesis ondersoek die funksionering van 'n vervloeidegas stuwer. So 'n stuwer kan gebruik word om sekondêre aandrywing aan 'n mikro-satelliet in 'n wentelbaan te verskaf. 'n Algemene oorsig oor die behoeftes van stuwers vir mikro-satelliete word voortgesit in die inleiding. Redes vir die gebruik van 'n vervloeidegas stuwer word bespreek. Onlangse ontwikkelinge in die veld van mikro-satelliet aandrywing word bespreek asook die toepaslikheid daarvan. Fundamentele teoretiese agtergrond verbonde aan die ontwikkeling en analise van so 'n stuwer stelsel word ook gegee. Rekenaarprogramme is geskryf om die gedrag van so 'n stuwer stelsel te simuleer. Eksperimentele werk is gedoen om die stelsel vanuit 'n praktiese oogpunt te analiseer. Aandag word ook gegee aan die metings- en kalibrasietegnieke soos toegepas vir die eksperimentele werk. Eendimensionele volle eksplisiete wiskundige modelle is ontwikkelom die oorgangsgedrag van die stuwer-stelsel te simuleer met beide lug en butaan as dryfmiddel. Hierdie modelle is geïnkorporeer in die rekenaar programme om die stuwer stelsel te simuleer. Alhoewel dit beoog word om butaan as die dryfmiddel aan boord die satelliet te gebruik, is lug ook gebruik vir simulasie weens sy gerieflikheid as 'n vloeier uit beide 'n teoretiese en 'n praktiese oogpunt. 'n Eksperimentele model van die stuwer stelsel is ontwerp, gebou en getoets met beide lug en butaan as dryfmiddels. Die model is hoofsaaklik uit perspex gebou sodat die twee-fase gedrag van die butaan uitgebeeld kon word. Vrylik beskikbare butaan aansteker vloeistof IS gebruik VIr butaan toetsing. Selfvervaardigde verhittingselemente is gebruik om hitte aan die dryfmiddel te verskaf. Toetse is gedoen deur verskeie omgewingsdrukke varieërend van 100 kPa af tot 20 kPa in 'n vakuumtenk te gebruik. Goeie ooreenstemming tussen die teoretiese en eksperimentele resultate vir die toetsing van lug is verkry. Die teoretiese resultate neig om die piek stukrag 15 % hoër te voorspel as die eksperimentele resultate vir 'n stelsel wat tot 'n omgewingsdruk van 100 kPa by die uitlaat. Piek stukragte van meer as 0.2 N is gekry vir vakuum toetse wat gedoen is by 'n omgewingsdruk van 20 kPa. Tydens eksperimentele toetsing met butaan teen 'n aanvanklike druk tussen 270 en 290 kPa, in atmosferiese toestande, is piek stukragte van ongeveer 50 mN behaal. Tipiese gemiddelde stukragte van tussen 20 en 30 mN is waargeneem vir butaan toetsing teen 'n aanvanklike druk tussen 200 en 300 kPa. Piek stukragte van meer as 0.1 N is behaal vir vakuum toetse met 'n absolute druk van 20 kPa. 'n Vergelyking om die gemiddelde stukrag, wat eksperimenteel bepaal is, as 'n funksie van puls tydsduur en aanvanklike druk te korreleer, is ontwikkel. Die meeste eksperimentele data se afwyking van die korrelasie-vergelyking was minder as 25 %. Teoretiese resultate vir butaantoetse het piek stukragte binne 20 % van die eksperimenteel metings korrek voorspel vir aanvanklike drukke tussen 200 tot 300 kPa. Weens die feit dat die projek 'n oorhoofse ondersoek in In vervloeidegas stuwer behels, is aandag ook gegee aan addisionele aspekte wat verband hou met sulke stelsels. Die effek van die vloeistof-dryfmiddel se onstabiele beweging in sy tenke is in ag geneem en voorstelle vir die ontwerp en plasing van die dryfmiddel tenke is gemaak. Die gebruik van hitte pype as 'n alternatief vir elektriese verhittingselemente is ondersoek. Verskeie ontwerp aspekte word grafies voorgestel. Die bestuur van die vloeistof-dryfmiddel deur van oppervlak spannings apparaat gebruik te maak, is kwalitatief ondersoek. Voorstelle vir verdere navorsing in die veld van eenvoudige, lae-koste stuwer stelsels vir mikro-satelliete is gemaak. Meer spesifiek is hierdie voorstelle gerig op die termo-vloeidinamiese modellering van die dryfmiddel, verdere eksperimentele navorsing, tegnieke om klein stukragte te meet en vakuumtenk toetse.

Using the Debye-Huckle approximation, the effects of Coulomb interactions on the equilibrium, frozen, and nonequilibrium flow of an ionized gas have been investigated. The gas is assumed to be monatomic, electrically neutral, and thermal equilibrium (i.e., a one temperature fluid); but the composition of the gas is arbitrary, that is, multiple ionization of any degree is allowed. The thermodynamic variables are derived starting from the appropriate expression for the Helmholtz free energy. Using Boltzmann statistics and assuming that the velocity distribution functions are given by their Maxwellian values, the rate of ionization is derived for atom-atom, atom-ion, and atom-electron collisions. The resulting expressions are then employed in solving the quasi-one-dimensional flow in a converging-diverging nozzle for the equilibrium, frozen, and nonequilibrium cases. Numerical examples, using argon as the working substance, are discussed and the results presented graphically. The results of these calculations indicate that, for single ionization, the effect of Coulomb interactions on the performance of rocket engines is negligible; but that data obtained from hypersonic arc jet wind-tunnels can be significantly influenced by the presence of the interactions. / Ph. D.

The analysis and redesign of a centrifugal impeller for a rocket pump is presented in this thesis. A baseline impeller was designed by Rocketdyne for the NASA Marshall Pump Consortium. Initially, the objective was to reduce the circumferential exit flow distortion of the baseline impeller. Later in the study, the objective became raising the head coefficient of the impeller. The study presented in this thesis was also undertaken to demonstrate current CFD capabilities for impeller design. A literature review includes an overview of centrifugal impeller geometries and configurations. Centrifugal impeller performance and secondary flows are discussed, and a summary of studies on the effects of impeller exit and diffuser inlet velocity distortion on diffuser performance is also presented. The flow calculation details and the results of the baseline impeller flow calculations are described. Fourteen redesigned impeller geometries were analyzed using the Moore Elliptic Flow Program, and the results were compared to the baseline geometry in terms of head rise, losses, and exit flow distortions. A final geometry was chosen; this geometry will be built and tested by Rocketdyne. The results show that backward blade lean can be effective in red using the exit flow distortion of the impeller. Tip slots or holes were not beneficial because of the large inlet boundary layer. Also, it appears possible to raise the head coefficient of the baseline impeller without creating excessive flow distortion. The planned testing is necessary to verify the predictions of the flow code. / Master of Science