Liquid rocket engine injectors are complex components that significantly influence engine performance, combustion stability and manufacturing cost. They typically cycle through numerous iterations during development and achieving an injector design that successfully meets the target criteria often requires the application of prior experience, extensive simulation and testing, and ordinary trial and error. In an effort to reduce the heuristic nature of the process, this paper proposes a methodological approach to impinging-jet injector design for LOX/Kerosene engines. This structured approach considers the risks and limitations of early decisions, such as element and manifold types, and their downstream effects on the design process. As an illustration of the approach, the methodology is applied to the design of two different injectors with varying inputs for the same engine. Both of these injectors were manufactured and cold-flow tested to characterize fluid dynamic performance. Cold-flow tests were conducted with water as an analog to analyze the propellant distribution through each orifice and pressure drop. One injector was hot-fire tested in a technology demonstrator liquid rocket engine and performed nominally, with high c* efficiency and stable combustion. The application of this methodology ensures a more deterministic approach to the injector design process potentially reducing the time and cost of development.

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.

The Phoenix-1C sounding rocket, developed under the Aerospace Systems Research Institute’s Phoenix Hybrid Sounding Rocket programme, was flight tested for the first time in March 2023. The rocket featured a payload module specifically designed to accommodate three experimental payloads and to be recovered by parachute. This paper seeks to detail key aspects of the design, fabrication and testing of the payload module and its constituent subsystems, as well as report on its operation prior to and during the Phoenix-1C flight test.

This paper describes preliminary results from an ongoing experimental study on the ablation performance of compression-moulded ablative rocket engine materials under the development at the University of KwaZulu-Natal's Aerospace Systems Research Institute (ASRI). The study supports ASRI's larger objective of developing new techniques to manufacture rocket engine ablative liners more cost-effectively and rapidly using compression moulding. Although the tape-wrapping process yields ablative liners possessing excellent ablation performance, it is a significantly more complex and costly process to perform in comparison to compression moulding. In this study, we focus on compression molding of polymeric ablative composites comprising powdered novolac phenolic resin and short-strand fibrous materials as reinforcements. This process requires adequate mixing and pre-impregnation of the fibers with the resin prior to the moulding operation. The efficacy of dry and wet mixing of the composite constituents has been evaluated with water and Isopropyl alcohol (IPA), and adequate mixing was achieved with the use of wetting agents. Ablative test samples were manufactured via the conventional compression moulding process at low pressure and subjected to ablation testing under the steady-flow hot gas environment of the oxyacetylene torch. The results showed that the thermal insulation effectiveness of the samples increases with decreasing IPA content while the erosion resistance decreases with decreasing IPA content. Insulative index and erosion rate results from the ablation testing are presented and future plans for continued testing are outlined. Compression moulded ablatives showed promising ablation characteristics, and in the near future, more samples moulded at high pressure and using a vacuum-assist compression moulding process will be tested and compared with the results presented in this paper.

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.

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.

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.

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.

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.

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.

Paraffin wax has been identified as a hybrid rocket motor fuel, which offers enhanced regression rates and improved combustion performance. While various investigations into the performance of this class of fuels are being conducted around the world, the consideration of its structural performance is often overlooked. The research presented here establishes a simplified, yet accurate method of defining the structural performance of a paraffin wax hybrid fuel grain to be introduced early in the design phase of a motor. The use of the Johnson–Cook (J–C) material model has been verified to work within the “low speed” ignition range experienced in paraffin wax/N2O hybrid motors, and therefore is used to predict failure in a variety of motors. The resultant stress profiles within the grains indicate that the grain outer to inner diameter (OD/ID) ratio, as well as the outer diameter (OD) itself, play an important role in the grain ability to withstand the loading conditions applied. Additionally, the grain structural properties, and the stiffness of the combustion chamber affect the severity of the internal stresses in the grain. The feasibility of large-scale pure paraffin wax grains without structural enhancement additives is thus found to be poor. Fuel additives should be considered for structural enhancement.

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.