Chandrayaan-3 Lander propulsion system designed and developed by Liquid Propulsion Systems Center (LPSC) played a vital role in achieving successful soft landing near the south pole of moon on 23rd August 2023. Four number of 800 N lander engines reduced the lander orbital velocity of 1683 m/s to 0 m/s at the designated landing site within a span of 1131s through a series of complex thrust control maneuvers while 8 Nos. of 58 N attitude control thrusters oriented the lander along the desired flight path. The throttleable propulsion system consisting of lander engines, control thrusters, flow control valves, flow control electronics package, propellant supply and regulation system along with numerous flow control elements is a sophisticated bi-propellant propulsion system which has undergone extensive development, testing and qualification process over a span of 6 years. The lander engines alone have undergone hot firings for a cumulative duration exceeding 21720 s over 107 hot tests on three different hardware as part of the qualification program. A unique element of the propulsion system involves a real-time throttling algorithm which regulates the engine thrust to meet the guidance and navigation thrust requirements every 16 ms in open-loop. The onboard algorithm is validated extensively through 23 Nos. of system level engine hot tests in addition to numerous computer simulations. A detailed analysis of the thrust regulation algorithm indicated an additive 3σ dispersion of ±5% in thrust setting accuracy and ±3% mixture ratio regulation in open-loop. In order to accurately regulate the total thrust, the onboard throttling algorithm is augmented with a novel closed loop thrust control logic based on real-time lander acceleration feedback with an regulation accuracy < 1%. In Chandrayaan-3 mission, the lander propulsion system performance is demonstrated through a series of planned deboosting maneuvers on lunar orbit 2 days prior to the actual powered descent of the Chandrayaan-3 lander. In flight experience of the performance of lander propulsion system and throttling algorithms are found to be within the acceptable dispersion. The closed loop thrust regulation algorithm has accurately maintained the lander guidance acceleration requirements throughout the powered descent duration of ~19 min resulting in precise nominal successful soft-landing of Chandrayaan-3 lander near the designated landing zone.

The inducer is a special type of axial pump, which is especially designed to improve suction performance of a centrifugal pump. The usage of inducers allows designers to choose higher operating speeds for a typical centrifugal pump, thereby reducing the overall weight of the system. As far as rocket engine pumps are concerned, this aspect is of prime importance and most of the rocket engine pumps are equipped with inducers upstream of centrifugal impellers. Inducers are broadly classified into low head and high head inducer based on the head coefficient (ψ). This paper focuses on hydrodynamic design of a high head inducer having head coefficient (ψ) > 0.35. A mean line design is done initially based on the recommended design criteria available from literature, followed by a 3D CFD analysis to further optimize and fine tune the design. An accurate numerical model capable of predicting the performance of an axial inducer is of particular interest to rocket engineers in order to rapidly provide indications for the preliminary design of the machine. Commercial CFD solver ANSYS CFX is used for the numerical studies and meshing was done in TurboGrid. Grid independence study was done to ensure correctness of the mesh. The head rise and efficiency characteristics over a broad range of flow coefficients were determined. NPSH requirements for the nominal flow rate were also determined through CFD using Rayleigh–Plesset cavitation model. Among various design parameters, number of blades, radial tip clearance, and wrap angle were found to have a major influence on the performance of the inducer. Based on the results of the CFD analysis, initial design was optimized and found to meet the design requirements.

Liquid rocket engines are among the most complex machines ever created. High energy propellants are pumped by using high speed turbopumps and combusted to generate the required thrust for a launch vehicle. When the propellants used are cryogenically cooled, they pose concern due to the difficulty in handling the extreme thermal conditions. The mixing of the incompatible fluids could cause catastrophic failures to the launch vehicle. Hence, sealing systems are employed to prevent or minimize the interaction between such fluids. ISRO is presently in the process of developing a pump-fed liquid rocket engine employing liquid oxygen (LOX) and kerosene as propellants. The presence of large thermal gradients in the turbopump makes the sealing system quite complex and challenging, especially while designing with reuse capability. The present paper discusses the various dynamic seals that are employed in the turbopump.

Electric pump fed rocket engine cycle (EPFC) is gaining the attention of rocket designers across the globe in recent years. The idea of using electric motors to run propellant pumps instead of the conventionally used gas turbines seems attractive in many aspects ranging from development easiness to in-flight engine control. Though the successful flights of Electron rocket by Rocket Labs Inc., solely powered by electric pump fed Rutherford engine cluster, proves the feasibility of using this cycle in a practical rocket, its merits and demerits in comparison to conventional rocket engine cycles are yet to be well understood. In the study reported in this paper, a comprehensive comparison is made between Electric Pump Fed Cycle (EPFC) and Gas Generator Cycle (GGC) for their relative merits/demerits while used in the development of a micro-satellite launch vehicle. LOX-methane propellant combination is chosen as the single propellant combination for the vehicle and a full-fledged comparison with practical considerations in selection of design parameters and constraints is presented. The results obtained are discussed in detail with regard to applicability in a practical scenario. Quantitative estimation of performance parameters is used, to the extend possible, for comparison and qualitative conclusions are adopted only where quantitative estimation is not possible.

The Indian Space Research Organisation (ISRO) envisaged a launch vehicle to put 4 ton class satellites to Geosynchronous Transfer Orbit (GTO). GTO is Geosynchronous Transfer Orbit of 170 km perigee and 35,975 km apogee, where the satellites are injected. It is further raised to the Geostationary orbit by firing satellite thrusters. The upper stage of this launch vehicle is powered by a cryogenic engine of 200 kN nominal thrust working in gas generator cycle using liquid hydrogen (LH2) and liquid oxygen (LOX) as propellants. The engine has independent LOX and LH2 turbopumps, with their turbines operating in series mode. The turbopumps were successfully developed through a series of development tests, which included subsystem level tests for pumps, turbines, bearings and seals with simulated fluids. Subsequently, tests were carried out in the turbopump level with different pump fluids like water, liquid nitrogen, liquid oxygen and liquid hydrogen. Finally, tests were carried out in closed-loop mode (boot-strap) with gas generator and turbopumps operating together. The successful completion of these tests demonstrated the design validity of these turbopumps. Subsequently, turbopumps have undergone flight duration and extended flight duration tests as part of engine. The third development engine was used in the successful flight testing of the launch vehicle in 2017. The efforts towards the development of turbopumps are detailed in this paper.

Regenerative cooling of thrust chamber is the unique solution for the thermal management of high heat flux generated inside the combustion chamber of Cryogenic rocket engine. Heat is transferred from combustion hot gas to coolant through the channels provided on inner copper shell, thereby cools the inner wall of the nozzle. A novel technique of providing copper foam inside the channels will act as an infinite fin and also act as barrier for coolant stratification. This will improve the heat transfer to the coolant and reduce the nozzle wall temperature. Heat transfer improvement with copper foam inserts to the coolant channel is demonstrated through experiments with simulated fluids. Experiments are conducted with simulated hot gas chamber and coolant channels using water as the coolant. Copper foam with high porosity is selected to fill the channels. Hot tests are carried out with copper foam filled coolant channels and measured the coolant temperature rise and pressure drop across the channels. Tests are repeated with similar hot gas condition, but without inserting copper foam inside the channels. A substantial enhancement in heat transfer to the coolant is observed with copper foam inserts experiments, which will reduce the wall temperature. This gives a good handle on the life cycle improvement of multi-start cryogenic engines for future space transportation systems. This paper details the specification of copper foam, hardware design, experiments and measurements, and the application of the augmentation of heat transfer coefficient in operating cryogenic engines.

Cryogenic engine capable of delivering 200 kN thrust is being developed for the first time in the country by ISRO for powering the upper stage of GSLV Mk-III, the next generation launch vehicle of ISRO capable of launching four tonne class satellites to Geo-synchronous Transfer Orbit(GTO). Development of this engine started a decade ago when various sub-systems development and testing were taken up. Starting with injector element development, the design, realization and testing of the major sub-systems viz the gas generator, turbopumps, start-up system and thrust chamber have been successfully done in a phased manner before conducting a series of developmental tests in the integrated engine mode. Apart from the major sub-systems, many critical components like the igniter, control components etc were independently developed and qualified. During the development program many challenges were faced in almost all areas of propulsion engineering. Systems engineering of the engine was another key challenge in the realization. This paper gives an outlook on various technological challenges faced in the key areas related to the engine development, insight to the solutions and measures taken to overcome the challenges.

Cryogenic propulsion systems using liquid hydrogen and liquid oxygen propellant combination are used in satellite launch vehicles in view of the higher specific impulse (ISP). Due to extreme low temperature of cryogenic propellant and low density and explosive nature of LH2, the development of cryogenic propulsion system is very complex and time consuming. In ISRO an indigenous cryogenic upper stage powered by an engine developing a nominal thrust of 73.5 kN in vacuum with a propellant loading of 12.8 tonnes is developed and successfully flight tested for the first time in GSLV D5 flight on 5th January 2014. A mathematical model is developed for finalizing the engine start and shut off sequence, ensuring smooth and safe ignition, predicting performance of subsystems under transient phase, nominal and off nominal conditions, control systems parameter settings and for the performance estimation of engine in sea level and flight. This paper highlights the configuration and working of a cryogenic engine, mathematical modelling of the engine, applications of the model in the cryogenic engine development and comparison of predicted values with the test results.

The liquid propellant rocket engine combustion chamber represents one of the most difficult engineering flow systems in operation. The ignition of the propellants injected into a rocket combustion chamber and the subsequent propagation and anchoring of the flame is an important design consideration for all types of rockets. Reliable ignition has to be guaranteed and the initiated turbulent diffusion flame has to stabilize without over pressure or blow out. The control of ignition in a rocket engine is a critical problem for combustion chamber design. Delayed ignition may lead to high unsteady chamber pressure that can damage the engine (strong ignition) whereas early ignition may not sustain to reach steady state. Predictivity of the models and numerical tools to analyse the ignition transient has still limitations. In H2/O2 rocket combustors, the injected propellants are ignited by a stream of hot gas originating from the igniter device. The hot gas has to mix with the injected propellants to initiate combustion in a situation which is characterized by strong spatial inhomogeneties. This paper presents the numerical study of ignition characteristics of H2 & O2 propellant combination in cryogenic combustor. The main objective is to get an insight into the main processes involved in the ignition of cryogenic engines. The pressure, temperature, velocity profile and propellant mass fraction variation along the combustion chamber is addressed. The characteristics of diffusion flame, shear layer combustion, recirculation zone and flame propagation are also addressed. The reaction mechanism is studied using Eddy Dissipation Model/Finite Rate Chemistry. Numerical simulation results were compared with the experimental data.

This paper presents a study of a fully reusable earth-to-orbit launch vehicle concept with horizontal take-off and landing, employing a turbojet engine for low speed, and a rocket for high-speed acceleration and space operations. This concept uses existing technology to the maximum possible extent, thereby reducing development time, cost and effort. It uses the experience in aerial filling of military aircrafts for propellant filling at an altitude of 13km at a flight speed of M=0.85. Aerial filling of propellant reduces the take-off weight significantly thereby minimizing the structural weight of the vehicle. The vehicle takes off horizontally and uses turbojet engines till the end of the propellant filling operation. The rocket engines provide thrust for the next phase till the injection of a satellite at LEO. A sensitivity analysis of the mission with respect to rocket engine specific impulse and overall vehicle structural factor is also presented in this paper. A conceptual design of space plane with a payload capability of 10ton to LEO is carried out. The study shows that the realization of an aerial propellant transfer space plane is possible with limited development of new technology thus reducing the demands on the finances required for achieving the objectives.

Experimental studies have been conducted on a new type of injector which can be used very advantageously in rocket engines designed for throttling applications. The aim is to keep the drop size constant for varying chamber pressures. The injector uses very small quantities of a gas to atomize a straight liquid jet before discharging into the combustion chamber. The concept is demonstrated over an eightfold variation of injector pressure drop. A feedback control system, which can regulate the air pressure with respect to the sensed chamber pressure, based on the input transfer function, so as to keep the drop size constant, is suggested. The transfer function is obtained from the experimental data. The major advantage of this injector is that it uses a very small quantity of gas, thereby limiting the weight penalty to a minimum. This system is much less complicated compared to variable area injector and cavitating venturi valves used for this purpose.