Graduate dissertations
2021
Supervisor: Dr Jean-Francois Pitot de la Beaujardiere
Co-Supervisor: Prof Michael J Brooks
The African small satellite industry (micro and nano satellites in particular) continues to grow with developments in the miniaturization of satellite technology. However, the costs and delays involved with the traditional “piggy backing” satellite launch method is unsustainable for small sat developers and has thus created a niche market for dedicated small satellite launch services. Notably, there is no satellite launch capability whatsoever in Africa, meaning all of the continent’s launch requirements are serviced by foreign providers, incurring additional cost. As its primary objective, the University of KwaZulu-Natal’s (UKZN’s) Aerospace Systems Research Group (ASReG) seeks to enable the establishment of an indigenous small satellite launch capability in alignment with the South African Government’s goals. To this end, ASReG is currently developing the LOX/Kerosene SAFFIRE (South African First Integrated Rocket Engine) to propel a hypothetical two-stage orbital launch vehicle, termed Commercial Launch Vehicle 1 (CLV). The upper stage of the launch vehicle will use a vacuum-expanded variant of SAFFIRE called SAFFIRE-V. The upper stage for both cases must meet the design constraints of a 0.85 mass fraction and a 1.2 m outer diameter. CLV has been envisaged to deliver a 75kg payload to 400 km sun synchronous orbit. This thesis presents a high level analysis focusing on the upper stage of CLV, which intends to guide design decisions by comparing design options based on mass, and develop a methodology for upper stage vehicle design. One of the major design decisions is the type of propellant feed system the vehicle should use; in this regard, the analysis compares an electric pump feed system to a pressure fed system. Another is the selection of propellant tank material, given that the propellant tanks constitute most of the mass of a rocket. Stainless steel (301 and Duplex), aluminium alloy (7075), aluminium-lithium (2195), carbon fibre reinforced plastic (T700/Epoxy), as well as combinations of materials were compared. To perform the preliminary mass analysis, each of the major components/systems of the CLV upper stage were independently designed and the various design options available for each of the components/systems were compared based on mass. These systems and components include: fuel and oxidiser propellant tanks, the propellant pressurization system and the reaction control system. After the individual analyses of the variations of each component, the best suited architectures were modelled in SolidWorksÒ CAD software. The components were then assembled, in CAD. The analysis found that, on a preliminary basis, the Lithium ion (Li-Ion) based electric pump fed upper stages did not meet the mass requirements while Lithium polymer (Li-Po) based upper stages achieved the mass requirements. An upper stage employing stainless steel propellant tanks was found to meet the mass requirements, but only for a pressure fed upper stage. Overall, pressure fed upper stages had lower masses compared to electric pump vehicles. The mass reduction of thin walled, low pressurized, propellant tanks (resulting from using electric pumps) was offset by the mass of the battery packs required to power the pumps.
2020
Supervisor: Dr Jean-Francois Pitot de la Beaujardiere
Co-Supervisor: Dr Michael Brooks
The University of KwaZulu-Natal’s Aerospace Systems Research Group (ASReG) was formed in 2009 and has been developing hybrid sounding rockets since 2010 under the umbrella of the Hybrid Sounding Rocket Programme (HSRP). The HSRP was started with the goal of developing a sounding rocket platform capable of meeting the need of the South African and African scientific communities to access atmospheric data without using expensive foreign sounding rocket launch services. To reach this goal the HSRP has begun developing a series of hybrid propelled rockets as technology demonstrators, each with incrementally improved apogee and technology integration. [1]
The primary series of vehicles under the HSRP are the Phoenix series hybrid rockets. The Phoenix series rockets are technology demonstrators and form part of the HSRP’s roadmap to developing a sounding rocket which can reach an apogee of 100 km. The next rocket in the Phoenix Series is the Phoenix-1B Mk. II. This rocket aims achieve an apogee of 35 km and to achieve this apogee, composite materials will be integrated into the airframe structure, replacing the previously utilised aluminium construction. The primary mass reduction on the Phoenix-1B Mk. II came from the integration of a filament-wound composite pressure vessel for use as the nitrous oxide storage (oxidiser) tank, which was designed in this thesis.
This thesis details the creation of a methodology capable of generating and analysing filament wound pressure vessels suitable for use as oxidiser tanks on rockets. The methodology was based upon analytical methods but progressed to verify the design on a computational platform. The methodology was executed to generate a structure for use on the Phoenix 1B Mk. II rocket. The structure was constructed using a liner made from a combination of un-plasticised polyvinyl-chloride and aluminium. The liner was then filament wound to create the load bearing structure of the oxidiser tank and was bonded to supplementary coupling structures to allow for coupling with the fore and aft sections of the rocket. After the manufacture of the proposed oxidiser tank, it was qualified using a variety of different tests. The first were proof and destructive pressure tests which ensured that the tank would operate safely. The next phase of testing involved a cold flow and hot fire test of the rocket motor. The tank passed these tests and only remains to be flight tested.
The resulting oxidiser tank offered a conservatively estimated 15 % mass saving in comparison to the previously utilised aluminium oxidiser tank on the Phoenix-1B Mk. I hybrid sounding rocket.
2019
Supervisor: Ms Kirsty Veale and Dr Clinton Bemont
Slab motors are used to determine and investigate the regression rate characteristics of hybrid rocket propellant combinations. This information is fundamental to the overall design and thus used to determine the payload, altitude and thrust parameters of a rocket. The Phoenix Hybrid Sounding Rocket Programme in the University of KwaZulu-Natal’s (UKZN) Mechanical Engineering Department uses paraffin wax and nitrous oxide in their series of hybrid sounding rockets. The regression rate behaviour of paraffin wax with nitrous oxide has not previously been investigated in slab motors. This study focused on the regression rate behaviour and entrainment mechanism with regards to non-classical fuels including those with metal additives. This was used to gain a greater understanding of the increased regression rates associated with these fuels. The addition of metal additives, such as that of aluminium to fuel grains, was explored since the research suggested that it increases the regression rate of pure paraffin wax by 30%. A hybrid rocket slab motor visualisation test stand was developed to observe and obtain regression rate data. The stand includes a feed system, injector and a combustion chamber. All the components were manufactured using brass and stainless steel materials for their nitrous oxide compatibility, strength, and thermal resistance. Quartz glass windows were incorporated into the combustion chamber design for visualisation purposes. Due to the presence of quartz glass the use of finite element analyses became critical and more complex in order to ensure that the glass could withstand the operating conditions of the slab motor. A side-glass spacer was implemented to minimise the effects of side burning and to observe the influence of regression rate. Tests were conducted at a 130 g/s oxidiser mass flow rate and an atmospheric chamber pressure. A data acquisition system using LabVIEW software was implemented to obtain tank readings for the duration of the burn and to ensure safe motor operation. The regression rate of Sasolwax 0907 fuel was volumetrically determined and observed to be on average 3.74 mm/s. This shows a much higher regression rate than other paraffin wax compositions which have been found to regress at 1.5 mm/s. The characteristics of the entrainment process were validated for the investigated propellants, and the high regression rate mechanism of paraffin wax was observed in the liquid melt layer, droplet entrainment, and roll waves. Tests using aluminised wax fuel grains at atmospheric conditions proved to be unsuccessful with nitrous oxide as the oxidiser. A possible reason for this could be due to the aluminised fuel grains requiring increased heat transfer, therefore not producing sufficient vaporisation of the fuel. Moreover, decomposition of the oxidiser appeared to be inhibited by the combination of the oxidiser mass flow rate and the port area which prevented combustion.
2018
Supervisor: Ms Kirsty Veale
Co-Supervisor: Mr Jean-Francois Pitot de la Beaujardiere and Dr Clinton Bemont
The Aerospace Systems Research Group (ASReG) at the University of KwaZulu Natal is actively developing sounding rockets in the Phoenix Hybrid Sounding Rocket Programme, for use by the South African scientific community. These sub-orbital launch vehicles use nitrous oxide and paraffin wax as propellants. While paraffin wax offers large performance gains over typical polymeric fuels, due to its high regression rate, further performance gains can be achieved via the use of metal additives such as aluminium powder. The main advantage of using additives such as aluminium is the ability to create a smaller, more compact launch vehicle. This is due to a decrease in the optimal oxidiser-to-fuel ratio brought about by metallisation, which increases overall propellant density. Theoretically, an added advantage is the higher heat of combustion as a result of aluminium combustion. This added heat further increases the regression rate of the solid fuel grain. In order to realise these performance gains, various challenges need to be overcome. Some of these include delayed combustion due to the alumina layer that naturally coats the aluminium particles, slag formation and nozzle erosion. In this study, a laboratory scale hybrid rocket motor was developed to test aluminised paraffin wax fuel grains via a series of hot fire tests. A nitrous oxide feed system was developed, as well as a computer program and associated electronics to control the system remotely and capture data from an array of sensor equipment. Due to time constraints placed on the project, only pure paraffin wax and fuel grains comprising 40 % aluminium by mass were tested. Using specific impulse and characteristic velocity as performance metrics, preliminary data shows little to no gain in performance with aluminised fuel grains due to incomplete combustion of the aluminium. Substantial erosion of the copper nozzles that were used in the aluminium grain tests, due to localised melting, was also noted. Large amounts of aluminium and alumina slag was also found on the nozzles converging face. In order to seek maximum performance gains from aluminium as an additive, it was recommended that the particle size be reduced and stripped of its oxide layer before addition into the solid fuel grain. This will ensure more complete and rapid combustion of the particles before being ejected from the combustion chamber.
Supervisor: Dr Michael Brooks
Co-supervisors: Prof Graham Smith, Dr Glen Snedden (CSIR)
South Africa has a fledgling satellite industry but lacks the ability to launch spacecraft into low Earth orbit. As a result, the University of KwaZulu-Natal’s (UKZN’s) Aerospace Systems Research Group (ASReG) began the development of the South African First Integrated Rocket Engine (SAFFIRE). SAFFIRE aims to be a versatile, small scale, liquid rocket engine capable of being clustered for use on small-satellite (‘small-sat’) launch vehicles. The propellants for the proposed engine are Rocket Propellant-1 (RP-1) and liquid oxygen (LOX), which are fed into the combustion chamber via the injector. The uniqueness of SAFFIRE lies in the use of electrically driven pumps (‘electropumps’) as opposed to the conventional turbopump design. The electropump system has the fuel and oxidiser pumps independently housed and driven by brushless DC motors, which draw power from a lithium-polymer battery pack. A hypothetical launch vehicle was proposed to validate design specifications for the SAFFIRE engine, from which the hydrodynamic requirements of the electropump system were obtained. A meanline design algorithm was developed, using conventional design methods for centrifugal pumps. The algorithm was constructed to simultaneously meet the hydrodynamic system requirements of the engine, minimize the potential of cavitation at the fuel pump inlet and maximize the operational speed to minimize the overall pump weight. The hydrodynamic requirements of the system result in a low specific speed design, thus placing the pump in the region between full emission centrifugal pumps and positive displacement pumps. The low specific speed presented unique problems, not commonly encountered via the conventional pump design method, such as excessively small blade exit widths that are sensitive to dimensional variations. The Barske pump was investigated as a potential solution; it is a partial emission pump with the meanline design being governed by vortex theory. A comparative analysis between the conventional and Barske design was done using computational fluid dynamic techniques. The final hydrodynamic design is a hybrid between a Barske impeller and a scroll collection volute, which is typically found on a full emission pump. An investigation was done to determine an appropriate solution for mitigating the cavitation. It was found that the initial 3 bar tank pressure, suggested by literature, is applicable for an equivalent engine utilizing a turbopump system. The optimal tank pressure for the electropump system was found to be 9 bar. This increased available pressure head at the inlet of the pump eliminated any form of cavitation. The hybrid pump delivers 62.12 bar of pressure at a mass flow rate of 2.75 kg/s with a 62.12 % efficiency.
Supervisor: Dr Michael Brooks
Co-Supervisor: Mr Jean-Francois Pitot de la Beaujardiere
The Phoenix Hybrid Sounding Rocket Programme (HSRP) was started in 2010 with the primary aim of developing an indigenous sounding rocket service for South Africa. Two vehicles have been developed to date, with nominal design apogees of 10 km and 16 km respectively. This study describes the development of the hybrid propulsion system for the Phoenix-1B Mk II demonstrator rocket, with the primary mission of achieving a 35 km apogee. A particular emphasis of the project work was refining the motor from the baseline Phoenix-1B Mk I vehicle so as to double its apogee.
Initial design requirements specified to achieve this target apogee included the use of energetic metal fuel additives, improving motor performance and improving vehicle propellant mass fraction. The in-house developed Hybrid Rocket Performance Simulator (HYROPS) and NASA CEA™ were used extensively in an iterative manner to design the motor and vehicle. The motor utilises a propellant combination of nitrous oxide and paraffin wax with aluminium additive at 20% by mass. Magnesium additive was also considered for its relative ease of ignition and thus postulated higher combustion efficiency, but was abandoned due to lack of empirical regression data. The addition of aluminium into the fuel improves the density specific impulse and reduces the nominal oxidiser-to-fuel ratio from 6.8 to 5.4. This was found to effectively reduce the vehicle inert mass by 3.1% with an apogee increase of 1.7%. An associated decrease in combustion efficiency and increase in nozzle erosion and flame temperature were noted. A nominal thrust of 7250 N, or average of 5280 N, with a chamber pressure of 40 bar for a burn time of 14.2 s, or an average total impulse of 76.5 kNs, was found to propel a 76 kg vehicle to a 35 km apogee. Due to the motor calibre constraint, it proved challenging to arrive at an adequate motor design whilst remaining below the stable oxidiser mass flux limit set at 700 kg/m2 -s.
Analytical and numerical methods were employed to design each of the motor components, ensuring a safety factor of 1.5, focusing on the injector and nozzle designs. Analytical models and CFD analyses were used to predict the mass flow rate through the axial injector, which is nontrivial due to the two-phase flow nature of nitrous oxide. Cold flow testing showed that these modelling techniques under-predict the required injector flow area, necessitating a subsequent injector design revision, to increase the flow area by 44%. The composite ablatively-cooled nozzle was adapted from the Phoenix-1B Mk I nozzle and verified with a coupled thermal-structural analysis, with the Bartz equation used to obtain the temporal and spatial thermal loading. A hotfire test confirmed that the motor slightly under-performed with an average thrust and chamber pressure of 4920 N and 27.9 bar, respectively. A total impulse of 63.3 kNs was achieved, 17% below nominal. Combustion and specific impulse efficiencies of 79.4% and 77.6% were recorded. The motor was deemed qualified for flight and integrated with the Phoenix-1B Mk II for launch.
Advisors: Dr Mike Brooks
Co-advisors: Prof Graham Smith and Dr Glen Sneddon (CSIR)
The deployment of micro- and nanosatellites has greatly increased over the past few decades with advances in miniaturized electronics for communication, imaging and attitude control. The South African satellite industry is now also currently developing two microsatellites and nanosatellites for launch by foreign providers. The outsourcing of launch services to foreign providers is costly and can lead to unanticipated delays. In this context, the UKZN Aerospace Systems Research Group (ASReG), in conjunction with the Council for Scientific and Industrial Research (CSIR) has begun designing a modular and compact liquid propulsion engine (LOX/RP-1) named SAFFIRE (South AFrican First Integrated Rocket Engine). This dissertation details the design and analysis of the liquid oxygen pump that delivers the oxidiser to the SAFFIRE combustion chamber at high pressure, where the propellants are burnt and expelled, generating thrust. The pump is electrically powered as opposed to the conventional turbine-driven turbopump, to further simplify start-stop procedures and reduce the complexity of the engine. The pump’s operating conditions were determined by an engine performance analysis, with these results forming the initial conditions for the pump design process. The oxidiser pump is required to deliver a mass flow rate of 6.13 kg/s at a pressure of 62.8 bar. The pump was designed using conventional centrifugal pump design procedures, with special considerations taken due to the working temperature of liquid oxygen being -183°C. The final one-dimensional design for the impeller was developed using the commercial software PUMPAL™, which was provided by the CSIR. A 3D impeller geometry was developed by importing the one-dimensional design into AxCent™, where quasi-3D Multiple Stream Tube (MST) analysis and full 3D computational fluid dynamics (CFD) simulations were performed. The impeller design was refined multiple times until the parameters set by the engine performance analysis were met. The AxCent™ analyses determined that low-pressure zones occurred at the inlet of the pump impeller. Hence Star-CCM+™, which has a more robust computational solver and allows for a full transient, multiphase CFD to be performed, was employed to analyse any potential cavitation affects. The results from Star-CCM+™ and AxCent™ were compared and designs altered until a final design was realized that met the prescribed performance parameters. The final pump impeller has an outer diameter of 86 mm, delivering a mass flow rate of 6.13 kg/s at a pressure of 64.2 bar. The pump operates at an efficiency of 60.8% requiring a power input of 51.96 kW at a rotational speed of 26000 rpm.
Supervisor: Dr Michael Brooks
Co-Supervisor: Mr Jean-Francois Pitot de la Beaujardiere
Hybrid rocket motors produce thrust by reacting a solid fuel with a liquid oxidizer inside a combustion chamber. This approach has certain advantages over conventional solid propellant rockets including improved safety and the potential for thrust control, while also being less expensive than liquid propellant engines. Liquefying hybrid fuels, such as paraffin wax, regress at a faster rate than the conventional solid fuels like HTPB that are dominated by vaporization at the solid-gas interface. Non-classical theory is still in its infancy, however, and more work is required to validate performance models experimentally, especially where throttling of the oxidizer mass flowrate is incorporated. While hybrid motor throttlabilty remains a subject of considerable interest, there has been little investigation of throttling in motors that use high regression rate, liquefying fuels such as paraffin wax. This study proposes a closed-loop thrust control scheme for paraffin wax/nitrous oxide hybrid rocket motors using a low-cost ball valve as the controlling hardware element. There are a number of advantages to throttling hybrid rocket motors but the most important is to enforce a constant thrust curve throughout the burn. A test facility and laboratory scale hybrid rocket motor utilizing paraffin wax as fuel and nitrous oxide as oxidiser were used for experimental testing. Using a mathematical model of a laboratory-scale hybrid rocket motor, the controller constants for a PID controller were obtained and tested through experimental testing. Open-loop testing was first done in order to determine the control authority of the ball valve over the oxidiser mass flowrate, as well as characterize the oxidiser mass flowrate in relation to each valve angle value. Closed-loop testing was undertaken to verify and refine the controller constants obtained via the laboratory-scale model. The tests prompted a redesign of the injector and additions to the LabVIEW™ controller regime. Using results from the open-loop tests a feed-forward lookup table was developed to allow for the controller to move to a specified angle quickly and thereby remove nonlinearities present in flow control using ball valves. Three successful closed-loop tests were done where the controller causes the thrust of the motor to track a predetermined thrust or chamber pressure set point with a reasonable degree of accuracy. The set-point profile of the first test was a constant thrust throughout the burn while the second test had a ramp set-point profile. The final test used chamber pressure as the feedback variable and had a step-down set-point profile. This study demonstrates that thrust control can be exercised over a paraffin wax/nitrous oxide hybrid rocket motor, using a low-cost ball valve as the control element to modulate the oxidiser mass flowrate.
2017
Advisor: Dr Mike Brooks
Co-advisors: Mr Jean-Francois Pitot de la Beaujardiere and Ms Kirsty Veale
In August 2014, South Africa’s first university-based hybrid rocket, Phoenix-1A, was launched at the Overberg Test Range near Cape Agulhas. The vehicle suffered nozzle and parachute failures during flight which, together with a reduced oxidiser load, reduced the nominal design apogee of 10 km to 2.5 km. The aim of this research was to improve on the design and performance of the prototype demonstrator and thereby develop a workhorse hybrid sounding rocket, named Phoenix-1B, to serve as a reliable platform for future hybrid rocket research at the University of KwaZulu-Natal (UKZN). Analysis of Phoenix-1A shortcomings served as the starting point for the new design, which utilises a paraffin wax and nitrous oxide propellant combination. The focus of this research was the propulsion system, with specific attention being paid to the nozzle and injector designs. In addition, an aerodynamic study was applied to the 1 m long ¾ parabolic nose cone and four tapered swept fins. Final design of the aluminium oxidiser tank and combustion chamber bulkheads incorporated finite element analyses to ensure an operational safety factor greater than 1.5. The oxidiser tank and combustion chamber assemblies were pressure tested to 80 and 60 bars respectively. A key output of the present work is an analysis of the effect of aluminium loading in the paraffin wax fuel grain, which indicated a potential rocket mass reduction of 23 kg when transitioning from a pure paraffin grain to one containing 40% aluminium by mass. The analysis also indicated that combustion temperature rises with aluminium loading, increasing from 3300 K for pure paraffin to 3600 K for 40% aluminised fuel. Consequently, an iterative transient thermo-structural analysis was conducted on the nozzle, resulting in an optimised design able to sustain the higher operating temperatures as well as mitigate the risk of failure as seen with Phoenix-1A. The final manufactured composite nozzle has a throat diameter of 32 mm, an expansion ratio of 6.38, and a length of 156 mm. The nozzle has a steel casing which provides structural support to the silica phenolic insulation and graphite throat insert. A two phase CFD analysis, coupled with analytical mass flow rate models, was used to configure the axial injector and reduce the potential for combustion instabilities associated with the nitrous oxide flow. The Phoenix-1B motor has a design thrust of 5 kN to propel the fully loaded vehicle, with a mass of 70 kg, a length of 4.3 m and a diameter of 164 mm, to an altitude of 16 km.
2015
Advisor: Prof Graham Smith
Co-advisors: Dr Mike Brooks, Prof Jeff Bindon and Dr Glen Snedden (CSIR)
This dissertation presents the validation of a universal impeller test rig, designed and constructed at the University of KwaZulu-Natal (UKZN). The research was conducted as part of UKZN’s Aerospace Systems Research Group’s (ASReG) work into liquid rocket propulsion. The rig will be used to evaluate the performance of impellers for use in commercial launch vehicle fuel turbopumps. Head rise versus flow rate characteristics, as well as cavitation performance are to be assessed by the rig. The power requirements of the impellers necessitated the reduction in rotational speed and geometric size of the test case. Scaling laws and dimensionless numbers were used to predict performance of a test impeller. Validation of the rig and testing procedures was performed using a standard industrial KSB ETA 125 – 200 centrifugal pump, by comparing the experimental results with those of the supplier. Head rise characteristics were determined by measuring the change in pressure between the inlet and discharge of the pump and then plotted against the flow rate for varying system heads. Cavitation performance was assessed by decreasing the inlet pressure while maintaining a constant flow rate. This was performed at various flow rates within the range of operation. Head breakdown, vibration and noise levels, both in the time and frequency domains, were used to assess the cavitation performance. The head rise versus flow characteristics of the pump, determined on the rig, showed good agreement with the supplier’s data. Cavitation performance, determined by head breakdown, was also in accordance with the supplier. It was found that both the vibration and general noise levels increased, indicating the presence of cavitation, before any head breakdown was detected. By monitoring the level of the high frequency noise (> 10 kHz) the presence of cavitation was detected at a significantly higher inlet pressure than would be suggested by the head breakdown approach.
2014
Advisor: Dr Mike Brooks
Co-advisors: Mr Jean-Francois Pitot de la Beaujardiere and Prof Lance Roberts
In this dissertation, UKZN’s in-house Hybrid Rocket Performance Simulator (HYROPS) software is used in the design of Phoenix-2A, a proposed hybrid rocket for delivering a 5 kg instrumentation payload to an apogee altitude of 100 km. As a benchmarking exercise, HYROPS was first validated by modelling the performance of existing sub-orbital sounding rockets similar in apogee to Phoenix-2A. The software was found to approximate the performance of the published flight data within 10%. A generic methodology was then proposed for applying HYROPS to the design of hybrid propellant sounding rockets. An initial vehicle configuration was developed and formed the base design on which parametric trade studies were conducted. The performance sensitivity for varying propulsion and aerodynamic parameters was investigated. The selection of parameters was based on improving performance, minimising cost, safety and ease of manufacturability. The purpose of these simulations was to form a foundation for the development of the Phoenix-2A vehicle as well as other large-scale hybrid rockets. Design chamber pressure, oxidiser-to-fuel ratio, nozzle design altitude, and fin geometry were some of the parameters investigated. The change in the rocket’s propellant mass fraction was the parameter which was found to have the largest effect on performance. The fin and oxidiser tank geometries were designed to avoid fin flutter and buckling respectively. The oxidiser mass flux was kept below 650 kg/m2s and the pressure drop across the injector relative to the chamber pressure was maintained above 15% to mitigate the presence of combustion instability. The trade studies resulted in an improved design of the Phoenix-2A rocket. The propellant mass of the final vehicle was 30 kg less than the initial conceptual design and the overall mass was reduced by 25 kg.
Advisor: Dr Mike Brooks
Co-advisors: Dr Graham Smith, Prof Jeff Bindon and Dr Glen Snedden (CSIR)
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. 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. 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. 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.
2013
Advisor: Mr Jean Pitot
Co-advisors: Dr Mike Brooks and Prof Lance Roberts
This work describes the development of a hybrid rocket propulsion system for a reusable sounding rocket, as part of the first phase of the UKZN Phoenix Hybrid Sounding Rocket Programme. The dissertation details the development of the Hybrid Rocket Performance Code (HRPC) together with the design, manufacture and testing of Phoenix-1A’s propulsion system. The Phoenix-1A hybrid propulsion system utilises SASOL 0907 paraffin wax and nitrous oxide as the solid fuel and liquid oxidiser, respectively. The HRPC software tool is based upon a one-dimensional, unsteady flow mathematical model, and is capable of analysing the combustion of a number of propellant combinations to predict overall hybrid rocket motor performance. The code is based on a two-phase (liquid oxidiser and solid fuel) numerical solution and was programmed in MATLAB. HRPC links with the NASA-CEA equilibrium chemistry programme to determine the thermodynamic properties of the combustion products necessary for solving the governing ordinary differential equations, which are derived from first principle gas dynamics. The combustion modelling is coupled to a nitrous oxide tank pressurization and blowdown model obtained from literature to provide a realistic decay in motor performance with burn time. A targeted total impulse of 75 kNs for the Phoenix-1A motor was obtained through iterative implementation of the HRPC application. This yielded an optimised propulsion system configuration and motor thrust curve.
Advisor: Dr Mike Brooks
Co-advisors: Prof Lance Roberts and Mr Clinton Bemont
The Loop Heat Pipe (LHP) is a passive, two-phase heat transfer device used, most commonly, for thermal management of aerospace and aeronautical electronic equipment. This research had two aims. Firstly, to create and validate a robust mathematical model of the steady-state operation of an LHP for terrestrial high heat flux electronics. Secondly, to construct an experimental LHP, including a sintered porous wick, which could be used to validate the model and demonstrate the aforementioned heat exchange and gravity resistant characteristics. The porous wick was sintered with properties of 60% porosity, 6.77×10-13 m2 permeability and an average pore radius of 1 μm. Ammonia was the chosen working fluid and the LHP functioned as expected during horizontal testing, albeit at higher temperatures than anticipated. The heat load range extended to 60 W, 50 W and 110 W for horizontal, gravity-adverse and gravity-assisted operation respectively. Because of certain simplifying assumptions in the model, the experimental results deviated somewhat from predicted values at low heat loads. Model accuracy improved as the heat load increased. The experimental LHP behaved as expected for 5° and 10° gravity-assisted and gravity-adverse conditions, as well as for transport line variation, in which performance was assessed while the total tubing length was increased from 2.5 m to 4 m.
Advisor: Dr Mike Brooks
Co-advisors: Mr Jean-Francois Pitot de la Beaujardiere and Prof Lance Roberts
This dissertation describes the development of the Hybrid Rocket Performance Simulator (HYROPS) software tool and its application towards the structural design of the reusable, 10 km apogee capable Phoenix-1A hybrid sounding rocket, as part of the UKZN Phoenix programme. HYROPS is an integrated 6–Degree of Freedom (6-DOF) flight performance predictor for atmospheric and near-Earth spaceflight, geared towards single-staged and multi-staged hybrid sounding rockets. HYROPS is based on a generic kinematics and Newtonian dynamics core. Integrated with these are numerical methods for solving differential equations, Monte Carlo uncertainty modeling, genetic-algorithm driven design optimization, analytical vehicle structural modeling, a spherical, rotating geodetic model and a standard atmospheric model, forming a software framework for sounding rocket optimization and flight performance prediction. This framework was implemented within a graphical user interface, aiming for rapid input of model parameters, intuitive results visualization and efficient data handling. The HYROPS software was validated using flight data from various existing sounding rocket configurations and found satisfactory over a range of input conditions. An iterative process was employed in the aerostructural design of the 1 kg payload capable Phoenix-1A vehicle and CFD and FEA numerical techniques were used to verify its aerodynamic and thermo-structural performance.
Advisor: Dr Mike Brooks
Co-advisor: Prof Lance Roberts
In this study a semi-passive pulse thermal loop (PTL) was designed and experimentally validated. It provides improved heat transfer over passive systems such as the loop heat pipe in the moderate to high heat flux range and can be a sustainable alternative to active systems as it does not require an electric pump. This work details the components of the engineering prototype and characterizes their performance through the application of compressible and two-phase flow theory. A custom LabVIEW application was utilized for data acquisition and control. During operation with refrigerant R-134a the system was shown to be robust under a range of heat loads from 100 W to 800 W. Operation was achieved with driving pressure differentials ranging from 3 bar to 12 bar and pulse frequencies ranging from 0.42 Hz to 0.08 Hz. An evolution of the PTL is also proposed that incorporates a novel, ejector-based pump-free refrigeration system. The design of the pulse refrigeration system (PRS) features valves at the outlet of two PTL-like boilers that are alternately actuated to direct pulses of refrigerant through an ejector. The design of the ejector was carried out using a one-dimensional model implemented in MATLAB that accounts for compressibility effects with NIST REFPROP vapor data sub-routines. The model enables the analysis of ejector performance in response to a transient pressure wave at the primary inlet.