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Development of liquid propellant tanks for a suborbital launch vehicle.

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dc.contributor.advisorPitot de la Beaujardiere, Jean-Francois Philippe.
dc.contributor.advisorBrooks, Michael John.
dc.contributor.authorMchunu, Vulinhlanhla Salvation.
dc.date.accessioned2023-07-25T19:00:25Z
dc.date.available2023-07-25T19:00:25Z
dc.date.created2022
dc.date.issued2022
dc.descriptionMasters Degree. University of KwaZulu-Natal, Durban.en_US
dc.description.abstractHaving developed several hybrid sounding rockets, the Aerospace Systems Research Group (ASReG) at the University of KwaZulu-Natal is currently investigating the development of an indigenous launch vehicle for micro satellites. As part of this effort, a liquid propellant rocket engine called the South African First Integrated Rocket Engine (SAFFIRE) is in the advanced design phase. SAFFIRE combusts liquid oxygen and kerosene propellants to generate thrust. Following ground tests, the first version of the SAFFIRE engine will be incorporated into a single-stage launch vehicle to test its flight performance during the suborbital flight. This study aimed to develop a design procedure and subsequent engineering designs for propellant tanks suitable for use by this launch vehicle, known as the Suborbital Test Vehicle (STEVE). The liquid oxygen and kerosene propellant tanks were designed according to NASA propellant tank design guidelines based on the mechanical properties of half-hard 301L stainless steel – the alloy selected as the material of construction for both tanks. This material has various advantageous characteristics, including its compatibility with liquid oxygen, its high strength once work-hardened and its increased strength and ductility at cryogenic temperatures. As part of the study, comprehensive material testing was conducted to establish a suitable tank welding procedure and to evaluate the achievable weld efficiency. For the best performing welding procedure assessed, mean weld efficiencies with respect to yield strength and tensile strength were determined to be 70 % and 81 %, respectively. A trial propellant tank was fabricated using the selected welding procedure and subsequently subjected to a destructive hydrostatic pressure test. The outcomes of this test provided a clear indication of the modes and progression of tank failure and served to inform final design work. In terms of propellant tank layout, a tandem configuration with the liquid oxygen tank positioned above the kerosene tank was selected, in order to improve the stability characteristics of the vehicle and to minimise total feedline length. To reduce vehicle drag and mitigate aerodynamic heating effects, the liquid oxygen feedline was configured to pass coaxially through the kerosene tank via a tunnel tube. The incorporation of elliptical tank ends in both tank designs was dictated by pre-existing tooling made available by the tank end manufacturer. Based on these design characteristics, the anticipated tank loading conditions and the mechanical properties of the as-welded 301L stainless steel alloy, a minimum wall thickness requirement of 2 mm was determined for both tanks via finite element analysis.en_US
dc.identifier.urihttps://researchspace.ukzn.ac.za/handle/10413/21962
dc.language.isoenen_US
dc.subject.otherHybrid sounding rockets.en_US
dc.subject.otherHydrostatic pressure.en_US
dc.subject.otherCryogenic stainless steel.en_US
dc.subject.otherTIG Welding.en_US
dc.subject.otherRocket propulsion.en_US
dc.titleDevelopment of liquid propellant tanks for a suborbital launch vehicle.en_US
dc.typeThesisen_US

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Masters Degrees (Mechanical Engineering)