Proceedings of the National Aerospace Propulsion Conference

Gas turbine engines are widely used in the aviation industries because of high power to weight ratio. The gas turbine engine consists of the three main components: compressor, combustion chamber, and turbine. The design of the axial flow compressor depends on various design parameters such as blade stagger angle, blade height, blade chord, rotor tip clearance, and stator tip clearance. Blade stagger angle is a critical parameter which plays vital role in performance of the gas turbine engines. In this study, a steady state analysis has been carried out to understand the effect of rotor and stator blade stagger angle on the aerodynamic performance of single stage axial flow transonic compressor through three dimensional viscous analysis using ANSYS CFX 19.2 software. The analysis was carried out for various stagger angle configuration varying from 24° to 36° for rotor and 06° to 18° for stator. The study concluded that as the rotor stagger angle is increased from 24° to 36°, the mass flow is reduced by 16.5%, peak pressure ratio by 2.8%, and increase peak efficiency by 6.8%. The effect of change of stator stagger angle from 06° to 18° is insignificant.

In the quest of maximizing the pressure ratio and the efficiency of the gas turbine engine, designing a compressor that can generate the required pressure ratio with a minimum number of stages is one of the challenges. A single blade, if designed with a higher diffusion factor, has an inherent problem of flow separation. Tandem or slotted blades with the help of nozzle-gap phenomenon have shown promising results in terms of higher pressure ratio and efficiency, but it has mechanical complexity and lower stall margin. The hybrid rotor called part-span tandem has been designed to get the benefits of both a single rotor as well as tandem rotor blades. Design methodology as well as parallel comparison of results of a low speed hybrid rotor, tandem rotor, and a single rotor is included in the paper. CFD results such as pressure distribution at different span locations, total pressure rise, static pressure rise, contours of Mach number, entropy, the behavior of the tip leakage flow, and performance curve are included in this paper. Full span tandem rotor has higher pressure rise, and the hybrid rotor has a better stall characteristic.

The present computational study focusses on the flow physics analysis of an aero-derivative inter-compressor duct with struts along with the low-pressure (LP) compressor for a generic small turbofan engine. The transition duct connects the transonic fan to LP compressor. For numerical analysis, four inflows under consideration are Inflow-A (uniform inflow), Inflow-B (fan exit condition with uniform flow at fan inlet), Inflow-C (fan exit condition with radial hub distortion at fan inlet) and Inflow-D (fan exit condition with circumferential distortion at fan inlet). A pair of streamwise counter-rotating pair of vortices forms at the duct-struts-hub corner region for Inflow-A. But, for all other inflows, the vortex formation is deflected on the leeward side of the strut owing to the tangential component generated by the upstream stage. Besides, the exit conditions of the duct are imposed on the LP compressor to estimate its performance and the phenomena that lead to stall. For each inflow, the flow separation at the tip side of the stator leads to compressor stall. In general, the paper briefly elucidates how different inflows ingested by the transonic fan affect the duct and subsequently the LP compressor in a gas-turbine engine.

Aero-engine compressor designers have a challenging task of developing compressors that have a higher-pressure ratio and better efficiency with a lower number of stages. Such designs would require blades with high diffusion factor and hence the inherent risk of flow separation. Tandem blade is an interesting concept, which possibly addresses this problem. In tandem blading, the forward blade and the aft blades are arranged in such a manner that a converging nozzle flow path is created between the two blades. The flow accelerates through this nozzle, energizes the suction surface flow, and thereby prevents the early onset of flow separation. This paper presents the steady computational analysis of a tandem rotor stage and baseline stage in low-speed axial flow compressor at design and off-design condition using ANSYS CFX. The study is further extended to analyze the effect of radial distortion on the performance of tandem rotor and the single rotor.

Compressor test rigs are commonly used for evaluating the performance of axial and centrifugal compressors at different operating conditions. Axial flow compressors are tested in rigs by driving the test compressor and throttling the exhaust flow, thus creating a higher back pressure or loading to the compressor. Collector box is an important unit which collects the exhaust air from compressor outlet and directs the flow to throttling valves present at downstream location. It plays a major role in compressor testing by providing intended aerodynamic loading at the exit of compressor. This paper discusses in detail about the design of collector box for a given compressor configuration and the influence of various geometrical parameters on compressor loading. A baseline collector box configuration was studied and compared with other similar configurations. Static pressure distortion and total pressure loss coefficient were considered as major design parameters. 3D RANS analysis was carried out to estimate the performance of collector box. Further, the effect of collector box length, outlet configuration, and outlet duct area on collector box performance is studied. Greitzer ‘B’ parameter is evaluated for all cases to estimate the dynamic instability effects of flow inside the ducting system. Finally, a new configuration with variable outlet duct volume was designed which gave better static pressure uniformity at the compressor exit.

Small gas turbines have been used in many unmanned aerial vehicle (UAV) applications. For UAV, the design of gas turbine engine is driven by its simplicity, cost-effectiveness and reliability, and moderate fuel efficiency. This paper describes the aerodynamic design and salient features of a single-stage axial turbine designed for powerplant of UAV application. Profile, secondary and tip clearance losses are dominant due to reduced blade heights, Reynolds number, and manufacturing constraints. The transonic axial turbine is designed with low-aspect ratio and moderate stage loading. The design point pressure ratio is 2.3, and the target efficiency is 86%.

Scalloping of turbine wheel blisk of a radial turbine is done to reduce the centrifugal stress on the disc portion of the blisk, reduce weight and to decrease turbo lag in turbochargers. An attempt has been made to model the effect of scalloping of the turbine wheel on the performance of the radial turbine. The scallops have been modelled as hub clearances and analyses with various depths of scalloping and hub clearances have been analyzed using commercial 3D NS solver. The decrease in efficiency with increased scalloping and hub clearances has been plotted, and the results have been qualitatively compared with experimental results of similar but different radial turbines with and without scalloping. The results of the analyses are found to follow the same trend as those of the experiments.

The gas turbine engine performance can be improved in off-design condition by using variable area nozzle turbine (VANT) concept as mass flow rate varies in off-design conditions. In order to analyze the flow field through VANT, second-stage LPT nozzle geometry was selected from EEE proposed by Pratt and Whitney, and its endwalls were modified. In the present study, meshing of the nozzle domain was done in ICEM CFD® due to the limitation of TurboGrid® which is discussed. For the numerical analysis, two-passage and one-passage fluid models were analyzed. It was found from two-passage analysis that due to the 3D geometry of the vane at LE, incidence effect was not properly captured. Hence, flow physics is not identical in two passages. Further, one-passage numerical analysis with and without varying the periodic surfaces at −5°, 0°, and + 5° vane setting angle is performed. The results were analyzed in terms of static pressure field in the part clearances and total pressure loss coefficient. It is found that mass flow averaged total pressure loss coefficient changes within 3% range with change of different periodic surface concepts.

The function of the turbine aerofoil is to turn the flow to a certain angle and provide smooth flow acceleration to a certain Mach number with minimum pressure loss. The flow field in the blade passage is complex, unsteady, transonic, and three dimensional. So, to aerodynamic designer, it is necessary to evaluate the profile losses and other aerodynamic performance parameters to improve the efficiency of the turbine stage. In the present study, CFD simulation has been carried out for the high-pressure turbine nozzle guide vane (NGV) profiles for design and off-design conditions. Parameters such as profile loss coefficient, surface Mach number distribution, and flow deflections are investigated. The results obtained from numerical analysis are compared with experimental data from transonic cascade tunnel.

Compact high speed turbomachines are complicated in aerodynamic design, mechanical construction, and fabrication. Design choices made during the aerodynamic design are strongly coupled to the mechanical integrity due to high rotational speeds and thermal gradients. In the present paper, aerodynamic design of a turbine with a low pressure ratio of 2:1 is presented. The turbine designed is a compact single stage machine, intended to be used in a 1 kN thrust small gas turbine engine. For the present work, the flow path is a constant hub and shroud radius rotor design, with a hub-tip ratio of 0.72. Overall guidelines used for parameter selection in velocity triangles, mean-line design and blade geometry are discussed. The turbine is designed for a mean section reaction of about 25%. Performance evaluation and flowfield analysis for the turbine geometries are carried out using a three-dimensional RANS solver. Performance characteristics of the turbine are generated for a range of pressure ratio at design speed. The intended efficiency of the turbine stage at design point is 88%. Challenges during the design process to obtain blade geometries with wide blade passage throat are put forth. The choices made in aerodynamic design which affect the ease of machining of the stator and rotor components are also brought out.

Performance evaluation of gas turbine engine on the test bed is of prime importance during the engine developmental stage. It is prudent to study the rotor dynamics, lube system, fuel system, vibratory characteristics of the engine along with its performance on the test bed. Micro Gas Turbine (MGT) Engine of 50 N thrust is being under development at CSIR-NAL. To carry out performance tests of MGT, a test bench has been developed. The test bench is equipped with (i) Lubrication system, (ii) Fuel system, (iii) Starting system, (iv) Data acquisition and measurement system. The test bed is designed to measure thrust directly using movable bed and a ‘S’ type load cell of 100 N capacity. A miniature fuel pump is used to supply fuel to engine. Air-oil lubrication system is used to lubricate the engine bearings. External compressed air supply system is developed to assist the starting of engine. DAQ and instrumentation layout has been designed and developed specific to the existing module. Test bench is configured so as to cater the measurement and online condition monitoring requirements with required sensors to measure thrust, temperature, speed and vibration. Sensors are integrated to the National Instrument data acquisition hardware. DAQ system software is developed on the LabVIEW platform.

In this paper, strategy for development of high-speed rig for performance evaluation of gas turbine driven alternator is discussed. In aeronautical systems, only high-speed alternator is the power generating unit, which is directly or through gear train coupled with variable speed gas turbine engine. Performance evaluation of this type of alternator needs to be conducted on bench to validate its rating. Development of high-speed (30,000–60,000 rpm) rig is really demanding to meet this performance evaluation requirement. Apprehension and challenges for development of high-speed rig, instrumentations, and its rig characterizations are presented in details.

With the evolving technology and extensive software usage, aircrafts have become a software embedded flying contrivance. Navigation system, landing gear system and propulsion system are some of the major subsystems of the aircraft. Propulsion system is one of the vital sub-systems with demarcated purpose to propel the aircraft. Earlier, the control unit of the engine was completely controlled by mechanical means but with the technical advancements it has been automated by software embedded control unit. Software has become so important these days that its safety, complications and risks cannot be ignored. The embedded software in digital engine control unit is a safety critical software as its failure can lead to hazardous state that can cause loss of property, damage to environment and even loss of human life. Therefore, intensive care needs to be taken while ensuring the safety and reliability of such software. The traditional testing approach needs to be fortified with more firm and rigid standardized methodology in order to diminish the probability of failure of the system. This paper throws light on the Independent Verification and Validation process followed to ensure safety, reliability and robustness of aero engine propulsion system software.

Design of mechanical structures for aeronautical applications is mainly aimed towards maximizing the strength-to-weight ratio and alleviating the vibratory response to the maximum extent so that the stresses are well below the endurance limit. The present paper aims to bring out features that fulfil these requirements. An innovative approach is the introduction of a chiral lattice in the structure that can enhance the damping in structure through intentional deformations in the chiral webs and that can also be tuned to act as vibration absorbers to reduce the overall vibratory response of the structure. The design features suggested can be successfully implemented in components like aerofoils, engine carcass, blisks and disks. This is possible due to the advent of 3D printing technology. The paper aims to bring out these aspects by considering a simple cantilever beam with a chiral lattice internal to it. Further parametrization of geometric designs is undertaken to understand the vibratory response characteristics.

With the increase in demand of light and durable material, composite materials are replacing conventional materials in aerospace industries. Efficiency of aeroengine can be increased by increasing turbine entry temperature of an engine. Composites are replacing conventional materials for low- and high-temperature applications due to their high-specific strength. It consists of reinforce fiber along with suitable matrix. Unlike monolithic material, the mechanical behavior of composite depends on the properties of its main constitutes, i.e., fiber and matrix. This paper uses micro and mesoscale modeling techniques to develop fatigue behavior of composite material. A novel technique has been developed to estimate fatigue life of composite components using progressive damage by propagation of crack through matrix and fiber under cyclic loading.

Mechanical and electrical designs for a micro-pulsed plasma thruster have been presented here, which is to be accommodated in the KIIT University’s Nanosatellite for technological demonstration in space. Design-related calculations based on dimensional limitations, power constraints, and optimal thruster performance have been done and laid out. Mechanical design includes accurate computer-aided design (CAD) and physical calculations of thruster casing, miniature spark plug, and electrodes. The electrical design includes the printed circuit board (PCB) design and circuit-related calculations. The thruster being considered here uses a flared electrode and casing in order to achieve optimal thruster performance. Thrusters are well known for their issues of erosion which hamper the thruster lifetime, therefore, the correct choice of materials for electrode, casing, and spark plug has been considered after a wide survey in order to limit the erosion as much as possible. Analysis of both the mechanical and electrical system has been done using SOLIDWORKS and PROTEUS software, respectively, for critical validation. A space grade micro-pulsed plasma thruster weighing only 130 g capable of surviving loads of a polar satellite launch vehicle (PSLV) by Indian Space Research Organization (ISRO) and providing a specific impulse of 600 s has been concluded.

Present study deals with the study on the effect of Nytrox oxidizer which is the combination of both gaseous oxygen and gaseous nitrous oxide with the wax as a fuel on regression rate, combustion efficiency, and sliver loss. The oxygen was injected through the axial or showerhead injector, and for the nitrous oxide, two different injectors were used, i.e., radial and swirl injector. The study had been carried out with three different combinations of the N2O and oxygen, i.e., 50:50, 70:30, 80:20 ratio. It was further compared with the GOX and N2O. It has been observed that the pure N2O gave lower regression rate compared to the oxygen unless swirl injector was used with N2O. Among the injectors used, most effective in enhancing performance was swirl, then the radial and last was the showerhead injector. Among the combination of all the three oxidizers used, the 50:50 ratio was the most promising in all respect of performance improvement.

In the present study, attempts were made to investigate the effect of the protrusion configuration on the performance and combustion stability behavior of the hybrid rocket motor. The protrusion configurations used for study include circular, star, multi-hole and fin. All the protrusion configurations were tested by placing at 0.7X/L location in a hybrid rocket motor. The wax and gaseous oxygen were used as a fuel and oxidizer, respectively. The regression rate and combustion efficiency were improved upon inserting the protrusion in hybrid rocket motor. Among the protrusion configurations, multi-hole case showed the highest improvement in regression rate and combustion efficiency. The hybrid rocket with/without protrusion case exhibited the first and second longitudinal acoustic modes which are dominant throughout the combustion. Upon inserting the protrusion, the stability of the combustion was improved significantly. Among all the protrusion cases, the star case exhibited the higher stable combustion. The frequency range of longitudinal acoustic modes was changing while changing the protrusion configuration.

The base flow of a truncated plug nozzle exhibits a phenomenon called Wake Structure Transition (WST), where in an open base wake develops into a closed wake as the pressure ratio is increased. From design point of view, it is important to predict this transition. There are empirical models available for predicting WST but they are inaccurate for routine use in design. One physics-based model proposed by CAd lab of IISc, shows a potential to do the same with a great accuracy. However, its utility is limited by the requirement that accurate base pressure at two operating conditions, one in open and other in closed wake are to be known. The inherent limitations of RANS (or URANS)-based CFD solvers in predicting base pressure for large separated flows, limits the utility of aforementioned models. The present attempt explores the use of high-fidelity DES tool for the simulation of base flow and thereby WST prediction. The preliminary results obtained are very encouraging both from the view point of accuracy of the predicted base flow and its use in the physics-based model for predicting WST.

Experimental and computational analysis has been already carried out by many researchers on supersonic flow past cavities, but detailed analysis of computational results still needs some insight. For this purpose, an open rectangular cavity with a length to depth ratio of 2 ( $$L/D = 2$$ L / D = 2 ) and inlet Mach number 1.71 was considered for an unsteady computational analysis in ANSYS FLUENT, using SST $$k-\omega $$ k - ω turbulence model. The two dimensional structured grids were generated in Pointwise grid generation software. FFT using Power Spectral Density (PSD) was carried out on the unsteady pressure data for 10,000 time-steps, with a total flow time of 10 ms. Many modes were observed, with dominant frequency at 10.5 kHz. The mode frequencies obtained were validated from experimental results and from the corresponding Rossiter’s Modes. Correlation between the unsteady pressure data was also found to analyze the flow dynamics. Many flow visualization techniques were employed such as density gradient-based numerical schlieren, which revealed many flow features associated with the flow. Vortex Shedding Visualization was carried out in terms of the lambda 2 criterion, where the vortex core ( $$\lambda _2 < 0$$ λ 2 < 0 ) can be observed moving downstream in the shear layer. Lastly in the acoustic pressure contour, an acoustic wave can be observed moving within the cavity. The analysis was extended for different shapes of subcavities on the front and aft wall. As the front wall subcavity act as a passive control device, reducing the overall sound pressure level inside the cavity, whereas the aft wall subcavity acts as a passive resonator with distinct harmonic fluid-resonant modes. A more detailed analysis on these configurations with different shapes will give a comparative and better understanding on the flow features, mode frequencies, Rossiter’s coefficients, and fluid-resonant oscillations in a supersonic cavity. Also, the applicability of Rossiter’s Modes has been compared with the Closed-Box acoustic model for different configurations.

Blockage of strut injector in a supersonic blowdown wind tunnel was studied numerically. Blockage of a model depends mainly on model cross-sectional area, Mach number, shockwave boundary layer interaction, displacement of the boundary layer, and total pressure loss. Simulations were carried out for Mach number of 2 for different chamber pressures while keeping all the other parameters fixed. The blockage phenomenon was simulated using commercial software ANSYS Fluent with SST K- $$\omega $$ ω turbulence model. Results showed that increasing the chamber pressure did not help in avoiding blockage. Flow field was also not affected by the change in chamber pressure. Due to the influence of the boundary layer, blockage had appeared for all the cases even though the total pressure recovery was better than the theoretical value.

Results of numerical investigation of flow and performance characteristics of a rectangular mixed-compression intake with boundary-layer bleed are presented. A flush slot configuration applied at the intake throat has been studied to evaluate its effect on the compression performance. The viscous flowfield has been obtained by solving Favre-averaged Navier–Stokes equations with SST k-ω turbulence model. The analysis has been carried out at a range of freestream Mach numbers of 1.7–2.8 and 0° angle-of-attack, corresponding to a unit Reynolds number of approximately 1.8–3.0 × 107 m−1. The results show that with the use of bleed, the low energy boundary layer and the secondary cross flows are successfully removed near the throat as well as the terminal shock gets stabilized at its entrance. Consequently, the exit flowfield uniformity as well as the critical and peak total pressure recovery improves considerably compared to the intake without bleed system. Furthermore, the critical area-averaged intake exit Mach number reduces by about 0.04 in the complete operating regime. As well as the overall mass capture is found to increase at all operating conditions prior to “unstart because of SWBLI”, since the throat separation bubble is reduced and hence the cowl shock is pushed downstream.

The new detonation-based combustors for hypersonic and supersonic missiles could reduce the launch mass (up to 3–4 times) and body length (up to 2 times) of such systems considerably and are most sought after by researchers worldwide. Near future practical detonation-based engines will be based on liquid hydrocarbon fuels. A promising synthetic jet fuel used in many military applications is JP-10. We are interested in improving the performance of JP-10 by using ignition promoters like O3 and H2O2 in trace amounts for applications in detonation-based hypersonic and supersonic missiles. The effect of these promoters is to enhance combustion rates and energy release rates simultaneously. Stable propagation of detonation waves near its limits without the danger of attenuation or failure for continuous operation of detonation-based combustors is an ongoing problem of interest for military and air force researchers worldwide.

Rotating detonation wave engine (RDE) would be the futuristic engine for air-breathing missile systems and gas turbine systems (aero and stationary applications). The continuous operation of RDE with an operating frequency of 3–15 kHz is attractive for propulsion systems based on rocket, ramjet and turbojet engines. The cell size is the characteristic dimension of the cellular pattern of a propagating detonation wave. The cell size is found to depend strongly on the choice of fuel and oxidizer, its equivalence ratio, initial temperature and initial pressure. Some empirical relations based on cell size are used to design the present detonation combustor. Hydrogen is chosen as the fuel, while air is selected as the oxidizer. The RDE hardware has been realized, and the test facility is being modified at NCCRD, IIT, Madras to carry out static tests. The analysis of the RDE combustor without the nozzle is carried out using the “pressure history model” reported in literature. As the hydrogen fuel and air are entering as two different streams perpendicular to each other, a simple mixing analysis has been carried out to evaluate the mixture properties ahead of the detonation wave. The Chapman–Jouguet (CJ) detonation computations are carried out using the detonation tool box runs in conjunction with Cantera software assuming chemical equilibrium. The fuel-based specific impulse resulted from the present analysis for our configuration using H2-air is 4733 s compared to a value of 4706 s reported in literature for a stoichiometric composition. The same model has been used to evaluate the C2H4-air system. The specific impulse of our study is 2111 s compared to 1975s reported in literature for the fuel-based equivalence ratio of 0.5. This has given credence to the results of the analytical work.

In a rotating detonation wave engine (RDE), a unidirectional detonation wave could be created and the exhaust gases are expanded through an annular plug nozzle producing thrust. Empirical relations reported in open literature based on detonation cell size are used to design the combustor. Hydrogen is chosen as the fuel, while air is selected as the oxidizer. The RDE hardware has been realized, and the test facility is being modified at National Combustion Centre for Research and Development (NCCRD), IIT Madras, to carry out static tests. In order to obtain necessary increment in thrust for propulsion applications, a plug nozzle is designed based on simple wave theory under stoichiometric condition. The inlet conditions to the nozzle are established based on “axial flow model” of RDE reported in open literature. As the hydrogen and air are entering as two different streams perpendicular to each other, a simple mixing analysis is carried out to evaluate the mixture properties ahead of the detonation wave. The Chapman Jouguet (CJ) detonation computations are carried out using the shock and detonation toolbox runs in conjunction with Cantera software assuming chemical equilibrium. The modelling of the flow field downstream of the detonation wave is established using the solution of integral mass, momentum and energy equations written for the streamline flow from detonation wave to the exit of the combustor. These conditions are used to evaluate the propulsion parameters at different fuel-based equivalence ratios (0.7–1.3) as a result of expansion through the plug nozzle. The increment in fuel-based specific impulse resulted from the present analysis for our configuration using H2–air is 18% due to the presence of plug nozzle for the stoichiometric composition at a nozzle entry stagnation pressure of 6.9 bar. The fuel-based specific impulse based on the “axial flow model” reported in the literature for the stoichiometric hydrogen–air mixture at the combustor exit static pressure of 1 bar without plug nozzle is 5383 s. For the present combustor at the same condition, it is 5474 s, which appears to be close while modelling the complex processes using simplified model equations. Several input conditions and combustor–plug nozzle combined performance parameters would be utilized for setting the conditions for the experiments.

Electric propulsion is rapidly entering the aviation sector. The main drivers are lower carbon footprint and lower noise emission. The present work examines the feasibility of hybrid propulsion for two-seater light aircraft, such as the Hansa 3 aircraft of CSIR-NAL. Both parallel and series arrangements of electric motors and internal combustion engines are considered. A general-purpose code is developed to determine the aircraft performance for different flight profiles and specific combination of power plant units. The state-of- the-art rotary (Wankel) engines are considered. Electrical hardware includes state-of-the-art motors and currently available lithium-ion batteries. A comparison is made between specific case of all-electric propulsion with various hybrid-electric propulsion system (HEPS). It is found that significant improvement in the performance of hybrid combinations with respect to all-electric case is observed. The work suggests that it is now an opportune time to enter the electric propulsion through the hybrid route.

Flame-synthesized carbon nanoparticles of varying sizes were produced in premixed stretch-stabilized stagnation ethylene-oxygen-nitrogen flames under various sooting conditions. The experimental setup consists of a burner with an aerodynamically-shaped nozzle, a stagnation surface/sampling probe assembly, and a scanning mobility particle sizer. The pseudo-one-dimensional formulation was invoked to simplify the stagnation flow field. A modified version of the OPPDIF code was used to compute the gas-phase species, temperature, and velocity profiles using USC Mech II as the reaction kinetics model, which consists of 111 species and 784 reactions. Thermophoretic velocities and particle residence times were calculated for each flame configuration. Soot volume fractions and particle size distribution functions were measured in a series of five atmospheric pressure ethylene-oxygen-nitrogen flames with maximum temperatures ranging from 1969 to 2132 K. The UV–Visible absorption spectral analysis was conducted in the 190–1400 nm range for the flame-synthesized nanoparticles to evaluate the optical bandgap from their resulting Tauc spectra. The results from the present study suggest a strong dependence of the optical bandgap on particle size. It is shown that quantum confinement effects, which arise due to the finite size of the particles, play an essential role in determining the final absolute value of the optical bandgap. In the present study, flame-synthesized carbon nanoparticles (CNPs) are found to exhibit quantum confinement behaviors.

Self-sustained propagation of detonation waves near limits is essential for the successful operation of detonation-based combustors since they suffer from high-velocity deficits near limits due to geometric constraints. This can potentially lead to its failure or attenuation near limits. The failure or attenuation of a detonation wave under such circumstances could lead to the failure of a detonation-based engine altogether. Existing models like Fay’s model reasonably predict detonation velocity deficits for only stable mixtures. The present work focuses on estimating velocity deficits for both stable and unstable mixtures. The proposed model is similar to Fay’s model with the modified reaction zone thickness calculated using $$x = c\left( {\Delta_{i} + \Delta_{r} } \right)$$ x = c Δ i + Δ r . The value of c is found to be 33.2, 8.6, and 19.5 for H2–air, CH4–O2 (unstable mixtures), and H2−O2−Ar mixtures (stable mixture) using existing experimental data. The proposed model predicts velocity deficits better than other existing models for both stable and unstable mixtures over a range of pressure ratios and tube diameters and also near the limits. The addition of O3 and H2O2 at modest concentrations was shown to reduce the velocity deficits near propagation limits. The present work shows that the use of ignition promoters in trace amounts could help in the widening of detonation limits for detonation-based combustors.

Thrust augmentation of military aircraft engine is carried out during critical phases of flight with the help of an afterburner. In advanced aeroengines, afterburner entry temperatures are beyond material allowable temperature limits. Thus, it is vital to cool V-gutter even when afterburner is not in operation, to lower its metal temperatures and to achieve increased creep life. Present study is preliminary design of three cooling configurations for a given V-gutter geometry at design condition of operation of afterburner. 1D code is developed to arrive at geometry of each configuration after a parametric analysis. Effect of thermal barrier coating on the V-gutter metal temperature is studied. 3D conjugate heat transfer analysis (CHT) is carried out using FloEFD commercial software to study the local coolant flow field and its effect on metal temperatures.

The main objective of the present work is to study the effect of lip thickness and the lip cross section on the fluid flow behavior and effectiveness in tangential film cooling in a combustor liner. A three-dimensional jet exiting from a circular hole of diameter 2 mm is considered as coolant and the hot mainstream as a co-flow. The coolant and the mainstream are separated by a thick slab called the lip. A three-dimensional, steady-state numerical study is performed for different lip thicknesses (t) ranging from 0.25D to 2.5D and a blowing ratio of 0.5, 0.9, 1.5, and 2.5. Results are presented for a dimensionless stream-wise distance between 0 and 35. At all blowing ratios, a lip thickness of 0.25D (thin lip) performs better than the other lip thicknesses under consideration due to the attached flow. Additionally, it is observed that a change in the cross section (tapered) of the lip shows improvement in the film effectiveness, which may be attributed to the Coanda effect.

This paper investigates the influence of concentration of boron nano-particles on the combustion behavior of Jet A1-based slurry droplets especially for high-particle loadings. The nanofuel was prepared using commercially available boron nano-particles. The particle loading ratio (by weight) was varied from 2.5 to 50%. Shadowgraphic visualization of the burning droplet and direct flame imaging was carried out to examine the burning characteristics of horizontally suspended slurry droplet. While for particle loading up to 20%, the droplet combustion was dominated by vigorous oscillations and micro-explosions, beyond 20% loading, significant differences in the combustion behavior were observed which are discussed. The normalized surface area plot supported the observations and depicted clear departure from classical model for droplet burning.

Interaction of hydrocarbon fuel drops with high-temperature substrates is commonly encountered in combustion chambers, and such dynamics are relevant in design and optimization of fuel spray systems. Dynamics of fuel drops impacting on a heated solid surface in the splashing regime is investigated through high-speed imaging experiments. The effect of surface temperature in altering the impacting drop morphology and quantitative trends in splashing is studied in detail. The variation of splash behavior of drops impacting at three distinct drop Weber number (We) at different surface temperatures (TS) is considered. For a fixed We, an increase in the surface temperature causes a shift in the impact dynamics from splashing to spreading, which indicates higher threshold We at higher TS. The dynamics of the ejected liquid sheet and the spreading lamella post-impact are analyzed from the high-speed images to quantify the observed transitions.

This paper presents an experimental study to demonstrate transitions between luminous and blue whirling flames. We have designed a fixed-frame four-sided square aluminum test apparatus where liquid fuel is injected vertically upward from the center of the base plate, which reacts with self-entrained tangential air. Our experiments show evidence of the transition of a highly luminous non-premixed fire whirl into a blue whirling flame above the base surface. The formation of a blue whirling flame depends on the enhanced mixing and fuel spread above the metal surface. The smooth surface boundary leads to the formation of Ekman type boundary layer, which is shown to affect the flame. The transition from a fire whirl to a blue whirling flame is found to be dependent on the fuel spread over the metallic surface which depends on the fuel flow rate. This paper also shows the visual observation of flame patterns at different fuel flow rates, gap size (for air entrainment), and surface conditions (to identify the key parameters that are responsible for the formation of a blue whirling flame). The flame images show that a conical whirling flame with a laminar bluish base could be an essential (but not enough) condition to produce the transition to blue whirling flame.

One of the most important milestones in gas turbine burner technology was the incorporation of swirling flows for flame stabilization. The objective of present work is the design and development of a generic fuel flexible multiple swirl burner with enhanced flashback resistance and low emissions. The burner design will allow operation in premixed and non-premixed modes with liquid and gaseous fuels. The investigated burner consists of 3 annular co-rotating swirlers: an outer radial swirler stage and two concentric axial swirler stages. Insights from the first isothermal and reactive numerical simulations for premixed methane–air combustion are being presented here. Results based on the characterization of the flow fields, temperature distribution, streamwise and azimuthal shear layer dynamics, and turbulence characteristics are presented. The velocity profiles obtained from isothermal numerical simulations are also validated by experimental results. Flame stabilization and flashback propensity are discussed with respect to the features of vortex breakdown, specifically the central recirculation zone (CRZ).

Gelled propellants for rocket propulsion applications offer the advantage of safer storage and handling in comparison with liquid fuels. Sheet formation and break-up study of non-reactive gel simulant prepared with Carbopol 934 in de-ionized water was conducted for understanding atomization of gels by impinging jets configuration. Material properties of gels prepared and their flow behavior estimated. The simulant is injected through an orifice of 0.413 mm up to a range of 20 bar injection pressure is studied by analyzing high-speed shadow graph images of liquid jets impingement, sheet formation, and disintegration. The prominent effect of gelling agent concentration in sheet formation and break-up is revealed. Sheet break-up is occurring in two different modes for five different gel concentrations. Waves generated from impingement point caused break-up of sheets for low-concentration gels and high-velocity jets while tearing and hole formation in sheets led to their break-up for mostly high-concentration gels and particularly for low jet velocity modes. Droplet trajectories and their velocity at various locations in the periphery of sheets were measured for two different jet velocities for five gel concentrations.

This paper reports an experimental study of spray characteristics in rotary slinger atomizers using different laser diagnostic tools. The main objective of the present study is to investigate the effect of size of the slinger orifices on liquid breakup dynamics and droplet size distribution over a wide range of rotational speed (5000–45,000 rpm) and liquid flow rates (0.2–2 lpm). Three different slinger discs having the same number of orifices but different sizes (d0 = 1, 1.5, and 2 mm) are considered. The primary liquid breakup visualization is achieved using volumetric laser-induced fluorescence (VLIF) technique; whereas, the droplet size is measured by interferometric laser imaging for droplet sizing (ILIDS) technique. The results demonstrate critical role of orifice size on liquid breakup mode and droplet size up to rotational speed about 30,000 rpm, beyond which the aerodynamic force dominates the atomization process such that neither the orifice size nor rate of rotation has strong influence.

Injector plays a pivotal role in meeting requirements of combustion performance in terms of combustion efficiency, flame stability, ignition, lower emissions, etc. In a multi-swirler injector configuration, air flow field inside injector is mainly dictated by primary swirler. Present CFD studies have been attempted to characterize flow field of a conical nozzle fitted with a radial swirler. Embedded LES-based hybrid model has been used where computational domain is divided into three zones which are seamlessly connected by capturing the interface fluid dynamics. In LES zone, both time and spatial scales have been resolved based on the results of a precursor RANS analysis. Analysis is carried out with CFL no. around 2, time step of 1 μs. The analysis is reasonably able to capture various unsteadiness (PVC, CTRZ, frequencies, TKE useful for the atomization of liquid fuel) which are not possible to be captured using URANS models.