This work describes the development of an injector test rig to characterize the behaviour of rocket propellants in gelled form. The study included the design and manufacture of the test rig's propellant/simulant feed system, injector inserts, manifolds, frame, and propellant/simulant supply tank. Two converging injection orifices were used to form spray sheets at 45 degrees, 60 degrees and 90 degrees impinging angles. In addition, MATLABTM image processing tools were used to binarize spray sheet images and quantify the resulting fluid structures generated by the equipment. This work falls under a broader research program into gelled propellants conducted by the Aerospace Systems Research Institute (ASRI) at the University of KwaZulu Natal, Durban, South Africa.

A hybrid rocket launch campaign was run by the University of KwaZulu-Natal’s Aerospace Systems Research Institute (ASRI) at the Denel Overberg Test Range (OTR), South Africa in March 2023. Two vehicles were launched seawards over the southern Indian Ocean: the Phoenix-1C (P-1C) and Phoenix-1D (P-1D), each reaching apogees of 7.9 km and 13 km, respectively. The P-1D featured significant mass reductions over the Phoenix-1B Mk IIr vehicle launched in 2021 by replacing metal components with composite parts. It also had an improved video telemetry system. The P-1C served as a payload test bed and carried various instruments supplied by third parties. The intension was to test a recovery system, but unfavourable wind conditions and safety considerations prevented its deployment. The preliminary results showed that the P-1C exceeded predictions and the P-1D underperformed.

Africa is playing an increasing role in the global space economy through the establishment of national space agencies in numerous countries and the participation of many states in satellite development programs. Despite this surge in activity, all African satellite programs rely on foreign launch providers since the continent possesses no local launch capability. South Africa is a prime candidate to become an African leader in space operations due to its well-developed infrastructure and unique geographical location on the continent. The ABLE liquid rocket engine was developed as a technology demonstrator in the first step towards the development of a propulsion system for a sovereign launch capability. This paper describes the design, manufacture, and testing of the thrust chamber for the ABLE engine.

South Africa is one of numerous developing countries invested in the growth of aerospace technologies such as micro- and nanosatellite manufacture. Due to the lack of launch capability it is reliant on foreign launch providers. Within this context, the Aerospace Systems Research Institute (ASRI) based at the University of KwaZulu-Natal (UKZN) began conducting research and development into the South African First Rocket Engine (SAFFIRE). This compact liquid engine produces thrust through the combustion of Liquid Oxygen (LOx) and kerosene. The propellent pumps intended for SAFFIRE are electrically driven as opposed to conventional turbine driven pumps which greatly reduces the complexity and start-up procedure of the system. The SAFFIRE engine programme is the first step in the development of a Commercial Launch Vehicle (CLV) capable of delivering a payload into Low Earth Orbit (LEO). This paper presents an outline for the design methodology and validation of a one-dimensional, low specific speed pump design algorithm intended for use in the SAFFIRE propulsion program. The validation includes results of a dynamically scaled test article with design points of 14.3 bar at 25000 RPM and 1 kg/s.

For more than a decade, researchers at the University of KwaZulu-Natal (UKZN) Aerospace Systems Research Group (ASReG), now the Aerospace Systems Research Institute (ASRI), have pursued the development of liquid and hybrid propulsion systems for suborbital and orbital rockets. This paper addresses progress in these efforts, which have as their driving goal the creation of a sovereign commercial satellite launch capability for South Africa, and the African continent. Recent achievements include setting a new African altitude record for hybrid rockets at 17.97 km by a Phoenix vehicle and the static testing of a liquid oxygen (LOX)/kerosene SAFFIRE ABLE liquid propellant prototype engine of 18 kN nominal thrust. Details of the design and performance of both systems are addressed, along with a review of past successes and failures, and an overview of current and future research directions in ASRI.

The Aerospace Systems Research Institute (ASRI) is developing electric turbopumps for the South AFrican FIrst Rocket Engine (SAFFIRE) program. This aims to provide an indigenous pathway to microsatellite deployment. Dynamically scaled, water analog versions of these oxidizer and fuel pumps are important subsystems of this project that have been tested to acquire preliminary experimental data. This research details the numerical validation and surrogate optimization of the kerosene water-analog, scaled pump's performance. Benchmark simulations of the impeller and volute were first conducted in FINE (TM)/Open. Impeller optimization followed the objectives of improving hydraulic efficiency and cavitation reduction using the commercial software Numeca (TM) and Minamo (TM). The impeller geometry was first parametrized with the blade leading edge meridional profile and coordinates chosen as the free variables. Simulations were conducted on impeller geometries that varied according to the Latinized Centroidal Voronoi Tessellation (LCVT) sampling method. Tuned Radial Basis Function Networks (TRBFNs) were used to replicate and extend this design space. Multi-Objective Genetic Algorithms (MOGAs) were implemented to create optimized designs from the TRBFN surrogate model. Pareto-assisted techniques identified the globally optimum impeller design from the improved samples that increased hydraulic efficiency and minimum static pressure by 0.51% and 400.45 Pa, respectively. Manual modifications of the cutwater were made to attain short and long tongue configurations to further improve the global pump performance.

This paper reports on a mobile liquid rocket engine test facility developed by the University of KwaZulu-Natal’s Aerospace Systems Research Institute (ASRI) to quantify the performance of an 18kN LOX/Kerosene liquid rocket engine (LRE). The ablatively-cooled engine was developed as a technology demonstrator under ASRI’s SAFFIRE liquid rocket engine development programme, which seeks to develop propulsion units for a South African small satellite launch vehicle. Performance and stability metrics of the engine were required to validate the design methodology of the injector and combustion chamber. Prior to the facility being commissioned, South Africa did not have the infrastructure required to test LREs and quantify their performance. This study documents the test facility’s propellant feed system design and the cold flow testing methodology used to characterise the system before hot fire testing. The facility used a regulated blowdown system with cavitating venturis to control the propellant mass flow rates. Cold flow tests were conducted to determine venturi discharge coefficients, the system losses at nominal mass flow rates, the thermal conditioning required for the cryogenic oxidiser, and the valve timings for the engine start-up sequence. The facility control system used National Instruments (NI) hardware and LabVIEW software to monitor the system pressures and temperatures. An analytical model was used to estimate the required tank pressures for the nominal mass flow rates during hot fire tests. The cold flow testing showed that the cavitating venturis effectively controlled the propellant mass flow rate throughout the test duration. The injector experienced cavitation-related problems during cold flow testing to atmospheric pressure, but the testing methodology nevertheless enabled a successful characterisation of the system, leading to multiple successful hot fire tests with on-target mass flow rates.

The South African First Rocket Engine (SAFFIRE) program by the University of KwaZulu-Natal's (UKZN) Aerospace Systems Research Institute (ASRI) aims to develop a flight-weight liquid rocket engine (LRE) to be used on an indigenous South African launch vehicle. The first step towards a flight-weight engine design was developing and testing the Ablative Blowdown Liquid Engine (ABLE), a prototype LOX/Jet A-1 engine, designed to operate at 12 bar and producing 17.6 kN of thrust. This paper provides an overview of the SAFFIRE ABLE ground tests conducted at South Africa’s Denel Overberg Test Range (OTR) and includes preliminary data from ignition, engine start-up and steady hot-fire tests.

The development of an indigenous launch capability is the next logical step for the African continent where space technologies play an increasing role in economic growth and development. This paper presents a preliminary design for a flight weight regulator designed to control the pressurization of the liquid oxygen (LOX) tank in a suborbital launch vehicle currently in development by the Aerospace Systems Research Institute (ASRI), South Africa. The vehicle is propelled by a pressure-fed engine. With a prescribed LOX mass flow rate of 7.05 kg/s from the main tank, the regulator must supply helium to the tank at a constant pressure of 48 bar. Through analytical design it was determined that a spring-loaded pressure reducing regulator with an orifice diameter of 5 mm would be capable of delivering helium at the prescribed pressure to the LOX tank for all pressure ratios of the helium tank blowdown configuration. Additionally, the regulator will supply pressurant in the mass flow range of 0.085 - 0.152 kg/s for all pressure ratios with a maximum predicted stroke of 1.24 mm.

Slosh occurs when fluid in partially filled enclosures is excited, causing the fluid to move irregularly. When designing launch vehicles and spacecraft, the sloshing phenomenon must be considered as the forces and moments exerted by the fluid can damage the vehicle's structure and even alter its trajectory. Analytical, numerical, and experimental methods can all be used to study slosh, which manifests as rotational, chaotic, simple planar, symmetric, and other types of motion. Importantly, the fluid's disturbance and the tank's geometry determine the type of motion, while the fluid's fill depth, properties, frequency of disturbance, and the tank's geometry impact the slosh's amplitude and force on the tank. This paper describes the development of multiple additively manufactured baffle designs to mitigate slosh in a model spherical and cylindrical propellant tank.

Currently, most space propulsion systems use hydrazine and its derivatives, a toxic, carcinogenic monopropellant. These systems entail logistic and environmental difficulties. Significant efforts are underway to develop "green" propellants which do not have these undesirable characteristics. Among those considered include ionic liquid compounds such as HAN and ADN. These blends offer similar or better Isp values to hydrazine but offer reduced costs in handling and lower environmental effects. Gel propellants offer a number of advantages over their liquid and solid propellant counterparts, including higher theoretical specific impulse, restart and throttling capabilities, and safer transport and handling properties. These advantages are offset by a number of drawbacks including inadequate atomization and large feed pressure requirements to force the gelled propellant through the injector at the required flow rates. This present work seeks to investigate optimal "green" propellant formulations based on their stability, homogeneity, sensitivity, along with an array of rheological properties in order to characterize the flow behavior of the gelled propellant at every stage in its life cycle.

Professor Essam E. KhalilProfessor Essam E. Khalil On Wednesday, September 7, 2022, we were deeply saddened to hear that Professor Essam E. Khalil passed away at the age of 74. Professor Khalil was born on July 12, 1948, in Scotland, United Kingdom, and through his extensive career, he became one of the most respected educators/researchers in Egypt in the field of mechanical engineering. He received his B.Sc. and M.S. degrees in July 1971 and December 1973, respectively, from the Department of Mechanical Power Engineering of Cairo University. He then joined the Imperial College of Science and Technology where he received his Ph.D. in February 1977 under the supervision of Professor James Hunter Whitelaw. His dissertation is titled “Flow, Combustion and Heat Transfer in Axisymmetric Furnaces.” After finishing his Ph.D., Professor Khalil obtained a postdoctoral fellowship with the United Kingdom Atomic Energy Research Establishment in Harwell, UK. After his fellowship, he returned to Egypt as an assistant professor in the Department of Mechanical Power Engineering at Cairo University. He was promoted to Associate Professor and Professor in 1982 and 1988, respectively. During his academic career, Professor Khalil supervised many master’s and Ph.D. students in the areas of heat transfer, turbulent combustion, air distribution in buildings, and sustainability. Several of his past Ph.D. students became professors at different universities. Professor Khalil published over 950 papers in journals, conference proceedings, book chapters, and books. Some of his publications are listed in Refs [1–5]. His 1983 paper with Professors D. Brian Spalding and James H. Whitelaw titled “The Calculation of Local Flow Properties in Two-Dimensional Furnaces” was his most cited paper [6].During his academic career, Professor Khalil served as an editorial board member for Advances in Mechanical Engineering and the International Journal of Thermal and Environmental Engineering. He was a Fellow of the American Society of Mechanical Engineers (ASME); the American Institute of Aeronautics and Astronautics (AIAA); and the American Society of Heating, Refrigerating, and Air-Conditioning Engineers (ASHRAE). He was the Regional Director for ASHRAE in the Middle East and North Africa and served on the ASHRAE Board of Directors as Director-at-Large from 2016 to 2019. He served as a Convener for ISO/TC205 WG2 on Energy Efficiency and as a Convenor for ISO/TC163 WG4 on Thermal Performance and Energy Use in the Built Environment. He was an ASHRAE Distinguished Lecturer and an Authorized Instructor for ASME. He also served as an ABET program evaluator (PEV). He was a member of the AIAA Green Engineering Program Committee as well as the AIAA Thermophysics Technical Committee, the AIAA Terrestrial Energy Systems Technical Committee, and the AIAA Gas Turbine Engines Technical Committee.The scientific contributions of Professor Khalil spanned many areas and have been recognized by many honors and awards. These include the Decor of Science and Arts of the First Order from former Egyptian President Anwar Sadat (1981), the National Award for Scientific Achievement in Engineering Sciences (1981), Member of L’Institut D’Egypte (2007), ASHRAE Presidential Award of Excellence (2009), ASHRAE Chapter Service Award (2009), ASME Westinghouse Gold Medal (2009), AIAA Energy Systems Award (2010), ASHRAE Regional Award of Merit (2010), AIAA Sustained Service Award (2011), ASHRAE Distinguished Service Award (2011), ASME James Harry Potter Gold Medal (2012), University of Wisconsin Milwaukee Distinguished Lecture Award (2013), ASME Egypt Achievement Award (2013), Syndicate of Egyptian Engineers Award of Excellence (2016), ASHRAE Exceptional Service Award (2017), and multiple Awards of Excellence from Cairo University.It is difficult to find appropriate words to express our emotions at this wistful moment. The scientific community will remember Professor Khalil as a prolific and exceptional scientist, mentor, teacher, and a highly respected colleague. He will also be remembered for his exceptional character as a human being who cared about his family, friends, students, and professional colleagues. Figures 1 through 5 show some memorable pictures of Professor Khalil with different colleagues, students, and friends. He will be sorely missed by all of us.

The thrust produced by a hybrid rocket motor (HRM) can be controlled by varying the oxidizer flow rate to the combustion chamber. This feature is useful in shaping motor thrust profiles and optimizing a vehicle flight trajectory, but propellant throttling in solid-fuel hybrids is limited to the liquid component only, complicating the control scheme and potentially destabilizing combustion in the motor. While hybrid motor throttling ability 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 article describes the development and implementation of a closed loop thrust control scheme for a laboratory-scale paraffin wax/nitrous oxide HRM using a low-cost ball valve as the controlling hardware element. A model of motor performance is first developed from which proportional-integral-derivative (PID) controller constants are obtained through experimental testing. The control scheme is demonstrated through closed loop hot fire tests of a laboratory-scale motor in which thrust tracks a set-point value with feedback provided through a load cell. Upon reaching the setpoint, the motor remains throttled within ± 2.4% of the maximum thrust of the motor. Constant and ramping thrust profiles are demonstrated.

In order to assess the capabilities of South Africa as a launch site for commercial satellites, an optimal control solver was developed. The developed solver makes use of direct Hermite-Simpson collocation methods, and can be applied to a general optimal control problem. Analytical first derivative information was obtained for direct Hermite-Simpson collocation methods. Typically, a numerical estimate of the derivative information is used. This paper will present the solver algorithm, and the formulation and derivation of the analytical first derivative information for this approach. A sample problem is provided as validation of the solver.

Metal additive manufacturing involves manufacturing techniques that add material to produce metallic components, typically layer by layer. The substantial growth in this technology is partly driven by its opportunity for commercial and performance benefits in the aerospace industry. The fundamental opportunities for metal additive manufacturing in aerospace applications include: significant cost and lead-time reductions, novel materials and unique design solutions, mass reduction of components through highly efficient and lightweight designs, and consolidation of multiple components for performance enhancement or risk management, e.g. through internal cooling features in thermally loaded components or by eliminating traditional joining processes. These opportunities are being commercially applied in a range of high-profile aerospace applications including liquid-fuel rocket engines, propellant tanks, satellite components, heat exchangers, turbomachinery, valves, and sustainment of legacy systems. This paper provides a comprehensive review of metal additive manufacturing in the aerospace industry (from industrial/popular as well as technical literature). This provides a current state of the art, while also summarizing the primary application scenarios and the associated commercial and technical benefits of additive manufacturing in these applications. Based on these observations, challenges and potential opportunities are highlighted for metal additive manufacturing for each application scenario.

A hybrid rocket design methodology is dependent on a number of factors, such as mission requirements and resource availability and thus no single solution exists. An in-house software called the Hybrid Rocket Performance Simulator (HYROPS) was used as the foundation of the design process followed in this paper. The software is based on a one dimensional, unsteady combustion flow model that predicts the performance of a hybrid rocket motor as well as a generic 6-DOF flight dynamics simulator that determines the flight performance of a rocket. Parametric studies using the HYROPS software were conducted to investigate the effects of various design parameters on the performance of a sub-orbital hybrid rocket. The propellant combination used in this trade study was paraffin wax fuel with nitrous oxide. The various design parameters under analysis include chamber pressures, O/F ratio, and grain characteristics. Performance gain and mass saving obtained from the performance simulations reduce costs associated with the development of a hybrid rocket.

The performance of a South African parabolic trough solar collector (PTSC) module has been characterised using the ASHRAE 93-1986 standard. The collector is designed for component testing and development in a solar energy research programme. Low-temperature testing was performed at Mangosuthu Technikon’s STARlab facility using water as the working fluid. Both an evacuated glassshielded receiver and an unshielded receiver were tested, with which peak thermal efficiencies of 53.8% and 55.2% were obtained respectively. The glass-shielded element offered superior performance at the maximum test temperature, desensitising the receiver to wind and reducing the overall heat loss coefficient by half. The collector time constants for both receivers indicate low thermal inertia and the measured acceptance angles exceed the tracking accuracy of the PTSC, ensuring the collector operates within 2% of its optimal efficiency at all times. Off-sun thermal loss results and the behaviour of the PTSC under increased angles of incidence are described. A description of the test system components is given.