LOX/Methane propellant combination is one of the most suitable solutions for future liquid rocket engines, due to good performances achievable in terms of specific impulse combined with operation advantages, such as storability, low toxicity, availability and production cost. Another reason to pursue the development of LOX/Methane propulsion systems is the potential for in-situ resource utilization during inter-planetary missions. Due to these advantages, ISRO is developing LOX/Methane engine for its future launch vehicles.

For cooling the thrust chamber, Methane is to be used for both regenerative and film coolant. However, Methane is an inferior coolant as compared to hydrogen and methane properties shows large variations in its operating regime due to two phase flow in the coolant channels.

This calls for detail modelling (considering proper reaction steps, species and reaction rates) of two phase flow of coolant in regenerative coolant channels, participation of film coolant combustion (effective film coolant layer thickness) and conjugate heat transfer from combustion products to chamber inner wall through film cooling layer.

Deliverables to ISRO: Validated flow & thermal code (including the source code) has to be submitted to ISRO.

B 2.3
004

Clearance seals are used in Liquid Hydrogen/Liquid Oxygen during chilling phase and in vacuum conditions during operation of turbo pumps. The vibration, friction and wear characteristics of the seal/runner combination plays a vital role in the turbopump performance. Detailed study in this area is required to optimize the existing seal designs/configurations.

Control of combustion instability in liquid engines.
Combustion instability is a phenomenon that sometimes occurs in liquid engines and can lead to damage/destruction of the hardware.  It can be controlled by passive techniques such as slots, baffles and resonators.  To widen the range of operating conditions under which control is effective, active control techniques such as anti-sound and secondary fuel injection can also be used.  It is necessary to characterize the performance and stability of passive/active control techniques to evaluate suitability for liquid engines.

Supersonic film cooling of nozzle divergent
The Nozzle divergent of ISRO upper stage engines are cooled by dump cooling. Dump cooling needs double wall construction, which is having higher weight penalty. Moreover overall Isp of the engine is reduced as the dump coolant is expelled at low temperature. The existing dump cooling method can be upgraded by introducing supersonic film cooling, in which coolant will be injected in the nozzle divergent in the form of a thin slit. This method of cooling needs only single wall structure. Also the coolant will expand through main nozzle at high temperature, which will result in higher Isp.

Theoretical modeling of atomization in cryogenic injectors
Development of the primary atomization model to predict the droplet size as well as to have the secondary droplet formation, vaporization etc. for the complex gas-liquid flow existing in the cryogenic engine injectors of CE20 and CUS, where the engines are operating in supercritical conditions of the propellant.  Some common approach could be of generation of spray model by full Lagrangian approach, or the phenomenological approach of the Ω-Y model or the hybrid approach using VOF-Lagrangian coupling. The model generated can be validated through experimental data in actual hardware.

Development of throttleable injector element for liquid engines
Throttleable injectors are necessary for missions requiring soft landing or stage recovery.  Various types of swirl configurations as well as pintle type of injectors which will have control over the orifice opening area for flow control and thus control over thrust developed by engine are employed.  This could find application in LOX-Methane engine design concept which requires soft landing type of operations for future space missions. Modeling and experimental characterization of the injectors for various operating conditions are areas to be studied.

Development of a theoretical model to determine the characteristic frequency of feed line coupled-oscillations in liquid rocket engine
This project is intended to study the vibration response of metallic pipelines conveying propellants to a liquid engine. The pipelines will have the added mass of the fluid inside it. During operation of the engine (hot firing) it will generate vibrations predominantly in the axial direction. The vibratory response of these lines is to be investigated.

Regenerative cooling analysis with Kerosene for Semi cryogenic thrust chamber to study coking characteristics
For Semi-cryogenic engine the coolant used for regenerative cooling is refined version of kerosene (equivalent to RP1). The Kerosene is a mixture of many hydrocarbons. During passing through the coolant channels the temperature of Kerosene increases. However, when Kerosene comes in to contact with high temperature walls, it decomposes and leaves behind a sticky rubber like substance called coke. This coke can clog the injector element holes, which are very small in size (~0.8 mm). Hence the coking characteristics are to be studied thoroughly for different operating conditions.

Prediction model for vibration in Turbopumps considering the effects of unbalance, constraints, fluid forces, seals, internal clearances (housing/shaft/bearing) etc.
Turbopumps are highly critical systems due to its high power, high speed operation and hazardous propellants being handled. Condition monitoring of turbopumps is essential for the safe shutdown and preventing catastrophic damage. Vibration assessment is a very effective tool for condition monitoring of turbopumps. A prediction model for vibration in turbopumps can be effectively used for condition monitoring of turbopumps during its operation.

Mathematical modeling of liquid migration under Zero ‘g’ condition and the associated heat transfer with warm tank wall and pressurant gas is essential to predict the rate of pressure build up in LH2 tank
In cryogenic propulsion Stage residual liquid migration in LH2 tank is generally observed after engine shut down.  This causes higher tank pressure due to mixing of liquid hydrogen with warm pressurant gas and heat transfer with warm tank wall.

Development of diagnostic tools for measurement of plasma/plume parameters of stationary plasma thruster and pulsed plasma thruster
Plasma parameters like electron temperature, density, energy distribution function and plasma potential are essential parameters to evaluate to understand the performance of plasma thruster. Since near field region of Hall thruster is so hot, measurements cannot be carried in low frequency mode. So that high temporal resolution probes are required to serve the purpose. However, measurement using electric probes like Langmuir probe and Emissive probe is challenging in high frequency range. Hence Langmuir and Emissive probes with special measurement circuit is needed to special attention to serve the purpose.

1. Langmuir probe and emissive probes compatible to high temperature plasma
2. Measurement circuit with high frequency measurement compatibility
3. Program (code) to analyze and deduct the results from raw data

Model for the prediction of thruster’s life (LPSC)
Non-contact type measurement of thruster anode liner erosion Anode liner is a crucial part, which decides the life of Stationary Plasma Thruster (SPT) also known as Hall thrusters. This liner protects the magnetic circuit from hot plasma and also play role in thruster performance. The high energetic ion beam impregnates the liner material and causes erosion and affects the life. Since the physical presence of probes will disturb the thrusters plasma and also and can damage the probe, non-contact type measurements are preferable. On this ground, non-contact measurement of liner erosion and prediction of the thrusters life is very important.

1. Develop and demonstrate a non-contact experimental technique to measure the anode liner erosion.
2. 2. Model for the prediction of thrusters life.

The complete thermal modeling of the thruster
The monopropellant hydrazine thrusters are used in reaction control system of IRS projects.  The monopropellant hydrazine thrusters use principle of dissociation of hydrazine using catalyst to produce the exhaust gases.  These exhaust gases are expanded through the nozzle to produce thrust.  The complete hydrazine dissociation model for the monopropellant thruster is required for thruster design and optimization.  Based on the dissociation model, the complete thermal modeling of the thruster to be carried out. 

Alternate green propellants
The monopropellant hydrazine is highly corrosive, carcinogenic and not environmental friendly.  The alternate green propellants such as Ammonium Di Nitramide (ADN), Hydroxyl Ammonium Nitride (HAN) based monopropellants are under studies.  The green propellant formulation and its detail properties, dissociation phenomenon is essential to replace the existing hydrazine system.  Development of suitable catalyst for the green propellant.

Life cycle prediction of thrust chamber for reusable, regeneratively cooled liquid engines
This project is to study the cyclic life of double walled regeneratively cooled thrust chambers of liquid rocket engines. Theoretical and experimental investigations are required to study the thermomechanical behavior of different thrust chamber materials in the parent metal and welded forms for this at different temperatures and strain rates. Damage mechanics has to be incorporated in the studies.

Liquid film cooling study of thrust chamber with kerosene for LOX/Kerosene Semi-cryogenic Engine
The 2000kN Semi-cryogenic engine is a high thrust engine and generates high thermal load on the thrust chamber wall. Hence the thrust chamber is cooled by film cooling along with regenerative cooling. The film coolant, after injection in to the thrust chamber takes part in combustion, hence film coolant layer depletes gradually. For safe operation of engine, a positive film coolant (without film layer breakage) is an essential requirement. This necessitates detail modelling of film coolant phase change considering proper reaction steps, species, and reaction rates through CFD.

Modelling and analysis of throat film cooling for semi cryogenic Engine Thrust Chamber
The throat region of thrust chamber is subjected to maximum heat flux.  The conventional film coolant injected near the injector end is not sufficient to keep the throat wall temperature below safe temperature limit. Hence additional film cooling provided near the throat region. Isrosene (fuel used for Semi cryo engine: equivalent to RP1) is used for both film cooling as well as regenerative cooling in liquid state.

For safe operation of engine, a positive film coolant is an essential requirement. Hence detail modelling of phase change of film coolant and participation in combustion is required considering proper reaction steps, species and reaction rates. Based on the CFD results an empirical correlation is also to be formulated for predicting the film coolant thickness in the thrust chamber for varying operating parameters.

Modelling of anomalous electron transport in Hall thruster
Anomalous electron transport is one of the prominent mechanism in Hall thruster. Various explanations like cross-field diffusion, Bohm diffusion, near wall conductivity and azimuthal electric fields are attempted to understand the electron transport. The real cause likely to be combination of all. The complex behavior of electrons in cross-field hall thruster can be handled by categorizing in the following way: (1) Electron mobility in the discharge channel (in the near-anode, ionization and acceleration zones), (2) Azimuthal Hall current (electron propagation around the channel) and (3) electron mobility in the plume. Since these regions are characterized by different magnetic field strength and orientation, each region has to be handled differently. The first part of the study reveals the efficiency of ionization and thrust generation mechanism, second part gives the information of electron confinement in the channel and last part gibes the electron transport from external cathode to the discharge channel across and along the magnetic field. The following are expected deliverables:

1. Simulation of electron mobility in discharge channel of Hall thruster with different magnetic field topology.
2. 2.Simulation of Azimuthal Hall current with different magnetic field topology.
3. 3.Simulation of electron transport in plume along and perpendicular to the magnetic field.
4. 4.Prediction of suitable magnetic configuration for better operation of Hall thruster from the simulation results

Modeling of film cooling/sweat cooling in Liquid Rocket Engines
All high thrust liquid rocket engines  employ film cooling along with regenerative cooling to reduce the heat flux at throat. As the film coolant undergoes phase change and gradually takes part in combustion, it is difficult to predict the overall effectiveness in different operating conditions. Also higher film cooling rate reduces engine specific impulse. Hence a detail CFD analysis is required for all ISRO liquid engines to optimize film cooling flow rate.

Combustion modelling& combustion instability modelling of liquid rocket Engines
Numerical steady state flow model can be carried out with the consideration of the various species for the Cryo/Semi cryo thrust chamber and this can be perturbed to give unsteady results which give signatures of the dominant acoustic modes.  The initial guess of the droplet size of the propellant will decide the vaporization, mixing and combustion in the chamber, therefore an experimental assessment of the droplet size will help in accurate predictions.  Stability assessment can be done for ranges of droplet sizes and the stability boundaries can be defined.  Design evaluation of the present Cryo/semi cryogenic engine with stability enhancing mechanisms can be taken up for studies.

Numerical modeling of nonlinear thermo-acoustic instability in liquid rocket Engine
Numerical modelling of nonlinear thermo-acoustic instability is required as instability is inherently non-linear. Work has already been done in computational modelling using Unsteady Reynolds Averaged Navier Stoke technique (URANS). The use of Large Eddy Simulation (LES) and Direct Numerical Simulation (DNS) techniques is to be investigated to improve accuracy.

Optimization of passive suppression devices for thermo-acoustic instability in liquid rocket chamber
Passive suppression devices for suppressing thermo-acoustic instability such as baffles and resonators are being used in rocket motors worldwide. LPSC needs to investigate the damping characteristics of baffles, resonators. Etc. for application to our engines Semi cryogenic engines.

Two phase flow modeling in cryogenic propellant feed lines
Cryogenic engines make use of propellants such as liquid hydrogen and liquid oxygen at very low temperatures in order to obtain a high specific impulse and controllable thrust. The feed lines are to be chilled to cryogenic temperature prior to the start of engine operation to avoid undesirable flow oscillation. The flow of cryogenic fluids through a feed line is complex due to two phase flow, heat in leak in the feed line and boils off at the source. The flow of cryogenic fluids is complicated because surface tension makes all the dynamics nonlinear. Moreover, the density of the two phases differs considerably and compressibility becomes important due to the large change in density. Flow induced pressure drop can lead to further change of phase. In addition various types of instabilities may develop. Methods are available for modeling of two phase flow using numerical techniques such as volume of fluid method, level set method, front tracking and the Lattice-Boltzmann methods. However, modeling of two phase cryogenic flow incorporating properly phase change and heat transfer as well as fluid dynamics is still a developing field. Specialized models for cryogenic engines will have to be developed considering the actual fluids and operating conditions. New methodologies may have to be investigated to accurately capture the flow behavior.

Conjugate Heat Transfer Analysis in Cryogenic Engine systems

1. Development of liquid film cooling model for Cryo and non-Cryo engines. Detailed model for helical chemical regenerative cooling in LOX-LH2 propellant along with film cooling
2. Ablative throat charring analysis in conjugate model.

Experimental evaluation and constitutive modeling to simulate structural behavior and failure criteria for dissimilar weld joints
This study plans to address the structural behavior and failure of dissimilar material weld joints commonly encountered in liquid rocket engine combustion chambers. Material combinations such as SS- Copper, Copper-Nickel, SS- Nickel have to be addressed. Different thicknesses have to be accounted for. FEA based simulations have to be done to predict the failure of such joints.

Spray interaction effects in a multi-element injector head of a liquid rocket engine
Multi-element swirl injectors are used in high thrust liquid rocket engines. The intra-element characteristics of the swirl injector is mainly influenced by geometrical and flow parameters. The injector elements are arranged in a specific pattern based on the thrust per injector element of the rocket engine. In addition, the conical spray from an injector element interacts with the spray formed in its neighboring elements. The performance and stability of liquid rocket engine is influenced by both intra-element and inter-element spray characteristics. Spray interaction in multi-element injector head depends on both the intra-element spray characteristics as well as combustion chamber operating conditions. For simulating the spray interaction in a multi-element injector, experiments and analyses need to be carried out at different operating conditions.

Studies on deflagration to detonation transition
Combustion can be classified into deflagration and detonation based on whether the flame is travelling at subsonic or supersonic speeds respectively with respect to the unburned medium.  Deflagration is associated with low overpressure whereas detonation is explosive in nature having high overpressure associated with a shock wave. In conventional liquid rocket engines, deflagration occurs whereas in scramjets and pulse detonation rocket engines, detonation can occur. Dynamics of deflagration and detonation waves have been investigated extensively.  However, under certain conditions a deflagration wave may accelerate and transform into a detonation wave. This mechanics of this transition are not well understood. It is necessary to obtain a deeper understanding of the deflagration to detonation transition in scramjets and detonation based rocket engines using numerical and experimental techniques. The data would be required for design and optimization of these systems.

Modeling of atomization of coaxial injectors, impinging jet injectors
Coaxial injectors are mainly selected for gas-liquid propellant combinations in both the Cryo and Semi cryo engines.  The impinging injector finds application in the earth storable engines where the injectors are operating in liquid-liquid mode. Flow modeling of atomization of coaxial or impinging jets and parameters affecting the atomization, mixing, vaporization of the propellant is to be studied. Theoretical studies with experimental correlation can be carried out.

Effect of acoustics on spray characteristics of swirl coaxial injectors
Combustion instability is characterized by large pressure perturbations with attendant large thermal stresses and is one of the most important challenges for liquid rocket engine design.  Low and medium frequency combustion instability is believed to be caused by the dynamic processes in supply system or combustion. Dynamic processes with specific reference to atomization are to be studied & modeled as it plays important role.

Finite element simulation of non-linear, high strain forming processes of metals like deep drawing, flaring etc.
This study is for evaluating the state of stress and strain in sheet metal work undergoing high plastic strains as in deep drawing and flaring. FEA based simulations and tests are required to assess the structural integrity of the work piece so as to optimize the process parameters.

Transient chill down analysis of regeneratively cooled thrust chambers
The Thrust chambers of cryo engines are preconditioned by low and high flow rate chill down before engine firing. This is a highly transient phenomenon involving conjugate heat transfer, radiation and phase change. A proper flow analysis is essentially required to optimize the coolant flow rate during the chill down as this coolant is drawn from the fuel tank during flight. Optimizing chill down flow rate can significantly improve the payload carrying capability as it will reduce the fuel loading from the upper stage engine.

Film cooling breakage studies under unstable combustion conditions (Boundary Layer breakage)
During combustion, instability can lead to multiple high pressure zones inside the thrust chamber. This in turn breaks the film coolant layer near the boundary layer. As a result the heat transfer to chamber wall can increase significantly, which may lead to catastrophic failure of thrust chamber. Hence a proper study of effect of various modes of combustion instabilities on film coolant layer is to be carried out.

Evaluation of damage cretiria for AA2219 welds under bi-axial stress field from experimental and simulation (LPSC)
This project plans to address the damage modelling and failure prediction of AA12219 GTA welds and parent metal under biaxial state of stress. Both tests and simulations based on FEA are required for this. Different biaxial stress rates have to be addressed. Tests have to be conducted for different specimen thickness commonly encountered in liquid propellant tanks.

Prediction of anode liner erosion of Hall Effect thruster
Anode liner is a crucial part, which decides the life of Stationary Plasma Thruster (SPT) also known as Hall thrusters. This liner protects the magnetic circuit from hot plasma and also play role in thrusters performance. The high energetic ion beam impregnates the liner material and causes erosion and affects the life. Since the life of the SPTs is of the order of 1000s of hours, it is not practical to measure the erosion pattern on full scale. Hence we need a technique to estimate the erosion rate and life of the thrusters based on suitable modelling.

1. Model to predict the anode liner erosion of Hall thrusters.
2. 2. Estimation of Hall thruster’s life.

Design of packaging/interfaces for MEMS based fabricated valves & actuators
Development and testing of a MEMS basedPiezo valve is successfully demonstrated. However the next challenge in development is to integrate the microsystem (valve) to the mechanical fluid system. Interfacing of the silicon wafer etched micro valve to the stainless steel tubing with proper interface sealing is yet to be developed.

Development of a mathematical model for characterizing the dynamic behavior of a check valve under different operating conditions
Check valves sometimes exhibit chattering under flow conditions which is not desirable.  For double poppet check valves which are two check valves in a series mode this phenomenon is complex and hence needs to be studied.
Modelling of double poppet check valves preferably using specialized software tools like AmeSIM  to

1. Characterize check valve chattering under different input conditions for the given Design and
2. 2. Characterize and optimize the valve design parameters for eliminating chattering. Check valve chattering

Design & Development of solenoid coils for Liquid Helium applications
Liquid helium storage under pressurized condition and on-board isolation will call for a fast response electromagnetically actuated solenoid valve. At these temperatures because of the low viscosity of liquid helium sealing is a concern.  The valve envisages the usage of superconducting winding wires wherein the current carrying capacity is amplified many times due to a drastic drop in coil resistance accompanied by minimal increase in power at around liquid helium temperatures. The valve design can make use of magneto-strictive actuator. Development of valve with super conducting coil andachieving leaktightness for Liquid Helium application is a challenge.

Estimation of torque co-efficient and load distribution in threaded joints
This project is to study the mechanical behavior of bolted joints at Cryo, ambient and elevated temperatures under the combined action of bolt pre-torque, internal pressure and axial loads. Experimental and numerical studies have to be carried out considering different nut factors and bolts/flange materials.

Development of a mathematical model for propellant tank pressurization system chain for cryogenic application
Complex thermodynamic processes occur in the ullage Propellant volume of cryogenic tank during pressurization process and propellant outflow from tank. Hence analysis of cryogenic tank pressurization system requires development of a mathematical model for the pressurization process of cryogenic tank addressing heat transfer between cold tank wall and warm ullage gas, heat transfer between liquid free surface and ullage gas, enthalpy of pressurant gas supplied for tank pressurization, tank geometry, propellant outflow rate during engine operation etc. This thermal model computes temperature gradient in tank wall and ullage gas column with respect to liquid expulsion time and thereby predicting mass flow rate variation of pressurant gas with liquid expulsion time. This input is necessary for designing the pressurization system including the gas bottle for pressurant gas storage.

Modelling of Magneto Plasma Dynamic Thruster
Magneto Plasma dynamic thruster (MPD) is a plasma thruster where the ions accelerated with magnetic force. MPDs are mainly operates at
very high power in the range of 10 kW to one MW. Though it is possible to realize MPD's less than 10kW range, the performance of the thruster will not be as good as Gridded Ion thruster or Stationary Plasma thruster.  Pulsed Plasma thruster, VASMIR are the examples of MPD. There are mainly two types of MPD thruster.

1. Self-induced MPD
2. 2.  External magnet MPD.

MPDs with external magnets are the devices where strong magnetic field is generated which compliments the induced magnetic field. In this type of thruster also the thrust production mechanism is being the same, modelling of the magnetic field generated and applied magnetic field along with the modelling of plasma and its acceleration is important for the design of the thruster and understanding the thruster behavior. Last but not the least, to understand the main technical challenge associated with MPD i.e. erosion of cathode a very good model of MPD is essential.

C 2.31
005

Development of foil bearings for high speed Cryogenic turbopumps
The turbopumps and propulsion systems for space transfer vehicles require long-life turbopumps with high shaft speed capability. Rolling-element bearings used in current turbopumps do not have sufficient life for these applications. Process fluid foil bearings generally have long life and high speed capability, with exceptional reliability, over a wide range of temperatures and fluids. Obtaining high radial stiffness is often a challenge in the development of foil bearings.  The research will be focused on the development of foil bearings with high stiffness, speed capability and life for cryogenic turbopump applications (Liquid hydrogen and Liquid Oxygen). This also involves mathematical modeling for designing and estimation of performance parameters of foil bearings. Advanced foil designs namely leaf type, bump type, tape type etc. will be studied for use in turbopumps.

Expected deliverable will be the foil bearings & its design, which work with the cryogenic process fluid and possess high radial stiffness, speed capability and life. A mathematical model for the design of the bearings and estimation of performance parameters will also be developed.

C 2.32
006

Development of damper seals for turbopumps
The proposed research is for the development of damper seals for cryogenic turbopumps. The damper seals, in addition to leak control are required for dampening and controlling the excess rotor response which occurs during critical speed crossings in high speed turbopumps. Damper seals of various geometry namely honeycomb, pocketed, hole pattern etc. are currently in experimental phase worldwide.  We need to develop a seal of this category, experimentally demonstrate its performance and optimize its geometry for use in high speed cryogenic turbopumps.

Expected deliverable will be the optimized damper seal design which works in cryogenic fluids (Liquid Oxygen & Liquid Hydrogen). A mathematical model for evaluating the seal parameters like damping coefficients, stiffness etc. will also be developed.  Experimental setup for evaluating the performance of damper seals.

Non-contact type measurement of thruster anode liner erosion and prediction of thruster life
Non-contact type measurement of thruster anode liner erosion Anode liner is a crucial part, which decides the life of Stationary Plasma Thruster (SPT) also known as Hall
thrusters. This liner protects the magnetic circuit from hot plasma and also play role in
thruster performance. The high energetic ion beam impregnates the liner material and
causes erosion and affects the life. Since the physical presence of probes will disturb the
thruster’s plasma and also can damage the probe, non-contact type measurements are
preferable. On this ground, non-contact measurement of liner erosion and prediction of
the thruster’s life is very important.
Deliverables:
1. Develop and demonstrate a non-contact experimental technique to measure the anode
liner erosion.
2. Model for the prediction of thrusters’ life.

C 2.47
010

Multi plume interaction studies of Clustered Engines
Clustering is favorable because of several merits including a cheaper manufacturing cost, less demanding requirement from test facilities, more robustness and an ability to tolerate failure of single thrusters. The performance of a thruster in a cluster may be different from a standalone situation. One interest is to investigate the plume interactions, especially in the complex and important near field locations. To accurately simulate the plasma plumes from a cluster of Hall thrusters requires an accurate modelling of the complex physical plume mechanism on three-dimensional meshes. Traditionally, the computational simulation of plasma plume flows into vacuum is performed with a hybrid particle-fluid approach. The direct simulation Monte Carlo (DSMC) method models the collisions of the heavy particles (ions and atoms) while the Particle In Cell (PIC) method models the transport of the ions in electric fields. This study is intend to simulate the detailed three-dimensional plume structures and plume interactions. The following are expected deliverables:

1. Modelling of single thruster plasma plume.
2. Model to investigate near field plume interaction in cluster configuration.
3. Prediction of plasma parameters, electric potential and beam interaction in cluster configuration.

C 2.48
015

Characterization of Heat transfer parameters in Gel Propellant Engines
Gelled fluids are a homogeneous mixture of a base fluid and a gelling agent. They possess highly non-Newtonian rheological characteristics. Gel propellant rocket engines involve use of special additives to alter the rheological properties of the liquid propellant so as to convert the propellants into the gel state. Gel propellants offer advantages of solid propellants in that they are easy to store and handle. They also have the advantages of liquid propellants in that they liquefy when subject to high shear stresses in the injector. Gel propellants will not slosh, will not spill through leaks and have reduced vapour pressure compared to the base fluid. The gel propellants need a high pressure to drive them through the feed system and expel them through the nozzle of the injector. They may also be impregnated with certain metallic particles to increase the density impulse. The flow of gel propellants under high pressure through the regenerative passages, with and without metallic additives, is to be investigated. The heat transfer parameters such as heat transfer coefficient, specific heat and thermal conductivity are to be determined under the relevant conditions. Heat transfer parameters of gel propellants relevant to film cooling can also be investigated. The role of heat transfer parameters in ignition and sustainability of the combustible mixture in steady state is to be studied.

The deliverables of project shall include (i) Study of heat transfer parameters in regenerative cooling (ii) Study of heat transfer parameters in film cooling (iii) Characterization of gel propellant droplet combustion in steady state and transient operating conditions.

Evaluation of damage criteria for AA2219 welds under bi-axial stress field from experiments and simulation
This project plans to address the damage modelling and failure prediction of AA2219 GTA welds and parent metal under biaxial state of stress. Both tests and simulations based on FEA are required for this. Different biaxial stress rates have to be addressed. Tests have to be conducted for different specimen thickness commonly encountered in liquid propellant tanks.

C 2.67
007

Modelling of plasma and its interaction in Vacuum chamber during electric thruster firing
Hall thrusters are being used for in-space propulsion functions on spacecraft. The ground experiments of electric thrusters are to predict the behavior of the thruster in space. Exact space conditions simulation in laboratory is a high end task. Though a standard low pressure conditions are set as a scale for operation of thrusters, the interaction of plasma plume with background gas and chamber boundaries is inevitable. In this regard, the effects of back ground pressure and chamber boundaries (facility effects) on the Hall thruster performance attains crucial role in estimating the actual behavior in space. The background pressure increases the neutral collisions and hence plasma plume divergence and the electron transport are affected. Hence the simulations of interaction of plasma plume and effects on electrons transport with background pressure and chamber boundary is needed. The following are expected deliverables:

1. Simulation of facility effects on plasma plume and electron transport using hybrid particle-fluid approach.
2. Study of plume divergence and thrust obtained with facility background pressure.
3. Effects on electron transport from external hollow cathode to discharge channel by facility background pressure.

C 2.73
014

Development of a mathematical model for estimation of crimping loads for different material and design configurations
It is planned to develop material models for metals and non-mechanical materials used for crimping. It will be based on mechanical tests done on the materials at the applicable temperatures (Cryo, room, elevated etc.). These models have to be implemented in FEA codes like ANSYS.

C 2.74
016

Theoretical Investigation of pressure waves generated by heat addition in a gaseous medium
Heat addition in a combustible mixture involves various processes. The first process is heating of the mixture. When the temperature reaches the auto-ignition temperature, the mixture ignites, and combustion occurs resulting in deflagration or detonation depending on whether the flame propagation velocity is subsonic or supersonic respectively. Both these processes can generate pressure waves which should be investigated theoretically to ensure they are not detrimental to the system. Once the mixture reaches the steady state, under certain conditions, combustion instability may set in. This may involve the constructive interaction of the heat release of the flame and the corresponding pressure perturbations which are amplified by the acoustics of the combustion chamber. Alternatively, the heat release fluctuations may generate entropy waves which can subsequently be converted to pressure waves when they are accelerated in a nozzle. The characteristics of pressure waves generated during the thermo-acoustic and entropy-acoustic combustion instability is to be theoretically investigated in a detailed and comprehensive manner.

The deliverables shall include (i) Model of detonation and deflagration waves in a combustible mixture and their transition (ii) Detailed model of thermo-acoustic instability for a rocket combustion chamber and different propellant combinations (iii) Detailed model of entropy-acoustic instability for different combustion conditions and nozzle geometries.

Modelling of two phase flow heat transfer of Liquid Methane in regenerative cooling channels of LOX/Methane rocket engines with Methane film cooling.
In future LOX/Methane engines are essentially needed for Mars missions where in fuel refilling is possible.  A good thermal model for two phase flow will be of much application for design of LOX/Methane engines.

Development of ceramic material with higher electrical insulation at high temperature (LPSC)
Ceramic materials with high electric insulation at high temperature of 18000C are required for Hollow cathode. Hollow cathodes serve as electron source for discharge as well as beam neutralization on electric propulsion thrusters. To get the desired emission currents, emitter material in the cathode has to be heated for high temperatures. An insulation material at 18000C is needed for the purpose. Presently, these ceramic materials have mostly imported. Based on the overwhelming demand of electric propulsion thrusters in the near future, the indigenous material development is essential.
Deliverables:
1.Ceramic materials with high electric insulation at 18000C.

Development of ceramic coating to prevent metal burning in high temperature and oxygen rich environment
Metal burning in hot oxygen environment is an important issue which is not yet solved. Ceramic coating containing oxides is one of the solution for preventing metal burning. Adhesion of the coating to the metal substrate is an important aspect. Addition of metal particles to ceramic will help to improve the ductility of the coating so that it will not fracture under tensile loads.

Development of new thermal barrier coating to reduce heat flux in Semi Cryogenic Engine Thrust Chamber
Thermal barrier coating (TBC) for Semi-cryogenic thrust chamber is mandatory requirement to bring down the coolant channel temperature below the coking limit of the coolant (fuel). Heat flux to thrust chamber material (copper alloy) can be minimized by TBC. TBC can be any material which has got conductivity less chamber material (copper alloy)

Development of coating materials used in high temperature environment
For high temperature application, two kinds of coatings are required . One is TBC for super alloys including Co base or Ni base alloys which has got melting point less than 1400°C. Other is oxidation protection coating  which is needed for refractory alloys like C103 or Mo/Ta alloys

Material behavior at hot Hydrogen environment
Gaseous Hydrogen embrittlement of metals and alloys are possible at ambient to 200°. In order to avoid this, new alloy design which is compatible with hydrogen is required. Existing alloys shall be given coating which is not permeable for hydrogen

Fracture behavior studies of rocket engine materials for cryogenic application
This study plans to address the fracture behavior of Aluminium and titanium alloys used in pressure vessel materials in rocket engines. The behavior of this sections containing part through cracks under tensile stresses is to be explored at ambient and Cryo temperatures through tests and simulations.

Physical property measurement at low temperature up to 20K
Physical properties like modulus, thermal, electrical and expansion/contraction measurement are mandatory requirement for all materials used up to 20K

Development of materials/alloys including coatings for high pressure Oxygen environment
New alloys or coatings for existing materials are required handling high pressure oxygen. Coatings can be of noble metals whose oxides are unstable or ceramic coatings which will be never ignited in high pressure oxygen. New alloys should have very high ignition temperature so that it will ignite in normal operating temperature.

Development of thermal barrier coating with Nano materials
Using Nano materials, TBC can be made with superior mechanical properties as compared to conventional powders. Even with air plasma spraying Nano powdered TBC shows mechanical properties similar EB-PVD

E 5.10
013

Laser Ultrasonic for online EBW evaluation of Ti alloys
Spacecraft propellant tank is a Ti6Al4V alloy construction. The Electron beam welds of thickness 3-4 mm are to be evaluated for porosity, cracks and LOF. The present technique of conventional PA-UT is proposed to be replaced by LASER based ultrasound generation. The research shall be carried out on generation of ultrasound, optimizing the parameters in terms of power, pulse width etc. in order to achieve detectability of LOF of a/2c=0.1, a=0.5 and better and porosities of 0.3 mm or smaller.
The technique shall be in-situ inside the EB chamber with or without vacuum and the set-up shall be portable to be able to move across EB machines. The change in properties of the test article in the ablation regime for ultrasound generation (if considered) shall also be studied.

The deliverables at the end of research and study shall be the LASER source, suitable interferometer and the processing systems.

Development of nano-technology based pressure & strain sensors
Various nano material and nano structures can be used for strain/pressure sensing. Selection of suitable substrate, nano material/structure, synthesis, characterization, packaging etc. are envisaged.  Sensors based on nano technology is to be developed for measuring pressure of propellant and gases in the range 1 to 1000 bar under harsh environment, cryo temperature etc..

Experimental evaluation of damping in fluid conveying pipelines immersed in fluid environment (both theoretical empirical relation &experiments)
Most of the pipelines carrying fluids are immersed in fluid environment. To design these properties, the damping is very important parameter. In order to evaluate the effect of added mass on frequency viscosity on damping, it is envisaged to design a vibration test setup to carry out experiments. By doing experiments it is proposed to evaluate dynamic behavior of pipelines immersed in fluid.

Crack growth studies in propellant tanks through experiments & theoretical modeling
This study plans to address the fracture behavior of Aluminum and titanium alloys used in pressure vessel materials in rocket engines. The behavior of this sections containing part through cracks under tensile stresses is to be explored at ambient and Cryo temperatures through tests and simulations.

Modelling of plasma and its dynamics inside hollow cathode in Hall thruster.
Hollow cathode is one of the most important component of Hall thrusters which is an electron source for plasma discharge and beam neutralization. The life and performance of hollow cathode directly resembles thrusters life and performance. The hollow cathode can be divided into orifice region, insert region and plume region. Plasma density and temperature inside the hollow cathode decides the discharge current that can extract from the cathode. The insert temperature of the cathode is provided by orifice heating, ion heating and electron heating in heater less operation mode. Hence models that quantitatively describe the trends of plasma parameters with varying operating conditions and to simulate discharge parameters (discharge voltage and cathode temperature) is needed.
The following are expected deliverables:

1. Thermal model of Hollow cathode using orifice heating, ion heating and electron heating phenomena.
2. Qualitative description of trends of plasma parameters with varying operating conditions and simulate discharge parameters (discharge voltage and cathode temperature)

Characterization and development of new thermionicmaterial for hollow cathode of electric thruster which cannot easily get poisoned.
Hollow thermionic material is a critical element in the development of cathode for electric propulsion. Cathodes are electron source for plasma generation and neutralization in EP thrusters. Thermionic material made of inorganic refractory material. ie, Lanthanum hexaboride LaB6 with work function of about 2.6  e V emit electron as bulk material without any chemical reaction. Further it is less sensitive to impurities and sir exposure. LaB6 cathode has long life in thruster application because of low evaporation rate. At present this material imported.
Hence indigenization material required with well understanding of surface properties, emission characteristic and temperature profile to operate the cathode for various thruster discharge current.

Expected deliverables are: Material development, Characterization of emitter material for work function, surface properties & mechanical properties.