The performance of the gas turbine engines can be improved using variable area nozzle turbine (VANT) in the off-design condition by actively controlling the mass flow rate for various applications of engine. However, VANT demands part clearances to be provided near both the hub and tip of the vane for its mechanical movement. This clearance induces leakage flow and formation of the leakage vortex which interacts with the secondary flow in the vane passage. In this paper, the numerical study is performed on high endwall angle vane geometry of 2 nd stage LPT from Pratt and Whitney E 3 engine using commercially available CFD tool ANSYS CFX. The effect of free stream turbulence at 0.5%, 5% and 10% is studied along with the change of incidence angle from −10 to +10 on the annular cascade nozzle flow field. The static entropy contour, vorticity contour and total pressure loss coefficient are used to analyzed the flow. The free stream turbulence affects the structure of the vortices within the flow passage. However, no significant change on total pressure loss coefficient is present as turbulence intensity is changed from 0.5% to 5%, but total pressure loss coefficient increases as turbulence intensity is changed to 10%.

The paper presents the results of numerical simulation of the effect of flow nonuniformities at the engine inlet on the working process of the engine fan. Flow nonuniformities is created by pushing the interceptor into the flow path of inlet device like as it is often done during field tests. The authors have created a numerical model capable of considering non-stationary processes in the fan using nonlinear harmonic analysis. As a result, qualitative and quantitative estimates were obtained of the influence of overlapping of the inlet duct by the interceptor on the main parameters of the fan workflow. It is shown that the more the duct is blocked, the more its parameters are deteriorated. Moreover, the deterioration is not linear, but according to the dependence of the 2nd order.

The paper presents the main ideas of the virtual test bench concept for rapid obtaining of the reliable characteristics of compressors based on a multi-level mathematical model with a two-step identification using data obtained from mathematical models with a high order of accuracy. One of the possible identification algorithms and the results of its successful testing are given on the example of a centrifugal compressor stage developed and tested at NASA.

In this paper the influence of various parameters of the primary and secondary zones of a small-sized combustion chamber on its environmental characteristics was studied. The study of the environmental characteristics of the combustion chamber was carried out in two ways. The first method consisted of two steps. The first step was a 3D simulation of combustion processes using FGM combustion model. The second step was based on results of the first step using reactors network implemented in the ANSYS Fluent 18.2 software. The construction of the reactors network in this approach occurred automatically according to a temperature and mixture fraction. The number of reactors in the result was about 500. The second method was a simple model of a reactors network in which the primary zone was simulated by a perfectly stirred reactor, and the secondary zone was a plug flow reactor. Methane was used as a fuel. The influence of the residence time of the mixture and fuel-air equivalence ratio in each zone on the emission of CO and NO x at the combustion chamber exit was studied. The residence time and fuel-air equivalence ratio for the first method were changed using the design parameters of the combustion chamber. For a simple reactors network, these parameters are set as input data, so this method can be used at the preliminary design stage. As a result of the work, a method was obtained that allows to find the parameters of the primary and secondary zones of the combustion chamber in order to minimize pollutant emissions at the preliminary design stage.

In this research, heat transfer and pressure drop from a tube with vortex generators (VGs) insert are numerically investigated. The effects of heights, attack angles of VGs inside a tube on heat transfer and flow behavior are investigated. CFD simulations, with and without VGs insert, are done for an air flow range (Reynolds numbers 6000 to 33000) and for a constant heat flux on the tube model surface. Four VGs are fitted in a circular pattern on the inner surface of the tube. We studied the characteristics of the delta winglet VGs for different attack angles and blockage ratios. The Nusselt number and friction factor results show the influence of the VGs insert on heat transfer and frictional factor. The maximum Nusselt number increment (Nu/Nu 0 ) was achieved to be 1.75 while the maximum friction factor increment (f/f 0 ) was 3.21. In order to understand the flow behavior and different vortices, path lines released by the VGs surface and details of the vortices are also studied.

Due to the significant increase in the human population, satisfying its energy demand becomes a major challenge at present and in the future. Hydropower is a form of renewable energy that converts the potential or kinetic energy of water to useful mechanical power. Considering the low speed of streaming water, hydraulic Savonius turbines can be effectively used for such conditions. Such turbines are robust and very simple to manufacture. However, Savonius turbines suffer from a poor efficiency. For this purpose, an optimization process was done to improve the power density of a hydraulic Savonius turbine while keeping its simple structure. Here in this study, the two main geometry parameters, i.e., 1) the overlap ratio, 2) the gap ratio, are served. Furthermore, the optimal operating range for both parameters is presented. To achieve this goal, many different geometries of the turbine are simulated by using the computational fluid dynamics (CFD) code Star-CCM+. In the optimization process, all simulations are driven by the in-house software OPAL++. The output power coefficient (Cp) is considered as a single target function to be maximized. The performance of the optimal geometry is compared with the standard design over the whole range of operation, leading to an improvement in the performance of the turbine by about 8%.

Various studies have been carried out related to the labyrinth seals and reported in the open literature using the different seal arrangements at the stator-rotor seal cavity region. In the present study, numerical analysis has been carried out for the static and rotational effects of labyrinth seals at various flow and geometrical, parametric conditions for the optimized leak flow using straight and steeped seal configurations. And, an experimental data has been generated for the straight through seals, and the numerical data of the same case is validated with the experimental data. The k-omega SST turbulence model is considered with 5% turbulence intensity for the CFD analysis. At a particular seal clearance, as the number of teeth increases the leakage flow is found to be decreased. The leak flow is found to be lower with the stepped labyrinth seals in comparison to the straight through seals. The leak flow amount is found to be lower at a rotational condition in comparison to the stationary condition. From the overall results, it is observed that the stepped seal with a lower clearance at a compressor bleed air temperature and rotational conditions have shown better performance with the lower leak air mass flow.

Future developments have been indicated for further research and development in the Aeroacoustics of the components of Unmanned Aerial Vehicles (UAVs). In which the implementation of multi-rotor UAVs in the complex applications is quite high but it has the drawback of high drive-line noise levels, which is one of the prime radar detectability factors. As a result, an idea is emerged to design and test the quite UAV, in which the noise from propellers plays a major role. in order to the successful completion of the design study, a complete is conducted, in which the design parameters and various noise reduction methodologies in the rotating components have been noted and included in the final design. To minimize the noise signature issue in UAV, the idea finalized that to minimize the decibel of small Unmanned Aircraft System (UAS) propellers via leading-edge modifications. A computer-aided design of base propeller and three different versions propeller with leading-edge modifications are generated with the help of CATIA for Computational Fluid Dynamics (CFD) simulations. Comparative noise variation simulations between the existing and the propellers with modification are performed, in which dynamic conditions play a predominant to initiate the analysis and thereby the analyses are carried out with the help of ANSYS Workbench Fluent 16.2. Especially, to make an acceptable solution, the Moving Reference Frame (MRF) approach is used in order to capture the propeller rotation in an effective manner. Finally, a propeller with airfoil cut at the leading edge has induced the low noise.

Can type combustors are robust, with ease of design, manufacturing and testing. They are extensively used in industrial gas turbines and aero engines. This paper is mainly based on the work carried out in designing and testing a can type combustion chamber which is operated using JET-A1 fuel. Based on the design requirements, the combustor is designed, fabricated and tested. The experimental results are analysed and compared with the design requirements. The basic dimensions of the combustor, like casing diameter, liner diameter, liner length and liner hole distribution are estimated through a proprietary developed code. An axial flow air swirler with 8 vanes and vane angle of 45 degree is designed to create a re-circulation zone for stabilizing the flame. The Monarch 4.0 GPH fuel nozzle with a cone angle of 80 degree is used. The igniter used is a high energy igniter with ignition energy of 2J and 60 sparks per minute. The combustor is modelled, meshed and analysed using the commercially available ansys-cfx code. The geometry of the combustor is modified iteratively based on the CFD results to meet the design requirements such as pressure loss and pattern factor. The combustor is fabricated using Ni-75 sheet of 1 mm thickness. A small combustor test facility is established. The combustor rig is tested for 50 Hours. The experimental results showed a blow-out phenomenon while the mass flow rate through the combustor is increased beyond a limit. Further through CFD analysis one of the cause for early blow out is identified to be a high mass flow rate through the swirler. The swirler area is partially blocked and many configurations are analysed. The optimum configuration is selected based on the flame position in the primary zone. The change in swirler area is implemented in the test model and further testing is carried out. The experimental results showed that the blow-out limit of the combustor is increased to a good extent. Hence the effect of swirler flow rate on recirculation zone length and flame blow out is also studied and presented. The experimental results showed that the pressure loss and pattern factor are in agreement with the design requirements.

This paper deals with the numerical investigations on the effect of trailing edge crenulation on the performance of a transonic axial compressor rotor. Crenulation is broadly considered as a series of small notches or slots at the edge of a thin object, like a plate. Incorporating such notches at the trailing edge of a compressor cascade has shown beneficial effect in terms of reduction in total pressure loss due to enhanced mixing in the wake region. These notches act as vortex generators to produce counter rotating vortices, which increase intermixing between the free stream flow and the low momentum wake fluid. Considering the positive effects of crenulation in a cascade, it was hypothesized that the same technique would work in a rotating compressor to enhance its performance and stall margin. However, the present CFD simulations on a transonic compressor rotor have given mixed results. Whereas the peak total pressure ratio in the presence of trailing edge crenulation reduced, the stall margin improved by 2.97% compared to the rotor with straight edge blades. The vortex generation at the crenulated trailing edge was not as strong as reported in case of linear compressor cascade, but it was able to influence the flow field in the rotor tip region so as to energize the low momentum end-wall flow in the aft part of the blade passage. This beneficial effect delayed flow separation and allowed the mass flow rate to be reduced to still lower levels resulting in improved stall margin. The reduction in pressure ratio with crenulation was surprising and might be due to increased mixing losses downstream of the blade.

The main objective of this work is the validation of Computational Fluid Dynamics (CFD) code used for analysis of transonic axial compressors. NASA Rotor 35 is used here as test case for validation. In this work, computations are performed using parallelized RANS code, to predict the transonic axial compressor rotor flow characteristics. Advection Upstream Splitting Method (AUSM) scheme has been used. A Multiple Frame of Reference approach has been used to model the rotor passage. Spalart-Allmaras turbulence model is used to model turbulence. Multiblock Structured mesh is used. Performance characteristics for the entire range of operation, from maximum mass flow rate till maximum pressure ratio, have been simulated. The results obtained are comparable with experimental data within 5–10% error. Investigations have been carried out to study the effect of varying tip clearance in NASA Rotor 35. The present work is intended to study the clearance flow trajectory as a function of varying tip clearance. The effects of shock/vortex interaction in tip clearance region are also studied. The effects of tip clearance size on the generation and evolution of the end-wall vortical structures are discussed by investigating their evolutionary trajectories. By this study, it is observed that as tip clearance reduces, clearance flow trajectory moves downstream. From this it can be concluded that if tip clearance increases, tip clearance vortices expand. This may help in casing-treatment or tip-treatment to mitigate the loss in the performance, if the tip clearance increases.

The effect of solid thermal conductivity on conjugate heat transfer from a flat plate with combined impingement and film cooling is studied both experimentally and computationally. One side of the plate is exposed to hot mainstream, while the other side experiences impinging air jets. A geometric configuration with multiple staggered rows of cylindrical film holes and a matrix of impingement holes is considered. In the experiments a thermo-chromic liquid crystal (TLC) technique is used to measure the surface temperature. For the purpose of computations, the physical domain is meshed with very fine sized hexahedral and tetrahedral grids suitable for the finite volume method. Grid refinement is carried out using GCI method and SST κ-omega RANS equations are used for turbulence modeling. Three blade materials: A - with k = 0.2W/m K, B - with k = 1.5W/m K and C - with k = 15W/m K and three blowing ratios, M = 0.6, 1.0 and 1.6 are considered for the conjugate heat transfer study. A good agreement of effectiveness distribution on the interaction surface is found between the measured and the computations. On the interaction surface, heat flux values are the lowest for material A and progressively increase with increasing thermal conductivity. However the effectiveness values vary significantly for material A in the stream-wise direction. The Nusselt number on the other hand differs only slightly among different materials.

Casing Treatment (CT) is one of the passive methods to increase the stability margin of the compress and hence that of the aircraft jet engines. In this paper, a novel J-shaped axial CT slot geometry is designed and numerically analysed for its effect on the performance of a single stage NACA transonic compressor. The predicted performance of the isolated rotor was validated by comparing with the published experimental results. The predicted efficiency of the baseline transonic rotor agreed well with experimental data, but the total pressure ratio was under predicted over the entire operating range. The J-shaped CT slots, with 100% axial coverage over the rotor tip chord, were able to extend the stall mass flow rate by almost 19.45% compared to the baseline rotor, accompanied with a slight reduction in rotor efficiency by 1.42%. The high pressure air entered the slots at rotor exit and flowed back through the slots and the plenum, and ejected at the rotor inlet to energise the low momentum end wall flow. The interaction of main inlet flow and the ejected flow having large tangential component of velocity, had favourable effect on the rotor incidence angle, and hence on rotor stall margin.

In the present work, the performance of Steam Ejector System in High Altitude Test (HAT) facility is numerically studied, in the absence of the condenser. Steam is used as secondary fluid to eject the burnt gases into atmosphere. Experimental visualization of mixing of burnt gas and steam and subsequent flow pattern is difficult, hence numerical simulation using FLUENT was done and the resulting flow stream lines, static and total pressures, shock patterns are computed along the ejector system to understand the physics of the problem. Three burnt gas flow rates of 9.17, 27.5 and 45.8 kg/s corresponding to lower, mid and upper limits of ejection from the HAT facility with the steam flow rate of 50 kg/s from Ejector I and 130 kg/s from Ejector II are studied. This corresponds to three cases of Entrainment Ratios for each of the ejector. Results show that for a burnt gas flow rate of 27.5 and 45.8 kg/s with the given dimensions of the HAT facility provided by ASL, DRDO, the gas and steam start mixing in the converging duct, pass through the mixing tube and attains atmospheric pressure at the exit of the HAT facility. For the burnt gas flow rate of 9.17 kg/s, reverse flow is observed in the Ejector II, indicating the malfunction mode of the system for the given design parameters.