Computation of Particle Laden Turbulent gas Jet Flows Employing the Stochastic Separated Flow Approach

The dispersion of particles in the presence of Turbulent gas flow is studied theoretically using a stochastic separated flow model and the results compared with the available experimental data. As the particle loading in the jet is of the order of 0.1-0.4 per cent, the particles are assumed to have negligible effect on the mean and the turbulent gas phase properties (one-way coupling). The particle-turbulent eddy interactions are calculated by paying attention to the energy containing eddies, characterised by the integral length scale. The fluctuating velocities are sampled randomly from Gaussian distribution, and the particle trajectories are obtained using a procedure similar to random-walk computation. A large number of particle trajectories are averaged to obtain the statistical nature of the turbulent gas-particle jet. It is seen that the particles with less inertia, which are characterised by the Stokes number, tend to diffuse more. The turbulent diffusivities of the particles are in agreement with the available experimental data, when the time-averaged velocities of gas and particles are the same, obtained by the stochastic separated flow model

Modelling of Ram-Accelerator Flow Fields

Dynamic phenomena in 'ram-accelerator', a ramjet-in-tube concept for accelerating projectiles to ultra high velocities, have been investigated analytically and compared with the experimental investigations reported in open literature. The projectile resembles the centrebody of a conventional ramjet, but travels through a stationary tube filled with a mixture of gaseous fuel and oxidizer. The energy release process travels with a projectile inside the accelerator tube. The characteristics of subsonic combustion, thermally-choked mode of propulsion, which is capable of increasing the velocity up to Chapman-Jouguet (C-J) detonation velocity of the propellant mixture used in ram-accelerator tube, have been studied. The ram-accelerator with a fixed diffuser area ratio operates with different initial velocities for different propellant mixtures. Propellant mixture with CO/sub 2/ as diluent is used for velocity range ~770-1150 m/S; propellant mixture with nitrogen as diluent is used for velocity range ~ 925-1450 m/s and that with helium as diluent is used for velocity range ~ 1500-2000 m/s. Mixtures of propellants with different diluents in varying degree of proportions, giving rise to different acoustic and C-J detonation speeds, have been investigated to evaluate their suitability in the ram-accelerator divided into several segments

Design and Testing of Lab-scale Red Fuming Nitric Acid/Hydroxyl-terminated Polybutadiene Hybrid Rocket Motor for Studying Regression Rate

This paper presents the design of a hybrid rocket motor and the experiments carried out for investigation of hybrid combustion and regression rates for a combination of liquid oxidiser red fuming nitric acid with solid fuel hydroxyl-terminated Polybutadiene. The regression rate is enhanced with the addition of small quantity of solid oxidiser ammonium perchlorate in the fuel. The characteristics of the combustion products were calculated using the NASA CEA Code and were used in a ballistic code developed for predicting the performance of the hybrid rocket motor. A lab-scale motor was designed and the oxidiser mass flow requirements of the hybrid motor for the above combination of fuel and oxidiser have been calculated using the developed ballistic code. A static rocket motor testing facility has been realised for conducting the hybrid experiments. A series of tests were conducted and proper ignition with stable combustion in the hybrid mode has been established. The regression rate correlations were obtained as a function of the oxidiser mass flux and chamber pressure from the experiments for the various combinations.Defence Science Journal, 2011, 61(6), pp.515-522, DOI:http://dx.doi.org/10.14429/dsj.61.87

Hybrid Rocket Technology

With their unique operational characteristics, hybrid rockets can potentially provide safer, lower-cost avenues for spacecraft and missiles than the current solid propellant and liquid propellant systems. Classical hybrids can be throttled for thrust tailoring, perform in-flight motor shutdown and restart. In classical hybrids, the fuel is stored in the form of a solid grain, requiring only half the feed system hardware of liquid bipropellant engines. The commonly used fuels are benign, nontoxic, and not hazardous to store and transport. Solid fuel grains are not highly susceptible to cracks, imperfections, and environmental temperature and are therefore safer to manufacture, store, transport, and use for launch. The status of development based on the experience of the last few decades indicating the maturity of the hybrid rocket technology is given in brief.Defence Science Journal, 2011, 61(3), pp.193-200, DOI:http://dx.doi.org/10.14429/dsj.61.51

Evaluation of Kerosene Fuelled Scramjet Combustor using a Combination of Cooled and Uncooled Struts

The scramjet combustor a vital component of scramjet engine has been designed by employing fuel injection struts. Several experimental studies have been carried out to evaluate the propulsive performance and structural integrity of the in-stream fuel injection struts in the connect-pipe test facility. As the mission objective of hypersonic demonstrator is to flight test the scramjet engine for 20 s duration, in-stream fuel injection struts which are designed as heat sink devices encounter hostile flow field conditions especially in terms of high thermal and high convective loads in the scramjet combustor. To circumvent these adverse conditions, materials like Niobium C-103 and W-Ni-Fe alloys have been used for the construction of struts and a number of tests have been carried out to evaluate the survivability of the in-stream fuel injection struts in the scramjet combustor. The results thus obtained show that the erosion of leading edges of the Stage-II fuel injection struts in the initial phase and subsequently puncturing of the fuel injection manifold after 10-12 s of the test are noticed, while the other stages of the struts are found to be intact. This deteriorating leading edges of Stage-II struts with respect to time, affect the overall propulsive performance of the combustor. To mitigate this situation, Stage-II struts have been designed as cooled structure and other Stages of struts are designed as un-cooled structure. Material of construction of struts used is Nimonic C-263 alloy. This paper highlights the results of the static test of the scramjet combustor, which has been carried out at a combustor entry Mach number of 2.0, total temperature of 2000 K, with an overall kerosene fuel equivalence ratio of 1.0 and for the supersonic combustion duration of 20 s. Low back pressure has been created at the exit of the scramjet combustor using ejector system to avoid flow separation.Visual inspection of the fuel injection struts after the test revealed that all the Struts are found to be thermo-structurally safe in the combustor environment except for minor erosion of the leading edges of the struts. Stage-II struts made of two-passage cooled configuration are found to be thermo-structurally safe. Although other stages of struts used in the test are of un-cooled configuration, they too are found to be safe and intact. This demonstrates the fact that they experience thermally benign flow conditions compared to Stage-II struts in the scramjet combustor.Defence Science Journal, 2014, 64(1),  DOI:http://dx.doi.org/10.14429/dsj.64.273

Experimental Investigations of Hydrocarbon Fueled Scramjet Combustor by Employing High Temperature Materials for the Construction of Fuel Injection Struts

ABSTRACT For the Hypersonic Technology Demonstrator Vehicle (HSTDV) programme half-width strut based scramjet combustor has been designed, developed and tested for the short durations (5 s) as well as for the long durations (20 s) using various materials for the construction of fuel injection struts. Extensive experimental investigations have been carried out to identify suitable material for the long duration (20 s) tests. Niobium C-103 alloy and W-Ni-Fe alloy materials have been used for the construction of fuel injection struts and they have been employed in two different tests. In the first test struts made of Niobium alloy is used and in the second test struts made of W-Ni-Fe alloy is used. It is inferred from the results of the static tests for the 20 s test duration that the leading edges of the struts are eroding due to high thermal load, shear force and oxidizing environments in the five-strut scramjet combustor configuration. The failure of the struts is noticed in the Stage-II injection of the scramjet combustor. The thermo-structural failure of the stage-II fuel injection struts in the scramjet combustor in both the tests has detrimental effect on the performance of the combustor. In the case of Niobium C-103 alloy struts, erosion of the leading edges is found to be severe compared to W-Ni-Fe alloy struts. Hence, the total pressure loss in the former is found to be more compared to the latter. In the first test (Niobium struts used) the flow separation is occurring earlier compared to the second test (W-Ni-Fe struts employed). This is indicative of the onset of the severe leading edges erosion of Niobium C-103 alloy struts compared to W-Ni-Fe alloy struts resulted in more skin friction drag and hence the flow separation at a shorter length. Struts made of W-Ni-Fe alloy seem to be promising candidate material compared to Niobium C-103 alloy. Subsequent tests carried out by employing struts made of W-Ni-Fe alloy divulged that the powder metallurgy route to realise the W-Ni-Fe alloy plate is unable to deliver/impart consistent mechanical properties in all the directions of the plate i.e., anisotropy is prevailing. On this front, it is found that the material developed at this juncture is found to be unsuitable for the scramjet application. To circumvent such scenario two strategies have been proposed for the realization of fuel injection elements. Hypersonic, Scramjet, Strut, NOMENCLATURE H = height of the combustor P = pressure T = temperature = equivalence ratio SUBSCRIPTS f = fuel i = inlet t = stagnation condition w = wall wd = wedge Keywords

Supersonic Jet Interactions in a Plenum Chamber

Understanding thè supersonic jet interactions in a plenum chamber is essential for thè design of hot launch systems. Static tests were conducted in a small-scale rocket motor ioaded with a typical nitramine propellaiit to produce a nozzle exit Mach number of 3. This supersonic jet is made to interact with plenum chambers having both open and closed sides. The distance between thè nozzle exit and thè back piate of plenum chamber are varied from 2. 5 to 7. 0 times thè nozzle exit diameter. The pressure rise in thè plenum chamber was measured using pressure transducers mounted at different locatìons. The pressure-time data were analysed to obtain an insight into thè flow field in thè plenum chamber. The maximum pressure exerted on thè back piate of plenum chamber is about 25-35 per cent. of thè maximum stagnation pressure developed in thè rocket motor. Ten static tests were carried out to obtain thè effect of axial distance between thè nozzle exit and thè plenum chamber back piate, and stagnation pressure in thè rocket motoron thè flow field in thè open-sided and closed-sided plenum chambers configurations

Development and Demonstration of Control Strategies for a Common Rail Direct Injection Armoured Fighting Vehicle Engine

The development of a controller which can be used for engines used in armoured fighting vehicles is discussed. This involved choosing a state of the art reference common rail automotive Diesel engine and setting-up of a transient engine testing facility. The dynamometer through special real-time software was controlled to vary the engine speed and throttle position. The reference engine was first tested with its stock ECU and its bounds of operation were identified. Several software modules were developed in-house in stages and evaluated on special test benches before being integrated and tested on the reference engine. Complete engine control software was thus developed in Simulink and flashed on to an open engine controller which was then interfaced with the engine. The developed control software includes strategies for closed loop control of fuel rail pressure, boost pressure, idle speed, coolant temperature based engine de-rating, control of fuel injection timing, duration and number of injections per cycle based on engine speed and driver input. The developed control algorithms also facilitated online calibration of engine maps and manual over-ride and control of engine parameters whenever required. The software was further tuned under transient conditions on the actual engine for close control of various parameters including rail pressure, idling speed and boost pressure. Finally, the developed control strategies were successfully demonstrated and validated on the reference engine being loaded on customised transient cycles on the transient engine testing facility with inputs based on military driving conditions. The developed controller can be scaled up for armoured fighting vehicle engines

Analysis of Rotating Detonation Wave Engine

274-282Rotating Detonation Wave Engine (RDE) due to its promising potential as a propulsive and power generation device has been researched worldwide based on both numerical and experimental investigations. The thermodynamic analysis has been of importance prior to the commencement of the experimental investigations as the set conditions could be established with ease. The flow field behind the detonation wave has been quite complex due to oblique shock wave, contact surface between combustion products of detonation wave and shocked combustion products and the expansion waves. The simultaneous establishment of the flow parameters has been of importance to the success of understanding the RDE. The enthalpy values at different states have provided the energy conversion to kinetic energy as a result of expansion of the product gases in the RDE flow field. Stability of the oblique shock wave attached to the detonation wave has been crucial for obtaining optimum performance of RDE. The intersection of oblique shock polar and the Prandtl – Meyer expansion characteristics has given the conditions under which the oblique shock remains attached to the detonation wave and be a part of the triple point. Under all the set conditions, the stability of the oblique shock has been ascertained. In the present analysis, the specific thrust for the present configuration using H2–air is 1374 Ns/kg compared to a value of 1347 Ns/kg reported in the literature for a stoichiometric composition. The marginal difference has been due to the different input conditions ahead of the detonation wave. This has given credence to the results of the analytical work based on gas dynamic and thermodynamic relationships. The practical implications of this analytical work have been brought out