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.

Chandrayaan-3 became the first spacecraft to land on the lunar surface near the south pole of the moon on 23rd August 2023. The spacecraft was inserted into an elliptical orbit by the launch vehicle, following which the spacecraft went under multiple orbit raising maneuvers using a LAM engine. Chandrayaan-3 spacecraft composite consisted of two modules viz., Propulsion Module and Lander Module. The Propulsion Module of Chandrayaan-3 employed a bi-propellant chemical propulsion system propelled using a LAM engine and eight 22N RCS thrusters. LAM Engine incorporated in Chandrayaan-3 was an AR-250 class engine. LAM engine has a heritage of over 50 flights and was critical to all the earth bound and lunar spacecraft maneuvers. Reliable and repeatable performance of the propulsion system and the LAM engine are critical for the optimized operations of mission. This paper highlights the test and flight performance analysis of the LAM engine used in Chandrayaan-3.

In the age of space exploration, the need of robust engines to perform a variety of operations is inexorable. Engines must cater to a variety of operations like space rendezvous, orbital manoeuvring, planetary entry and descent, and hovering and hazard avoidance during landing operations. All these operations call for a robust engine with throttling capability. This paper summarizes the development of a throttling engine of 800N thrust and its performance in Chandrayaan-3 mission.

Future human-space exploration efforts aim to achieve maximal synergy between human and robotic missions, where in human-scale robots shall supplement astronaut exploration activities and also undertake robotic precursor missions. Towards this, design and development of a flying Smart Space Robot (SSR) is initiated. The Smart Space Robot (SSR) is a space flying robot which will be tethered to the unmanned orbital platform on the fourth stage (PS4-UOP) of Polar Satellite Launch Vehicle (PSLV) for micro g experiments. SSR is of nano-satellite class with dimensions of ~350 × 350 × 350 mm, weight of ~12 kg and power of ~30 W. The nanosat has to undergo different phases of operations which include deployment, station keeping, retrieval and docking. These manoeuvres calls for a robust system capable of operating for an extended duration with multiple restart and pulsing capability. This paper details out the development and design process of a cold gas system within the constraints of space, volume, mass, voltage and power as designated by the nanosat specification which meets the mission requirements.

Thrust developed in a rocket is mainly a function of the chamber pressure developed in the combustion chamber and thrust amplification factor (Cf) due to expansion of hot gases in the nozzle as long as the ambient pressure remains constant. For a liquid propellant engine, chamber pressure can be varied by altering the feed pressure, and its applications are in the blowdown or throttling mode of operations. Cf will also vary with chamber pressure as flow separation in divergent nozzle gets altered. A study has been carried out to develop a model to predict the thrust of a hypergolic bipropellant engine for varying operating chamber pressure at the same ambient pressure condition. Engine considered for the study is of medium thrust range. A mathematical model was developed from first principles to predict the thrust of the engine for varying chamber pressure for the same ambient pressure condition. Prediction was made from 100 to 50% of the maximum thrust. The model was validated based on sea level hot tests for 10 s each with thrust measurement, and the variation in thrust was achieved by altering the feed pressures. The results were found to be in close agreement with the predicted performance, and the maximum deviation was found to be 3.0%. This paper gives the details of the study carried out.

Hypergolic bipropellant radiation cooled thrusters utilizing Mono-methyl Hydrazine and Nitrogen Tetroxide are commonly used in spacecraft missions for attitude and orbit control. The performance index of a rocket engine is the specific impulse which is a function of combustion efficiency, known as C* efficiency, and the nozzle efficiency. An experimental investigation is carried out to evaluate the effect of combustion chamber design parameters on performance(C*efficiency) for varied injector spray and atomisation characteristics as well as injection pressures. Analytical model with empirical correlation available in the literature is used as a tool for understanding the physical process in the combustion chamber and predicting C* efficiency which was validated with experiments. Design variables considered are characteristic length and the contraction ratio. Cold flow evaluation of the injector using simulant water was done to evaluate the droplet size and injection velocity, which is normalised to the propellant flow conditions. Hot test for 10s using the stainless steel chamber was done at sea level with instrumentation for chamber pressure, mass flow rate of propellants and throat temperature. Results show that for a given injection and operating conditions, there exists a range of L* and contraction ratio where C*efficiency will be optimum and less sensitive. Trends in throat temperature measured are also evaluated. This paper presents the details of the investigation.