## 1 Introduction  
Thrust regulation in liquid propelled rocket engines (LPRE) is applied to control the mass flow rate to the gas generator (GG) to keep the turbine power at the level necessary to achieve the required thrust chamber pressure. It is carried out by using thrust regulators and secondary combustion in the GG. Thrust regulation through afterburning inside the engine nozzle is of key importance for rocket and jet engines. For space applications, afterburning can be accomplished by injecting propellants (fuel or oxidizer or a combination of both) inside the nozzle divergent section of a rocket engine, which results in much-needed thrust regulation. Though thrust regulation has not been fully applied practically, it has the potential to provide several benefits to expendable and reusable space launch vehicles (ELVs and RLVs) [1, 2]. This concept can be applied to ELVs and RLVs in the form of the thrust augmented nozzle \"TAN\". Thrust augmentation can be used to re-energize the jet flow, avoid flow separation inside the nozzle, obtain a higher thrust-to-weight ratio, and keep the nozzle optimized throughout the mission, even in a vacuum. Furthermore, this study can also be applied to thrust vectoring in a launch vehicle, which will not only improve its flight performance but will also permit high-g flight maneuvers at reentries and vertical landings. The future of space exploration looks promising, as government and private businesses plan frequent space flights for Moon and Mars expeditions, for which this study could be a cost-effective solution. [3, 4]. In rocket engines, secondary combustion in GG has been under study with solid propellants. The interior ballistic stability and ejection efficiency of a rocket are analyzed by conducting secondary combustion in GG of solid-propellant air-turbo rockets through solid propellants. Three-dimensional planar numerical simulation was performed on the secondary combustion and ring-cavity structures by [PERSON]; furthermore, the combustion effect on the loads and interior ballistic stabilization during ejection was studied [5; 6]. Moving to fuel-rich combustion results in more receptiveness to instability and makes it more difficult to achieve the required exhaust temperature. Two-zone GG with ethanol as a secondary fuel gives a stable performance in terms of exhaust temperature. Fuel and oxidizer mass flow rates of the main chamber are controlled through the varying throat cross-sectional area of the GG nozzle, and thrust vector control is achieved [7; 8; 9]. Secondary combustion in a hybrid rocket engine minimizes the loss of performance by shifting the oxidizer to fuel ratio. [PERSON] modelled secondary combustion with gaseous secondary propellant in CFD and compared its results with the experiment. LPRE lacks research in secondary combustion of GG with gaseous oxygen as a secondary oxidizer, and, hence, thrust regulation of the main engine through turbine power. Thrust regulation is done to maintain the performance of the rocket engine. This paper describes the design of a gas generator which can perform thrust regulation for the main rocket engine designed for a lab testing unit. Thrust regulation is achieved by doing secondary combustion in the divergent section of the GG nozzle. Secondary combustion in GG leads to Mach enhancement; hence, giving more power to the turbine shaft, which results in a pressure rise of the pump and an increase in the propellant mass flow of the main combustion chamber. Hence, the overall thrust increment of the main engine is achieved [10; 11; 12]. For achieving better thrust regulation, in this research, two secondary injection techniques, normal and tangential, were analyzed through CFD and a better performing technique was selected for the prototype development. The experiment was performed using gaseous oxygen (GOX) as a secondary injection propellant, which was obtained by heating the main oxidizer, liquid oxygen (LOX).