Design and Analysis Software for Propellers

Recent interest in the field of micro and nano scale air vehicles attracted the attention of many researchers all over the world. The challenge associated with these classes of vehicles is to develop efficient miniaturized components. There are different types of micro and nano air vehicles out of which fixed wing micro air vehicle is one of them. Propulsion system for most of the fixed wing MAVs is propeller driven by an electric motor powered by a battery. The endurance of the MAV mainly depends on the performance of these two components. Hence there is a scope to improve the performance of the propeller and motor. Efficient propeller design and its performance analysis are an iterative process and time consuming. In the present study, to ease the process of propeller design and analysis NALPROPELLER code has been developed using MATLAB. This code is based on minimum induced loss theory presented by E.E.Larrabee to generate planform, blade element momentum theory along with Prandtl hub-tip loss model for overall performance analysis and the performance plots could be viewed in the GUI windows. The code consists of three modules namely single airfoil design, multi airfoil design and analysis module. This code is compared with one of the propeller design and analysis code available in the internet JavaProp by Martin Hepperle, which is also based on minimum induced loss method. From literature Eppler 193 airfoil show high lift to drag ratios at low Reynolds numbers [16]. Eppler-193 airfoil is used in the evaluation of propeller performance. A four inch diameter, two bladed, fixed pitch propeller is designed and analysed using this code. The design is compared with one of the design software JavaProp available online as an open source. A poly urethane casting propeller is fabricated based on the design. The performance comparison of the NALPROPELLER code, JavaProp and 3D CFD analysis is presented and discussed.

Life Prediction of Directionally Solidified Air Cooled HPT Gas Turbine Blade Used in a Supersonic Aircraft Using FEM

Life estimation of Directionally Solidified (DS) MARM-247 HPT gas turbine blade used in a turbofan engine of a supersonic aircraft is presented. These blades were drafted into the engine as a replacement for the polycrystal (NIMONIC) blades since a more efficient, reliable and durable material with high strength and temperature resistance was required to further enhance the life of the turbine blade and the efficiency of the power generation process. The supersonic aircraft is having a repeated mission cycle of a fast acceleration from idle, a 1hr cruise at Mach 1.5 and a fast deceleration to idle. The mission cycle which is a repetition of acceleration, cruise and deceleration cycles can produce wide variety of complex loading conditions which can result in HCF, LCF and creep damage of the turbine blade. Empirical equation of the universal slope developed by Manson was used to estimate the damage component due to LCF. The cumulative stresses and strains due to creep as a function of time was determined using Time hardening rule. Creep data for MARM-247 was correlated using LMP to predict the lives to 1% of creep strain at worst possible combination of temperature and stress value. Damage due to creep per mission cycle was determined using Life fraction Rule proposed by Robinson and Taira. The vibration characteristics of the turbine blade were predicted using Modal analysis. Campbell diagram was plotted to ascertain whether any nozzle passing frequency fall within the working range of the blade. Harmonic analysis was carried out to evaluate the magnitude of the alternating stresses resulting from the blade vibrations at resonance during the acceleration and deceleration cycle. HCF life of the turbine blade was assessed using Goodman diagram. The total damage of the turbine blade per mission cycle due to the above loading was assumed as the combination of the individual damage due to fatigue and creep. Time to failure under combined creep and fatigue damage was estimated using linear damage rule. Non linear features of FEA tool ANSYS12.0 was exploited to calculate the stress distribution, creep, plastic and the total strain encountered by the turbine blade as a function of mission cycle time. The loading spectrum associated with the mission cycle which includes the temperature, gas pressure and the speed profiles were obtained from a sophisticated engine ground test facility which was configured to simulate actual engine operating conditions. The proposed method of cyclic life estimation using FEM was validated by performing various component and engine level tests. A good agreement was observed between the calculated and observed blade lives.

Drop Size Characteristics of Forward Angled Injectors in Subsonic Cross Flows

Cross flow fuel injection is a widely used approach for injecting liquid fuel in gas turbine combustors and afterburners due to the higher penetration and rapid mixing of fuel and the cross flowing airstream. Because of the very limited residence time available in these combustors it is essential to ensure that smaller drop sizes are generated within a short axial distance from the injector in order to promote effective mixing. This requirement calls for detailed investigations into spray characteristics of different injector configurations in a cross-flow environment for identifying promising configurations. The drop size characteristics of a liquid jet issuing from a forward angled injector into a cross-flow of air were investigated experimentally at conditions relevant to gas turbine afterburners. A rig was designed and fabricated to investigate the injection of liquid jet in subsonic cross-flow with a rectangular test section of cross section measuring 50 mm by 70 mm. Experiments were done with a 10 degree forward angled 0.8 mm diameter plain orifice nozzle which was flush mounted on the bottom plate of test section. Laser diffraction using Malvern Spraytec particle analyzer was used to measure drops size and distributions in the near field of the spray. Measurements were performed at a distance of 70 mm from the injector at various locations along the height of the spray plume for a reasonable range of liquid flow rates as in practical devices. The sprays were characterized using the non dimensional parameters such as the Weber number and the momentum flux ratio and drop sizes were measured at three locations along the height of the spray from the bottom wall. The momentum flux ratio was varied from 5 to 25. Results indicate that with increase in momentum flux ratio the SMD reduced at the specific locations and an higher overall SMD was observed as one goes from the bottom to the top of the spray plume. This was accompanied by a narrowing of the drop size distribution.

Design and Analysis of a High Pressure Turbine Using Computational Methods for Small Gas Turbine Application

This paper describes the aerodynamic design and analysis of a high-pressure, single-stage axial flow turbine suitable for small gas turbine engine application using computational methods. The specifications of turbine were based on the need of a typical high-pressure compressor and geometric restrictions of small gas turbine engine. Baseline design parameters such as flow coefficient, stage loading coefficient are close to 0.23 and 1.22 respectively with maximum flow expansion in the NGV rows. In the preliminary design mode, the meanline approach is used to generate the turbine flow path and the design point performance is achieved by considering three blade sections at hub, mean and tip using the AMDC+KO+MK+BSM loss models to meet the design constraints. An average exit swirl angle of less than 5 degrees is achieved leading to minimum losses in the stage. Also, NGV and rotor blade numbers were chosen based on the optimum blade solidity. Blade profile is redesigned using the results from blade-to-blade analysis and through-flow analysis based on an enhanced Dawes BTOB3D flow solver. Using PbCFD (Pushbutton CFD) and commercially available CFD software ANSYS-CFX, aero-thermodynamic parameters like pressure ratios, aerodynamic power, and efficiencies are computed and these results are compared with one another. The boundary conditions, convergence criterion, and turbulence model used in CFD computations are set uniform for comparison with 8 per cent turbulence intensity. Grid independence study is performed at design point to optimize the grid density for off-design performance predictions. ANSYS-CFX and PbCFD have predicted higher efficiency of 3.4% and 1.2% respectively with respect to targeted efficiency of 89 per cent.

Numerical Studies on the Effect of Tip Clearance and Tip Cooling Flow on Endwall Losses of an Unshrouded Axial Turbine Rotor

This paper deals with the numerical investigations on a low pressure axial turbine stage to assess the effect of variation in rotor tip clearance and tip coolant ejection rate on the end wall losses. The rotor, along with the NGV, was modeled to represent the entire turbine stage. The CFX TASCflow software was used to perform steady state analysis for different rotor tip clearances and different tip coolant ejection rates. The locations of the cooling slots were identified on the blade tip and the coolant ejection rate was specified at these areas. The simulations were carried out with tip clearances of 0%, 1% and 2% of blade height and ejection flow rates of 0.5%, 0.75% and 1% of main turbine flow rate. It is shown that the size and strength of the leakage vortex is directly related to the tip clearance. The reduction in efficiency is not in linearity with the tip clearance owing to the effect of boundary layer growth on the end walls. Introduction of the tip coolant flow shows increased total–total efficiency compared to that of the uncooled tip. This is attributed to a reduction in the strength of the leakage vortex due to reduced cross-flow over the tip clearance from pressure surface to suction surface. At a coolant flow rate of 0.75% of the main flow rate, there is significant increase in efficiency of about 0.5%. Optimum tip clearance and coolant flow rate are obtained based on the results of the present analysis.

Coolant Gas Injection on a Blunt-Nosed Re-Entry Vehicle

Very high thermal loads are expected in re-entry vehicles traveling at hypersonic Mach numbers due to severe aerodynamic heating. In the present study, numerical investigations are carried out to analyze the use of film cooling technology for a fully reusable and active thermal protection system of the re-entry vehicle. Simulations are done to examine the fundamental flow phenomenon and the performance of blunt body film cooling in hypersonic flows. Simulations are conducted for a blunt -nosed spacecraft flying at Mach numbers varying from 4 to 8 and 40 deg angle of attack. Film cooling holes are provided on the bottom of the blunt-nosed body. Standard values at an altitude of 30 km are used as in flow boundary conditions. The dependency of blowing ratios, stream-wise injection angle and inlet Mach number on the film cooling effectiveness are investigated. It is observed that the film cooling effectiveness reduces with increase in coolant injection angle. The film cooling performance is found to be decreasing with increase in Mach number. The results could provide useful inputs for optimization of an active thermal protection system of re-entry vehicles.

Influence of Pre-History and Leading Edge Contouring on Aero-Performance of a 3D Nozzle Guide Vane

Experiments are conducted to investigate the effect of the pre-history in the aerodynamic performance of a three-dimensional nozzle guide vane with a hub leading edge contouring. The performance is determined with two pneumatic probes (5 hole and 3 hole) concentrating mainly on the endwall. The investigated vane is a geometrically similar gas turbine vane for the first stage with a reference exit Mach number of 0.9. Results are compared for the baseline and filleted cases for a wide range of operating exit Mach numbers from 0.5 to 0.9. The presented data includes loading distributions, loss distributions, fields of exit flow angles, velocity vector and vorticity contour, as well as, mass-averaged loss coefficients. The results show an insignificant influence of the leading edge fillet on the performance of the vane. However, the pre-history (inlet condition) affects significantly in the secondary loss. Additionally, an oil visualization technique yields information about the streamlines on the solid vane surface which allows identifying the locations of secondary flow vortices, stagnation line and saddle point.

Identification of Multiple Fault Parameters in a Rigid-Rotor and Flexible-Bearing-Coupling System: An Experimental Investigation

Rotating machineries are very common and widely used in the modern industrial world. A breakdown of the rotating machine may result in economic losses and even worse, in the damage of human beings. So it needs accurate and reliable prediction of its dynamic characteristic. The multiple fault identification algorithm developed in [1] identifies the bearing and coupling dynamic parameters along with residual unbalances in a rigid-rotor and flexible-bearing-coupling system has been validated experimentally in the present article. An indigenously developed test rig was used for the experimentation. After estimating multiple fault parameters the accuracy of estimated parameters has been checked by checking the consistency of the estimated parameters and by performing the standard impact test. It has been observed that with more number of measurements consistencies of the estimated parameters are very good.

Design and Testing of a Can Combustor for a Small Gas Turbine Application

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.

Parametric Study of Moss-Brookes (MB) and Moss-Brookes-Hall (MBH) Model Constants for Prediction of Soot Formation in a Turbulent Hydrocarbon Flames

In the present work, two equation soot models proposed by Moss-Brookes (MB) and Moss-Brookes-Hall (MBH), available in ANSYS FLUENT14.5, are used to study the soot formation in a turbulent kerosene-air flame. The model constants in the original works of MB and MBH model were primarily tuned for the methane-air or other lower hydrocarbon flames. In this work, the emphasis has been given on the applicability of these models in modeling the soot formation in heavy hydrocarbon fuels. The current work is primarily focused on the parametric study of the various modeling constants for calculating the soot inception and oxidation rates. A parametric study is performed to calculate the soot inception rates by considering different soot precursors like C2H2, C2H4, C6H6 and C6H5. Steady laminar flamelet approach with a detailed chemical reaction mechanism (Jet_SurF_2.0), is used for modeling gas phase combustion. The current numerical predictions are compared with experimental results of Young et al. [1] and earlier published numerical results of Wen et al. [2]. The study is further extended to understand the role of chemical reaction mechanism on soot predictions considering detailed versus reduced (JP10revC) chemical mechanisms.